UH-60M Test PIlot Course Fam Guide [PDF]

Zitiervorschau

UH-60M

UH-60M Developmental Test Pilot Systems Familiarization Training

FOR TRAINING PURPOSES ONLY Publication Date: September 2003

Draft: 4 September 2003 TK46036

8 September 2003

SUBJECT: UH-60M Developmental Test Pilots Systems Familiarization Manual ATTACHMENT: UH-60M Developmental Test Pilots Systems Familiarization Manual, Initial Version September 2003 TO: Recipient of Manual The content of this manual reflects UH-60M Aircraft No. 1 and 2 configured for first flight and early developmental flight-testing. This manual is intended to familiarize Government Developmental Test Pilots and other personnel with the basic systems configurations of the UH-60M as equipped for aircraft first flight. Currently, no plans exist for the update of this manual. This manual provides only an overview of the UH-60M. The intended purpose of this manual is for aircraft familiarization in a training environment only. The following sections of this manual have not been reviewed or validated by Sikorsky Engineering personnel: Section 2-5 Section 2-10 Section 3-1 Section 3-2 Section 3-3

Aircraft Systems - Automatic Flight Control System Aircraft Systems - Flight Display System Avionics Systems - Flight Management System Avionics Systems - Communication Systems Avionics Systems - Navigation Systems

Please provide comments or observation related to these training materials to: Paul Robinson Army Programs Training Manager. Sikorsky Aircraft Corp. V 203-384-7123 F 860-998-4812 [email protected]

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UH-60M Developmental Test Pilot Course Index Unit 1 - Introduction

Unit 4 – End of Course Administration

Section 1 – Introduction to Course

Section 4-1 – Course critique Stuff

Unit 2 – Aircraft Systems

4-1.1

Glossary

Section 2-1 – Introduction Section 2-2 – T-700-701D Section 2-3 – Fuel System Section 2-4 – Mechanical Flight Control Systems Section 2-5 – Dual Digital AFCS Section 2-6 – Powertrain Systems Section 2-7 – Main and Tail Rotor Systems Section 2-8 – Electrical Systems Section 2-9 – Aircraft Lighting Systems Section 2-10 – Flight Display System Section 2-11 – Active Vibration Control System Unit 3 – Avionic Systems Section 3-1 – Flight Management System Section 3-2 – Communication Systems Section 3-3 – Navigation Systems

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Section 1-1 Introduction to Course

Food Service Location/Hours

Course Scope

Sikorsky Cafeteria – Building 1 – First Floor

The UH-60M Developmental Test Pilot Familiarization Course is designed to provide required knowledge of the UH-60M aircraft and it’s systems to United States Army Aviators or Department of the Army Civilians (DAC) qualified as rotary wing aviators assigned as member of the UH-60M Combined Test Team (CTT). The course provides general information pertaining to aircraft systems that have been impacted by the UH-60M modernization. Information contained in the course is reflective of Aircraft 1 as it is configured for first flight. In some cases, data related to systems or equipment modes or functionality that is not active for first flight has been included for informational purposes.

Schedule Scheduled presentation times will start at 8:00 am each day and conclude at 4:00 pm. Teaching periods will generally last 50 minutes. In some cases periods may be longer, as required. One hour will be provided for lunch.

Course Completion Criteria Participants must be present in the classroom for at least 75% of total class time to receive a Certificate of Attendance.

Important Phone Numbers Sikorsky Security Gate (561) 775-5511/-5512 (ext.5-5511/5-5512)

Classroom Fire/Medical Emergency

Breakfast Lunch

6:00 – 8:30 AM 10:30 – 12:30 PM

Vending Machines are available in the cafeteria.

Approved Smoking Areas Sikorsky Aircraft is a non-smoking facility. Smoking is permitted only in the approved areas outside/away from the buildings and hangars. Dispose of used butts in the provided containers in the approved smoking areas.

Facility Access All visitors must be cleared for access through Sikorsky Aircraft Security. Send all visitor requests To the Attn: DFC Kaz Lott Tel. (561) 775-5240 Fax (561) 775-5414.

Flight Line Access Unauthorized personnel are to remain clear of the Flight Line/Hangars/Aircraft. Flight Line/Hangar/Aircraft Access must first be cleared by Sikorsky Aircraft Hangar Supervision and escorted to the Flight Line/Hangar/Aircraft by a Sikorsky employee.

Safety Glasses Eye protection is required in the hangars and on the Flight Line. Aircraft Ground/Flight operations (If access is authorized) require eye protection and hearing protection.

Photography/Recording Devices Cameras/videos and recording devices are not permitted on the facility unless first cleared by Sikorsky Security Manager (561) 775-5222.

(561) 775-5555 (ext. 5-5555)

9/3/2003 Page 1 of 2 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

NOTES

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Section 2-1 UH-60 Modernization Introduction UH-60M Program Overview The UH-60 Black Hawk is Army aviation’s primary utility aircraft. With more than 1,500 fielded, UH60s account for the largest percentage of the Army’s annual flying hours. Serving honorably since 1978, the Black Hawk is a common platform that is capable of performing multiple missions, including air assault, medical evacuation, command and control and special operations. In production between 1978 and 1988, 960 UH-60A models have been flying with an average age of more than 18 years. UH-60L production started in 1989. UH-60L models currently have an average age of eight years. As a result of the climbing fleet age and declining mission capability rates coupled with increased operating expenses, the U.S. Army’s Utility Helicopter Program Manager has established the UH-60M modernization program to ensure that the Black Hawk remains a viable part of the Objective Force and relevant on future battlefields. The modernization program of the UH-60 will enhance the Objective Force commander’s ability to conduct non-linear, simultaneous, fully integrated operations in order to decisively mass the effects of the Unit of Employment or Unit of Action’s warfighting assets. The UH-60M, enhanced by digital connectivity as well as improved lift, range, deployability, and survivability, will increase the commander’s ability to conduct operations across the entire spectrum of the battlespace. The modernization program will allow for an 8000 flight-hour airframe service life extension as well as allow the insertion of new technologies into the UH-60. Among them, systems and equipment that increase pilot efficiency, increase mission safety and effectiveness, provide a digital communications architecture, and enhance survivability. Just as aircraft systems have been enhanced to allow increased operational flexibility, improvements in Reliability, Availability, and Maintainability (RAM) have also been designed into the UH-60M to reduce Operational and Support (O&S) costs and allow for future system growth.

development and production approach resulting in a fleet with mixed performance capability. The first tier will provide life extension, digitization, and other enhancements to make the UH-60 capable and effective on the modern battlefield. The second tier will focus primarily on performance related capabilities needed to support the Army Objective Force requirements that cannot be achieved today within acceptable cost, schedule, and risk constraints. UH-60A and UH-60L model aircraft will be retrofitted to the UH-60M configuration using kits. New production UH-60M aircraft will be built under future multi-year H-60 production contracts. The UH-60M will begin being fielded in the 2006/07 time frame. The modernization program will result in a fleet average age of 10 years.

UH-60M Specifications General The UH-60M is a twin-engine, single-rotor, four bladed utility helicopter. It is designed to be joint forces capable and operate 24 hours a day including operations at night an in adverse conditions. It is used to support the following types of operations: • • • • •

Air Assault Air Cavalry Troop and Equipment Transport Command and Control Medical Evacuation

The UH-60M carries a crew of four and up to 11 additional troops.

Dimensions Principle dimensions of the helicopter are based on the cyclic stick and tail rotor pedals being centered and the collective stick being in its lowest position. All dimensions are approximate.

The UH-60 modernization and enhancement program will be achieved through the two-tiered

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1

3

2

4

5

6 18

17

16

15 14

13

12

1. 2. 3. 4. 5.

10

9

20

19

25

11

24

5

23

Pitot Cutter Backup Hydraulic Pump No. 1 Hydraulic Pump and No. 1 Generator Upper (Rotor Pylon) Cutter Infrared Countermeasure Transmitter Provisions 6. Aft Maintenance Light Receptacle 7. Tail landing Gear Deflector 8. Chaff Dispenser Provisions 9. APU Exhaust Port 10. Rear Laser Detection Sensor Provisions 11. Pneumatic Port 12. Pressure and Closed Circuit Refueling Ports

8 21

7 21

8

21

22 UMHV001

13. No. 1 Engine 14. Main Landing Gear Deflector/Cutter 15. Forward Laser Detector Sensor Provisions 16. Step and Extension Deflector 17. Landing Gear Joint Deflector 18. Door Hinge Deflector 19. Right Position Light (Green) 20. Fire Extinguisher Bottles 21. Formation Lights 22. Tail Position Light (White) 23. APU 24. Left Position Light (Red) 25. Pitot Tubes

UH-60M General Arrangement 9/3/2003 Page 2 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

26

28

27

39

38

35

36

37 47

34

33 32

29

30

31

26

40 40 41

42 43 45

41 42 43 44 UMHV002

26. Upper Anticollision Light 27. Tail Drive Shaft 28. No. 2 Hydraulic Pump and No. 2 Generator 29. Pylon Cutter 30. No. 2 Engine Inlet 31. External Electrical Power Receptacle 32. No. 2 Engine Inlet 33. Ice Detector 34. Ambient Sense Port 35. Engine Fairing/Work Platform (Same Both Sides) 36. Gravity Refueling Port (Same Both Sides)

37. Tail Pylon Fold Hinges 38. Tail Pylon Service Ladder (Same Both Sides) 39. Stabilator 40. HIRSS 41. Windshield Post Deflector 42. Windshield Wiper Deflector 43. Static Air Temperature Sensor 44. Avionics Compartment 45. OAT Sensor 46. Ice Detector 47. Pylon Cooling Air Intake

UH-60M General Arrangement UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

9/3/2003 Page 3 of 34

WIDTH WITH ESSS AND EXTERNAL EXTENDED RANGE TANKS INSTALLED 21 FEET FUSELAGE WIDTH WITH HOVER IR SUPPRESSORS INSTALLED 9 FEET 8 INCHES

FUSELAGE WIDTH 7 FEET 9 INCHES 20O

8 FEET 9 INCHES 5 FEET 1 INCH 3 FEET 9.5 INCHES TREAD 8 FEET 10.6 INCHES MAIN LANDING GEAR 9 FEET 8.6 INCHES STABILATOR WIDTH 14 FEET 4 INCHES

TAIL ROTOR DIAMETER 11 FEET

12 FEET 4 INCHES MAIN ROTOR DIAMETER 53 FEET 8 INCHES

2.8 INCHES

9 FEET 5 INCHES

WHEEL BASE 29 FEET 7 FEET 7 INCHES

6 FEET 6 INCHES

1 FOOT 7 INCHES LENGTH ROTORS AND PYLON FOLDED 41 FEET 4 INCHES

FUSELAGE LENGTH 50 FEET 7.5 INCHES OVERALL LENGTH 64 FEET 10 INCHES

UMHV003

9/3/2003 Page 4 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

TURNING RADIUS 41 FEET 7.7 INCHES

*TAIL ROTOR IS CANTED 20 o. UPPER TIP PATH PLANE IS 16 FEET 10 INCHES ABOVE GROUND LEVEL 7 FEET 7 INCHES ROTOR STATIONARY 16 FEET* 10 INCHES

12 FEET 4 INCHES 9 FEET 5 INCHES ROTOR TURNING

6 FEET 6 INCHES

12 FEET 1 INCH

11 FEET 4 INCHES

WHEELBASE 29 FEET

12 FEET 5 INCHES UMHV004

9/3/2003 Page 5 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Fuselage Sections The UH-60M fuselage sections are identical in nomenclature and purpose as the UH-60A/L. Significant changes have been made to the major structural components that constitute the fuselage to extend the life of the recapitalized UH-60A/L airframe 8000 hours by zero-timing the airframe. Additionally, the modification effort incorporates advances in materials and manufacturing processes to reduce weight and parts count as well as enhance structural integrity. From a maintenance perspective, the airframe changes facilitate a decrease in operations and support costs and increase mission availability by enhancing reliability and maintainability.

MAIN ROTOR PYLON

TAIL ROTOR PYLON

TAIL CONE

TRANSITION SECTION

CABIN

COCKPIT

UMHV009

9/3/2003 Page 6 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

In the nose avionics bay, the forward bulkhead has undergone modification to allow the mounting of the two Sealed Lead Acid batteries. The deck or floor has been structurally enhanced to support the precision EGI units mount plates above and the active vibration control components located below. The upper e-bay shelf has also been

Cockpit While maintaining the same physical appearance of the UH-60A/L cockpit section, the UH-60M cockpit incorporates numerous changes to enhance its functionality and survivability. The cockpit section has been completely refurbished and in that process much of the structure has been replaced or revised to accommodate equipment that is part of the UH-60M core configuration. 24

1

2 3

3 4

23 22

5

21 6 7

7 8

20 #1 ENG OUT

#2 ENG OUT

MASTER

#1 ENG OUT

RADIO CALL

CAUTION

ON

ON

VID

VID

BRT

BRT

VID

ON

VID

OFF

OFF

12

LOW ROTOR RPM

PRESS TO RESET

FIRE

ON

#2 ENG OUT

MASTER CAUTION

LOW ROTOR RPM

PRESS TO RESET

FIRE

BRT

BRT

4 22 20 00 0

18 0

OFF

16 0

10

OFF

3 2 1

10

7 13 6 6 5

1 2 3

12 0

9

28 .82

10 0

03

55

60

5

50

10

45

10

BK LT

15 ET C

40

VID

35 BK LT

VID

BRT

CON

ATT

FMS

ADC

HDG

DCU

REV

REV

REV

REV

REV

BRT

CON

NAV SRC

OBS

REF SEL

BRG 1

BRG 2

REF ADJ

11

LOC CAP

GS CAP

GA

DECL CAP

RALT 250

ALTP 2500

P-SYNC

P-SYNC

ALT 1500

IAS 120

HDG 240

P-SYNC

P-SYNC

VS 500

SEL

CPLD ****

30

20

BK LT

BRT

ON

HVR P-SYNC

VID

25

CTRL

OF F

ENGINE IGNITION

NAV SRC

OBS

REF SEL

BRG 1

BRG 2

REF ADJ

CON

LOC CAP

GS CAP

GA

DECL CAP

RALT 250

ALTP 2500

P-SYNC

P-SYNC

ALT 1500

IAS 120

HDG 240

P-SYNC

P-SYNC

VS 500

HVR P-SYNC

BK LT

VID

BRT

CON

CPLD **** ATT

FMS

ADC

HDG

DCU

REV

REV

REV

REV

REV

10 11

13 ST LI P K E C MA CH A & GE T A DATOW S

CH DA ECK T ST A & LIST OW MA AG P E

12 13

14

19

1. 2. 3. 4. 5. 6. 7. 8.

UPPER CONSOLE PILOT S COCKPIT UTILITY LIGHT FREEAIR TEMPERATURE GAGE NO. 2 ENGINE FUEL SELECTOR LEVER NO. 2 ENGINE POWER CONTROL LEVER NO. 2 ENGINE OFF / FIRE THANDLE WINDSHIELD WIPER INSTRUMENT PANEL GLARE SHIELD

18

17

9. 10. 11. 12. 13. 14. 15. 16.

16

INSTRUMENT PANEL VENT / DEFOGGER ASHTRAY PEDAL ADJUST LEVER MAP / DATA CASE ELT CABIN DOME LIGHTS DIMMER PARKING BRAKE LEVER

15

17. 18. 19. 20. 21. 22. 23. 24.

CHAFF RELEASE SWITCH LOWER CONSOLE LOWER CONSOLE UTILITY LIGHT STANDBY (MAGNETIC COMPASS) NO. 1 ENGINE POWER CONTROL LEVER NO. 1 ENGINE OFF / FIRE THANDLE NO. 1 ENGINE FUEL SELECTOR LEVER COPILOT S COCKPIT UTILITY LIGHT

UMHV005

9/3/2003 Page 7 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

modified to allow new equipment to be mounted. The nose bay door has been painted using conductive paint and been equipped with new seals for Electromagnetic Interference (EMI) hardening purposes. The nose section is equipped with a lower avionics bay. Equipment found in the space are components of the Active Vibration Control System, the Digital Intercommunications System as well as equipment for aircraft data bus systems. Sweeping changes have been made within the cockpit as well. The lower console enclosure, which was made of fiberglass in the UH-60A/L, is now an all aluminum assembly for EMI hardening purposes. Additionally, the kick panels located at the pilot and copilots feet that attach to the forward lower console, formerly made of Kevlar, have been replaced with aluminum panels for EMI hardening. The lower console floor has been changed to allow added avionics equipment support and electrical wiring. An avionics cooling fan has been located on the forward right side of the lower console enclosure to provide active cooling of installed lower console components.

include mounting brackets and other necessary hardware to allow the rotor brake assembly, hydraulic components and lines to be installed. All doors and exterior surfaces of the cockpit have been painted using conductive paint for EMI hardening purposes. Cockpit Doors The purpose and general configuration of the crew compartment or cockpit doors is identical to the UH-60A/L. The doors incorporate single piece window transparencies with a circular pop-out vent located in the lower forward corner of the window. Each door is equipped with a window emergency release handle. The emergency release is accomplished by using a latch handle to allow unlatching the door from either inside or outside the cockpit. Emergency release handles are on the inside frame of each door. They allow the cockpit doors to be jettisoned in case of an emergency. There is an emergency release pulltab on the inside forward portion of each cockpit door window for pilot egress. Data compartments are on each cockpit door.

The structure that mounts the instrument panel has been enhanced to allow the new narrow profile instrument panel to be installed. Modifications to sheet metal throughout the cockpit tub assembly, while not readily visible, enhance maintainability due to different and larger penetrations or feedthroughs for wiring and other equipment. The windows that surround the operators have also been impacted by changes as well. Windows in the doors incorporate the changes made to UH60L Lot 21 and subsequent Black Hawks. The windshields, which have been made of tempered glass, are now made of a polycarbonate material that greatly enhances its integrity and protection of pilots flying the aircraft. The left hand upper observation window has been changed due to the addition of an additional circuit breaker panel. Shelving found in both seatwells has been changed to permit the location of multiple new components. The seatwell covers, formerly made of fabric, have been replaced with metalized NOMEX covers to create an EMI shield for the avionic components located in the seatwells. In right side of the cockpit overhead, provisions for a rotor brake can be found. The provisions

9/3/2003 Page 8 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

A noticeable feature of the instrument panel assembly profile is that it has been reduced from the UH-60A/L’s 58-inch wide panel to only 52.4 inches. This reduction in instrument panel size is due to the clean arrangement of the MFDs and associated control panels. This narrow profile panel enhances safety as a result of the increased pilot visibility both looking forward and down.

Instrument Panel The instrument panel provides visual displays on aircraft system to both operators and maintenance personnel. The primary instruments located on a single piece instrument panel that is tilted back 30°. The primary flight instruments are comprised of four Multifunction Displays (MFDs), Flight Director/Display Control Panels (FD/DCP), and Reversionary Control Panels. The Electronic Standby Instrument System (ESIS) display and key ignition switch is located in the center section of the instrument panel. The master warning panels are mounted on the upper instrument panel below the glare shield to inform the pilots of conditions that require immediate attention.

Flight Display System (FDS) The Flight/Mission Display System (FDS) provides integrated control and display of essential flight and mission information. The FDS components in the cockpit include: 1. Four Multifunction Displays (MFDs). 2. Two Display Control Panels (DCPs). 3. One Reversionary Switch Panel per pilot. 4. One Electronic Standby Instrument System Display

NON SECURE RADIOS WILL NOT BE KEYED

RA D IO C A LL 24432

WHEN USING ANY SECURE RADIO OR THE INTERCOM FOR CLASSIFIED COMMUNICATIONS #1 ENG OUT

#2 ENG OUT

MASTER

#1 ENG OUT

CAUTION PRESS TO RESET

FIRE

#2 ENG OUT

MASTER CAUTION

LOW ROTOR RPM

ON

ON

VID

BRT

160

10

10

3 2 1

10

1 2 3

7 136 6 5 10

N

03

B

3

A RO

28.82in

100

M

55

60

5

50

10

45 BK LT

VID

BRT

CON

15 ETC

40 35

CON

FMS

ADC

HDG

DCU

REV

REV

REV

REV

REV

NAV SRC

OBS

REF SEL

BRG 1

BRG 2

REF ADJ

LOC CAP

GS CAP

GA

DECL CAP

RALT 250

ALTP 2500

ALT 1500

IAS 120

HDG 240

VS 500

SEL

CPLD ****

OFF

0

120

BRT

VID

BRT

40 22 20 00

180

VID

ON

OFF

OFF

BK LT

LOW ROTOR RPM

VID

BRT

VID

BRT OFF

ATT

PRESS TO RESET

FIRE

ON

30

20

CTRL

OFF

ON

HVR P-SYNC

P-SYNC

P-SYNC

P-SYNC

P-SYNC

BK LT

VID

BRT

CON

25

NAV SRC

OBS

REF SEL

BRG 1

BRG 2

REF ADJ

ENGINE IGNITION

LOC CAP

GS CAP

GA

DECL CAP

RALT 250

ALTP 2500

ALT 1500

IAS 120

HDG 240

VS 500

HVR P-SYNC

P-SYNC

P-SYNC

P-SYNC

BK LT

VID

BRT

CON

CPLD **** ATT

FMS

ADC

HDG

DCU

REV

REV

REV

REV

REV

P-SYNC

UMAV001

UH-60M Instrument Panel

9/3/2003 Page 9 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Upper Console All cockpit electrical controls are on the upper and lower consoles and instrument panel. The upper console, overhead between pilot and copilot, contains engine controls, fire emergency controls, heater and windshield wiper controls, internal and external light controls, electrical systems, and miscellaneous helicopter system controls. The rear portion of the upper panel contains the dc essential bus circuit breaker panels. Provisions for a rotor brake are located to the outboard side of the pilot’s side of the overhead console.

NO. 1 DC ESNTL BUS NO. 1 ENG

LIGHTS

AUX CB

IFF

CABIN ICS

5

5

FIRE DET

SEC

NO.1 STAB

SAS

ECP

CARGO HOOK

5

2

7

1 2

PWR

35

5

4

7

1 2

SENSE

JTSN OUTBD

NO. 1 VHF

HOIST CABLE

ESSS JTSN

5

5

5

EMER RLSE

FM

SHEAR

7

1 2

INBD

EMERG REL NORM

EXT LTS MODE

5

4

O F F

BATT

N O R M

2 1

LOWER

O F F

LTD SW

OFF

BRT

OFF

DIM

BRT

CABIN DOME LT

PLT ICS

2

2

2

MSTR WARN

PRI

PLT CDU

NO. 2 ADC

BACKUP HYD

5

2

5

PNL SPLY

WINDSHIELD WIPER PARK

OFF

LOW

PLT MWC

DIM

OFF

BRT

SENSE

PLT MFD

PLT FD/

7

1 2

DCP

EMER RLSE TEST LT

FUEL BOOST PUMP NO. 1

ENG NO. 1

BATT LOW

OIL HOT

ACC LOW

APU FAIL

APU ON

APU

7

INBD

1 2

3

PRI

HEATER MED O F F

O F F

ANTI-ICE NO. 2

ON

ON

OFF

PITOT LEFT O F F

O F F

O F F ON

APU FIRE

ICS ICU

NO. 2 DCU

25

VENT BLOWER

APU

BATT GOOD

HI

HEAT RIGHT O F F

ON

ON

FUEL BOOST PUMP O F F

NO. 2

WINDSHIELD COPILOT CTR

ANTI-ICE PILOT

FUEL IND

ON

BRT

O F F

O F F

O F F ON

LWR CSL

BRT

NO. 2 ESNTL

35

HI

LIGHTS UPR CSL

VOR/ ILS

5

O F F

FLASH

BRT

ON

SECONDARY

AUX CB

5

ON

ENG

ON

RESERVE

STEADY

LIGHTS

BRT

ON

VHF AM

2

CONTR

AIR SCE HI/STRT

O F F

MAIN

NIGHT

DIM

ON

APU TEST

NO. 2 EGI

ON

O F F BLUE

R O E F S F E T

BOOST PUMP

CONT

FIRE EXTGH

POSITION LIGHTS

DIM

CPLT MWC

ON

R O E F S F E T

NO. 2 TEST

APU O F F

O F F

O F F

ON

R O E F S F E T

NO. 1 TEST

PLT

ON

DAY

SECONDARY WHITE

ON

ON

IR

ANTI COLLISION LIGHTS

B O T H

O F F

O F F

O F F

O F F

O F F

OFF

UPPER

NO. 2 TEST

NO. 1 TEST

STBY INST TEST

ARMED

GENERATORS

RESET

3

ARM SAFE

ALL

SHORT

ON

EXT PWR

CARGO HOOK LT

CONTR CKPT

O P E N

ARM

FORMATION LTS

NO. 2 DC ESNTL BUS

CARGO HOOK

VIB CONT

ESSS

2

PNL SPLY

BOOST

NO. 1 ESNTL

ON

ON

TEST

INST PNL

OFF

BRT

FIRE DETR TEST OPER 1

WHITE

O F F

LAMPS 2

TEST

BLUE

ENG OVSP TEST A NO. 1 B A NO. 2 B

ENG SPD TRIM DECR O F F INCR

UMAV002_1

UH-60M Overhead Console 9/3/2003 Page 10 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Lower Console The lower console extends rearward through the cockpit between the pilot and copilot, is easily reached by either pilot. The console is arranged with communication panels, navigational panels, mission equipment panels, and flight attitude/stability controls. The console also houses the battery bus and battery utility bus circuit breaker panel.

9/3/2003 Page 11 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Pilots Seats The pilot and co-pilot seats in the UH-60M replace those found in the UH-60A/L. The new seats are common the U.S. Navy MH-60 model aircraft. The seats are designed to accommodate 5th percentile female through 95th percentile male pilots as well as incorporate enhanced safety and comfort features. The UH-60M crew station seats function and general configuration remains the same. The seats are intended to provide a man-machine interface as well as ballistic protection.

9/3/2003 Page 12 of 34

UH-60M Pilots Seats UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Variable Load Energy Absorber Adjustment

Seat Adjustment Controls

An additional control has been added that allows the adjustment of the crash load attenuation capability of the seats. A Variable Load Energy Absorber (VLEA) is located on the back, center of side of the seat back. The adjustable VLEA is equipped with a control lever, which is located on the right side of the seat. The adjustment lever permits the seat load to be varied between 140 and 250 pounds in 20-pound increments. The VLEA adjustment lever has seven detented positions that can be selected using the liner lever. The VLEA minimum setting requirement is 140 lbs. In the event of the seat stroking due to hard landing or crash loads, red bands appear on the outside of the VLEA housing.

Vertical and forward/rearward seat adjustment controls of the UH-60M seat are common with the UH-60A/L. Levers located on the bottom left and right sides of the seat buckets are used to adjust the seat height and forward or rearward positions. The levers are spring loaded and return to the locked position when released.

Emergency Tilt Release Levers Common with the UH-60A/L, the emergency tilt release levers are on each side of the seat support frame. The seat may be tilted back into the cabin for removal or treatment of a wounded pilot. Seat tilting can be done from the cabin, only with the seat in the full down and aft position by pushing in on the tilt handles, and pulling the seat top rearward.

Seat Adjustment Controls Emergency Vertical Release Lever

Variable Load Energy Absorber

The emergency vertical release lever permits the seat to drop to the lowest adjustment point for tilting. The release lever is on the upper center back of the seat and is actuated by pulling right on the lever. Seatbelts Operator’s seats shoulder harnesses, seatbelts, crotch straps, and buckle assemblies are all common in purpose, configuration, and utilization. Use of the belts and straps adjustment fittings, and single-point release mechanism are identical with the UH-60A/L.

VLEA Adjustment Lever

9/3/2003 Page 13 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Protective Armor The UH-60M seat provides improved pilot armor protection due to manufacturing changes. Just as with the UH-60A/L, armor protection is provided for the body of the seat occupant against 7.62 mm rounds from the side and from the back and below. The UH-60M seat armor bucket uses a contoured monolithic tile design that eliminates the 18 individual tiles found on the UH-60A/L seat armor system. The new design also provides increased coverage as a result of eliminating the large cutouts found on the legacy seat armor package.

UH-60M Monolithic Armor Bucket

9/3/2003 Page 14 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

UH-60A Tiled Armour Bucket

Cabin Section The cabin section of the UH-60M is functionally identical to that of the UH-60A/L. The new production cabin section permits service life to be extended to 8000 hours. Significant analysis has been conducted of the wear and usage of fleet airframe components in effort to target required improvements and design criteria to achieve them. Cabin Structure The cabin structure has been manufactured using multiple new technologies that allow for weight reduction, increased corrosion resistance, as well as a configuration that permits a great degree of flexibility for the incorporation of mission equipment kits.

Improved Firewalls Integral ESSS/Frames Improved Step Design Enhanced Water Integrity

HSM Cabin Upper CH-60 Upper Structure

HSM Cabin Tub Structure (80% HSM) SOA Penetrations “Flexibility for Future Growth” Built-in Provisions for Electrical Clips

Stub Wing Beef-up

Side Framing Structure (66% HSM)

9/3/2003 Page 15 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

High Speed Machining Many of the primary structural elements used in the construction of the cabin section use a process known as High-Speed Machining (HSM). HSM is a process through which components are made from or machined as a single component. The principle advantages afforded by the use of HSM in aircraft manufacturing are the enhancement of structural rigidity or strength, reduction of total parts counts, a decrease in the time required for manufacturing and ultimately cost.

An example of the use of HSM components can be found in the Station 295 and Station 308 frames. The External Stores Support System (ESSS) support fittings previously were piece built or made up of ribs, caps, doublers, and clips. The ESSS fittings are now made of two single piece components that span the width of the cabin section. This construction method allows greater airframe strength and decreased risk of cracking.

HSM Station 295/308 Frame Rough Cut

HSM Station 295/308 Frame Finished Product

9/3/2003 Page 16 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Examples of High Speed Machined Monolithic Parts

“Before” and “After” HSM Tail Components

9/3/2003 Page 17 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Troop/Cargo (Cabin) Doors

Door Locks

The aft sliding doors located on each side of the troop/cargo compartment are functionally identical to those found on the UH-60A/L. Operation of the single-action door latches that allow the doors to be latched in the fully open or fully closed positions remains unchanged.

Key door locks are installed on each of the cabin, cockpit, avionics compartment, and transition access doors. A common key is used to lock and unlock the doors from the outside to secure the helicopter. Each crew chief/gunner sliding window is locked from the inside only.

Crew Chief/Gunner Windows

Cabin Interior - Soundproofing

The crew chief/gunner windows on the UH-60M are identical in function and use to the UH-60A/L. Both crew chief windows are forward sliding hatch windows that are split vertically into two panels.

Cabin interior panels have been revised to allow for a better fit as well as to accommodate added EMI shielding requirements. Forward cabin flight control panels or “broom closet covers” as well as the cabin ceiling panels incorporate metalized fabric, conductive seals, and conductive paint on the interior surfaces.

Broom Closet Cover • Conductive Paint on Inside Surfaces

Ceiling Panels •Conductive Material on Upper Surface •Conductive Seals

Soundproofing Panel Improvements

9/3/2003 Page 18 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Cabin Overhead Equipment Arrangement Many of the aircraft systems components located in the cabin overhead have been removed and replaced by new equipment or relocated to new positions.

AVC Mechanical Units

Dimmer

Blade De-Ice Unit (relocated)

Troop Commander ICS Generator Control Unit (relocated)

Potentiometer (relocated)

Dimmer ESIS Air Data Computer

Air Data Computers

Forward Cabin Overhead - Lkg Fwd, Up CVR/FDR

Blade DeIce Eqpt. (relocated)

Generator Control Unit

Aux Fuel Signal Conditioner

Aux Fuel Relay Panel

Data Concentrator Units Aft Cabin Overhead - Lkg Aft, Up

9/3/2003 Page 19 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Crew Chief/Gunner Seats

Troop Seats

The two seats located in the forward cabin area that are outward facing seats, are for the crew chief/gunner. Each seat faces a window. Each seat steel tube supported assembly with a fireresistant, high strength fabric seat and backrest. Each seat contains two lower energy attenuators designed and oriented to reduce personal injury in a crash. Each seat has a complete lap belt and dual torso-restraint shoulder harness attached to a dual action rotary release buckle. The shoulder harness is connected to inertia reels on the seat back and bottom. This gives the wearer freedom to move about his station. The restraint system is equipped with a single action rotary release buckle with a guard. A release plate must be pressed to allow rotation of release, preventing inadvertent handle rotation from contact with equipment, etc. The inertia reel lock control is replaced by a shorter push/pull manual locking control. Push in and the inertia reel is manually locked in place. When the control is pulled out, the reel will lock on sudden pull.

The UH-60M troop seats have been changed to enhance safety as well as increase usability by eliminating cabling and associated mounting hardware. 13 seats can be installed in the cabin.

Top Cross Beam

Upper Vertical Guide Tubes

Each troop seat has a belt and shoulder harness for body restraint. The backs and seat pans are attached by means of tubular supports to the cabin ceiling and floor. Each seat contains two lower energy attenuators designed and oriented to reduce personnel injury. Just as with the UH-60A/L, the troop seats can be removed from the cabin and stowed in the stowage compartment or hung along the back cabin wall.

Top Fitting Attenuation Rollers Attenuation Tubes

Secondary Tubes Cross Braces Quick Release Pins

Fabric Cover Bottom Fittings (Quick release)

Lower Vertical Guide Tubes UH-60M Troop Seat 9/3/2003 Page 20 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Head Rest 4 Point Harness Adjusters

Rotary Quick Release

Hinged Rigid Seat Pan

MA16 Inertial Reels

Stand Up Harness

MA16 Inertial Reel

Rotary Quick Release

UH-60M Crew Chief/Gunners Seat

UH-60M Cabin Seating Arrangement 9/3/2003 Page 21 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Transition Section The transition section interconnects the cabin section and the tail cone. The transition section contains the main fuel cells, storage areas or provisions for mission equipment kit installations and avionic and electrical equipment. It’s physical arrangement and function remains unchanged from the UH-60A/L. Use of transition section space has been greatly increased using available overhead space in the forward transition section as well as additional space made available by the installation of racks and shelving. As with the Cabin section structural components, the transition section is new production as well. The new transition increases service life by 8000 flight hours.

Many transition section changes are not immediately noticeable, while several are. The majority of the enhancements involve building additional structural strength into transition structure as well as drive shafting brackets, IHIRRS and work platform interface with the transitions section.

Replace Drive Shaft Brackets Added Stringer BL – 0 Firewall Reinforcement Added Angles, Doublers & Straps

Replace Tailcone Attachment Replace Vapor Barrier Attachment Angle

Fuel Cell Structure/Foam Replacement

Transition Structural Improvements 9/3/2003 Page 22 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Transition Avionics Bay Access Door Located on the right side of the transition section, the transition avionics bay access door allows greatly improved access to the avionic equipment located in the tail cone, transition section, and the rear side of the fuel cells. The door is secured by two latches and is held in the open position by a

lockable support strut. The door is made of composite material with an integral layer aluminum mesh for EMI hardening. Additionally, the door is equipped with EMI gasketing to provide proper electromagnetic sealing and protection of avionic equipment housed in the transition section.

Transition Access Door 9/3/2003 Page 23 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Transition Avionics Shelves Shelving located in the aft transition section allows the installation of multiple avionic and electrical processors and control boxes. The shelving is conductive and care should be taken when removing and installing so as to ensure proper electrical boding of equipment mounts, shelving and surrounding airframe structure.

IFF Transponder

New IDM Unit

ALQ-144 Converter space and weight

AFCS Rate Gyro (qty 2)

Stormscope Data Bus Couplers

KY-100

AFCS Computer

ARC201D Low Pass Filter

ARC201 Amp.

APR-39 Processor

APR-39 Blanking Unit

HF Amp

RT-1749/URC HF Radio AVR-2 Comparator

Transition Equipment Installation

9/3/2003 Page 24 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Tail Cone The tail cone contains the tail rotor control cables, electrical and avionic components. The tail cone also provides the structural interface for the tail drive shafting and the tail landing gear assembly. The UH-60M tail cone has been recapitalized to allow service life extension. The tail cone has had numerous structural enhancements to provide greater reliability and decrease maintainability requirements.

Tail drive shaft support brackets have been constructed with reinforced doublers. Stringers throughout the tail cone structure have been reinforced by either as a result of a thicker build or through the use of reinforced doublers. Highspeed machined frames are also used in the tail cone at Stations 485 and 505.

Thicker Driveshaft Brackets With Reinforced Doublers

Thicker Stringers With Reinforced Doublers Thicker Stringers With Reinforced Doublers Lower Shear Deck

Tail Cone - Top

Thicker Lower Stringers & Plating

Added Shear Ties On Station Frames 505, 525, 545, AND 565

Tail Cone - Bottom

Recapitalized Tail Cone 9/3/2003 Page 25 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Tail Rotor Pylon The tail rotor pylon provides mounting for the intermediate gearbox (IGB) and tail gearbox (TGB), associated drive shafting as well as antennae, a cambered fin, and the stabilator. The recapitalized tail rotor pylon incorporates numerous structural upgrades for service life extension. Doublers have been added to the tail rotor pylon in several locations to reduce the likelihood of structural fatigue and cracking. The tail rotor gearbox mount fitting has also been replaced.

Whip Antenna Mod New Tail Rotor Gear Box Fitting

Replace/Add Crescent Doubler

New Reinforced Camlocks (9 PLCS) And Hinges Common To #1 VHF/FM Antenna Fairing

Replace With Thicker Doubler Mod Bushing Kit For Folding Stabilator

Reinforced Doublers & Stringers

Recapitalized Tail Rotor Pylon

9/3/2003 Page 26 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Folding Stabilator The most noticeable modification to the tail rotor pylon and its associated components is the stabilator. The stabilator is comprised of three sections: two outboard foldable panels and a center section or center box. The outboard panels are configured with an equal planform area and can be folded upward on hinges mounted to the center box.

The stabilator center box provides the mechanical interface with the tail rotor pylon. The center box is also equipped with locking mechanical locks and position switches The switches are wired in series to provide an indication that the stabilator panels are in the down and locked position.

MH-60K Folding Stabilator

Existing UH-60 Stabilator

Stabilator Actuator Attachment Clevis

Limit Switch and Position Transmitter Attachment Clevis

Stabilator Elastomeric Bearings

Stabilator/Airframe Attachment Points

Stabilator Center Box Assembly 9/3/2003 Page 27 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Lockpin Assembly

Lockpin “Keeper”

Stabilator Panel Locking Hardware

Lockpin Switch

Lockpin “Keeper”

9/3/2003 Page 28 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Main Rotor Pylon The main rotor pylon is an aerodynamically shaped structure that houses all mechanical flight controls, main rotor transmission, engines, auxiliary power unit (APU) and other utility equipment systems. The main rotor pylon general configuration and purpose is identical to the UH60A/L. All access panels, doors, and fairings are identical to the UH-60A/L. All composite fairings have been refurbished. The existing paint was removed and a new anti-static paint has been applied. Firewalls located in the aft section of the main rotor pylon feature improved lower parts count assemblies common with later lot UH-60L aircraft.

UMHV009_2

Main Rotor Pylon

9/3/2003 Page 29 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Landing Gear System – Weight-onWheels (WOW) Switch Functions

Landing Gear System – Parking Brake Handle Controls

The landing gear system incorporates changes made to the landing gear shock strut and drag beam assembly. These changes allow the UH60M to match the landing gear system configuration found in Lot 21 UH-60L aircraft.

The function, physical configuration, operation, and caution/advisory indications of the wheel brake system and parking brake remain unchanged from the UH-60A/L. The parking brake handle has been relocated to a control panel mounted in the center mid-section of the lower console to permit easy access for both pilots.

The Lot 21 modifications upgrade or refurbish the shock strut assembly, replace drag beams, and replace seven lug wheels found on early UH-60A aircraft. Additionally, the Lot 21 configuration increases upper shock strut nitrogen pressure to 1000 psi. The function of and physical location of the WOW switches are identical to the UH-60A/L.

Tail Landing Gear – Tail Wheel Lock The function, physical configuration, and operation of the tail landing gear system remain unchanged from the UH-60A/L. The TAIL WHEEL locked switch and indicator has been relocated to the center of the lower console to permit easy access for both pilots. The indicator provides a LOCK indication when the tailwheel lockpin has been extended into the tail wheel fork. The UNLKD indication is presented when the lockpin has been retracted from the fork.

TAIL WHEEL

UNLKD LOCK

PARKING BRAKE

UMAV003_16

Parking Brake/Tail Wheel Lock Control Panel

Main Landing Gear Assembly

9/3/2003 Page 30 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Airframe Electromagnetic Environmental Effects (E3) Hardening Electromagnetic Effects General Information Many weapons systems have experienced increasing problems involving performance degradation or damage of electrical, electronic, or avionic equipment, systems and subsystems due to inadequate consideration of the intended operational Electromagnetic Environment (EME). In some cases, electromagnetic effects may be temporary and normal system operation returns when the emissions are reduced or removed. In other cases, the effect on equipment may be permanent requiring the replacement of aircraft systems or components.

Electromagnetic Environmental Effects or E3 is the impact of the EME upon the operational capability of military forces, equipment, systems, and platforms. It encompasses all electromagnetic disciplines, including Electromagnetic Compatibility (EMC)/EMI; Electromagnetic Vulnerability (EMV); Electromagnetic Pulse (EMP); Electronic Protection (EP); Hazards of Electromagnetic Radiation to Personnel (HERP), Ordnance (HERO), and volatile materials such as fuel (HERF); and natural phenomena effects of lightning and precipitation static (p-static).

Electromagnetic Environment

9/3/2003 Page 31 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

The EME in which military platforms/systems such as the UH-60M and its subsystems/equipment must operate is comprised of a multitude of natural and manmade sources. Natural sources consist of: • • • • •

Atmospheric noise Solar noise P-static Lightning Electro Static Discharge (ESD)

Manmade sources can be either friendly or hostile emitters, both intentional and unintentional. Examples of intentional emitters include, but are not limited to the following types of subsystems/equipment: • • • •

Communications Navigation Meteorology Radar, Weapon, and Electronic Warfare (EW)

Unintentional emitters can be any item that uses, transforms, or generates any form of electromagnetic energy. Therefore, any electrical, electronic, electromechanical, or electro-optic device can be an unintentional emitter. Examples of unintentional emitters include the following: • • • • • • • •

Intentional radiators emitting other than the intended emission Computers and associated peripherals Televisions, cameras, and video equipment Microwave ovens Radio and radar receivers Power supplies and frequency converters Motors and generators Electrical hand tools.

• • • • •

Burnout or voltage breakdown of components, antennas Performance degradation of receiver signal processing circuits Erroneous or inadvertent operation of electromechanical equipment, electronic circuits, components, ordnance Unintentional detonation or ignition of ordnance and flammable materials Personnel injuries

UH-60M Electromagnetic Effects Protective Design Features The UH-60M design incorporates multiple measures to provide E3 protection for system, subsystems, and equipment by creating shielded volumes throughout the airframe as well as using grounding and bonding to meet the requirements of ADS-37A. Shielded volumes have been created using several methods. In some locations, panels previously made of composite materials have been replaced with aluminum panels. In other locations, aluminum mesh has been bonded to the composite material and used in conjunction with numerous conductive materials. The following is a list of equipment or components that have been modified or changed in effort to enhance the UH-60M electromagnetic effects protective capabilities: • • • •

Nose Door- (Paint Door and Install Gaskets) Cockpit Doors- (Paint Doors and Replace Grounding Pads) Center Well Panels- (Composite with Aluminum Mesh) Center Console Skins- (Aluminum Sheet Metal)

Undesired electromagnetic energy may degrade the performance of an item temporarily, in which case the item may operate in a degraded mode when sufficient electromagnetic energy is present. Alternatively, the electromagnetic energy may cause permanent damage, in which case the item will not operate until it is either repaired or replaced and the E3 problem has been resolved. Examples of the effects that can be caused by undesired electromagnetic energy, depending on the victim, are:

9/3/2003 Page 32 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

• • • • • •

Center Console Kick Panels- (Aluminum Sheet Metal) Glare Shield- (Composite with Aluminum Mesh) Seatwell Covers- (Nickel Nomex Fabric (metalized)) Doppler Antenna Cover Plate- (Aluminum Sheet Metal) Tub Buttline-0 Vent- (Aluminum Mesh or Honeycomb Panel)

• • • • • •

Floor Panels- (Paint Underside of Panels) Interior Panels- (Nickel Nomex Fabric (metalized)) Station 398 Barrier- (Nickel Nomex Fabric (metalized)) Transition Avionics Bay Door- (Composite with Aluminum Mesh and Gaskets) Transition Fuel Vent and Drain Line Enclosure- (Aluminum Sheet Metal) Aircraft Top Coat Paint Trade Study Decision (Change to Anti-Static Top Coat Paint)

UH-60M Shielded Volumes

9/3/2003 Page 33 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

NOTES

9/3/2003 Page 34 of 34 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Section 2-2 T700-GE-701D Engine and Related System Differences Description The T700 engine used on the UH-60M incorporates new components to modify the engine to the T700-GE-701D configuration. The upgrade of the 701C engine permits an increase both in the power of the engine and its maintainability by allowing for an increase in time between engine removals. The higher-powered 701D engine increases the flight performance of the UH-60M. The result is increased internal and external load carrying capacity.

The T700-GE-701D engine external configuration and general arrangement is identical to the 701C. The 701D engine is a front drive turboshaft engine with modular construction. The engine locations on the aircraft as well as engine major modular components are identical to the UH-60A/L and T700-GE-701C. The 701D engine modification was developed as part of the Army’s Component Improvement Program (CIP). The CIP focuses on extending and improving the lives of parts used in Army

Engine Section Impacted by 701D Improvements

9/3/2003 Page 1 of 24 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

engines. As part of the CIP, the current 701C engine has been upgraded with parts that are composed of higher temperature capability materials. These improved materials are already being used in other General Electric engines such as the CT7-5, CT7-8, and CT7-9 engines and provide longer lives or increased power capability for the 701D engine. The 701D powerplant system modifications impact the following engine modules and/or system functionality: • • • • •

The modifications specified for the 701D engine are used in other engines produced by GE, such as the CT7-9B. Nine out of the 11 new parts in the 701D engine are already in production and only require assembly into a current 701C engine. The two remaining parts are made from existing common forgings. The changed parts use a combination of new materials and/or coatings or coating processes that have not previously been used on military versions of the T-700 engine.

Hot Section Module Power Turbine Section Digital Engine Control Software Engine Indication Bias Cockpit Controls and Displays

CT7-9 Combustor w/TBC

New Stage 1 Solid Shroud

CT7-9 Stage 2 Nozzle

CT7-8 Stage 2 Blade w/ 4 TE Slots

CT7-5A Stage 3 Blade

CT7-9 Stage 1 Nozzle CT7-8 Stage 1 Blade w/TBC

CT7-9 Interstage seal

CT7-9 S2FCP

CT7-9 Short rotor bolts

Redesigned S2ACP (New)

Specific Engine Components affected by 701D Improvements

9/3/2003 Page 2 of 24 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

T-700-GE-701D Engine Characteristics Type of Engine: Turboshaft Engine Weight (Dry): 456.3 Lbs. Engine Length: 46.12 inches Engine Diameter: 15.55 inches Output Power: • Max Continuous:1716 shp @ 20,900 rpm • Intermediate: 1902 shp @ 20,900 rpm • Maximum (10 Minute): 1994 shp @ 20,900 rpm • Contingency (2.5 minute): 2000 shp @ 20,900 rpm • RPM – 100% Ng: 44,700 rpm • RPM – 100 Np: 20,900 rpm Type of Compressor: Five stage axial, singlestage centrifugal flow Number of Compressor Stages: 6

Variable Geometry: Inlet Guide vanes, Stage 1 and Stage 2 Type of Combustion Chamber: Annular through flow Turbine Stages: Two stage air cooled GG Turbine, 2 stage uncooled PT Direction of Rotation: Clockwise – viewed from aft end of engine Power Turbine Inlet Temperature (T4.5): 852°C @ MRP. T4.5H: (Max Measured) 852° C. Max allowable: 879° C Fuel Type: MIL-PRF-5624T, JP-4, JP-5 and JP-8 +100 Lubrication Oil: MIL-PRF-7807, MIL-PRF-23699 Oil Temperature: 132° C (270° F) – 149° C (300° F) Oil Pressure: 85 psid @ intermediate power (Max) – 45 psid (Min)

T-700-GE-701D Turboshaft Engine

9/3/2003 Page 3 of 24 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

T-700-GE-701D Engine Components/Subsystems The following internal engine components have been modified or changed as a result of the 701D engine upgrade.

Improved Combustor with Thermal Barrier Coating (TBC)

Stage 1 Blade with Thermal Barrier Coating (TBC)

An improved combustor with a low Peak Temperature Factor (PTF) liner has been incorporated to improve hot section durability. This combustor is currently being used on the CT7-9B engine.

The Stage 1 blade from the CT7-8 engine uses a Thermal Barrier Coating (TBC) to improve durability in high temperature environments. An extended plenum (recessed pocket) has been added to the tips of the blades. This prevents blade coolant exit holes at the bottom of the plenum from being covered by smeared metal in the event of a rub with the shroud. 

 Stage 1 Nozzle Inner band cooling has been added to the Stage 1 nozzle to improve the durability of the part. This is similar to the CT7-9B engine. 

9/3/2003 Page 4 of 24 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Stage 2 Blade with 4 TE Slots

Redesigned Second Stage Aft Cooling Plate (S2ACP) The Stage 2 aft cooling plate has been redesigned. Material has been added to the bore to reduce stress. Cooling holes have been added to cool the Stage 3 disk.

The Stage 2 blade in the CT7-8 engine is manufactured out of a new material and uses an improved cooling scheme for greater durability.



Second Stage Forward Cooling Plate (S2FCP)

Stage 2 Nozzle

The Stage 2 forward cooling plate is made of an improved material and it is a lower stress design. This cooling plate is currently being used in the CT7-9B engine.

A different cooling scheme and upgraded material have been applied to the Stage 2 nozzle to improve durability. This nozzle is currently being used in the CT7-9B engine.

9/3/2003 Page 5 of 24 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Short Rotor Bolts New short rotor bolts from the CT7-9B gas generator rotor are used. The shorter bolts contribute to enhanced performance by allowing greater LCF life capability.

Power Turbine Section Module

Interstage seal - CODEP Coated The interstage seal from the CT7-9B engine consists of a new material, which provides greater corrosion and oxidation resistance and has been CODEP (Co-Deposition) coated. The new material and coating help to improve the component durability.

Stage 3 Blade The Stage 3 blade material has been changed to new material in order to improve component durability. This blade is used on the CT7-5A engine.

Hydromechanical Unit (HMU) The accel and VG 3D cam in the engine 3D Hydromechanical Unit have been changed to improve the engines accel schedule in high altitude, hot day conditions. New Stage 1 Solid Shroud The 1st Stage Solid Shroud is made of a new material that replaces earlier ceramic coatings. The new solid shroud replaces the current ceramic solid shroud to improve component durability in a sand environment.

9/3/2003 Page 6 of 24 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Improved Hover Infrared Suppressor System (IHIRSS) The UH-60M is equipped with the Improved Hover Infrared Suppressor System (IHIRSS). It incorporates enhanced missile threat survivability features over those provided by the IHIRSS system found on the UH-60A/L aircraft. The IHIRSS is designed to channel exhaust gasses through the three-stage core and inner baffle to induce the flow of cooling air from the engine bay and the inlet scoops. The IHIRSS system, as with its predecessor, has no moving parts. It contains a three-stage removable core that reduces metal surface and exhaust gas temperature radiation and prevents line-of-sight viewing of hot engine surfaces. The three-stage core and inner baffle cold surfaces are coated with low-reflectance material. To maximize exhaust cooling, engine exhaust is ducted outboard and downward by the engine, away from the helicopter, by the exhaust deflector, where additional cooling air is provided by the main rotor downwash. The following IHIRSS components have been impacted by the changes made to the system.

Deswirl Assembly The total quantity of deswirl vanes have decreased from 20 to 14. The physical installation and mounting of the deswirl assembly remains unchanged. Core Assembly Coatings on the ducting and the three stages of the core assembly have been changed to enhance their heat resistive abilities. Additionally, core assembly vanes have been removed to improve core airflow. The installation and mounting of the IHIRRS core is identical to the UH-60A/L HIRSS components. Baffle Assembly Baffle assembly coatings have been changed for improved thermal performance. Baffle gussets have also been modified. The baffle assembly mounting and installation is unchanged from the UH-60A/L HIRSS configuration. For logistics and support purposes, all IHIRSS parts will be direct replacement parts for current HIRSS parts of same name.

Core Assembly (Duct - Stage 1,2 & 3)

Baffle Assembly

Deswirl Assembly

•

Quantity of vanes reduced from 20 to 14

• •

Coating changed Vanes removed

• •

Coating added Gussets modified

9/3/2003 Page 7 of 24 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

T-700-GE-701D Engine Control System The UH-60M engine control system is identical in purpose and general configuration to that of the UH-60A/L. The control system consists of an engine control quadrant, Digital Engine Control Unit (DECU), speed control, and load demand system.

Engine Control Quadrant With the exception of a change in the Lucite inserts found on the No. 1 and No. 2 ENG EMER OFF T-handles, the control quadrant is identical with the UH-60A/L. The T-handle Lucite inserts have been modified to satisfy Night Vision Goggle (NVG) lighting requirements.

UMFC003

9/3/2003 Page 8 of 24 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Digital Engine Control Unit (DECU) The DECU controls the electrical functions of the engine and transmits operational information to the cockpit. The DECU has been modified on the UH-60M helicopter due to Turbine Gas Temperature (TGT) bias and limiter change. The TGT bias is now –112°C on the T700-GE-701D engine compared to –71°C on the T700-GE-701C engine. The bias has been changed due to the increase in temperature in the hot section of the engine and to ensure similarity between engine display limits in the event that a T700-GE-701C and a T700-GE-701D engine are installed in the same aircraft.

No. 1 and No. 2 Engine Overspeed Protection System Test Switches The No. 1 and No. 2 ENG OVSP TEST A and TEST B switches have been moved. Previously located on the center panel of the overhead console, the switches have been relocated to the bottom portion of the left panel of the overhead console.

Engine Speed Trim Control Switch The engine speed control switch has been relocated from the collective stick grips to the bottom portion of the left panel of the overhead console. Use of the engine speed control switch is identical with the UH-60A/L. FORMATION LTS

NO. 1 DC ESNTL BUS NO. 1 ENG

LIGHTS

IFF

5

5

5

FIRE DET

SEC

NO.1 STAB

7

1 2

PWR

AUX CB

CABIN ICS

35

4

PNL SPLY

SAS

ECP

5

2

CARGO HOOK

BOOST

NO. 1 VHF

NO. 1 ESNTL

ESSS

2

7

1 2

SENSE

JTSN OUTBD

HOIST CABLE

ESSS JTSN

7

5

5

5

EMER RLSE

FM

SHEAR

1 2

INBD

EMERG REL NORM

EXT LTS MODE

5

4

CARGO HOOK LT

N O R M

2

BATT

O F F

1

DIM

OFF

BRT

CABIN DOME LT

OFF

BRT

PLT CDU

NO. 2 ADC

5

2

PLT MWC

DIM

OFF

BRT

AUX CB

VOR/ ILS

35

2

PNL SPLY BACKUP HYD

5

WINDSHIELD WIPER

ENG

PARK

OFF

LOW

O F F

PLT MFD

1 2

FUEL BOOST PUMP

BRT

NO. 1

BATT GOOD

ENG NO. 1

BATT LOW

OIL HOT

ACC LOW

APU FAIL

APU ON

O F F

APU

7

INBD

ICS ICU

1 2

1 HEATER

ON

ON

OFF

PITOT LEFT O F F

O F F

UPPER B O T H

ON

NO. 2

WINDSHIELD COPILOT CTR

ANTI-ICE PILOT

O F F

O F F

FUEL IND

ON O F F ON

ON

1

ON

SECONDARY WHITE

BRT

DIM

STEADY

BRT

FLASH

O F F

O F F LOWER

TEST

POSITION LIGHTS

DAY

NIGHT

INST PNL

OFF

IR

ANTI COLLISION LIGHTS

HI

HEAT RIGHT O F F

ON

CARGO HOOK LT O F F

OFF

MED

FIRE DETR TEST OPER

WHITE

O F F

N O R M

2

3

PRI

FUEL BOOST PUMP O F F

EXT LTS MODE

5

3

O F F

O F F

ANTI-ICE NO. 2

ON

APU FIRE

2

NO. 2 DCU

25

DCP

VENT BLOWER

HI APU

TEST LT

PLT ICS

2

ON

AIR SCE HI/STRT

EMER RLSE

NO. 2 ESNTL

SENSE

PLT FD/

7

CONTR

ON

LWR CSL

BRT

APU TEST

FLASH

ON DIM

UPR CSL

5

ON

LIGHTS SECONDARY

ON

ON

MAIN BRT

LTD SW

BRT

ON

RESERVE O F F

LIGHTS

DIM

R O E F S F E T

O F F

FIRE EXTGH

STEADY

O F F BLUE

ARMED

NO. 2 TEST

BOOST PUMP

O F F

O F F NIGHT

CPLT MWC

ON

R O E F S F E T

APU CONT

RESET

POSITION LIGHTS

DAY

O F F LOWER

SECONDARY WHITE

ON

NO. 1 TEST

VHF AM

5 PRI

ON

ANTI COLLISION LIGHTS

UPPER

ON

R O E F S F E T

EXT PWR

ON

IR

OFF

B O T H

NO. 2 TEST O F F

NO. 2 EGI

2 MSTR WARN

PLT

GENERATORS

NO. 1 TEST O F F

O F F

O F F

ARM SAFE

ALL

SHORT

ON

STBY INST TEST

ARM

FORMATION LTS 3

CONTR CKPT

O P E N

O F F

4

NO. 2 DC ESNTL BUS

CARGO HOOK

VIB CONT

LIGHTS CPLT MWC

LTD SW

PLT MWC

LAMPS 2

TEST

BLUE

ENG OVSP TEST A NO. 1 B A NO. 2 B

BLUE

ENG SPD TRIM

DIM

DECR

BRT

DIM

BRT

DIM

BRT

O F F

LIGHTS

INCR

SECONDARY

UPR CSL

LWR CSL

INST PNL

UMAV002_1

Overhead Console OFF

BRT

CABIN DOME LT

OFF

BRT

OFF

BRT

OFF

BRT

FIRE DETR TEST OPER 1

WHITE

O F F

LAMPS 2

TEST

BLUE

ENG OVSP TEST A NO. 1 B A NO. 2 B

ENG SPD TRIM DECR O F F INCR UMAV002_6

O/H Console – Left Panel 9/3/2003 Page 9 of 24 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

two tape groups to the right of the center tape group contain engine 2 data. The mid-center tape group contains transmission data. The lower middle tape group contains fuel data. The engine related instrument displays are labeled:

T-700-GE-701D Indications Flight Display System Engine Indications The flight display system engine operating performance indications are presented in two fashions: Using the Engine Instrument Crew Alerting System (EICAS) or as an element of the Primary Flight Display (PFD). Engine Instrument Crew Alerting System (EICAS) Engine Data Display

• • • • • •

Engine oil temperature – T Engine oil pressure – P Gas generator speed – NG Turbine gas temperature – TGT Engine torque – Q Power turbine speed – NP 1 and NP 2

The EICAS displays engine data from the selected data concentrator units (DCUs). The top center tape group contains rotor speed and engine power turbine speeds. The two tape groups to the left of the center tape group contain engine 1 data. The

No. 1 Engine Tape Groups

Engine Torque Indications

No. 2 Engine Tape Groups

9/3/2003 Page 10 of 24 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Engine / Fuel Drive Train A429 Fault Miscompare

Engine / Drive Train / Fuel Instrument CPCI Engine / Drive Train / Fuel

A429 Fault

Power Pod & EICAS Display

Miscompare Condition

ARINC 429 Receiver

MFD Miscompare CPCI

FMS-1 Reversion Switch Panel

ARINC 429

DCU-1

DEC Fault Codes Chip Detector IBIT DCU Status AVCS Status Fuel Quantity Discrete

DCU-2

Analog Engine / Drive Train / Fuel

Chip Power

ARINC 429 DEC Fault Codes Chip Detector IBIT DCU Status AVCS Status Fuel Quantity

FMS-2

Discrete Lamp Test

Aircraft Engine / Drive Train / Fuel

Engine Oil Temperature

Engine Oil Pressure

The outer tape group on each side of the display contains the engine oil temperature T data. Units for the readouts are °C. Each engine has an oil temperature sensor that sends a signal through each DCU independently to the T tape; and to an ENG OIL HOT caution.

The outer tape group on each side of the display contains the engine oil pressure P data. Units for the readouts are PSI. Each engine has an engine oil pressure transmitter, downstream of the oil filter, which sends a signal through each DCU independently to the P tape and to an ENG OIL PRESS caution.

9/3/2003 Page 11 of 24 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Gas Generator Speed

Engine Power Turbine/Rotor Speed

The tape group second from the outside contains the NG data. Units for the readouts are percent (%). Each engine alternator sends a signal through each DCU independently to the NG tape. Turbine Gas Temperature The tape group second from the outside contains the TGT data. Units for the readouts are °C. The TGT indicating system consists of thermocouples transmitting through the DECU to each DCU independently to the TGT tape.

The top center tape group contains rotor speed NR and power turbine speed NP1 and NP2 data. Units for the digital readouts are percent (%). A speed sensor on the right accessory module transmits a signal through each DCU independently to the NR tape. Speed sensors on each engine exhaust frame transmit a signal through each DCU independently to the NP1 and NP2 tapes.

Torque The tape group second from the outside contains the torque Q data. Units for the readouts are percent (%). The torque system shows the amount of power the engine is delivering to the main transmission. A torque sensor mounted on the exhaust case measures the twist of the power turbine shaft, and transmits this signal to the DEC and through each DCU independently to the Q tape. Under single engine operating conditions, the torque limits will adjust from dual engine to single engine operating limits. Refer to Chapter 9 for single engine operation details. When the airspeed changes to greater than 80 Knots Indicated Airspeed (KIAS), the continuous torque limit indication will automatically change from 120% to 100%.

Engine Out Indications Engine out indications are displayed when an engine out condition is detected by the DCUs. The engine out indication is a large red “X” displayed across the two engine specific icons, which overwrites the engine tape symbology within the icons. The engine out symbol is displayed below the readout data, however, providing the pilot ability to continue to monitor the engine data, even though the engine is inoperative.

9/3/2003 Page 12 of 24 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Engine/Propulsion Parameter Comparison

Miscompare thresholds are listed below:

The EICAS parameter values from the two DCUs are compared to detect possible faulty data. The comparison is always between DCU 1 and DCU 2 data, regardless of the selected EICAS data source for display. It is normal to see slight differences between the pilot and copilot EICAS readouts due to the processing of parameters done separately by the DCUs.

EICAS Parameter

Miscompare Threshold

Engine Oil Pressure P

10.0 PSI

Engine Oil Temperature T

13.0 °C

NG

4.44%

TGT

7.3 °C

Torque Q

5.64%

NP

1.02%

NR (NR,70%)

5.00%

NR (NR>=70%)

1.73%

Main XMSN Oil Pressure

10.0 PSI

Main XMSN Oil Temperature

13.0 °C

Main Fuel Quantity

10.0 lbs

Each DCU supplies both engine 1 and engine 2 data, but there is no comparison of engine 1 and engine 2 data. When a miscompare condition is detected, the text label for the affected parameter turns to black text on a yellow background, and a yellow EIC flag miscompare annunciation with a double arrow line appears above the torque (Q1) readout on the Primary Flight Display (PFD) power pod. This cues the pilot to view the EICAS display to determine the specific miscompare parameter. The flag remains displayed as long as the miscompare condition is detected. Data values are not altered because of miscompare. Pilots must manually evaluate parameter data between the two DCU data sources and other available resources to assess the validity of each source’s data.

9/3/2003 Page 13 of 24 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Primary Flight Display (PFD) Engine Data Display – Power Pod The PFD provides a grouping of engine performance data in what is called the “power pod”. The power pod is located in the lower left corner of the display. The engine data displayed on the PFD provides information required for normal performance, monitoring, and flight operations as well as quick recognition/diagnosis of abnormal conditions. This allows the pilots increased flexibility to use cockpit Multi Function Displays (MFD) for display of mission situational awareness data rather than dedicating one or more displays full time to an EICAS format.

This also serves to reduce pilot scan and minimize display format, and includes the following: • •

• • •

Engine data miscompare annunciation “Triple Tach” display – providing an integrated indication of rotor speed (NR) and engine No. 1 and No. 2 power turbine speed (NP) Engine No. 1 and No. 2 torque (Q) display Engine No. 1 and No. 2 turbine gas temperature (TGT) display (under abnormal conditions only) Engine No. 1 and No. 2 gas generator speed (NG) display (under abnormal conditions only)

Engine Instrument “Power Pod”

9/3/2003 Page 14 of 24 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Power Turbine/Rotor Speed (NR/NP) Indication The integrated NR/NP display combines the three EICAS NR/NP tapes into a single tape with pointers. The tape in the center represents NR. The two pointers that bracket each side of the NR tape indicate NP 1 and NP 2. The display is marked with lower/upper RED danger limits, minimum/maximum Y Y E O W YE ELLLLLLO OW W continuous limits, and 100% GREEN line. A digital NR readout is located above the tape.

The red limit lines represent the boundaries between the cautionary ranges and critical operational ranges (above upper cautionary range and below lower cautionary range). The NR and NP limit lines are separate; with the NR limit lines displayed in the center of the scale and the NP limit lines displayed along the left and right sides of the scale. This independence is necessary as the rotor and power turbine limits do not match in all cases. The limit lines mark the points on the scale at which the parameter indications will transition color (color coded to reflect the nature of the operational region). The tapes, readouts, indicators are green in the normal operational range between the yellow limit lines, yellow in the cautionary regions, and red in the critical regions. These limit values are as follows:

NR/NP Scale And Limit lines NR/NP scale provides an indication range of 20% to 120% for NR and NP. The scale consists of 3 separate linear ranges between 20% and 90% NR/NP, between 90% and 110% NR/NP, and between 110% and 120% NR/NP. The scale tick marks vary by range – there are tick marks at 20% and 90% in the first range, at 1% intervals in the 90%-110% range, and then at 115% and 120%. The tick marks displayed at 5% and 10% values are longer, to provide another aid in quickly determining where the NR/NP indications are on the scale.

• •

Lower red limit NR – 91% NP – 91% Lower yellow limit Nr – 95% NP – 95%

NR/NP Tapes

Numeric indications are included on the scale at 20%, 90%, 100%, 110%, and 120%. The scale is white with a black outline (to distinguish the scale markings from the NR tape that moves behind the scale). The scale is displayed all the time; regardless of the NR and NP input statuses.

Limit lines are displayed on the scale to identify the operational ranges for each of the parameters – including the normal operational range and operational ranges, which exceed maintenance and/or performance related limits. The limit line colors identify the severity of the consequences of exceeding the limit.

• •

Upper yellow limit (max continuous limit) NR – 101 - 110% NP – 101 - 104 % Upper red limit NR – 110% NP – 105%

The limit and reference lines are displayed at all times. These limit and reference lines are displayed on top of the scale (and tape/indicators), and have black outlines to make them more visible against the background scale, tape, and indicators.

Yellow limit lines represent the boundary between the normal operational range and cautionary operational ranges (above and below desired range).

9/3/2003 Page 15 of 24 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Torque (Q) Indication Torque (Q) tapes Q1/Q2 and readouts function identically as the EICAS torque tapes. The operating range the displayed torque value falls within determines the Q digital readout and tape color. The tapes are marked with upper RED danger and maximum Y Y E O W YE ELLLLLLO OW W continuous limit lines.

The red limit line represents the boundary between the cautionary range (below it) and critical operational ranges (above it). These limit lines mark the points on the tape at which the tape/readout will transition color (color coded to reflect the nature of the operational region). The torque tape and readout are green in the normal operational range, yellow in the cautionary region, and red in the critical region. These limit values are as follows:

Torque Tape and Limit Lines The torque tapes (engine 1 and 2) provide an analog indication of the associated engine torque output. Each torque tape includes a white outline, tape fill that represents the current torque output from the associated engine, and limit lines that define the boundaries between normal, cautionary, and critical operating ranges. The tape scaling/tape height is linear from the bottom of the outline that represents 0% torque) to the top of the outline that represents 150% torque).

•

Yellow limit (max continuous limit) IAS ≤ 80 kts – 120% IAS > 80 kts – 100% Split Q – 135%

•

Red limit IAS ≤ 80 kts – 144% IAS > 80 kts – 144% Split Q – 144%

The display determines which set of limits to use for display and color determination based on the source selected airspeed and torque.

The torque limit lines are displayed at all times Engine Torque There are no scaling corresponding to whichever set Indications indications or numerics on the of limits is currently active. The outline, just limit lines limit lines are displayed on top displayed at the boundaries of the torque tape (overwriting it), and have a between the operating regions. black outline to make them visible against the tape fill/outline in the background. The engine 1 and 2 tape outlines are displayed all of the time; regardless of the associated Q input Dual engine toque limits are as follows: status. The tape is removed when the status of the associated Q input is invalid. ≤ 80 Knots >80 Knots Gr. ≤ 119% ≤ 99% Limit lines are displayed on the scale to identify Yellow 120 – 143% 100 – 143% the torque operational ranges– including the Red ≥ 144 % ≥ 144 % normal operational range and operational ranges, which exceed maintenance and/or performance related limits. The limit line colors identify the severity of the consequences of exceeding the limit. The yellow limit line represents the boundary between the normal operational range (below it) and cautionary operational ranges (above it).

9/3/2003 Page 16 of 24 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

NG Readouts Gas Generator Speed Display (NG) Indication The NG symbology is displayed only under abnormal engine conditions. A condition, which causes NG for one engine to be displayed, will cause the NG for both engines to be displayed. • • •

Either engine NG at or exceeding the yellow limit (99%) The difference between engine 1 and 2 NG exceeds 5% Either engine is out (engine out indications come from the DCU)

NG Readout Limits Each readout includes a label above the readout, “NG1” for the engine 1 readout and “NG2” for the engine 2 readout. The operating range the displayed numeric value falls within determines the NG readout color. These readout limit values are as follows: • • • •

NG readout < yellow limit value – green NG readout ≤ yellow limit value – yellow NG readout < red limit value – Yellow NG readout ≤ red limit value Red

9/3/2003 Page 17 of 24 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

TGT1 2.5 Minute Limit Ball TGT1 10 Minute Limit Ball TGT1 30 Minute Limit Tape TGT1 Readout / Label Turbine Gas Temperature Display TGT Limit Exceedance Indications Turbine Gas Temperature (TGT) Indication The TGT tapes operate slightly different than the EICAS TGT tapes. The power pod TGT includes a 30-minute limit line and color-coded balls above the tapes to indicate TGT limit exceedance The TGT symbology is displayed only under abnormal engine conditions. A condition, which causes TGT for one engine to be displayed, will cause the TGT for both engines to be displayed. • • •

The TGT limit exceedance indications include multiple symbols, which are displayed individually or in additive combinations based on the current displayed TGT value. Each engine has its own set of limit exceedance indications, displayed above the corresponding TGT readouts. All the TGT limit exceedance indications for an engine are disabled when the status of the TGT input for that engine is invalid.

Either engine TGT at or exceeding the yellow limit (810 - 850°C) The difference between engine 1 and 2 TGT exceeds 75°C Either engine is out (engine out indications come from the DCU)

30-MINUTE LIMIT TAPE

9/3/2003 Page 18 of 24 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

30-Minute Limit Tape When an engine TGT falls in this range, the 30minute limit tape for that engine is displayed. The tape includes an outline representing the full range of the tape (810 degrees at the bottom of the outline to 850 degrees at the top), and a tape fill whose height is scaled linearly within the tape outline to represent the engine TGT. If the engine TGT rises above the upper end of the 30-minute limit range, the 30-minute limit tape remains displayed, filled to the top of the outline. The tape is removed when the TGT falls below the lower end of the 30-minute limit range. The 30-minute limit tape is also removed when the displayed TGT value rises up to the 10-second limit region (at or above 903 degrees) – in this case it is replaced by another indication.

10-MINUTE LIMIT BALL 10-Minute Limit Ball When an engine TGT falls in this range, the 10minute limit ball for that is displayed. The ball includes an outline and fill, the outline being displayed fully filled whenever the TGT value is in this region (or above the region). If the engine TGT rises above the upper end of the 10-minute limit range, the 10-minute limit ball remains displayed. The ball is removed when the TGT falls below the lower end of the 10-minute limit range. The 10-minute limit ball is also removed when the displayed TGT value rises up to the 10-second limit region (at or above 903°) – in this case it is replaced by another indication.

9/3/2003 Page 19 of 24 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

2.5-MINUTE LIMIT BALL (CONTINGENCY POWER ARMED)

2.5-MINUTE LIMIT BALL (ACTIVE)

2.5-Minute Ball The 2.5-minute limit region (contingency power) symbology includes a ball similar to the 10-minute limit region ball, except that the outline and fill are displayed independently. The ball outline for an engine is displayed without fill when contingency power is armed (an input from the DCU), while the displayed TGT value for that engine is below the 2.5-minute limit range. When an engine TGT value falls in this limit range, the 2.5-minute limit ball outline and fill is displayed for that engine. The ball is removed when the TGT falls below the lower end of the 2.5-minute limit range and contingency power is not armed. If the TGT value drops below the 2.5-minute limit range, but contingency power is still armed, the ball outline remains displayed. The 2.5-minute limit ball is also removed when the displayed TGT value rises above this region (up to the 10 second limit region - at or above 903°) – in this case it is replaced by another indication.

10-SECOND LIMIT BAR 10-Second Limit Bar When an engine TGT falls in this range (or above it), the 10-second limit bar for that engine is displayed, replacing the other limit exceedance symbology. The bar includes a RED outline and a RED and Y Y E O W YE ELLLLLLO OW W barber pole fill, the outline being displayed fully filled whenever the TGT value is in this region

9/3/2003 Page 20 of 24 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Indication will extinguish when the generator speed is above 55%.

Check EICAS Caution Indication The DCU will illuminate the CHECK EICAS caution when the No. 1/No. 2 Engine Oil Temp or No. 1/No. 2 Engine Oil Pressure parameters exceed or decrease below the maximum or minimum continuous limits OR the cross channel DCU indicates the Check EICAS Caution is true. The parameters are: •

Engine 1 or Engine 2 Oil Temperature is at or exceeds 135° C

•

Engine 1 or Engine 2 Oil Pressure is at or exceeds 100 PSI

•

Engine 1 or Engine 2 Oil Pressure is at or below 26 PSI

Engine Out Indications Engine out indications are displayed when an engine out condition is detected by the DCUs. The engine out indication is a large red “X” displayed across the engine data for the specific engine that is inoperative. The engine parameter symbology overwrites the engine out indication and black backgrounds are displayed behind the Ng and TGT readout numbers when the engine out indication is displayed but the parameter data can still be read when the engine out indication is displayed.

Master Warning Indications #1 ENG OUT - Indicates when the No.1 Engine Gas Generator Speed (Ng) is at or below 55%. Indication will extinguish when the generator speed is above 55%. # 1 ENG OUT FIRE

MASTER CAUTION PRESS TO RESET

# 2 ENG OUT LOW ROTOR R. P. M.

MASTER WARNING PANEL #2 ENG OUT - Indicates when the No.2 Engine Gas Generator Speed (Ng) is at or below 55%.

Caution/Advisory Indications The Caution/Advisories provided by the Data Concentrator Units/Engine Indication Crew Alerting System (EICAS) system are similar if not identical to UH-60A/L Caution/Advisory Panel indications. In some case there have been minor syntax changes. They are listed here for general informational purposes. ENG 1 OIL HOT - Caution Indicates that the No. 1 engine oil temperature has exceeded 150 ° C ENG 2 OIL HOT - Caution Indicates that the No. 2 engine oil temperature has exceeded 150 ° C ENG 1 OIL PRESS - Caution Indicates that the No. 1 engine oil pressure has dropped below 22 PSI ENG 2 OIL PRESS - Caution Indicates that the No. 2 engine oil pressure has dropped below 22 PSI CHIP ENG 1 - Caution Indicates that the No. 1 engine chip detector has detected a chip or excessive metallic particle buildup in the engine scavenge oil system.

9/3/2003 Page 21 of 24 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

CHIP ENG 2 - Caution Indicates that the No. 2 engine chip detector has detected a chip or excessive metallic particle buildup in the engine scavenge oil system. ENG 1 OIL BYPASS - Caution Indicates the No. 1 engine oil filter has excessive pressure differential across filter and has been bypassed. ENG 2 OIL BYPASS - Caution Indicates the No. 2 engine oil filter has excessive pressure differential across filter and has been bypassed. ENG 1 STARTER ON - Caution Indicates that the No.1 engine start circuit has been activated. ENG 2 STARTER ON - Caution Indicates that the No.2 engine start circuit has been activated.

AIR VEHICLE STATUS PAGE

ENG 1 ANTI-ICE ON - Caution The active state indicates that the No. 1 engine anti-ice start/start bleed valve is open. ENG 2 ANTI-ICE ON - Caution The active state indicates that the No. 2 engine anti-ice start/start bleed valve is open.

Flight Management System (FMS) Status And Test Status The powerplant system components monitored by the FMS are the Digital Engine Controls (DECs). Selecting the STS fixed function key to access the MAIN STATUS page on the FMS can check the status of the DECs. By selecting soft key (SK) 4 from the MAIN STATUS PAGE accesses the AIR VEHICLE status page. Selecting SK 3 or SK 8 accesses air vehicle equipment status for DEC 1 and DEC 2. This menu selection will allow a user to view the status of the engines. All faults and failures will be displayed under the DEC if a fault or failure should occur.

DEC STATUS PAGE

9/3/2003 Page 22 of 24 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Test The Test Menu is accessed by the function key and provides the ability to conduct and review tests of all systems and subsystems or line replaceable units (LRUs) that can conduct an Initiated Built In Test (IBIT). By selecting soft key (SK) 4 from the MAIN TEST PAGE accesses the AIR VEHICLE status page. Selecting SK 3 or SK 8 accesses air vehicle equipment test for DEC 1 and DEC 2. This menu selection allows the operator to initiate an IBIT of the DECs. All faults and failures will be displayed under the DEC if a fault or failure should occur. .

DEC TEST PAGE

AIR VEHICLE TEST PAGE

9/3/2003 Page 23 of 24 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

NOTES

9/3/2003 Page 24 of 24 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Section 2-3 Fuel System Differences Description

Fuel System Components/Subsystems Main Fuel Tanks

The fuel system of the UH-60M helicopter has been unchanged relative to component function and operation. No changes have been made to the configuration of the fuel supply, prime or refuel/defuel systems. Changes have been made in the material composition and construction of the main fuel cells as well as the location of fuel quantity system components and fuel quantity indications used by the pilots.

The main UH-60M fuel tanks have been subjected to minor modifications to decrease weight as well as to install different internal components. The fuel tanks are now equipped with a breakaway fuel inlet fitting on the right hand main fuel tank “racetrack”. The larger inlet fitting is used on the MH-60K.

The UH-60M fuel system has been designed to incorporate the Crashworthy Extended Range Fuel System (CEFS) and its installed equipment and controls. When installed, CEFS components will be located in the cabin and transition overhead areas.

The main fuel system fuel quantity signal conditioner has been retained in the same location. The fuel quantity SDC output, representative of the main fuel cell quantity, is routed to the Data Concentrator Units (DCUs) which process the signals for display on the cockpit PFD and EICAS displays.

Fuel Quantity Signal Conditioner

UH-60M Main Fuel System 9/3/2003 Page 1 of 8 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Fuel System Controls

Engine Prime Boost Pump Control Switches

Auxiliary Power Unit Prime Boost Pump Control Switch

The Prime Boost Pump control switches provide control of fuel cell internal pumps that provide pressurized fuel to the fuel inlet ports.

The APU Prime Boost Pump control switch has been relocated on the UH-60M overhead console center panel. Previously located on the forward portion of the center panel, the control switch has been moved to a position in the middle of the center panel. The APU Fuel Prime Boost Pump control switch purpose and functionality remains unchanged.

The Prime Boost Pump control switches, previously located on a dedicated control panel on the lower console, have been relocated and integrated into the upper console center panel. Fuel Quantity Test Switch The Fuel Quantity Test Switch, previously located on miscellaneous switch panel on the lower console has been relocated and integrated into the upper console right panel.

APU Fire T-Handle The APU Fire T-Handle has been relocated on the UH-60M overhead console center panel. Previously located on the rear portion of the center panel, the T handle has been moved to a position towards the front of the center panel. The APU Fire T-Handle purpose and functionality remains unchanged.

The Fuel Indicator Test Control tests the fuel quantity indication system to ensure proper operation. This control is not illuminated.

CARGO HOOK

VIB CONT NO. 1 DC ESNTL BUS NO. 1 ENG

LIGHTS

IFF

5

5

5

FIRE DET

SEC

NO.1 STAB

SAS

ECP

5

2

7

1 2

PWR

AUX CB

CABIN ICS

35

4

PNL SPLY CARGO HOOK

BOOST

NO. 1 ESNTL

7

NO. 1 VHF

HOIST CABLE

5

5

5

EMER RLSE

FM

SHEAR

1 2

JTSN OUTBD

EMERG REL NORM

7

1 2

BATT

EXT LTS MODE

5

4

N O R M

1

CARGO HOOK LT

O F F

ANTI COLLISION LIGHTS

B O T H

DIM

STEADY

BRT

FLASH

OFF

BRT

CABIN DOME LT

OFF

DIM

BRT

5 PRI

PLT CDU

NO. 2 ADC

5

2

5

RESERVE

BACKUP HYD

5

PLT MWC

DIM

OFF

BRT

2

1 2

25

DCP

FUEL BOOST PUMP NO. 1

WINDSHIELD WIPER PARK

OFF

LOW

BRT

BATT GOOD

ENG NO. 1

BATT LOW

OIL HOT

7

INBD

1 2

BATT

ACC LOW APU ON

O F F

ANTI-ICE NO. 2

MED

ON

OFF

PITOT LEFT O F F

O F F ON

STBY INST TEST

HEATER

ON

HI

O F F

HEAT RIGHT O F F

ON

ON

FUEL BOOST PUMP O F F

ANTI-ICE PILOT

FUEL IND

1

ON

ON

R O E F S F E T

O F F

O F F

O F F

ON

ON

ON

R O E F S F E T

NO. 2 TEST

R O E F S F E T

ON

APU TEST

ON

EXT PWR

BRT

TEST

ENG OVSP TEST A NO. 1 B A NO. 2 B

ENG

O F F

O F F ON

ENG SPD TRIM

BOOST PUMP

CONT

RESET O F F

LAMPS

AIR SCE HI/STRT

APU

TEST

2

BLUE

O F F

NO. 1 TEST

ON

INST PNL

OFF

NO. 2 TEST

O F F

FIRE DETR TEST OPER

WHITE

O F F

ARMED

GENERATORS

NO. 1 TEST

ARM

NO. 2

WINDSHIELD COPILOT CTR

ALL

SHORT

ON

O F F

O F F

ARM SAFE

3

PRI

VENT BLOWER

HI APU

APU FAIL

APU APU FIRE

2

ICS ICU

NO. 2 DCU

ON

ENG

TEST LT

PLT ICS

2 SENSE

PLT MFD

PLT FD/

7

CONTR

O F F

EMER RLSE

NO. 2 ESNTL

VOR/ ILS

35 PNL SPLY

ON

LWR CSL

BRT

APU TEST

AIR SCE HI/STRT

ON

O F F

LIGHTS UPR CSL

ON

BOOST PUMP

FIRE EXTGH

ON

SECONDARY

R O E F S F E T

O F F ON

MAIN

LTD SW

BRT

NO. 2 TEST

2

ON

LIGHTS

DIM

ON

R O E F S F E T

APU O F F

ON

O F F BLUE

NO. 1 TEST

CONT

RESET

O F F NIGHT

CPLT MWC

ON

EXT PWR

POSITION LIGHTS

DAY

O F F LOWER

SECONDARY WHITE

ON

R O E F S F E T

AUX CB

VHF AM

NO. 2 EGI

PLT

MSTR WARN

ON

OFF

UPPER

NO. 2 TEST O F F

O F F

O F F

O F F IR

ARMED

GENERATORS

NO. 1 TEST

STBY INST TEST

ARM

FORMATION LTS 3 2

ARM SAFE

ALL

SHORT

ON

ESSS JTSN

INBD

CONTR CKPT

O P E N

O F F

CONTR CKPT

O P E N

O F F

NO. 2 DC ESNTL BUS

CARGO HOOK

VIB CONT

ESSS

2 SENSE

EMERG REL NORM

ON

O F F ON

APU

DECR O F F INCR

FIRE EXTGH RESERVE

UMAV002_1

O F F

Overhead Console

MAIN

FUEL BOOST PUMP O F F

NO. 1

APU APU FIRE

ON

EMER RLSE TEST LT

BATT GOOD BATT LOW

OIL HOT

ACC LOW

APU FAIL

APU ON

FUEL BOOST PUMP O F F

NO. 2

ON UMAV002_4

O/H Console Center Panel

O/H Console Right Panel 9/3/2003 Page 2 of 8 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

UMFC003

Engine Control Quadrant – Engine Fuel System Selector Levers The physical configuration and functional use of the engine fuel selector levers is unchanged from the UH-60A/L.

Engine Control Quadrant – Fire T-Handles The physical configuration and functional use of the engine fire T-handles are unchanged from the UH-60A/L. The No. 1 and No. 2 Engine Emergency Off Fire Warning Capsules have been modified to conform to Night Vision Goggle (NVG) lighting conventions.

9/3/2003 Page 3 of 8 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Fuel System Indications

MFD Indications

Fuel Quantity Indicating System

Fuel indications, along with engine, transmission and main rotor data, are available in two locations:

UH-60A/L Vertical Instrument Display System (VIDS) display of fuel quantity using the Central Display Unit (CDU) fuel quantity tapes and totalizer have been replaced by the Multifunction Display (MFD) fuel quantity indications. The fuel tank unit sensors, or probes and probe-mounted low-level sensors located within the fuel tanks retain the physical configuration and functional characteristics of the UH-60A/L fuel system. Additionally, the fuel quantity signal conditioner has been retained. In the UH-60M, the fuel quantity data generated by the SDC as well as low-level sensor data is passed to the Data Concentrator Unit (DCU) where it is further conditioned and transferred to the MFDs via an ARINC-429 data bus.

• •

EICAS PFD

Main And Auxiliary Tank Fuel Quantity EICAS Display Main Tank Fuel Quantity Display The primary EICAS display of fuel quantity consists of main tank fuel quantity tapes and readouts, left and right side auxiliary tank summary readouts, and total aircraft fuel quantity readout.

Fuel Quantity Indicator Tapes and totalizers 9/3/2003 Page 4 of 8 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Each main tank fuel quantity is represented on the display by a tape and a digital readout. The main tank fuel quantity display includes several labels – “FUEL” which serves as a label for all the fuel quantity symbology and “MAIN” displayed between the two tapes, and individual labels on each tape (“1” on the left tape and “2” on the right tape). The labels indicate that the tapes represent main tank fuel quantity and which tape represents which tank.

Main Fuel Quantity Indications The main tank quantity fuel tapes are scaled from 0 lbs at the bottom of the tape to the maximum main tank quantity (1200 lbs) at the top of the tape. There is a yellow low fuel quantity caution line displayed at the 172 lb position on each tape. The tape color matches the color of the readout.

input is failed or if the input is not received, then in addition, the readout is displayed as 3 red dashes, and the main tank label (“1” or “2”) is displayed as a fail flag (red background with black text). Auxiliary Tank Fuel Quantity Display The quantity of fuel contained in auxiliary (aux) tanks installed on the aircraft is represented on the display by individual numeric readouts for each installed tank. The display receives installation status information from the DCU for each possible aux tank, and displays only the symbology appropriate for the aircraft as configured The aux tank readout is not displayed if the that tank is not installed. The auxiliary tank fuel quantity display includes: • Labels identifying the left and right aux tank data, • 4 digit numeric readouts for each tank, • Labels identifying which tank the quantity data applies to • A box surrounding the readout and label (grouping that symbology and separating it from the other aux tank data that is potentially near by). The auxiliary tank readout positions are “stacked” on the left and right side of the main tank fuel display, correspond to the physical location of the installed tanks they represent and are labeled to indicate the tank they apply to.

Centered below each main tank fuel quantity tape is a numeric readout representing main tank quantity in lbs. The readout is a maximum of 4 digits and is normally shown in green. If the fuel quantity readout for a tank is at or below the low fuel caution value (172 lbs), the readout (and tape) color associated with that tank changes to yellow. The “MAIN” and “FUEL” labels and the labels associated with each individual main tank quantity (“1” and “2” displayed inside the tapes), are displayed all the time and normally are white with a black outline to be visible against the tape fill. If a miscompare condition exists for a main tank fuel quantity, the associated tank label (“1” or “2”) is displayed as a miscompare annunciation. If the input associated with a main tank fuel quantity is invalid, the tape fill is removed and the numeric readout is removed. If the status of that

Auxiliary Fuel Quantity Indications The auxiliary fuel data display labels are ”L AUX” and “R AUX,” displayed above the positions reserved for the left and right aux tank readout “stacks.” These labels are white and are displayed

9/3/2003 Page 5 of 8 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

when any aux tank is installed on that side of the aircraft.

Primary Flight Display Fuel Data Display

The auxiliary tank readouts have fixed positions on the display for each left/right inboard/outboard/external tank position. The tank fuel quantity readouts are each centered within a white box outline, each having a maximum of 4 digits not to exceed 3060 lbs.

The display of fuel data on the PFD includes the information required for normal performance, monitoring, and flight operations as well as quick recognition/diagnosis of abnormal conditions. This allows the pilots increased flexibility to use cockpit PFDs for display of mission situational awareness data rather than dedicating one or more displays full time to an EICAS format. This also serves to reduce pilot scan and minimize display format.

Each readout includes a label displayed above the numeric fuel quantity value within the readout. The label is offset (within the readout box) toward the center of the display if it is an inboard tank readout, toward the outside of the display if it is an outboard tank, and centered in the readout box if it is an internal tank. The readout labels are: • • •

“OB” – for outboard external tanks “IB” for inboard external tanks “INT” for internal tanks

The PFD fuel quantity display includes readout of total main tank fuel quantity and fuel a graphical indication of fuel quantity in each of the main tanks. The main fuel tank gage includes a fuel quantity scale, a low fuel level indication on the scale, and main tank 1 and 2 fuel quantity tapes The fuel quantity scale is white with a black outline and displays a range from 0 to 1200 lbs max fuel quantity for the main fuel tanks installed on the aircraft. A center mark divides the scale in half representing each half with 600 lbs. A white “FUEL” label is displayed above the scale to identify the basic function of this area of the display. The fuel quantity scale is displayed at all times, regardless of the status of the fuel quantity inputs.

Aircraft Total Fuel Quantity Display The total aircraft fuel quantity (including all main tank and auxiliary tank quantities) is displayed as a numeric readout. The readout is summaries of the two main tank fuel quantities and the two summary aux tank fuel quantities. The aircraft total fuel quantity readout is centered within a white box outline, and has a maximum of 5 digits. The readout includes a label displayed above the numeric fuel quantity value within the readout outline. The label is white, and designates the readout as the “TOTAL” aircraft fuel quantity.

PFD FUEL QUANTITY DISPLAY Fuel Quantity Limits A yyyeeellllllooow w w limit line is always displayed on the gage at the low level indicating the low fuel quantity limit (172 lbs). When the fuel quantity falls below the

9/3/2003 Page 6 of 8 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

low fuel level then tape and the readout becomes yyyeeellllllooow w w. When the fuel quantity is greater than the low fuel quantity level, the color on the tape and the readout is gggrrreeeeeennn. If both main tank fuel quantity inputs are failed, the readout is replaced by three red dashes.

Fuel Quantity Test Indications The Fuel Indicator Test Control tests the fuel quantity indication system to ensure proper operation. Pressing and holding the fuel indicator test switch on the overhead console, will blank out the vertical tapes of the No. 1 and No. 2 main fuel quantity display on the PFDs and the digital readouts display 0 lb.

Fuel System Operation Fuel Supply System FUEL QUANTITY FAIL

Engine and Auxiliary Power Unit fuel supply system operation remain unchanged from the UH60A/L.

Caution/Advisory Indications PRIME BST PUMP ON - Advisory Indicates that the Prime/Boost Switch is in the "Prime" or "Boost" position. BOOST PUMP 1 ON Indicates that the No.1 Boost Pump is on. BOOST PUMP 2 ON Indicates that the No.1 Boost Pump is on. FUEL 1 LOW - Caution Indicates that the left fuel tank level is about 172 pounds at normal cruise flight. FUEL 2 LOW - Caution Indicates that the right fuel tank level is about 172 pounds at normal cruise flight.

9/3/2003 Page 7 of 8 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

NOTES

9/3/2003 Page 8 of 8 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Section 2-4 Mechanical Flight Control System Differences Description The UH-60M flight control system is virtually identical to the UH-60A/L. The UH-60M mechanical flight controls allow control inputs from the pilot, automatic flight control system or flight director to be mechanically transferred from either the cockpit or pilot assist assemblies to the main and tail rotor systems. The flight control system is comprised of lateral, longitudinal and collective control subsystems. The routing of the mechanical linkages between the cockpit and rotor systems is identical to the UH-60A/L. The UH-60M powertrain system modifications impact the following components and/or system functionality: • • • •

Collective stick grips Cyclic stick grips Trim System – Collective Trim Servo and associated mechanical linkages

New Collective Trim Servo

• •

Flight Control Mixer Unit – Roll channel Swashplate Rotating Scissors assemblies

The mechanical flight control system changes accommodate the incorporation of the wide chord blade and collective trim servo into the main rotor system. The lateral control system changes allow increased control authority, which is required to achieve the full benefit of the new wide chord blade. These changes also eliminate interference conditions encountered in achieving the increase in motion. The changes in motion and hardware have required some minor changes to the flight control rigging procedures. The procedural change requires maintenance personnel to place the system in certain positions to verify that proper control stops are contacted and clearance exists. These changes are designed to ensure that UH60A/L mixer components are not mistakenly installed on the UH-60M. The addition of a collective trim servo gives the UH-60M automatic altitude hold capabilities not found on UH-60A/L model aircraft. The collective trim servo is driven by inputs from the Flight Director system when altitude hold mode is engaged.

Top Deck

Main Transmission

Controls “Closets”

Modified Mixer Assembly

Cyclic Sticks

Collective Friction Lock Deleted Pedals

Collective Sticks 9/3/2003 Page 1 of 12 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Mechanical Flight Control System Components Cyclic Stick The cyclic control stick is the interface between the pilot and the pitch and roll channels of the primary flight control system. Additionally, 10 control switches are located on the stick.

The cyclic stick assembly function and use is identical with the UH-60A/L. The UH-60M cyclic stick assembly replaces the UH-60A/L cyclic with one adapted from Naval Hawk aircraft. The two major changes made to the cyclic sticks are: the stick grip and the number and location of switches located on and about the grip. The cyclic stick grip contains the following switches:

Cyclic stick grip control switches allow activation without requiring the pilot to remove their hands from the cyclic stick grip.

CARGO REL

C/F DISP

WEPS REL

Y RMT SB

TRIM REL

TRIM

UMFC001_7

UH-60M Cyclic Stick Grip 9/3/2003 Page 2 of 12 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

CHAFF-FLARE DISP

ICS/RADIO

Pressing the CHAFF-FLARE DISP guarded switch will cause chaff cartridges to be ejected from M130 chaff dispensers located on the tail cone module.

The WEPS RELEASE switch initiates the launching of weapons including the air volcano system.

This 2-detent trigger switch provides the pilots with push-to-talk radio control for either the internal communication system or external radio communication. The first detent position activates the ICS radio while the second, fully depressed position activates the external radio. Additionally, a foot-activated switch on the inboard side of each pilot station will activate transmission on the Intercommunication system (ICS) or the external radio as selected on the ICS panel.

CARGO HK REL

STICK TRIM

The CARGO HK REL pushbutton initiates the normal electrical release of the cargo hook load beam.

A beep signal from the pilot/copilot’s cyclic grip is a momentarily active signal that increments the current trim settings of an aircraft attitude parameter. The parameter value is incremented continuously for as long as the pilot holds the switch. • Left beep: increases roll attitude to the left. • Right beep: increases roll attitude to the right • Aft beep: increases nose-up pitch attitude. • Forward beep: increases nose-down pitch attitude. • Center beep: stop and hover at current location.

WEPS RELEASE

GO-AROUND This pushbutton control will cause the automatic flight system to bring the aircraft to a vertical speed of 750 fpm up and accelerates to an airspeed of 70 kilometers per hour (kph). VOX-CAUT ACK This switch allows a Master Caution Reset signal to be sent to both of the Data Concentrator Units (DCU), which in turn will cause the Master Warning capsules on the master warning panels to be extinguished. These switches have the same effect as pressing the master warning press to reset switch on the master warning panels. REMOTE STNDBY The REMOTE STANBY pushbutton, when pressed places the active Flight Director mode into standby mode.

STABILATOR SLEW The Stabilator slew control trigger switch provides the pilot and copilot with rapid accessibility to stabilator slew up. When activated, the stabilator trailing edge will begin to move up and continue until the up limit stop is reached or the switch is released.

TRIM RELEASE When the trim system is engaged, the use of this switch will momentarily disengages, or releases, the cyclic trim system.

9/3/2003 Page 3 of 12 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Collective Stick

HOOK EMER REL

The Collective stick provides the interface between the pilot and the collective pitch channel of the primary flight control system. It is also the location of eight control functions or switches. Collective stick grip control switches permit operator inputs to be made without the pilot having to remove his/her hand from the flight control.

Using the guarded pushbutton initiates the emergency release of an external cargo load by activating an electrically fired explosive cartridge.

The collective stick assembly function and use is identical with the UH-60A/L. The UH-60M collective stick assembly incorporates two changes; the deletion of the pilot’s friction lock and the stick grip. The friction lock is no longer required due to the addition of the collective trim actuator and associated system trim functionality. The function and location of switches located on and about the stick grip has changed. The collective stick grip has the following switches:

HOOK EMER REL

LDG LT EXT/RETR - PUSH ON/OFF Pushing the switch will alternately illuminate or extinguish the landing light. The EXT position will cause the landing light to extend. Using the RET position will cause the landing light retract to the stowed position.

LDG

LT EXT PUSH ON OFF

HUD

CRSR SLEW PUSH-SELEC

PG

UP

DCLT

RA SE D L UP

BRT L

DIM

T

EXTD

R

PG DN

RETR DN

TR IM DN L R UP

UMFC001_6

UH-60M Collective Stick Grip 9/3/2003 Page 4 of 12 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

•

CRSR SLEW This control, a multi-directional control pad with a center pushbutton activates and controls the movement and direction of the cursor on the Multifunction Display (MFD) screens. Pushing the center of the control allows the selection of items via the cursor on the MFD. RADIO SEL The radio select switch on each collective stick grip provides four signals to the on-side CDU. • EXT NEXT - provides the same function as the NXT (next) buttons on the CDU keyboard, allowing cycling through a selected list on the CDU display. • COMM/STEP DN - allows the list of available radios to be cycled through in descending order. • EXT PREV - provides the same function as the PRV (previous) button on the CDU keyboard, allowing cycling through a selected list on the CDU display. • COMM/STEP UP - allows the list of available radios to be cycled through in ascending order.

R - Moves the searchlight to the right.

HUD CONTROL Each collective stick-grip switch is connected in parallel to the Heads Up Display (HUD) signal data converter and converter control unit. The switch outputs five HUD control signals to the HUD. They are: • • • • •

Page Up Page Down Declutter Brightness Increase Brightness Decrease

Trim Servos The flight control trim servos are electromechanical and electrohydraulic devices that reduce pilot workload as well as enhance aircraft controllability. The UH-60M flight control system modification incorporates an electromechanical collective trim servo that allows altitude hold capability.

TRIM BEEPER

Hands Off Flight

Collective stick beep switch allows for the collective trim setting to be incrementally changed or “beeped” after collective trim mode has been engaged. The beeper trim signal can be incremented continuously until the end of travel and as long as the pilot holds the switch.

When the trim system is engaged, trim servo internal clutches connect the control axis to an internal servo gear train that maintain flight control positions. The trim system allows hands-off flight control operation by maintaining cockpit control stick position. Control stick position will remain fixed unless repositioned by either the pilot or the Automatic Flight Control System (AFCS).

The four collective beep signals are defined as: • Left heading beep: increases yaw attitude to left • Right heading beep: increases yaw attitude to right • Up beep: increases collective up. • Down beep: increases collective down. SEARCHLIGHT CONTROL The four-way switch with center pushbutton provides directional control of the searchlight. Pressing the center of the switch activates or deactivates the searchlight. • • •

EXTD - Extends the searchlight. RETR - Retracts the searchlight. L - Moves the searchlight to the left.

Pilot Override Capability Springs located within the trim servo assemblies allow the cockpit flight control to return to the original position after pilot displacement. If the trim servo system gear trains bind, mechanical clutches will allow the pilot to override the affected axis by pulling or pushing on the appropriate control input. AFCS Control The AFCS provides drive signals for long-term attitude retention, autopilot and couples modes of flight. Input from the AFCS system will drive servo internal motors that reposition the control axis to maintain the desired attitude of the aircraft. This

9/3/2003 Page 5 of 12 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

drive is transferred to the control axis via the servo gear train and torque shaft in order to alter the axis position. As the attitude correction is made, the corresponding cockpit control stick will move unless resisted by the pilot. The trim servos displace control axes through 100% of normal axis positioning when engaged.

breakthrough the clutch is 22 pounds. The trim point can be repositioned using the TRIM REL trigger switch located on the underside of the collective stick. AFCS and Flight Director coupled modes cause the trim actuator to reposition the cockpit collective control at a maximum rate of 10% of pilot control authority per second.

Collective Trim Servo The collective trim servo, just as the roll and yaw trim actuators do, provide basic manual trim functions by maintaining the pilot’s desired collective control position. Collective trim servo operation is commanded by the Automatic Flight Control system when operating in a coupled mode with the altitude hold mode of the Flight Director engaged. The collective trim actuator supplies a force preload and gradient when a pilot/copilot control is displaced from its engaged trim point. The force required to

9/3/2003 Page 6 of 12

Collective Trim Servo Installation UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Mixing Unit The mixer unit provides control mixing functions that minimize inherent control coupling. Using rods, levers, bellcranks, and mechanical linkages individual flight control axis inputs are combined in order to provide the required output to the main and tail rotor servos in order to achieve controlled flight. Mechanical limiters within the mixer fix, or limit, the maximum amount of travel for certain combinations of control inputs and provide the aforementioned mechanical coupling from one control axis to another to reduce pilot workload. The types of mechanical mixing or coupling and their functions are common with the UH-60A/L. They are:

Mechanical • Collective to pitch • Collective to yaw • Collective to roll • Yaw to pitch

Left

Right

Mixer Assembly – Rear (facing forward) View

9/3/2003 Page 7 of 12 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Mixer Modifications Two components within the mechanical mixer have been changed to accommodate the use of the wide chord blade in the UH-60M main rotor system. The left lateral limiter and left lateral bell crank have been modified and can be seen to be physically different than those found in an UH60A/L mixer.

New Lateral Bell Crank

New Left Lateral Limiter

Mixer Assembly – Modified Components

9/3/2003 Page 8 of 12 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Left Lateral Limiter The left lateral limiter has been replaced for increased left lateral control range as a result of the wide chord blade. The increase in left lateral control authority occurs mostly in the upper end of the collective range. At low collective, left control motion is identical to the UH-60A/L.

Existing Limiter

Flat Surfaces

Same Radius UH-60A/L Left Lateral Limiter

UH-60M Modified Left Lateral Limiter

Left Lateral Bellcrank The left lateral bellcrank has been replaced to remove an interference condition that resulted from the increase in left lateral control motion within the mixer.

UH-60M Bellcrank Profile

UH-60A/L Bellcrank Profile

UH-60M Left Lateral Bellcrank 9/3/2003 Page 9 of 12 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Mechanical Flight Control System Controls Cyclic Stick The cyclic sticks provide lateral and longitudinal control of the helicopter by means of mechanical linkages made by pushrods, bellcranks, and servos. The mechanical redundancy features found in the UH-60A/L are also a design feature of the UH-60M. Mechanical stick stops and mixer limiters limit the overall range of the cyclic stick.

linkages made by pushrods, bellcranks, and servos. The mechanical redundancy features found in the UH-60A/L are also a design feature of the UH-60M. The telescoping stick feature, that allows the copilot to twist the copilot stick grip and slide it rearward for better cockpit ingress/egress feature is common with the UH-60A/L. Mechanical stick stops limit the overall range of the collective stick.

Yaw Pedals The yaw pedals provide directional control of the aircraft. The pedals are equipped with switches that, when pressed, disengage the heading hold feature of autopilot below 50 KIAS.

Collective Stick The collective sticks provide the ability to change the pitch of the main rotor blades that cause lift generated by the main rotor system to increase or decrease. Collective control inputs are transmitted to the rotor system by means of mechanical

The pedals are adjustable to allow adjustment of pedal location to accommodate pilot height. The pedal adjustment T-handles are located on each side of the instrument panel.

Top Deck Collective Trim Servo Main Transmission

Controls “Closets”

Mixer Assembly

Cyclic Sticks

Pedals

Collective Sticks

UH-60M Mechanical Flight Control System 9/3/2003 Page 10 of 12 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Indicates that 2nd primary flight control servo pressure is too low.

Miscellaneous (MISC) Switch Panel The Servo Off and Hydraulic Leak Test control switches have been relocated from other cockpit location to the Miscellaneous Switch panel.

TRIM FAIL - Caution Indicates that yaw, roll, or pitch trim actuators are not responding accurately to computer signals.

Servo Off Switch

T/R SERVO 2 ON - Advisory Indicates that 2nd stage Tail Rotor Servo pressure is above the specified minimum (normal operation).

This control switch allows either the first or second stages of the primary servos to be deactivated.

TAIL SERVO

M I S C

NORMAL

BACKUP HYD PUMP ON

A U T O

N O R M

ON

OFF

OFF

BACKUP

S W I T C H

HYD LEAK TEST

SERVO OFF 1ST STG +

NORM 2ND STG

UMAV003_4 Miscellaneous Switch Panel

Mechanical Flight Control System Indications The Caution/Advisories provided by the data Concentrator Units/Engine Indication Crew Alerting System (EICAS) are identical to UH-60A/L Caution/Advisory Panel indications. They are listed here for general informational purposes.

Caution/Advisory Indications BOOST SERVO OFF - Caution Indicates that there is no pressure at the Yaw Boost Servo; the Pitch and Yaw Boost Servos are not operating. PRI SERVO 1 FAIL - Caution Indicates that 1st stage primary servo pressure is too low. PRI SERVO 2 FAIL - Caution

9/3/2003 Page 11 of 12 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Mechanical Flight Control System Operation Normal Operation Operation of the mechanical flight control system is virtually the same as the UH/60A/L. The extents of the differences are to be found in the lateral cyclic channel due to the wide chord blade. These differences are charted in the accompanying graphic.

Collective / Lateral Stick Motions 10

Collective Stick Motion - Inches

9

Left Stick Stop

8 7

Right Stick Stop

UH-60A/L Limiter

6 5 4 UH-60M Limiter

3

Right Lateral Limiter

2 1 0 -6

-5

-4

-3

-2

-1

0

1

2

Lateral Stick Motion - Inches UH-60A/L Limiter

UH-60M Limiter

9/3/2003 Page 12 of 12 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

3

4

5

6

Section 2-5 Automatic Flight Control System (AFCS) System Description The UH-60M helicopter is configured with a Digital Automatic Flight Control System (DAFCS) that is comprised of Dual Digital Flight Control Computers (DDFCCS) and a folding stabilator to replace the existing components of the AFCS in the UH-60A/L helicopter. The AFCS enhances the stability and handling qualities of the helicopter and provides autopilot functions. It is comprised of five basic subsystems: • • • • •

Stabilator Stability Augmentation System (SAS) Trim Systems Flight Path Stabilization (FPS) Coupled Flight Director (FD)

The stabilator system improves flying qualities by positioning the stabilator by means of electromechanical actuators in response to collective, airspeed, pitch rate, and lateral acceleration inputs. The SAS provides short-term rate damping in the pitch, roll, and yaw axes. Trim system provides control positioning and force gradient functions as well as basic autopilot functions with FPS engaged. The AFCS provides the following features: • Pitch, roll, and yaw stability augmentation • Stabilator control (STAB) • Cyclic, collective, and pedal trim. (TRIM) • Pitch and roll attitude hold. • Heading Select • Pitch and roll hover augmentation/gust alleviation. (DSAS) • Turn coordination • Fault detection and annunciations • Uncoupled FD integration

AFCS functions include pitch and roll attitude hold, airspeed hold, heading hold and turn coordination.

20 20 10 10

20 20 10 10

20 20 10 10

20 20 10 10

10 10 20 20

10 10 20 20

10 10 20 20

10 10 20 20

T

T

T

T

ARINC 429 •Bearing Pointers •Nav. Source Select •DH •Altimeter Set •RA Bug Set

Copilot’s FD/DCP

ARINC 429 •Mode Selection •Reference Selection

Pilot’s FD/DCP

ARINC 429 •Reference Display

#1 Flight Control Computer

#2 Flight Control Computer

AUTOMATIC FLIGHT CONTROL SYSTEM (AFCS) WITH DUAL DIGITAL FLIGHT CONTROL COMPUTERS (DDFCCs) 9/3/2003 Page 1 of 20 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Automatic Flight Control System Components Dual Digital Flight Control Computers The UH-60M AFCS is centered around two redundant FCCs. Each computer accepts signals from the pilot and copilot controls, motion sensors, and control panels, ARINC-429 avionics systems to compute commands, which are sent to the cockpit, trim actuators, SAS actuators, and stabilator actuators. The computer commands inner-loop SAS actuators and the outer-loop trim actuators in pitch, roll, and yaw control channels. The computers also provide self-monitoring, fault isolation, and failure advisory. The AFCS provides two types of control, identified as inner-loop and outer-loop. The inner-loop (SAS)

employs rate damping to improve helicopter stability. This system is fast in response, limited in authority, and operates without causing movement of the flight controls. The outer-loop (Trim) provides long-term inputs by trimming the flight controls to the position required to maintain the selected flight regime. It is capable of driving the flight controls throughout their full range of travel (100% authority) at a limited rate of 10% per second. Both inner and outer loops allow for complete pilot override through the normal use of the flight controls. The FCCs process incoming information from various sensors aboard the aircraft and stores this information in its memory. The sensor information is used by the computer Central Processing Unit (CPU) to compute required correction signals. Inner-loop correction signals are routed to the SAS actuators and outer-loop signals are routed to trim servos and actuators.

No. 1 Flight Control Computer

9/3/2003 Page 2 of 20 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

No. 2 Flight Control Computer

Digital Flight Control Computer 9/3/2003 Page 3 of 20 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

•

AFCS Flight Control Panel. The flight control panel located on the lower console controls the AFCS. The flight control panel permits communication with FCCs and flight controls through which AFCS flight modes of operation are selected. Control functions are as follows: • • •

The stabilator position control provides manual control and automatic positioning of the stabilator. The SAS function provides aircraft body rate damping in pitch, roll, and yaw. The trim system maintains stick trim in pitch, roll, yaw, and collective through the respective actuators.

STABILATOR MAN SLEW UP O F F

•

The FPS function maintains aircraft flight path through control of attitudes, heading, and lateral acceleration. The FD provides pitch, roll, and collective steering to various guidance commands. These commands are displayed on the PFD and are also sent to the FPS when the FD is coupled.

Hover Augmentation/Gust Alleviation. An additional feature of SAS, is hover augmentation/gust alleviation. It further improves aircraft stability at low airspeed using attitude retention and longitudinal and lateral acceleration to reduce drift.

CONTROL AUTO CONTROL

TEST

ON DOWN

SAS 1

ON

SAS/BOOST

ON

AUTO FLIGHT CONTROL TRIM SAS 2

ON

ON

R E S E T

FPS

ON

FAILURE RESET

CPTR 1

CPTR 2

UMAV003_8

9/3/2003 Page 4 of 20 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Folding Stabilator The UH-60M replaces the UH-60A/L stabilator with a folding stabilator and lockpin limit switches, which are mounted on each side of the center box to ensure that the stabilator is in the locked position and ready for flight. The Stabilator can be folded when the helicopter is being transported.

Stabilator Actuator Attachment Clevis

Stabilator Elastomeric Bearings

The angle of the stabilator can be controlled either automatically or by pilot control. Stabilator failures cause the STABILATOR caution to appear on the MFDs and generate an aural warning tone in the pilot’s and copilot’s headsets.

Limit Switch and Position Transmitter Attachment Clevis

Stabilator/Airframe Attachment Points

9/3/2003 Page 5 of 20 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Collective Trim Servo The collective trim servo, located in the forward top deck, provides the similar function as the yaw trim servo with the exception of engage/disengage signals. The Trim switch on the flight control panel, provides 28 vdc power to the collective trim servo amplifier and the collective stick’s Trim Release (REL) switch.

Collective Trim Servo Installation

9/3/2003 Page 6 of 20 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Collective Sticks

Cyclic Sticks

The UH-60M replaces the collective sticks with one that provides a TRIM REL and a beeper button function for the AFCS and removes the friction locks.

The AFCS controls on the UH-60M cyclic sticks differ from those on the UH-60A/L in the following areas: •

When either collective stick TRIM REL switch is pressed, 28 vdc is removed from the collective trim servo disengaging the trim servomotor from the flight control linkage. Force gradient springs in the collective servo permit the pilot to override the trim reference position without losing the trim reference point. The collective stick contains a four-way beeper to reposition collective and yaw trim. When FPS is engaged, activating the yaw portion of the switch adjusts the reference heading used by Heading Hold. The collective position of the switch adjusts collective Flight Director functions.

•

CARGO REL

C/F DISP

HOOK EMER REL

•

Five-Position Trim Beeper – The cyclic beeper can be used to incrementally adjust trim settings of the current aircraft attitude. Go-Around Switch – A cyclic mounted switch used to automatically bring the aircraft to a vertical speed of 750 fpm up and accelerate to an airspeed of 70 kilometers per hour. Standby Switch – When activated, the standby switch will decouple and disengage all FD modes.

LDG

LT EXT PUSH ON OFF

HUD

CRSR SLEW PUSH-SELEC T

PG

UP

DCLT

EXTD

RA SE D L UP

BRT L

DIM

R

PG DN

RETR

DN WEPS REL

TRIM

TR IM DN L R

Y RMT SB

TRIM REL

UP

UMFC001_7

CYLIC STICK UMFC001_6

TRIM BUTTON

COLL STK TRIM RELEASE

9/3/2003 Page 7 of 20 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Flight Director/Display Control Panel (FD/DCP) The Flight Director/Display Control Panel (FD/DCP) located in the pilot and copilot side of the Instrument panel, provides independent pilot and copilot control of both FD and display selections and settings. There are settings and selection control functions that are performed by the FCC providing the applicable settings/selections to the MFDs.

REF SEL LOC

CAP REF ADJ GA

GS CAP

RALT ----

VS ----

ALT ----

IAS ----

HDG 240

DECL CAP

CPLD **** HVR

PUSH-SYNC

PUSH-SYNC

PUSH-SYNC

PUSH-SYNC

PUSH-SYNC

Flight Director/Display Control Panel

9/3/2003 Page 8 of 20 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

limited to the range between –10°(trail edge up) and 45° (trailing edge down) rotation. If the stabilator angle exceeds these limits, the pointer remains parked at the nearest limiting value (-10 or 45). The stabilator pointer is displayed when the stabilator input status is normal.

Automatic Flight Control System Indications Stabilator Display The stabilator angular position is presented on the PFD in the upper left corner. It provides an angular indicator and a digital readout. The stabilator symbology is white when the stabilator is operating in automatic mode (position automatically adjusts based on airspeed) and yellow when operating in manual mode.

The stabilator readout is displayed above the stabilator indicator. The readout range and limits match the range and limits of the stabilator pointer. When the stabilator input status is failed, the stabilator indicator and readout are removed and a stabilator fail flag is displayed.

Stabilator Indicator The stabilator indicator is an airfoil shaped indicator, which rotates to indicate the current stabilator position. The indicator rotates around the center of rotation point. The pointer motion is

Stabilator Indicator

9/3/2003 Page 9 of 20 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Cautions and Advisories Cautions and advisories received by the Data Concentrator Units (DCUs) from the AFCS are via an ARINC-429 data bus. Each caution and advisory is received by each DCU from either or both FCCs. If either DCU receives a particular caution signal from either or both FCCs, that caution and/or advisory message is displayed on all MFDs. All detectable AFCS mode failures except for the stabilator, cause the AFCS FAIL caution to appear on the Engine Indication Crew Alerting System (EICAS) The caution/advisories provided by the DCU/Engine Indication Crew Alerting System (EICAS) are similar if not identical to UH-60A/L caution/advisory panel indications. In some cases there have been minor syntax changes. They are listed here for general informational purposes. AFCS FAIL – Caution Indicates that a significant AFCS failure occurred that caused the shutdown of the majority of AFCS functions. SAS OFF – Caution Indicates a failure that affects both the No.1 SAS and No. 2 SAS or low hydraulic pressure. FPS FAIL – Caution Indicates that one or more of the FPS functions has disengaged due to a failure in pitch or roll attitude hold, heading hold, or turn coordination. TRIM FAIL – Caution Indicates the yaw, roll, or pitch trim actuators are not responding accurately to computer signals.

Indicates that the FD was coupled but automatically decoupled due to a failure in the FD or related subsystems. SAS DEGRADED – Advisory Indicates that the pitch, roll, and/or yaw SAS channels from SAS 1 and/or SAS 2 has disengaged due to a fault. STAB DEGRADED – Advisory Indicates that the stabilator control has reverted to a condition of reduced speed and range functionality.

Coupled FD Annunciation A coupled FD annunciation indicates which FD the autopilot is currently coupled to. The green annunciation indicates an “engaged” (coupled) condition. The annunciation is displayed when the autopilot is coupled to an FD channel. The annunciation is removed when either the autopilot is not coupled to the FD or the information coming from the FD does not allow the coupled FD to be determined. REF SEL LOC

CAP REF ADJ GA

GS CAP

RALT ----

VS ----

ALT ----

IAS ----

HDG 240

PUSH-SYNC

PUSH-SYNC

DECL CAP

CPLD **** HVR

PUSH-SYNC

PUSH-SYNC

PUSH-SYNC

FD/DCP AFCS NOTE Due to software unavailability, most of the pushbuttons on the FDDCP will not be functional at first flight.

STAB UNLOCKED – Caution Indicates that the stabilator is in the folded position. STAB MANUAL MODE – Caution Indicates that the stabilator system is powered up but in manual mode. FD FAIL – Caution Indicates that one or more of the FD functions has disengaged due to a failure in the modes. FD COUPLE FAIL – Caution

9/3/2003 Page 10 of 20 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Flight Management System Status And Test Indications The FMS provides status and test status via ARINC-429 data bus from the AFCS. STATUS Pressing the STS button on the FMS and then pressing softkey (SK)-8 access the status page for the AFCS. When the Status page appears, an overall status of the AFCS will be displayed. Status for FCC-1 or FCC-2 can be selected from this page. All faults and failures will be displayed under the selected FCC if a fault or failure should occur.

AFCS STATUS TEST Pressing the TST button on the FMS and then pressing SK-8 access the test for the AFCS. When the Test page appears, an overall status of the AFCS will be displayed. Test for FCC-1 or FCC-2 can be selected from this page. An Initiated Built-In Test (IBIT) can be initiated from either FCC. All faults and failures will be displayed under the selected FCC if a fault or failure should occur.

MAIN STATUS PAGE

MAIN TEST PAGE

9/3/2003 Page 11 of 20 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

• •

•

Collective position sensors detect pilot collective displacement, which results in a change in stabilator angle to counteract the pitch changes. Provide pitch rate feedback to improve dynamic stability.

AFCS TEST

Automatic Flight Control System Operation Stabilator The primary function of the stabilator is to provide stability in the pitch axis in forward flight. The helicopter has a variable angle of incidence stabilator to enhance handling qualities. The stabilator is programmed to: •

• •

Align stabilator and main rotor downwash in low speed flight to minimize nose-up attitude resulting from downwash. Decreases angle of incidence with increased airspeed to improve static stability. Provide collective coupling to minimize pitch attitude excursions due to collective inputs from the pilot.

9/3/2003 Page 12 of 20 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

•

essential bus and No. 2 ac primary bus through circuit breakers marked STAB CONTR.

Provide sideslip to pitch coupling to reduce susceptibility to gusts.

Stabilator Actuators.

Stabilator/Flight Control Panel

The above features are provided via inputs to dual actuators, which position the stabilator. Failure of one actuator will restrict total

The stabilator flight control panel, on the lower center console, provides electrical control of the stabilator system. A slew switch on the

#1 ADC

EGI (Pitch Rate)

Motor Driver

Auto Mode Schedule

ARINC 429

* Not

All Criteria Shown

ARINC 429

Fault Detection

Cltv. Stick Posn

28 Vdc PWM (6A stall)

Auto/ Manual Logic*

Lateral Accel

Airspeed Stab. Angle

Independent Monitor

Each actuator is rate limited 0.7 inch/s (? 4°/sec. of Stab motion)

#1 Stabilator Actuator M

Stabilator Angle Transducer

Cross Channel Data Link

#1 Stab. Actuator Position Synchro

STAB IND

Cyclic Slew Switch Manual Slew Switch Up

DCU

Down #2 Stab. Actuator Position

M

MFD

#2 ADC Other Sensors

#2 Stabilator Actuator ARINC 429

#2 FCC #2 FCC has same functionality as #1 FCC, detail not shown for clarity

maximum movement of the stabilator to about 35° if failure occurs full down, or about 30° if failure occurs full up. The stabilator actuators receive power from the dc essential bus and No. 2 dc primary bus through circuit breakers marked STAB PWR. Since the battery powers the dc essential bus, it is possible to manually slew one actuator using battery power only. If the stabilator is slewed up, regain automatic control by manually slewing stabilator full down, then push AUTO CONTROL RESET twice. Otherwise, when only one actuator is slewed, it causes a very large mismatch between the two actuator positions. This is detected by the fault monitor and shuts down the automatic mode upon attempted engagement. Automatic control function sensors, airspeed sensors, pitch rate gyros, collective position sensor, and lateral accelerometer receive power from the ac

Stabilator Flight Control Panel is a single switch with redundant poles for both up and down control. The pilot and copilot each have a cyclic mounted slew switch that can command the stabilator in the up direction only. The No. 1 side of the switches is sent to FCC No. 1, and similar on the No. 2 side. Each FCC determines the correct mode for stabilator control (Auto or Manual). In manual mode, the FCC reads the status of the slew switches and commands its respective actuator. The actuators are electrically independent and mounted back to back to control the position of the stabilator. Each actuator gets its power from its respective FCC. Each channel required for manual control of the stabilator is powered from independent electrical sources. 9/3/2003 Page 13 of 20

UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

The Stabilator control panel on the lower console provides electrical control of the stabilator system. The panel contains a MAN SLEW switch, a TEST button, and AUTO CONTROL RESET switch with a push-to-reset feature.

WOW. When pressed, control of the stabilator should go to the manual mode. Automatic Mode The automatic mode of operation positions the stabilator to the best angle of attack for the existing flight conditions.

Manual operation is also restricted to these limits. If a malfunction occurs in the automatic Automatic mode will engage when power is mode, the system will switch to manual, ON will applied to the FCCs. The pilot can also press go off in the AUTO CONTROL window, and the the reset pushbutton on the stabilator flight STABILATOR and MASTER CAUTION PRESS control panel. Once TO RESET cautions engaged, each will appear and a STABILATOR CONTROL computer will use its beeping tone will be respective sensors to heard in the pilot’s AUTO MAN SLEW CONTROL UP compute a desired and copilot’s TEST R O angle of incidence. headphones. It may E F Each FCC then be possible to regain ON S F positions its respective the auto mode by DOWN E actuator to achieve the pressing the AUTO T AUTO FLIGHT CONTROL desired angle providing CONTROL RESET. If TRIM SAS 1 SAS 2 FPS that the other side the automatic mode actuator will be is regained, ON will extended just the same. appear in the AUTO ON ON ON ON CONTROL switch The automatic mode will window and the SAS/BOOST FAILURE RESET allow the stabilator to be cautions and automatically operated MASTER CAUTION CPTR CPTR ON from about 39° trailing PRESS TO RESET 2 1 edge down to 9° trailing cautions will edge up. disappear. The stabilator automatic mode is held in the Limited Automatic energized state within Mode the stabilator control UMAV003_8 In the event that a amplifier. failure occurs that prevents one FCC from positioning its respective actuator, the system On certain occasions during interruption of dc can revert from full Auto mode to Limited Auto power, such as switching of generators, it is Mode. The functioning Limited Auto Mode possible to have conditions where the stabilator channel uses the location of the failed channel’s automatic mode may shut down. If the automatic actuator to compute where it should command mode shuts down during flight because of an ac its own actuator. The stabilator will only be able power failure, the helicopter shall be slowed to to achieve half of its normal maximum rate 80 Knots Indicated Air Speed (KIAS) before because only one actuator is active. Available power is restored. In this case the AUTO range will depend on the position of the actuator CONTROL RESET switch may be pressed to of the failed channel. reengage the auto mode. If the automatic mode is not regained, the MASTER CAUTION PRESS TO RESET must be reset, which turns off the Manual Mode beeping tone, and the stabilator controlled Manual mode is provided in the event that, Auto throughout its range with the MAN SLEW switch. mode fail, if the pilot needs to reposition the When initial power is applied to the stabilator stabilator following a failure or if the pilot system, it will be in automatic mode. chooses to override Auto mode. The TEST switch is used to check the AUTO mode fault detector feature and is disabled by

9/3/2003 Page 14 of 20 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

In Manual mode, the stabilator will only move in response to direct pilot input. The pilot can slew the stabilator up or down using the slew switch located on the stabilator flight control panel in the center console or it can be slewed up using the cyclic slew switch located behind each cyclic grip. The stabilator can be slewed at full rate throughout its full range presuming both actuators are functioning.

Stability Augmentation System (SAS) The SAS enhances dynamic stability in the pitch, roll, and yaw axes. In addition, both SAS 1 and SAS 2 enhance turn coordination by deriving commands from lateral accelerometers which together with roll rate signals are sent to their respective yaw channels automatically at airspeeds greater than 60 knots (SAS 1)/50 knots (SAS 2). The SAS 1 operates on 28 vdc power from the dc essential bus through a circuit breaker marked SAS BOOST. Pitch, roll, and yaw rate; pitch attitude and roll attitude, are provided by EGIs. Longitudinal and lateral accelerations are provided by individual accelerometers in each axis. Loss of ac power to the EGI or SAS amplifier can cause erratic operation of SAS 1 due to loss of the reference for the ac demodulators. When this condition is encountered, the pilot must manually disengage SAS 1. If the malfunction is continuous, SAS 2 may be turned off. With SAS 1 or SAS 2 off, the control authority of SAS is reduced by one-half (5% control authority). Malfunction of the SAS 1 system may be detected by the pilot as an erratic motion in the helicopter without a corresponding failure advisory indication. If a malfunction is experienced, SAS 1 should be turned off. SAS actuator hydraulic pressure is monitored. In case of loss of actuator pressure, or if both SAS 1 and SAS 2 are off, the SAS OFF caution will appear on the MFD.

Trim When the TRIM is engaged on the flight control panel, the pitch, roll, collective, and yaw trim systems are activated to maintain position of the cyclic, collective, and tail rotor controls. Proper operation of the yaw trim requires that the SAS/BOOST on the flight control panel be on. The tail rotor and lateral forces are developed in the electromechanical yaw and roll trim actuators. Both yaw and roll trim actuators incorporate slip clutches to allow pilot and

copilot control inputs if either actuator should jam. The forces required to break through the clutch are 80 pounds maximum in yaw and 13 pounds maximum in roll. Electrohydromechanical actuator operated in conjunction with the FCC develops the longitudinal forces. When the pilot applies a longitudinal or lateral force to the cyclic stick with trim engaged, a combination detent and gradient force is felt. The pilot may remove the force by pressing the thumb-operated TRIM REL switch on the pilot/copilot cyclic grip. It should be noted that the trim actuators have only one input and thus, can only be driven by FCC #2. If that FCC should fail, the trim function is lost. The pedal gradient maintains pedal position whenever the trim is engaged. Below 50 KIAS, place feet on the pedals to press pedal switches and remove the gradient force (above 50 KIAS, the TRIM REL switch must also be pressed). The pedals may then be moved to the desired position and released. The pedals will be held at this position by the trim gradient. The pedal trim actuator also includes a pedal damper. The pedal damper is engaged continuously, independent of electric power and the TRIM switch on the flight control panel. The FCC continuously monitors operation of the trim system. In addition to the TRIM REL switch, a five-way trim switch on each cyclic stick establishes a trim position without releasing trim. With trim engaged, the trim position is moved in the direction of switch movement. The trim switch in one direction at a time moves the cyclic. When FPS is engaged, the TRIM switch changes the pitch and roll attitude reference instead of the cyclic stick position reference. The trim system release feature permits the pilot or copilot to fly the helicopter with light stick forces. The push-on/push-off TRIM switch on the flight control panel or the TRIM REL switches on the pilot/copilot cyclic and collective grips may be used to release trim. When the switch is on, the trim system provides gradient and detent holding force for pitch, roll, collective, and yaw. When turned off, the trim system is released and light cyclic and collective control forces are present.

9/3/2003 Page 15 of 20 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Flight Path Stabilization (FPS) The FPS function maintains helicopter pitch and roll attitudes, airspeed, and heading during cruise flight, and provides a coordinated turn feature at airspeeds >50 knots (kts). First engaging the control panel TRIM switch and at least one SAS switch and then pressing the control panel FPS switch engage the FPS function. When engaged, the switch legend ON is lit. The FCCs provide FPS command signals to the pitch trim servo, roll trim actuator, and yaw trim actuator, which in turn reposition the flight controls. The FD only repositions collective trim. The AFCS also provides command signals to the trim actuators to reposition the flight controls using the trim system. Proper FPS operation requires that the SAS/BOOST, TRIM, and SAS 1 and/or SAS 2 functions have been selected on the flight control panel. Although not required for proper operation, the FPS performance will be improved by the proper operation of the stabilator in the automatic mode. To use the FPS features, the pilot first assures that SAS/BOOST, SAS and TRIM are on and operating, and then turns the FPS switch on. The desired pitch and roll attitude of the helicopter may be established in one of these ways: 1. Pressing the STICK TRIM switch to slew the reference attitude to the desired attitude. 2. Pressing the TRIM REL switch on the pilot/copilot cyclic grip, manually flying the helicopter to the desired trim condition, and releasing the TRIM REL switch. 3. Overriding the stick trim forces to establish the desired trim condition, and then neutralizing stick forces by means of the TRIM switch. The trim attitude, once established, will be automatically held until changed by the pilot. At airspeeds greater than 50 knots, the pitch axis seeks to maintain the airspeed at which the trim is established by variation of pitch attitude. When pitch attitude is changed by means of the STICK TRIM switch, there is a delay from the

time that the STICK TRIM switch input is removed until the new reference airspeed is acquired. This is to allow time for the helicopter to accelerate or decelerate to the new trim speed. The yaw axis of the FPS provides heading hold at airspeeds less than 50 knots and heading hold or turn coordination at airspeeds greater than 50 knots. For heading hold operation at airspeeds less than 50 knots, the helicopter is maneuvered to the desired heading with feet on pedals. When trimmed at the desired heading, the pilot may remove feet from pedals, at which time the existing heading becomes the reference, which is automatically held. To change heading, the pilot may activate one or both pedal switches, trim up on the desired heading and remove feet from pedals. At airspeeds greater than 50 kts, heading hold will be automatically disengaged, and coordinated turn engaged under these conditions: • • •

STICK TRIM switches actuated in the lateral direction. TRIM REL switch is pressed and roll attitude is greater than 2.5° About ½ inch cyclic displacement and a roll attitude of about 1.0°. Heading hold is automatically reengaged and turn coordination is disengaged upon recovery from the turn when the lateral stick force, roll attitude, and yaw rate are within prescribed limits.

To make a coordinated turn, the pilot enters a turn in one of these ways: • • •

•

Changing reference roll attitude by pressing the STICK TRIM switch in the desired lateral direction. Pressing TRIM REL switch on the cyclic grip and establishing the desired bank angle with feet off pedal switches. Exerting a lateral force on the cyclic stick to achieve the desired bank angle, and then neutralizing the force with the STICK TRIM switch. Keeping a lateral force on the cyclic stick for the duration of the turn.

The FPS monitoring is automatic. If a malfunction is detected, the AFCS FAIL caution appears on the MFDs and the FPS will either continue to operate in a degraded mode, such as without heading hold or without attitude hold;

9/3/2003 Page 16 of 20 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

or may cease to function altogether. The pilot must take over manual flight of the helicopter and may either turn the mode off or evaluate performance to determine the degree and type of degradation and continue flight with the remaining features. To help evaluate the nature of the degradation, FCC alerts report status on the AFCS STATUS page in the Flight Management System (FMS). These tell the pilot which mode has experienced the failure. If an FCC light illuminates, the pilot may attempt to clear the indication of temporary malfunctions by pressing one of the switches. If the AFCS DEGRADED caution disappears, it may be assumed that normal operation is restored. Attitude Hold. Attitude hold is engaged with FPS at all airspeeds, and is commanded by changing the cyclic stick position with the stick TRIM switch. This causes the cyclic stick to move and the helicopter attitude to change approximately 5° per second in pitch and approximately 6° per second in roll. When cyclic movement is stopped, the autopilot stabilizes the helicopter around the new attitude. The roll channel autopilot holds roll attitude of the helicopter. Attitude information is supplied to the computer from the No. 1 and No. 2 EGIs. Heading Hold. The yaw channel of the autopilot provides the heading hold feature for hover and forward flight and is engaged when the AUTO PLT is illuminated. Heading hold is an outer-loop function operating through the yaw trim actuator, and therefore will only be operational when the yaw trim is engaged. Releasing all pedal switches at a given heading synchronizes the trim system to the established heading. The yaw autopilot also uses a collective stick position sensor to hold reference heading for yaw excursions caused by main rotor torque changes. The heading hold is re-engaged following a turn when the following conditions are maintained for 2 seconds: • •

Aircraft roll attitude is within 2° of wings level. Yaw rate is less than 2° per second. The weight-on-wheels (WOW) switch disengages the heading hold when the aircraft is on the ground. During coupled flight operation above 50 KIAS, heading

hold is achieved by the use of the roll axis. Turn Coordination. Automatic turn coordination is provided at airspeeds greater than 50 knots. Turn coordination allows the pilot to fly a coordinated turn with directional control provided by the AFCS. The AFCS uses lateral acceleration and roll rate to determine if the aircraft is out of balanced flight, and provides the yaw SAS and yaw trim with the inputs necessary to maintain an automatic coordinated turn. Automatic turn coordination is engaged and heading hold disengaged when roll attitude is greater than 1° and any of the following conditions exists: • • •

Lateral cyclic greater than 1.3 inch stick displacement from trim. Cyclic trim release is pressed. Roll attitude is beeped beyond 2.5° bank angle. Actuation of the HDG TRIM switch above 50 KIAS uncoupled in the lateral axis for more than one second provides a 1° per second coordinated turn.

Radar Altimeter (RADALT) Hold When RADALT hold is selected and engaged, the RADALT test function is still available. Selecting RADALT test causes the RADALT hold to be disengaged. When radar altitude hold mode is selected, the AFCS computer uses as a reference altitude the existing altitude from the radar altimeter. The computer commands both the collective SAS actuator and the collective trim actuator to maintain the reference altitude. The SAS actuator provides fast response, limited authority corrections, and the trim actuator provides a limited response (rate limited), full authority corrections. The AFCS computer uses altitude and rate from the radar altitude system and vertical acceleration to command the collective SAS and trim actuators. The computer also monitors engine torque to prevent dual engine torque from exceeding 140% when in RADALT hold mode. Radar altitude hold is engaged at any altitude from 0 to 1,500 feet Above Ground Level (AGL) and at any airspeed by pressing the RADALT with SAS and FPS engaged. Manually selecting the RADALT via the on-side FD/DCP 9/3/2003 Page 17 of 20

UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

will disengage the RADALT mode. Manually selecting the Remote Standby (RMT SBY) on the cyclic stick will also disengage the RADALT mode. If the RADALT mode is engaged on the ground, the RADALT mode will automatically disengage when the aircraft is in flight (weight off wheels). The RADALT hold mode uses the output of the RADALT to provide altitude correction signals. Engagement of the RADALT switch causes the selected FCC to hold the helicopter at the radar altitude engage point. To change to a new altitude, the collective stick TRIM RLSE switch must be pressed, the flight controls repositioned to change altitude, and then TRIM RLSE is released at the new altitude. The RADALT legend will flash as long as the TRIM RLSE switch is pressed. When the switch is released, the radar altitude mode is automatically reengaged and the legend will go on steady. Additionally, using the AFCS COLL switch on either pilot’s collective stick can change altitude. The rate of change is 4 feet per second (for the first 5 seconds of beep), then 16 feet per second (for longer than 5 second beep). If the RADALT becomes unreliable, the selected FCC will automatically disengage the radar altimeter mode. The RADALT mode may be used at any altitude below 1,500 feet. NOTE

Flight Director/Display Control Panel (FD/DCP) NOTE Due to software unavailability most of the pushbuttons on the FDDCP will not be functional at first flight. This is for informational purposes only. All FD modes are engaged via the FD/DCP. The FD provides pitch, roll, and collective steering cues on the PFD. In addition, any FD function can be coupled to the flight controls by pressing the “CPL” button on the FD/DCP. Except where otherwise noted, all FD functions are available at airspeeds greater than 50 kts. The FD has the following modes: Airspeed Hold (AS) AS hold is automatically engaged when transitioning from below to above 50 kts. Engagement status is displayed on both the PFD and FD/DCP. The mode will maintain an airspeed reference that is set when the mode is engaged. The reference can be changed using the cyclic beeper or the airspeed knob on the FD/DCP. The reference is displayed on both the PFD and FD/DCP. Barometric Altitude (BARALT) Hold

The pitch and roll cues on the MFD will be removed from view and it is recommended the pilot utilize the MFD hover display velocity bars for situational awareness. If the helicopter is flying with a tailwind, the airspeed can be less than 50 KIAS while the groundspeed is greater than 60 knots. Under these conditions if the TRIM REL switch is pressed while in the CLIMB mode, the helicopter will go into the VHLD mode and decelerate to 60 knots groundspeed, which is the maximum that can be commanded. Under these conditions, pitch up and down of helicopter attitude is normal while the helicopter stops accelerating, and then decelerates and captures 60 knots.

BARALT Hold engagement status is displayed on both the PFD and the FD/DCP. The mode will maintain an altitude reference that is set when the mode is engaged. The reference can be changed using ALT knob on the FD/DCP. The reference is displayed on both the PFD and FD/DCP. Radar Altitude Hold (RALT) RADALT Hold engagement status is displayed on both the PFD and the FD/DCP. The mode will maintain an altitude reference that is set when the mode is engaged. The reference can be changed using the collective beeper or the RADALT knob on the FD/DCP. The reference is displayed on both the PFD and FD/DCP. RADALT hold will automatically engage when hover hold is engaged.

9/3/2003 Page 18 of 20 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

maintain the heading indicated by the HDG Bug displayed on the PFD. The HDG Bug can be changed using the HDG knob on the FD/DCP. The reference value is displayed on both the PFD and the FD/DCP.

Vertical Speed Hold (VS) VS Hold engagement status is displayed on both the PFD and the FD/DCP. The mode will maintain a VS reference that is set when the mode is engaged. The reference can be changed using the collective beeper or the VS knob on the FD/DCP. The reference is displayed on both the PFD and FD/DCP.

VHF Omni Range (VOR) VOR engagement status is displayed on both the PFD and the FD/DCP. When the mode is first engaged, the engagement status is ARMed. When radial selected via Omni Bearing Select

Barometric Altitude Preselect (ALTP)

Prov. Long. Accel #1

Cltv. Stick #1

Lat. Accel #1

P. Rate Gyro #1

DCU

CDU

ADC

EGI

#1

#1

#1

#1

FD/ DCP #1

Copilot’s Inboard MFD

Copilot’s Outboard MFD

ARINC 429 ANALOG/DISCRETE

Stabilator Angle/Limit Switch Assy.

P. SAS Servo R. SAS Servo

# 1 FCC

Stabilator

Y. SAS Servo

Actuator

Stab/

#1

Flight Control Panel

Stabilator Actuator

# 2 FCC

#2

P. Trim R. Trim

Back Up

Y. Trim

Instrument

C. Trim

Radar Alt.

Cltv. Stick #2

Vert. Accel

Long. Accel #2

Lat. Accel #2

P. Rate Gyro #2

DCU

CDU

ADC

EGI

#2

#2

#2

#2

FD/ DCP #2

Pilot’s Inboard MFD

Pilot’s Outboard MFD

Prov.

ALTP engagement status is displayed on both the FD and the FD/DCP. When the mode is first engaged the engagement status is ARMed. When current altitude is within 10 ft of the target altitude, the status changes to CAPtured. When current altitude is within 10 ft of the target altitude, the mode automatically disengages and ALT engages. The reference can be changed using the ALTP knob on the FD/DCP. The reference is displayed on both PFD and the FD/DCP.

(OBS) is captured, the status changes to CAP. Instrument Landing System (ILS) Localizer (LOC) ILS LOC engagement status is displayed on both the PFD and the FD/DCP. When the mode is first engaged, the engagement status is ARMed. When ILS LOC is captured, the status changes to CAP. ILS Back Course (BC)

Heading (HDG) Select HDG Select engagement status is displayed on both the PFD and the FD/DCP. The mode will

ILS BC engagement status is displayed on both the PFD and the FD/DCP. When the mode is first engaged, the engagement status is ARMed. 9/3/2003 Page 19 of 20

UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

When ILS LOC is captured, the status changes to CAP. ILS Glide Slope (GS) GS engagement status is displayed on both the PFD and the FD/DCP. When the mode is first engaged, the engagement status is ARMed. When GS is captured, the status changes to CAP. ILS Deceleration (DCEL) DCEL engagement status is displayed on both the PFD and FD/DCP. DCEL mode will program airspeed to achieve 70 kts at 200 ft. Go Around (GA) GA engagement status is displayed on both the PFD and the FD/DCP. GA will command a climb of 750 ft per minute and airspeed of 70 kts. GA can be engaged at any airspeed below 70 kts. FMS Long Range Navigation (LNAV) LNAV engagement status is displayed on both the PFD and the FD/DCP. The mode provides a roll steering to maintain the FMS flight plan. FMS Vertical Navigation (VNAV) VNAV engagement status is displayed on both the PFD and FD/DCP. The mode provides collective stick commands to maintain a vertical flight plan during approach phase only (pseudo GS). Hover (HVR) Hold HVR engagement status is displayed on both the PFD and the FD/DCP. HVR Hold has two submodes of operation: position hold and velocity hold. When position hold is engaged, a position over the surface is maintained. Velocity HVR maintains a constant groundspeed over the surface. If the HVR button is pressed on the FD/DCP, Velocity HVR will be engaged at the current speed unless that speed is less than the threshold in which case Position HVR will be engaged. If the cyclic beeper center position is activated and the engagement criteria are met, Position HVR will be engaged.

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Section 2-6 Powertrain System Differences Description The principle components and configuration of the powertrain system are unchanged for the UH-60A/L. The modular composition and installation of the main transmission, input modules, and accessory drive modules are unchanged. The powertrain consists of inputs from two engines, a main transmission, intermediate gear box (IGB), tail gear box (TGB), and connecting drive shafting. Power from the engines is transmitted to the main transmission module through input modules. The main transmission is mounted on top of the cabin between the two engines. It mounts and powers the main rotor head, changes the angle of drive from the engines, reduces rpm from the engines, powers the tail rotor drive shaft, and drives the accessory module. The main transmission consists of five

Upgraded Input Module

modules: two input modules, the main module, and two accessory modules. The main transmission has a built-in 3° forward tilt. Power to drive the tail rotor system is transmitted via shafting from the main transmission tail take-off, through the main transmission oil-cooler assembly to the IGB. The IGB changes the angle of torque and routes it to the tail gearbox where the angle is changed again and drives the tail rotor. The UH-60M powertrain system modifications impact the following components: • • • • • • •

Main transmission housing Shorter section 1 drive shaft Longer tail takeoff gear and flange Larger TGB oil site gauge Chip detector retention clips Improved chip detector system fault monitoring Powertrain system indications

Pave Hawk Derived Improved Durability Main Module Rotor Brake Provisions

Upgraded Accessory Module

Upgraded Tail Gearbox

Upgraded Intermediate Gearbox New Sight Gage

New Section 1 Drive Shaft

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Powertrain System Components/Subsystems Main Transmission Assembly

The updated equipment includes hydraulic pumps, transmission sensors and wiring harnesses. The transmission assembly also includes a longer tail take off and an offset gust lock flange to allow for the installation of a rotor brake if mission requirements dictate.

The main module contains the necessary gearing to drive the main rotor and tail rotor systems. It provides a reduction in speed from the input module to the main module and the tail drive shaft. The main transmission assembly is replaced with an improved durability gearbox (IDGB) as part of the UH-60M modernization program. The IDG installed on the aircraft incorporate the latest versions of all external and internal transmission components.

Pavehawk Derived Improved Durability Main Module

Upgraded Input Module

Upgraded Accessory Module

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Rotor Brake Provisions

Input Module The input modules are mounted on the left and right front of the main module and support the front of the engines. They contain an input bevel pinion and gear and a freewheel unit. The freewheel unit allows engine disengagement during autorotation, or in case of a nonoperating engine, the accessory module will continue to be driven by the main rotor. The input module provides the first gear reduction between engine and main module. The input modules installed on UH-60M configured aircraft incorporate equipment to make the input module consistent with the UH60L Lot 21 configuration. The upgrade installs the latest version of the accessory drive gear as well as features that enhance the input modules heat resistance and lubrication system.

additional rotor speed sensor is mounted on the left accessory module, which provides input signals to the DEC for improved transient droop response. Like the main and accessory modules of the transmission assembly, the accessory module configuration match the UH-60L Lot 21 standard. This includes the installation of an improved VESPEL spline adapter for the main generator as well as an improved gear with associated spacers.

Main Transmission Lubrication System The lubrication system of the UH-60M transmission system is physically and functionally identical to the system found on the UH-60L. Transmission Oil Cooler

Accessory Module

The main transmission cooling system consists of an externally mounted oil cooler and mechanically driven blower assembly. Oil lines connect the main transmission to the oil cooler.

Main Transmission Chip Detector System The UH-60M transmission chip detector is configured and functions identically to the UH60A/L system. The transmission chip detector system consists of chip detectors on the left and right input modules, left and right accessory modules, as well as the main gearbox module (MGB). The detectors provide warning of chips in any of five areas of the main transmission system. Accessory Module One accessory module is mounted on the forward section of each input module. Each accessory module provides mounting and drive for an electrical generator and a hydraulic pump package. A rotor speed sensor is mounted on the right accessory module and provides signals for the Engine Indication Crew Alerting System (EICAS) and PFD power pod displays. An

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Tail Drive System Six sections of drive shaft connect the main module to the tail rotor gear box. The shafts drive the oil cooler blower and transmit torque to the tail rotor. Each shaft is dynamically balanced tubular aluminum. Multiple disc (flexible) couplings between sections eliminate universal

UH-60A/L Tail Take Off

joints. The shafts are ballistically tolerant if hit by a projectile and are suspended at four points in viscous-damped bearings mounted in adjustable plates and bolted to fuselage support brackets.

UH-60M Tail Take Off

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No. 1 Tail Drive Shaft Drive shaft segment No. 1 will be recognized to be different from the normal configuration of the tail drive shafting found on UH-60A/L aircraft. The shorter No. 1 drive shaft segment is intended to allow for the installation of a rotor brake assembly.

The IGB, when refurbished, is converted to a Lot 21 standard. The upgrade process incorporates improved chip detectors as well as the latest version of the intermediate gearbox oil sight gauge.

Tail Gear Box The oil-lubricated TGB at the top of the tail pylon transmits torque to the tail rotor head. The TGB changes angle of drive and gives a gear reduction and mounts the tail rotor system. It also enables pitch changes of the tail rotor blades through the flight control system. The gearbox housing is magnesium. The TGB may run at cruise flight for 30 minutes with loss of all oil.

UH-60M No. 1 Tail Drive Shaft

Intermediate Gear Box Mounted at the base of the pylon is the oillubricated IGB. It transmits torque and reduces shaft speed from the main gearbox to the TGB. The IGB may run at cruise flight for 30 minutes, with loss of all oil.

FR

ON

T

The TGB, when it’s refurbishment process is completed, is upgraded to the Lot 21 standard. The most noticeable change to the tail gearbox is the larger oil sight gauge. This sight gauge has been used on Naval hawk aircraft for some time. Also upgraded are the chip detectors as the use of new manufacturing processes and components that enhance the tail gearboxes reliability.

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Oil Sight Gage

Intermediate and Tail Gear Box Chip/Temperature Sensors The intermediate and tail gear boxes are equipped with identical chip/temperature sensors that provide a cockpit indication when the gear box temperature is too high, or a chip is present. The chip detectors incorporate a fuzz burn-off feature that eliminates false warning due to fuzz and small particles. When a chip is detected and will not burn off, the INT XMSN CHIP or TAIL XMSN CHIP caution will appear. The oil temperature sensor is a bimetal strip that reacts to temperatures. When the oil temperature reaches 140°C, the INT XMSN OIL HOT or TAIL XMSN OIL HOT caution will appear. Power to operate the oil temperature system is from the No. 2 dc primary bus through a circuit breaker marked MAIN XMSN.

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Powertrain System Indications Flight Display System Transmission Indications The Flight Display system provides transmission system indications on both the Engine Instrument Crew Alerting System (EICAS) Display and the Primary Flight Display.

above the readout defining the parameter the indicator represents. • •

T – Main transmission oil temperature (MTOT) P – Main transmission oil pressure (MTOP)

EICAS Data Display – Transmission Indications

The readouts provide a numeric indication of the current state. The range of the readouts varies with the lower limit typically being 0 and the upper limit typically being the maximum value the input data can assume.

The transmission indications consist of a tape to provide information and readout above the tape to provide precise values. A label is displayed

The readout color is determined by the displayed readout value to the operating range/limits defined for that parameter. • • •

Green – “Normal” operating region Y Y w Yeeellllllooow w – “Cautionary” operating region Red – “Warning” operating region

Transmission Indications Icon

UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

9/3/2003 Page 7 of 12

Check EICAS Caution Indication The DCU will illuminate the CHECK EICAS caution when the main transmission oil temperature or pressure parameters exceed or decrease below the maximum or minimum continuous limits OR the cross channel DCU indicates the Check EICAS Caution is true. The parameters are: •

Main Transmission Oil Temperature has reached or exceeded its maximum continuous limit of 105° C

•

Main Transmission Oil Pressure has reached or exceeded its maximum continuous limit of 65 PSI

•

Main Transmission Oil Pressure has reached or decreased below its minimum continuous limit of 30 PSI

Maximum Continuous Line The EICAS display draws a yellow line across the main transmission icon to facilitate quick interpretation of the transmission exceedance status. This line represents the maximum continuous operating limit for each of the parameters represented by the tapes. Under normal operating conditions all parameters should be on or below the line. If a parameter exceeds the maximum continuous limit, its tape and readout will turn yellow.

Caution/Advisory Indications The cautions and advisories provided by the Data Concentrator Units/Engine Indication Crew Alerting System (EICAS) are similar if not identical to UH-60A/L caution/advisory panel indications. In some case there have been minor syntax changes. They are listed here for general informational purposes. The powertrain system Caution/Advisory Indications are as follows: MAIN XMSN OIL HOT - Caution Indicates that main transmission oil temperature is greater than 140°C.

Transmission Indication Miscompare The indicator label is normally white. When a miscompare condition exists, the label is displayed as a yellow flag (black text on yellow background).

MAIN XMSN PRESS - Caution Indicates when oil pressure in the main transmission is less than 21 PSI. The indication will be cleared when oil pressure rises to above 22 PSI. TAIL XMSN OIL HOT - Caution Indicates that the tail transmission oil temperature is >284°F ±15°F.

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INT XMSN OIL HOT - Caution Indicates that the intermediate transmission oil temperature is >284°F ±15°F. CHIP IBIT FAIL - Caution Indicates that the DCU has detected a failed status for any chip detector.

Detector 2. This menu selection allows the operator to view the status of the detectors. All faults and failures will be displayed under the Chip Detector status page if a fault or failure should occur.

CHIP INT XMSN - Caution Indicates that a chip has been detected in the intermediate transmission. CHIP L ACC MDL - Caution Indicates that a chip has been detected in the left hand accessory module. CHIP L INPUT MDL - Caution Indicates that a chip has been detected in the left hand input module sump. CHIP MAIN MDL SUMP - Caution Indicates that a chip has been detected in the main module sump. CHIP R ACC MDL - Caution Indicates that a chip has been detected in the right hand accessory module.

AIR VEHICLE STATUS PAGE

CHIP R INPUT MDL - Caution Indicates that a chip has been detected in the right hand input module. CHIP TAIL XMSN – Caution Indicates that a chip has been detected in the tail transmission. GUST LOCK ENGAGED - Caution Indicates that the gust lock is engaged.

Flight Management System Status And Test Indications Status The FMS monitors the powertrain chip detector system. The FMS STS fixed function key is used to access the MAIN STATUS page to check the status of the Chip Detectors. Selecting soft key (SK) 4 on the MAIN STATUS PAGE accesses the AIR VEHICLE status page. Selecting SK 2 or SK 7 accesses air vehicle equipment status for Chip Detector 1 and Chip

CHIP DET STATUS PAGE

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Chip Detector Initiated Built In Test (IBIT) The chip detector can be tested as needed by initiating a test by accessing the test modes located on the Air Vehicle Test page of the FMS. The Test Menu is accessed by the function key and provides the ability to conduct and review tests of system line replaceable units (LRU’s). Selecting soft key (SK) 4 on the MAIN TEST PAGE accesses the AIR VEHICLE Test page. Selecting SK 2 or SK 7 accesses air vehicle equipment test for CHIP DET1 and CHIP DET 2. These menus allow an IBIT of the separate chip detector circuits. All faults and failures will be displayed under the DEC if a fault or failure should occur.

CHIP DETECTOR TEST PAGE Chip Detector System Continuous Built In Test (CBIT) The chip detector system BIT performs an automatic continuous test, controlled by the Master DCU of the following chip detectors:

AIR VEHICLE TEST PAGE

• • • • •

Left and right input modules, Left and right accessory modules, Intermediate transmission, Tail transmission, and Main module.

When ac power is applied to the system, the DCU will check for valid signals to the chip detectors. In the test mode, chip cautions are not indicated. If all chip detectors pass, no caution indications will appear. If one or more chip detectors fail the test, a CHIP IBIT FAIL caution indication will appear on the EICAS and the faults detected will be displayed on the DCHIP DET IBIT screen. In order to determine the specific failed chip detector, a chip detector IBIT must be performed. Test bit codes are displayed after

9/3/2003 Page 10 of 12 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

the IBIT test is completed (approximately two minutes). If a Chip Detector IBIT has already been selected and is in progress and a second test of the other chip detector system is selected, the second of the two tests will complete and display results.

Engine / Fuel Drive Train A429 Fault Miscompare

Engine / Drive Train / Fuel Instrument CPCI Engine / Drive Train / Fuel

A429 Fault

Power Pod & EICAS Display

Miscompare Condition

ARINC 429 Receiver

MFD Miscompare CPCI

FMS-1 Reversion Switch Panel

ARINC 429

DCU-1

DEC Fault Codes Chip Detector IBIT DCU Status AVCS Status Fuel Quantity Discrete

DCU-2

ARINC 429 DEC Fault Codes Chip Detector IBIT DCU Status AVCS Status Fuel Quantity

FMS-2

Analog Transmission Indications Block Diagram Engine / Drive Train / Fuel Chip Power

Discrete

Lamp Test

Aircraft Engine / Drive Train / Fuel

9/3/2003 Page 11 of 12 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

NOTES

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Section 2-7 Rotor System Differences Description The rotor system consists of a main rotor and tail rotor. Both main and tail rotor systems are driven by the engines through the transmission system, with pitch controlled by the flight control system. The main rotor system transmits pilot or automatic flight control system inputs to the main rotor blades. Forces developed by the blades are transmitted to the aircraft via the main rotor system. The tail rotor system

produces the forces necessary to counteract the torque generated by the main rotor system and provides directional control. Hardware changes to the rotor system are limited to the main rotor system. The main rotor system has been configured with new wide chord rotor blades. The new wide-chord blade allows an increase in performance in both lift and cruise parameters. The new blade has necessitated minor modifications to the rotor hub assembly. The extent of changes made to the UH-60M tail rotor system involve changes to tail rotor bias. Additional rotor system modifications may be required rotor hub configuration in the event that the rotor brake kit is installed.

UH-60 With Wide Chord Blades

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Rotor System Components

Main Rotor Parameters

Main Rotor System

Hub Type Number of Blades Direction of Rotation (Viewed from above) Rotor Location – aircraft Station

The main rotor system general configuration, function, and physical features remain common with the UH-60A/L. The main rotor system is comprised of four subsystems: main rotor blades, hub, flight controls, and the bifilar vibration absorber. Four all composite main rotor blades attach to spindles that are retained by elastomeric bearings contained in one-piece titanium hub. The elastomeric bearing permits the blade to flap, lead, and lag. Hydraulic dampers control lag motion and blade pitch is controlled through adjustable control rods, which are moved by the swashplate. When the rotor is not turning, the blades and spindles rest on hubmounted droop stops. Upper restraints called antiflapping stops retain flapping motion caused by the wind. Both stops engage as the rotor slows down during engine shutdown. Blade retaining pins can be pulled from the blade spindle joint and the blades folded along the rear of the fuselage.

Rotor Location – aircraft Buttline Rotor Location – aircraft Waterline Radius Chord Disc Area Blade Thickness (%Chord) Solidity Ratio Offset Blade Twist Airfoil Section Longitudinal Mast Tilt (-Fwd) Head Moment Constant (@257.7RPM)

UH-60M Main Rotor System 9/3/2003 Page 2 of 16 UH-60M AIRCRAFT 1 FIRST FLIGHT DATA

STD 4 CCW 341.215 in. 0 in. 315.0 in. 26.83 ft 1.92 ft 2261.8 ft2 9.5% 0.0911 15 in. -18° SC2110/ SSCA09 -3° 2852 ftlb/deg

Scissors Assemblies The UH-60A/L rotor system only requires 7.5° clearance between the swashplate and scissors assemblies. The changes made to the flight control system created a mechanical interference between the bottom of the scissors assembly and swashplate. To eliminate this interference, the scissors have been modified to allow a total of 8.5° of clearance to achieve the needed additional 16% in clearance required. The bottom of the scissors has been machined to reduce its profile and allow the clearance required for the increase in lateral cyclic clearance.

UH-60A/L Scissors Assembly

UH-60M Scissors Assembly 9/3/2003 Page 3 of 16 UH-60M AIRCRAFT 1 FIRST FLIGHT DATA

Bifilar Vibration Absorber The application and installation of the bifilar vibration absorber is unchanged on the UH-60M. It reduces rotor vibration at the rotor. The absorber is mounted on top of the hub and consists of a four-arm plate with attached weights.

Bifilar Assembly Main Rotor Dampers Main rotor dampers are installed between each of the main rotor spindle and the hub to restrain hunting (lead and lag motions) of the main rotor blades during rotation and to absorb rotor head starting loads. Each damper is supplied with pressurized hydraulic fluid from a reservoir mounted on the side of each damper. The reservoir has an indicator that monitors the reserve fluid. When the damper is fully serviced, the indicator will show full gold.

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Main Rotor Damper UH-60M AIRCRAFT 1 FIRST FLIGHT DATA

Main Rotor Blades The four main rotor blades produce the lift necessary for flight. The immediately noticeable features are its increased width or chord and the absence of the Blade Inspection Monitor (BIM) indicator. The UH-60M Main Rotor Blade design has a graphite epoxy spar, a 24.25 inch chord length, and an anhedral tip, all of which provide increased efficiency and capability over standard

designs. Each blade has a built in 16.3° static twist to optimize performance. The blade tips are swept back at 30°. The blade tips also feature an anhedral droop of 20°. The leading edge of each blade has a nickel and titanium abrasion strip, the outboard portion of which is protected by a replaceable nickel strip. Electro-thermal blankets are bonded into the blades leading edge for deicing.

Blade Balance Point

Swept Anhedral Tip Blade Cuff

Inboard Trim Tab

Balance Weights

Tip Cap

Outboard Trim Tab UMMR001

UH-60M Main Rotor Blade 9/3/2003 Page 5 of 16 UH-60M AIRCRAFT 1 FIRST FLIGHT DATA

Main Rotor Blade Construction Each of the main rotor blades is constructed with an all-composite structural spar fabricated from graphite and fiberglass tape that is preimpregnated with epoxy resin. The blade has a Nomex™ honeycomb core, fiberglass skin. Wire mesh is bonded to the blade surface to protect the blade in the event of a lightning strike. A leading edge sheath containing balance weights and abrasion strips is bonded to the leading edge of the spar to form the forward airfoil contour.

Main Rotor Blade Attachment The blades are attached to the rotor head by two quick-release expandable pins that require no tools to either remove or install. The UH-60M blade pins are larger than those found on the UH-60A/L due to the increased size of the blade cuff. To conserve space, all blades can be folded to the rear and downward along the tail cone. When mooring, the blades can be tied down with a fitting on the bottom of each blade.

Main Rotor Cuff/Expandable Pins 9/3/2003 Page 6 of 16 UH-60M AIRCRAFT 1 FIRST FLIGHT DATA

Rotor Brake The rotor brake system is currently intended to be a mission flexibility kit. Permanently installed provisions will allow the installation of the kit equipment when required by mission profile. The rotor brake data is included for informational purposes. UH-60M Rotor Brake System Provisions The following equipment constitutes the provisions installed on the aircraft to facilitate the installation of the rotor brake kit in the event that the complete rotor brake system is required for use. • • • • • • •

Shorter Section I Drive Shaft Longer Tail Takeoff Gear and Flange on Main Module Offset Gust Lock Flange (from Pave Hawk) Electrical Wiring Multifunction Display (MFD) Programming Brackets, Bulkhead Fittings, Standoffs, Hardware Quadrant Assembly Ground Idle Lock

The following equipment constitutes the kit items required for installation to complete the rotor brake system: • • • • • •

Brake Disc, Caliper, Pads Hydraulic Lines Master Cylinder, Accumulator, Relief Valve Pressure Switch Lag-Lead Rotor Head Modification Quadrant Ground Idle Solenoid

The rotor brake system is a self-contained manually actuated, hydraulically operated disc brake system. It provides a means of stopping and holding the rotor head during engine start, engine shutdown, or emergency rotor shutdown. It is designed to hold the rotor during engine start with both engines at idle and during shutdown. The system consists of a master cylinder, pressure gauge, manifold, accumulator, pressure switch, rotor brake solenoid, and the rotor brake disc assembly. The brake disc is mounted on the tail takeoff flange. The rotor brake is designed to stop and hold the main and tail rotor systems. The brake is applied by

actuating the brake handle on the rotor brake master cylinder located on the right side of the overhead console. Rotor Brake Master Cylinder The rotor brake master cylinder is located on the right side of the overhead console. It provides pressure, when the rotor brake lever has been applied, to the rotor brake assembly. When the rotor brake lever is in the off position, the system is vented back to the reservoir. A T-shaped rotor brake lever lockpin is used to prevent accidental brake release when the rotor brake has been applied. The pin is set by rotating it 90° and pushing it into the hole on the lever arm after the rotor brake lever has been moved forward. To release the pin, pull and rotate 90°. The rotor brake lever is then free to be released. Rotor Brake Pressure Switch The pressure switch monitors pressure in the rotor brake line. When the brake is applied, the pressure switch closes and activates the ROTOR BRAKE ON EICAS indication. It also deenergizes the engine control quadrant ground idle solenoid that is located in the engine control quadrant. The quadrant lock solenoid is part of a safety interlock that locks the engine power levers in the ground idle position preventing the engine controls from being advanced passed the idle setting with the rotor brake applied. Rotor Brake Accumulator The rotor brake accumulator is a cylinder containing a piston loaded with dual springs. The rotor brake accumulator compensates for thermal expansion or contraction of the hydraulic fluid due to changes in ambient air temperature. Backpressure is held in the accumulator by the springs. Accumulator pressure is maintained to the brake assembly until the brake handle is released or the spring’s bottom in the accumulator. The accumulator is capable of compensating for temperature changes. Rotor Brake Assembly The rotor brake assembly is a caliper-type system containing six brake pads that are pressed into contact with the rotor brake disc when the rotor brake is applied. The rotor brake assembly contains bleed valves for removing

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trapped air from the system as well as indicator pins to determine break pad wear.

Rotor Brake Operating Limits •

Rotor brake should NOT be applied with engines running and rotor turning except during an emergency stop. • Normal usage should be with engines off and Nr between 50 and 30%. • Maximum Nr for rotor brake application is 76%. Rotor Brake Assembly

mounting pad

Aft Transmission area without Rotor Brake Installed

Rotor Brake Assembly

Brake Disc

Rotor Brake Pressure Switch

Rotor Brake Accumulator Aft Transmission area with Rotor Brake Installed

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Tail Rotor System The tail rotor and quadrant systems on the UH60M are identical in function and configuration to the UH-60A/L tail rotor system. Information on the tail rotor system is provided for general review purposes. A cross-beam tail rotor blade system provides anti-torque action and directional control. The blades are of graphite and fiberglass construction. Blade flap and pitch change motion is provided by deflection of the flexible graphite fiber spar. This feature eliminates all bearings and lubrication. The spar is a continuous member running from the tip of one blade to the tip of the opposite blade. Electrothermal blankets are bonded into the blade leading edge for deicing. The tail rotor head and blades are installed on the right side of the tail pylon, canted 20° upward. In addition to providing directional control and antitorque reaction, the tail rotor provides 2.5% of the total lifting force in a hover. A spring-loaded feature of the tail rotor control system will provide a setting of the tail rotor blades for balance flight at cruise power setting in case of complete loss of tail rotor control. Tail Rotor Parameters Number of Blades Rotor Location – aircraft Station in. Rotor Location – aircraft Buttline Rotor Location – aircraft Waterline in. Radius Chord

4 732.0

Blade Twist Cant Angle Airfoil Aspect Ratio Root to tip Solidity Ratio

-18° 20° SC1095 6.79 9.5% 0.1875

The UH-60M tail rotor rigging has been modified to allow for higher values of tail rotor impressed pitch than are currently available of the UH-60A. The UH-60M tail rotor has been biased from the UH-60A’s -9° to +23° to -6° to +26°. The maximum tail rotor operational impressed pitch is limited to 23° above 80 kts forward airspeed, however, for airframe static design, the maximum impressed pitch is 26° for all airspeeds. A limit power-on RPM of 105% will be used for static design loads at sea level conditions. The additional 3° bias is required to maintain a 10% margin on all flight controls for hot day, high gross weight operation. Since the tail rotor thrust is offset by vertical pylon air loads at high speed, the need for the additional pitch bias is not required at speeds above 80 kts. Tail Rotor Quadrant/Warning The tail rotor quadrant contains microswitches to activate the T/R QUAD FAIL caution if a tail rotor cable becomes severed. Configuration and operation of the tail rotor quadrant is identical to the UH-60A/L.

0 in. 324.7 11.0 ft 0.81 ft

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Main Rotor Gust Lock The gust lock prevents the blades from rotating when the helicopter is parked. The gust lock is designed to withstand torque from one engine at IDLE, and thus allows engine maintenance checks independent of drive train rotation. The locking system consists of a locking handle at

the rear of the cabin, a GUST LOCK caution, and a locking device and teeth on the tail rotor takeoff flange of the main transmission. The lock shall only be applied when the rotor system is stationary; it can only be released when both engines are shut down.

Main Rotor Gust Lock Pawl

Main Rotor Gust Lock 9/3/2003 Page 10 of 16 UH-60M AIRCRAFT 1 FIRST FLIGHT DATA

The center tape group contains transmission and rotor speed data. The lower middle tape group contains fuel data.

Rotor System Indications Flight Display System Indications The Flight Display System provides rotor systems indications on both the Engine Indications Crew Alerting System (EICAS) Display and the Primary Flight Display “power pod”.

The rotor speed indication, found in the center tape group, is labeled: •

Rotor speed – NR

EICAS Data Display – Rotor Systems Indications The EICAS displays engine data from the selected Data Concentrator Units DCUs. Cautions and advisories are also displayed. The top center tape group contains rotor speed and engine power turbine speeds. The tape groups to the left and right of the center tape group contain engine data.

Rotor Speed Indication

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PFD Data Display (Power Pod) Rotor Systems Indications The PFD provides a grouping of engine performance data in what is called the “power pod”. The power pod is located in the lower left corner of the display. The PFD rotor speed data display provides information required for normal performance, monitoring, and flight operations as well as quick recognition/diagnosis of abnormal conditions. This allows the pilots increased flexibility to use cockpit MFDs for display of mission situational awareness data rather than dedicating one or more displays full time to an EICAS format.

Engine Instrument “Power Pod”

9/3/2003 Page 12 of 16 UH-60M AIRCRAFT 1 FIRST FLIGHT DATA

Rotor Speed (NR) The rotor speed indication is presented as part of a“Triple Tach” display – providing an integrated indication of rotor speed (Nr) and engines No. 1 and No. 2 power turbine speed (Np) The integrated Nr/Np display combines the three EICAS NR/NP tapes into a single tape with pointers. The tape in the center represents Nr. The two pointers, that bracket each side of the Nr tape indicates NP1 and NP2. The display is marked with lower/upper RED danger limits, minimum/maximum Y Y E O W YE ELLLLLLO OW W continuous limits, and 100% GREEN line. A digital Nr readout is located above the tape.

longer, to provide another aid in quickly determining where the Nr/Np indications are on the scale. Numeric indications are included on the scale at 20%, 90%, 100%, 110%, and 120%. The scale is white with a black outline (to distinguish the scale markings from the Nr tape that moves behind the scale). The scale is displayed all the time; regardless of the Nr input statuses. Limit lines are displayed on the scale to identify the operational ranges for each of the parameters – including the normal operational range and operational ranges, which exceed maintenance and/or performance related limits. The limit line colors identify the severity of the consequences of exceeding the limit.

Nr Readout

Nr Limit Lines

100% Nr

NR Scale and Limit Lines

Y Y w Yeeellllllooow w limit lines represent the boundary between the normal operational range and cautionary operational ranges (above and below desired range).

The red limit lines represent the boundaries Nr Tape between the cautionary ranges and critical operational PFD Power Pod Nr Indication ranges (above upper cautionary range and below lower cautionary range).

Nr scale provides an indication range of 20% to 120% for Nr. The scale consists of three separate linear ranges - between 20% and 90% Nr, between 90% and 110% Nr, and between 110% and 120% Nr. The scale tick marks vary by range – there are tick marks at 20% and 90% in the first range, at 1% intervals in the 90%110% range, and then at 115% and 120%. The tick marks displayed at 5% and 10% values are

The Nr and Np limit lines are separate. The Nr limit lines displayed in the center of the scale and the Np limit lines displayed along the left and right sides of the scale. This independence is necessary as the rotor and power turbine limits do not match in all cases.

9/3/2003 Page 13 of 16 UH-60M AIRCRAFT 1 FIRST FLIGHT DATA

The limit lines mark the points on the scale at which the parameter indications will transition color (color coded to reflect the nature of the operational region). The tapes, readouts, indicators are green in the normal operational range between the yyyeeellllllooow w w limit lines, yellow in the cautionary regions, and red in the critical regions. These limit values are as follows: • • • •

Lower red limit Nr – 91% Lower yellow limit Nr – 95% Upper yellow limit (max continuous limit) Nr – 101% Upper red limit Nr – 110%

The limit and reference lines are displayed at all times. These limit and reference lines are displayed on top of the scale (and tape/indicators), and have black outlines to make them more visible against the background scale, tape, and indicators.

Power Pod Nr failed Indications When the Nr input status has been lost, the readout is removed from the display while the “NR” label remains displayed. The Nr tape will also be removed in the event that rotor speed data is invalid. When the Nr input status is failed or the input is stale (not received), the readout is replaced with three red dashes and the “NR” readout label is displayed as a fail flag (red background, black text) as shown.

Caution/Advisory Indications GUST LOCK ENGAGED - Caution Indicates that the gust lock is engaged. T/R QUAD FAIL - Caution Indicates a tail rotor quadrant failure. ROTOR BRAKE ON – Advisory Indicates that the rotor brake is on.

PFD Power Pod Failed Nr Indications

9/3/2003 Page 14 of 16 UH-60M AIRCRAFT 1 FIRST FLIGHT DATA

Flight Management System (FMS) Status And Test STATUS The rotor system components are monitored by the FMS. Selecting the STS fixed function key to access the MAIN STATUS page on the FMS can check the status of the rotor overspeed system. By selecting soft key (SK) 4 from the MAIN STATUS PAGE accesses the AIR VEHICLE status page. Selecting SK 4 accesses air vehicle equipment status for the rotor overspeed. This menu selection allows the operator to view the status of the rotor overspeed. All faults and failures will be displayed under the rotor overspeed if a fault or failure should occur.

RTR OVRSPD STATUS PAGE Pressing SK-6 RESET removes all the overspeed faults from the displayed fault list on the FMS: • • • • • •

DCU 1 127% DCU1 137% DCU 1 142% DCU 2 127% DCU 2 137% DCU 2 142%

The RESET switch located in the forward nose compartment, left hand side, also removes the faults by resetting the switch. TEST AIR VEHICLE STATUS PAGE

There is no test indication for rotor overspeed.

9/3/2003 Page 15 of 16 UH-60M AIRCRAFT 1 FIRST FLIGHT DATA

NOTES

9/3/2003 Page 16 of 16 UH-60M AIRCRAFT 1 FIRST FLIGHT DATA

Section 2-8 Electrical System Differences Description The electrical system of the UH-60M provides three phase 115 vac and 28 vdc to 3 ac electrical busses and 8 dc electrical busses. The primary electrical system is comprised of 2 main ac transmission driven generators. Backup or alternate ac power is provided by an APU driven generator for in flight emergencies or ground maintenance power requirements. Ground maintenance power requirements may also be met by connecting a ground power source to the AC external Power receptacle. The dc electrical system is comprised of 2 ac power 400 amp dc converters and a battery electrical system comprised of two 5 amp-hour sealed lead-acid batteries. The main ac electrical system remains largely unchanged from the UH-60A/L system configuration. The dc and battery electrical systems however, have been impacted by the UH60M modernization.

The UH-60M dc electrical system has been revised to provide sufficient aircraft electrical power and power management provisions to support existing and new mission equipment packages. Existing mission equipment packages such as Air Volcano, Quickfix, and medical evacuation have always been able to interface with the Black Hawk electrical system without difficulty. New mission equipment systems such as Army Airborne Command and Control System (A2C2S), Land Warrior and Air Warrior will potentially push the limits of the electrical system capacity. To address these potential problems, higher amperage DC converters have been installed and a new battery system has been designed and implemented. Additionally some contactors, smaller but no less important elements of the electrical system, have been replaced with newer better equipment to accommodate the increased system capacities. Circuit breaker panels and various relays have been modified or revised to reflect changes in aircraft systems as well as new or different system power requirements.

10/2/2003 Page 1 of 16 UH-60M Aircraft 1 First Flight Training Data

Electrical System Components DC Power Supply System DC Converters Two 28 volt 400-ampere ac to dc unregulated converter (transformer-rectifier) are normally powered by the No. 1 and No. 2 ac primary buses to convert input ac power into dc power for use by aircraft systems. The converter output is applied to the No. 1 and No. 2 dc primary buses whenever ac power is applied to the ac primary buses. Just as with the UH-60A/L, the converters are capable of being cross-tied if one converter fails. The converters are located in the same location as the UH-60A/L converters.

28 Volt, 400 Amp DC Converter

10/2/2003 Page 2 of 16 UH-60M Aircraft 1 First Flight Training Data

Battery Electrical System The battery system has been revised to provide added reliability. The batteries also provide power directly to all buses to hold up bus power during power switching transient events. Its purpose is to provide uninterrupted emergency power to allow flight for 12 minutes with the following items: •

Manual Stabilator Slew Control (assumed to slew up & down once) • SAS Turn-off Valve Control • Fuel Shut-off Value Control • Crashworthy External Fuel System (CEFS) Transfer Valve Control • Standby Compass Light • Cockpit Utility Lights • APU System Control • APU and Main Engine Fire Detection Systems • APU and Main Engine Fire Extinguisher Systems • Battery System/ DC Essential Relay Contactors • Backup Hydraulic System/Pump Control Audible Warning System • ESSS Emergency Jettison System • Cargo Hook Emergency Release System •

• •

• • • • • • • • • • • •

•

Rescue Hoist Cable Shear Digital Clock Electronic Standby Instrument System powered by the independent 30-minute battery (not connected to the Aircraft Battery System during primary power failure) Standby Instrument Lighting Intercommunication System (ICS) VHF-AM (AN/ARC-186) with back-up control panel Pilot CDU IFF (AN/APX-118) No.2 EGI (Pilot) CVR/FDR No.2 DCU (Pilot) Pilot Inboard MFD (with internal heater off) VOR/ILS (AN/ARN-147) No.1 VHF-FM (AN/ARC-201) without Power Amplifier Radar Altimeter (AN/APN-209) Controllable Search Light (during landing only) Number 2 Air Data Computer

No. 1 Battery

No. 2 Battery

10/2/2003 Page 3 of 16 UH-60M Aircraft 1 First Flight Training Data

DC Bus Current Sensor The output of each converter is routed through a DC Bus Current Sensor. The current sensor controls the actuation of the associated Converter Line contactor and DC Essential Bus contactor. If the converter is not supplying output for greater than 250 msec (nominal), the associated Converter Line Contactor and DC Essential Bus contactor will open. The 250 msec time delay allows the batteries to hold up the DC Primary during transient switching events but will not allow the batteries to supply the DC Primary Buses indefinitely. DC Electrical Busses Eight DC buses distribute DC electrical power. They are designated: Sealed Lead Acid Batteries The UH-60M features a low-maintenance battery system. It consists of two 5 amp-hour sealed lead-acid batteries. The battery system provides power for ignition and control power to the fuel control sequencing valves for starting the APU on the ground.

•

No. 1 DC Primary Bus – Powers primary aircraft DC loads.

•

No. 2 DC Primary Bus – Powers primary aircraft DC loads.

•

No. 1 DC Essential Bus – Powers flight essential DC loads.

•

No. 1 DC Essential Bus – Powers flight essential DC loads.

•

No. 1 Battery Bus – Powers APU start and control power.

•

No. 2 Battery Bus – Powers fire detection control power.

•

No. 1 battery Utility Bus – Powers Fire Extinguishing control, No. 1 Battery control, and instrument secondary power.

•

No. 2 Battery Utility Bus – Powers Fire Extinguishing control, No. 2 Battery control, and Instrument secondary power.

Battery Low-voltage Sensing Relay Battery condition is monitored with an undervoltage sensor. The undervoltage sensor on each battery illuminates the BATT LOW status indicator and BATT LOW CHARGE caution when the battery voltage is initially below 21.5 VDC or subsequently falls below 20.5 VDC. Battery Contactor The battery contactor connects 24 vdc battery power to the No. 1 Battery, No. 2 Battery, No. 1 DC Essential and No. 2 DC Essential buses when the battery switches are selected to ON. The battery contactor in conjunction with it’s associated DC Essential Bus Contactor will also provide a path for the battery to supply voltage to the associated DC Primary Bus for a nominal 250 msec. This allows the equipment on the DC Primary Bus to be powered throughout normal switching transients that occur during power source switching. The converters supply charging power to the batteries whenever the converters are supplying power and the associated battery switch is ON.

10/2/2003 Page 4 of 16 UH-60M Aircraft 1 First Flight Training Data

DC Electrical Bus Diagram

10/2/2003 Page 5 of 16 UH-60M Aircraft 1 First Flight Training Data

AC Power Supply System No. 1 and No. 2 Main AC Generators The main generators are functionally and physically identical to the main generators found on the UH-60A/L. The two 30/45 KVA, brushless oil spray cooled, 12,000 rpm, three phase generators are mounted on the main transmission accessory gearboxes. Each generator operates in conjunction with a Generator Control Unit (GCU).

No. 2 Main Generator

Generator speed is maintained at 12,000-rpm (400 Hz) ± 5% throughout all flight modes of the UH-60M, therefore under-frequency protection in flight is not essential. Generator underfrequency protection is locked-out in flight by a switch actuated when the weight of the helicopter is off the landing gear. This design eliminates the risk of losing the generator when the rotor speed drops momentarily below 380 Hz after an engine malfunction is eliminated. On the ground, the under-frequency protection circuit removes the generator from the bus when the frequency drops below 380 Hz.

No. 1 Main Generator

10/2/2003 Page 6 of 16 UH-60M Aircraft 1 First Flight Training Data

AC Electrical Busses Three AC buses distribute AC electrical power. They are designated:

APU Generator The APU generator provides in-flight backup power and ground power. This generator is a 20/23.8 KVA, air-cooled, 12,000 rpm, and threephase generator, driven by the Auxiliary Power Unit (APU).

•

No. 1 AC Primary Bus – Powers primary aircraft AC loads including No. DC 1 Converter.

•

No. 2 AC Primary Bus – Powers primary aircraft AC loads including No. 2 DC Converter.

•

26 VAC Reference Bus – Powers 26 VAC loads.

There is no longer an AC Essential Bus. The 26 VAC Reference Bus is provided via a transformer on either AC Primary Bus.

Generator Control Units (GCU) The GCUs, each of which are solid-state units, provide voltage regulation, protection against under frequency, over and under-voltage, feeder faults for each of the generators. Underfrequency protection is only active during ground operation is provided for No. 1 and No. 2 generators. The location of the APU Generator Control Unit has changed due to equipment changes and space requirements in the cabin overhead. The APU GCU is now located in the aft cabin overhead on the left hand side.

10/2/2003 Page 7 of 16 UH-60M Aircraft 1 First Flight Training Data

Battery Junction Boxes The UH-60M has two battery junction boxes. Each battery junction box contains a battery contactor and three current limiters. Additionally, the No.2 battery box contains the battery tie contactor. The No. 1 battery junction box is located in the avionics nose tunnel. The No. 2 battery junction box is located in the pilot’s seatwell. Due to the physical distance between the No. 2 battery junction box and the No. 2 battery, an additional current limiter has been installed in line with the No.2 battery feeder for fault protection. No. 1 And No. 2 Main Junction Box The No. 1 and No. 2 main junction boxes are similar in physical configuration and function to those found on the UH-60A/L aircraft. Equipment changes in the junction box include the addition of the DC Bus Current Sensor, replacement of the existing DC Converter Line Contactor with a new style contactor, addition of the DC Essential Bus Contactor, and additional current limiters due to the new DC power system design.

No. 1 Battery Junction Box

No. 2 Battery Junction Box

10/2/2003 Page 8 of 16 UH-60M Aircraft 1 First Flight Training Data

DC and AC Circuit Breaker Panels The circuit breaker panels provide a central location for electrical power distribution and aircraft systems electrical protection. Breaker panels are located above and to the rear of each pilot and copilot, two are on the lower console, and two are on the upper console. There is also an auxiliary circuit breaker panel above and to the right of the copilot. The circuit breakers provide both ac and dc protection. As always, unnecessary recycling of circuit breakers, or using circuit breakers as a switch is discouraged.

NO. 1

AC PR

I BUS

UTIL

60 HZ AC

162

163

161

1

15

7 /2 30

30

RECP

CONVERTER

CPLT WSHLD

WSHLD WIPER

NO. 1

160

159

158

15

2

30

30

ANTI-ICE BACKUP PUMP 150

LEFT PITOT

INTRF

155

154

153

152

35

10

2

5

5

3O

0A

0A

0C

0C

CONVERTER

HEAT

BLKG

157

ESSS JTSN INBD OUTBD 149

148

146

147

145

144

1 7 /2

1 7 /2

NO.1 VHF FM

CNSL

1/ 2

NO.1 AFCS

CPLT MFD

125

124

123

122

1

5

2

2

0B

0B

XFMR

0B

0B

CMPTR

FAN

LIGHTS

DC PR

I BUS

OVSPD ANTI-COLL

NO. 1 ESH

UTIL RECP

ADVSY

ADVSY

LIGHTS LWR

LWR

RETR

CPLT WSHLD

AIR SOURCE HEAT/

BUS TIE

NO.1 DEN

TAIL RTR SERVO

142

141

140

139

138

137

136

135

134

133

132

131

130

129

128

127

2

1 7 /2

5

5

5

5

5

5

5

5

5

5

5

3

5

5

CSL 5V

CSL 28V

CSL 5V

LDG PWR

WARN

NO.1 EGI

ADF

CMD CSL

NO.1 AFCS

CABIN CPLT MFD

CPLT

CPLT MFD

CPLT

120

119

118

117

116

115

114

113

112

111

10

1

5

25

5

25

5

2

25

7 /2

AMP

COOLING CONTR

FD/DCP

INBD

CDU

OUTBD

PRI

121

NO. 1

151

143

WARN FUEL QTY

NO. 1 ENG

FUEL LOW

5

PWR NO.1 AUTO

156

NO.1 ENG ANTI-ICE

NO.1 SERVO

WARN

ANTI-ICE

START

CNTOR

WARN

WARN

CONTR

CHAFF

ICS

NO.1 DCU

NO.1 DTS

HUD

NO.1 ADC

LASER

RADAR

110

109

108

107

106

105

104

103

5

7 /2

1

5

2

5

5

SEC

PRI

WARN

WARN

1

1

7 /2

SET

DISP

1

126

102

101

UMEL003

Copilots Circuit Breaker Panel

.2 NO NO.2

AC PR

I BUS HEAT & VENT

262

263

35

7 /2

3o

30

264

1

CONVERTER

NO.2

IB DC PR

US

STAB 251

2 0B

IND MASK 226

227

1

5

NO.2 SERVO

BLOWERS WARN

201

202

NO.2 ESH

NO.2 GEN

BUS TIE

IDM

NO.2 ENG ANTI-ICE START

228

229

230

231

232

233

234

235

5

2

5

5

3

5

5

5

WARN

CNTOR

WARN

HEAT

NO.2 VHF FM

IRCM

NO.2 AFCS

205

206

207

208

5

5

2

7 /2

CONTR

203

204

VENT

CONTR

1

NO.2 STAB

26 AC REF NO.1 NO.2 DCU DCU 252

1

253

1

AFCS 254

2

0B

0B

0B

REF

REF

REF

LIGHTS RETR LOG

255

CARGO HOOK

CTR WSHLD

UTIL RECP

257

259

260

261

15

1 7 /2

1 7 /2

15

30

HF

30

30

30

ANTI-ICE

AMP HF

PLT WSHLD

AVC

ANTI-ICE

LTS CABIN

LTS LWR

NO.2 DTS

PLT WSHLD

CTR WSHLD

LIGHTS FORM

CARGO

PLT MFD

236

237

238

239

240

241

242

243

244

245

247

248

249

5

10

5

30

3

5

5

1

5

5

1

1

2

0B

0B

CONTR

PWR

CONTR

SCTY

DOME

CSL 5V

A-ICE

A-ICE

LV

HV

HOOK

FAN

PLT MFD

STORM

NO.2 ENG

RT PITOT

NO.2 AFCS

NO.2 AUTO

UHF AM

UHF AM

FIRE

246

0B

250

2 0B

POS

AVC

STBY

STBY

209

210

211

212

213

214

215

216

217

218

219

220

222

223

224

7 /2

1

5

3

5

5

15

5

25

2

7 /2

1

2

5

5

10

2

0C

0A

0B

0B

PWR

TRIM

XMSN

LTS

CMPTR

BATT

INSTR

OUTBD

SCOPE

SCTY

EXTGH

OVSPD

HEAT

CMPTR

XFMR

SPEED

MAIN

221

225

1

UMEL004

Pilots Circuit Breaker Panel 10/2/2003 Page 9 of 16 UH-60M Aircraft 1 First Flight Training Data

NO.1 DC ESNTL

NO.1 AC PRI ICS

NO.2 AC PRI

NO.1 AC PRI

CPLT MSTR

LIGHTS INSTR

LIGHTS ILLUM

AVION RLY

NO.1 ENG

DIGITAL

CPLT ICS

2

2

2

5

5

3

5

1

BKUP

WARN

PNL

PB

PNL

START

CLOCK

NO.2 DC PRI

CTR CNSL

ROTOR BRAKE

5

2

FAN

NO.1 DC PRI

NO.2 DC ESNTL TAIL WHEEL

IR

3

5

10

FDR

LOCK

CVR

SEARCH LIGHT WHITE CONTR

20

5

RAD

2 ALT

UMEL001

Copilots Auxiliary Circuit Breaker Panel

NO. 1 DC ESNTL BUS NO. 1 ENG

LIGHTS

IFF

AUX CB

CABIN ICS

NO. 1 ESNTL

5

5

5

35

4

2

FIRE DET

SEC

NO.1 STAB

SAS

ECP

CARGO HOOK

5

2

7

1 2

PWR

BOOST

ESSS

7

1 2

SENSE

JTSN OUTBD

NO. 1 VHF

HOIST CABLE

ESSS JTSN

5

5

5

EMER RLSE

FM

SHEAR

PNL SPLY

7

1 2

INBD

UMAV002_2

Left Overhead Circuit Breaker Panel

10/2/2003 Page 10 of 16 UH-60M Aircraft 1 First Flight Training Data

NO. 2 DC ESNTL BUS PLT

NO. 2 EGI

VHF AM

AUX CB

VOR/ ILS

NO. 2 ESNTL

PLT ICS

2

5

5

35

2

2

2

MSTR WARN

PRI

PLT CDU

NO. 2 ADC

BACKUP HYD

5

2

5

PNL SPLY PLT FD/

7

CONTR

SENSE

PLT MFD

1 2

DCP

25 INBD

NO. 2 DCU

7

ICS ICU

1 2

3

PRI

UMAV002_3

Right Overhead Circuit Breaker Panel

NO. 1 BATT BUS APU

NO. 1 ENG

APU CONTR

FUEL PRIME

5

2

5

5

CONTR

FIRE DET

WARN LIGHTS

INST

BOOST

NO. 1 EGI

NO. 1 BATT

BUS &

NO. 1 CEFS

5

2

5

5

SEC

BUS CONTR

CONV WARN

IFF

FIRE

NO. 1 DCU

CDU

APU CONTR

1

5

7.5

5

5

BKUP

EXTGH

SEC

SEC

INST

APU GEN

5

B A T T B U S

NO. 1 BATT UTIL BUS

UMAV003_13 Left Lower Console Circuit Breaker Panel

10/2/2003 Page 11 of 16 UH-60M Aircraft 1 First Flight Training Data

NO. 2 BATT BUS NO. 2 CEFS

NO. 2 ENG

NO. 2 ENG

UTIL BUS

5

2

5

5

WARN LIGHTS

FIRE DET

CKPT

B A T T B U S

NO. 2 EGI

FIRE

NO. 2 DCU

NO. 2 BATT

5

5

7.5

2

SEC

EXTGH

SEC

BUS CONTR

NO. 2 BATT UTIL BUS

UMAV003_18 Right Lower Console Circuit Breaker Panel

AUX CABIN

15 HEATER ICE-DET

A C

NO.2 FUEL

2

2 BOOST PUMP

NO.2 PRI BUS

RESQ HOIST

DE-ICE

ICE-DET

10

5

1 7 /2

CONTR

CTLR

NO.2 LTR

7A

5 LTS

PILOT

CREW 1

15

15

D C

MCU

MCU

NO.1 LTR

AUX HEATER

COPILOT

5

5

15

5

CONTR

MCU

MCU

LTS

NO.1 PRI BUS

CREW 2

D C

NO.1 FUEL

2 BOOST PUMP DE-ICE PWR

20 TAIL ROTOR

A C

CEFS

20 PUMP

UMEL002

10/2/2003 Page 12 of 16

Mission Readiness Circuit Breaker Panel UH-60M Aircraft 1 First Flight Training Data

Electrical System Controls

Generator Switches

Battery Switches

The location and function of the GENERATORS No. 1, NO. 2 and APU switches is identical to the UH-60A/L.

The control switches for the No. 1 and No. 2 battery systems are located on the center panel of the overhead console. The BATT NO. 1 and BATT NO. 2 switches are three position switches with TEST, OFF and ON positions. The function of each of the positions is as follows: •

OFF - Removes 28 vdc from battery contactor K501/K503 that in turn removes battery power from the respective battery bus and DC Essential Bus. TEST – In this position, a load of approximately four amps is placed on the associated battery. The BATT GOOD indication illuminates if the battery is capable of pulling in the undervoltage sensor with its additional load, Otherwise the BATT LOW indication will illuminate. The undervoltage sensor pulls in at 2`1.5 volts and drops out at 20.5 ON – Allows 28 vdc to energize battery contactor K501/K503 that in turn allows battery power to be connected to the respective battery bus. The battery tie contactor K502 will close if there is no converter power on only one contactor is supplying power with the battery switches to ON.

•

•

External Power Switch The location and function of the EXT PWR switches is identical to the UH-60A/L

CARGO HOOK

VIB CONT

EMERG REL NORM

LIGHTS

AUX CB

IFF

CABIN ICS

NO. 1 ESNTL

ESSS 1 2

5

5

5

35

4

2

FIRE DET

SEC

SENSE

JTSN OUTBD

NO.1 STAB

SAS

ECP

CARGO HOOK

NO. 1 VHF

HOIST CABLE

ESSS JTSN

5

2

5

5

5

EMER RLSE

FM

SHEAR

7

1 2

PWR

PNL SPLY

BOOST

7

7

1 2

INBD

EMERG REL NORM

EXT LTS MODE

5

4

CARGO HOOK LT

N O R M

1

O F F

STBY INST TEST O F F

NO. 1 TEST

ANTI COLLISION LIGHTS

UPPER

SECONDARY WHITE

DIM

STEADY

BRT

FLASH

BRT

CABIN DOME LT

OFF

DIM

BRT

ON

CONT

5

5

NO. 2 ADC

BACKUP HYD

5

2

5

BATT STBY INST TEST O F F

GENERATORS NO. 2 TEST

NO. 1 TEST O F F

O F F ARM

ON

ON

RESERVE

ENG

O F F

PLT MWC

DIM

OFF

BRT

2

BRT

FUEL BOOST PUMP NO. 1

TEST LT

WINDSHIELD WIPER PARK

OFF

LOW

1 2

APU

BATT GOOD

ENG NO. 1

BATT LOW

OIL HOT

ACC LOW

APU FAIL

APU ON

2

DCP

7

INBD

CONT

1

WHITE

ON

ON

APU TEST

ON

AIR SCE HI/STRT ENG

O F F

O F F ON

FIRE EXTGH RESERVE

3

PRI

O F F

O F F

HEATER MED O F F

O F F

ON

ANTI-ICE NO. 2

PITOT LEFT

APU

MAIN

FUEL BOOST PUMP

NO. 2

WINDSHIELD COPILOT CTR

ANTI-ICE PILOT

O F F

O F F

FUEL IND

ON O F F ON

EMER RLSE TEST LT

BATT GOOD BATT LOW

OIL HOT

ACC LOW

APU FAIL

APU ON

ON

ON

ON

HI

HEAT RIGHT O F F

O F F

O F F

OFF

ON

TEST

O F F

BRT

FIRE DETR TEST OPER

O F F

R O E F S F E T

2

INST PNL

OFF

NO. 2 TEST

BOOST PUMP

O F F

ICS ICU

1 2

FUEL BOOST PUMP O F F

ON

R O E F S F E T

PLT ICS

NO. 2 DCU

25

VENT BLOWER

HI

ON

APU FIRE

NO. 1 TEST

APU

RESET

SENSE

PLT MFD

PLT FD/

7

APU

EMER RLSE

R O E F S F E T

EXT PWR

ON

O F F ON

35 PNL SPLY

PRI

PLT CDU

CONTR

AIR SCE HI/STRT

O F F ON

APU TEST

MSTR WARN

ON

LWR CSL

BRT

R O E F S F E T

BOOST PUMP

FIRE EXTGH

LIGHTS

OFF

ON

R O E F S F E T

NO. 2 TEST

APU O F F

ON

UPR CSL

NO. 1 TEST

MAIN

LTD SW

BRT

PLT

NO. 2 ESNTL

VOR/ ILS

ON

LIGHTS CPLT MWC

ON

RESET

O F F NIGHT

DIM

SECONDARY

ON

R O E F S F E T

EXT PWR

O F F BLUE

O F F

O F F

POSITION LIGHTS

DAY

O F F LOWER

ARMED

ON

OFF

B O T H

NO. 2 TEST

ON

IR

ALL

GENERATORS

AUX CB

VHF AM

NO. 2 EGI

2

BATT

O F F

3 2

ARM SAFE

SHORT

ON

ARM

FORMATION LTS

CONTR CKPT

O P E N

O F F

ARMED

NO. 2 DC ESNTL BUS

CARGO HOOK

VIB CONT

ALL

SHORT

ON

ON NO. 1 DC ESNTL BUS

ARM SAFE

O P E N

O F F

O F F

NO. 1 ENG

CONTR CKPT

LAMPS

NO. 1

APU APU FIRE

FUEL BOOST PUMP O F F

NO. 2

2

ON

TEST

BLUE

ON

ENG OVSP TEST A NO. 1 B A NO. 2 B

ENG SPD TRIM

UMAV002_4

DECR O F F INCR

UMAV002_1

10/2/2003 Page 13 of 16 UH-60M Aircraft 1 First Flight Training Data

BATT LOW – Caution Yellow caution light that indicates a battery voltages is low. If the respective battery switches are on, illumination will occur if the batteries initially below 21.5 vdc or subsequently fall below 20.5 vdc.

Electrical System Indications Battery System Cold Start Status Indicator/Caution Light The replacement of the UH-60A/L Caution/Advisory panel by the multifunction displays has required that battery status indications, powered by the batteries themselves, be relocated. The Cold Start Status Indicators provide indication for Auxiliary Power Unit status, battery condition as well as indications for Emergency Release tests conducted on mission related equipment such as the cargo hook. One of the four cold start status indicator capsules is dedicated to providing battery state indications.

BATT GOOD – Advisory Green status light indicating battery is appropriately charged above 21.5 vdc and remains above 21.5 vdc.

CARGO HOOK

VIB CONT

EMERG REL NORM

CONTR CKPT

ARM SAFE

ALL

ARMED

O P E N

O F F

SHORT

ON

BATT STBY INST TEST O F F

GENERATORS NO. 2 TEST

NO. 1 TEST O F F

O F F ARM

ON

ON

R O E F S F E T

NO. 1 TEST

RESET

R O E F S F E T

ON

O F F

RESERVE O F F MAIN

FUEL BOOST PUMP NO. 1

APU APU FIRE

ON

O F F ON

FIRE EXTGH

ON

ENG

O F F ON

APU TEST

AIR SCE HI/STRT

BOOST PUMP

CONT

ON

O F F

NO. 2 TEST

APU

EXT PWR O F F

ON

R O E F S F E T

APU

EMER RLSE TEST LT

BATT GOOD BATT LOW

OIL HOT

ACC LOW

APU FAIL

APU ON

FUEL BOOST PUMP O F F

NO. 2

ON UMAV002_4

10/2/2003 Page 14 of 16 UH-60M Aircraft 1 First Flight Training Data

Caution/Advisory Indications BATT LOW CHARGE – Caution Indicates that the No.1 or No.2 Battery Bus voltage is less than 21.5 vdc or subsequently falls below 20.5 vdc. DC ESS BUS 1 OFF – Caution Indicates that the No.1 DC Essential Bus power is not powered. DC ESS BUS 2 OFF – Caution Indicates that the No.2 DC Bus power is not powered. CONV 1 FAIL – Caution Indicates that the No.1 Converter has no output power. CONV 2 FAIL – Caution Indicates that the No.2 Converter has no output power. GEN 1 BRG FAIL – Caution Indicates that the generator main bearing is worn or has failed and operating on auxiliary bearing. GEN 1 FAIL – Caution Indicates that the No.1 Generator is not supplying power to the bus. GEN 2 BRG FAIL – Caution Indicates that the generator main bearing is worn or has failed and operating on auxiliary bearing. GEN 2 FAIL – Caution Indicates that the No.2 Generator is not supplying power to the bus. AC REF BUS DEGRAD - Advisory Indicates that one of the two transformers supplying the dual redundant reference bus voltage for DCU operation has failed. Such a failure results in lost reference bus redundancy. APU GEN ON - Advisory Indicates that the APU Generator output is acceptable. EXT PWR CONNECTED - Advisory Indicates that an External Power Plug is connected to the External Power Connector.

10/2/2003 Page 15 of 16 UH-60M Aircraft 1 First Flight Training Data

NOTES

10/2/2003 Page 16 of 16 UH-60M Aircraft 1 First Flight Training Data

Section 2-9 Aircraft Lighting System Differences Description The UH-60M lighting systems, as with the UH6A/L, consist of external, internal, and cockpit lighting systems and related controls. The lighting systems provide illumination of the exterior and interior of the aircraft. Exterior lighting allows areas surrounding the aircraft to be lit for safety purposes as well as to provide navigation aids. The exterior lighting system consists of formation lights, anti-collision lights, position lights, and a cargo hook light. The interior lighting system provides a means of illuminating the cockpit and cabin sections of the aircraft. Cockpit light provides illumination of the instrument panel as well as upper and lower console control panels and switches in low and no-light flight operations. Cabin lighting illuminates the cabin section of the

aircraft for crew chiefs and troops. UH-60M exterior and interior lighting systems are Night Vision Goggle (NVG) capable by design. The function and location of exterior lighting remains largely unchanged on the UH-60M. Minor changes have been made in some components relative to the means by which they are mounted to the airframe. As with the exterior lighting systems, the interior lighting systems have not been greatly impacted by the UH-60M modernization. The most noticeable changes to the lighting systems can be seen in the location of lighting systems controls found on the upper console. The UH-60A/L lighting controls were distributed on both the left and right sides of the upper console. The UH-60M lighting controls have been centrally located on the copilot side of the upper console. This arrangement allows both pilots visual and operational access of the lighting controls but biases operation of lighting systems toward the copilot.

Anticollision Light

Tail Position Light

Anticollision Light

Right Position Lights

10/2/2003 Page 1 of 14 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

The WHITE position selects white light to be used as the cockpit secondary lighting. Using the BLUE setting position selects NVIS blue light to be used as the cockpit secondary lighting.

Lighting System Components/Controls Interior Lighting

FORMATION LTS

LIGHTS - SECONDARY WHITE/ BLUE The switch has been relocated from its former position on the center of the cockpit floodlight. This two-position toggle switch is used to select the color illumination of the cockpit secondary lighting.

N O R M

2 1

B O T H

IFF

AUX CB

CABIN ICS

NO. 1 ESNTL

5

5

35

4

2

FIRE DET

SEC

NO.1 STAB

7

1 2

PWR

PNL SPLY

SAS

ECP

5

2

SENSE

CARGO HOOK

BOOST

7

NO. 1 VHF

JTSN OUTBD

HOIST CABLE

5

5

5

EMER RLSE

FM

SHEAR

1 2

EMERG REL NORM

7

1 2

BATT STBY INST TEST O F F

EXT LTS MODE

5

4

CARGO HOOK LT

N O R M

1

O F F

LOWER SECONDARY WHITE

DIM

BRT

LTD SW

CABIN DOME LT

UPR CSL

DIM

BRT

AUX CB

VOR/ ILS

NO. 2 ESNTL

PLT ICS

5

35

2

2

2

MSTR WARN

PRI

PLT CDU

NO. 2 ADC

5

2

PNL SPLY BACKUP HYD

5

ENG

WINDSHIELD WIPER PARK

OFF

LOW

O F F

PLT MWC

DIM

OFF

BRT

PLT MFD

1 2

TEST LT

FUEL BOOST PUMP

BRT

NO. 1

APU

1 2

ENG NO. 1

BATT LOW

OIL HOT

ACC LOW

APU FAIL

APU ON

SECONDARY

3

O F F

ANTI-ICE NO. 2

1

BRT

DIM

BRT

PLT MWC

DIM

BRT

UPR CSL

LWR CSL

INST PNL

MED

O F F

O F F ON

OFF

HI

OFF

HEAT RIGHT

ON

ON

CABIN DOME LT

NO. 2

WINDSHIELD COPILOT CTR

ANTI-ICE PILOT

O F F

O F F

FUEL IND

ON O F F ON

BRT

OFF

BRT

OFF

BRT

OFF

BRT

O F F

ON

FIRE DETR TEST OPER 1

WHITE

TEST

O F F

BRT

FIRE DETR TEST OPER

WHITE

O F F

LTD SW

HEATER

ON

PITOT LEFT

INST PNL

OFF

CPLT MWC

PRI

FUEL BOOST PUMP O F F

FLASH

O F F

O F F

HI

BATT GOOD

ON

APU FIRE

7

INBD

BRT

LIGHTS

LIGHTS

ICS ICU

NO. 2 DCU

25

DCP

VENT BLOWER

APU

EMER RLSE

NIGHT

SENSE

PLT FD/

7

CONTR

ON

LWR CSL

BRT

VHF AM

5

STEADY

ON

FLASH

LIGHTS

OFF

APU TEST

AIR SCE HI/STRT

ON

O F F

ON

SECONDARY

R O E F S F E T

ON

NO. 2 EGI

2

ON

STEADY

BRT

BRT

NO. 2 TEST

O F F ON

RESERVE

LIGHTS

OFF

O F F

MAIN

NIGHT

CPLT MWC

R O E F S F E T

BOOST PUMP

FIRE EXTGH

O F F

DIM

ON

CONT

RESET

O F F BLUE

NO. 1 TEST

APU

EXT PWR

POSITION LIGHTS

DAY

O F F

R O E F S F E T

PLT

ON

ANTI COLLISION LIGHTS

UPPER

ON

ON

IR

OFF

B O T H

NO. 2 TEST O F F

ON

O F F

3 2

ARMED

GENERATORS

NO. 1 TEST O F F

ARM

FORMATION LTS

ARM SAFE

ALL

SHORT

ON

ESSS JTSN

INBD

CONTR CKPT

O P E N

O F F

POSITION LIGHTS

DIM O F F

DIM

NO. 2 DC ESNTL BUS

CARGO HOOK

VIB CONT

ESSS

ON

DAY

O F F

BLUE NO. 1 DC ESNTL BUS LIGHTS

IR

ANTI COLLISION LIGHTS

UPPER

LOWER

5

O F F

OFF

SECONDARY WHITE

NO. 1 ENG

CARGO HOOK LT

EXT LTS MODE

5

4 3

LAMPS 2

LAMPS 2

TEST

BLUE

TEST

BLUE

ENG OVSP TEST A NO. 1 B A NO. 2 B

ENG OVSP TEST A NO. 1 B A NO. 2 B

ENG SPD TRIM DECR O F F INCR

UMAV002_1

ENG SPD TRIM DECR

Overhead Console

O F F INCR UMAV002_6

Cockpit Secondary Light 10/2/2003 Page 2 of 14 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

LIGHTS - CPLT MWC The Copilot Master Warning Panel dimming control is a new lighting control for the Black Hawk. The CPLT MWC rotary dimmer control varies the brightness of the Copilot Master Warning Panel.

FORMATION LTS

EXT LTS MODE

5

4

The CPLT MWC lighting control is used in combination with the INSTR PNL lighting control. Moving the INSTR PNL dimmer control from the OFF position enables the CPLT MWC lighting control. Moving the co-pilots master warning caution rotary knob dimmer out of the BRIGHT position controls Copilot Master Warning Panel intensity. Clockwise rotation increases brightness while counterclockwise rotation dims the lighting to a minimum but does not turn it off.

CARGO HOOK LT

3 N O R M

2 1

O F F IR

ON

OFF

ANTI COLLISION LIGHTS

UPPER B O T H

LOWER

BLUE

DIM

STEADY

BRT

FLASH

O F F

O F F

SECONDARY WHITE

POSITION LIGHTS

DAY

NIGHT

LIGHTS CPLT MWC

DIM

LTD SW

BRT

DIM

BRT

PLT MWC

DIM

BRT

LIGHTS SECONDARY

OFF

BRT

CABIN DOME LT

UPR CSL

OFF

LWR CSL

BRT

OFF

BRT

# 1 ENG OUT

MASTER CAUTION

FIRE

PRESS TO RESET

# 2 ENG OUT LOW ROTOR R. P. M.

INST PNL

OFF

BRT

FIRE DETR TEST OPER 1

WHITE

O F F

LAMPS 2

TEST

BLUE

ENG OVSP TEST A NO. 1 B A NO. 2 B

ENG SPD TRIM DECR O F F INCR UMAV002_6

10/2/2003 Page 3 of 14 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

LIGHTS - LTD SW The LTD SW rotary knob allows control of the brightness of all upper and lower console illuminated pushbuttons and annunciator lights. The following control panel lighted switches are controlled by this control: •

Copilot’s Outboard Multifunction Display (MFD) Copilot’s Inboard MFD Copilot’s Flight Director Display Control Panel (FD/DCP) Copilot’s Reversionary Switch Panel Pilot’s Outboard MFD Pilot’s Inboard MFD Pilot’s Flight Director Display Control Panel Pilot’s Reversionary Switch Panel Copilot’s Interior Communications System (ICS) ICS Control Panel

• • • • • • • •

FORMATION LTS

EXT LTS MODE

5

4

• • • • • • • • •

Blade Deice Control Panel Blade Deice Test Panel Automatic Flight Control System (AFCS) Control Panel Parking Brake/Tail Wheel Lock Control Panel Chaff Dispense Control Pilot’s ICS Control Panel Left Hand Gunner ICS Control Panel Right Hand Gunner ICS Control Panel Troop Commander ICS Control Panel

Upper, lower, cabin and instrument panel illuminated pushbuttons and annunciator lights brightness is increased moving this control switch in a clockwise direction. Full rotation counterclockwise will dim the lighting to a minimum but does not turn it off.

CARGO HOOK LT

3 N O R M

2 1

O F F IR

ON

OFF

ANTI COLLISION LIGHTS

UPPER B O T H

DAY

LOWER

BLUE

DIM

STEADY

BRT

FLASH

O F F

O F F

SECONDARY WHITE

POSITION LIGHTS

NIGHT

LIGHTS CPLT MWC

DIM

LTD SW

BRT

DIM

BRT

PLT MWC

DIM

BRT

LIGHTS SECONDARY

OFF

BRT

CABIN DOME LT

UPR CSL

OFF

LWR CSL

BRT

OFF

BRT

INST PNL

OFF

BRT

FIRE DETR TEST OPER 1

WHITE

O F F

LAMPS 2

TEST

BLUE

ENG OVSP TEST A NO. 1 B A NO. 2 B

ENG SPD TRIM DECR O F F INCR UMAV002_6

10/2/2003 Page 4 of 14 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

LIGHTS - PLT MWC The Pilot Master Warning Panel dimming control is a new lighting control for the Black Hawk. The PLT MWC rotary dimmer control varies the brightness of the Pilot Master Warning Panel.

The PLT MWC lighting control is used in combination with the INSTR PNL lighting control. Moving the INSTR PNL dimmer control from the OFF position enables the PLT MWC lighting control.

Moving the PLT MWC rotary control clockwise increases the Pilots Master Warning Panel lighting intensity. Full counterclockwise rotation dims the lighting to a minimum but does not turn it off. FORMATION LTS

EXT LTS MODE

5

4

CARGO HOOK LT

3 N O R M

2 1

O F F IR

ON

OFF

ANTI COLLISION LIGHTS

UPPER B O T H

DAY

LOWER

BLUE

DIM

STEADY

BRT

FLASH

O F F

O F F

SECONDARY WHITE

POSITION LIGHTS

NIGHT

LIGHTS CPLT MWC

DIM

LTD SW

BRT

DIM

BRT

PLT MWC

DIM

BRT

# 1 ENG OUT

MASTER CAUTION

FIRE

PRESS TO RESET

LIGHTS SECONDARY

OFF

BRT

CABIN DOME LT

UPR CSL

OFF

LWR CSL

BRT

OFF

BRT

# 2 ENG OUT LOW ROTOR R. P. M.

INST PNL

OFF

BRT

FIRE DETR TEST OPER 1

WHITE

O F F

LAMPS 2

TEST

BLUE

ENG OVSP TEST A NO. 1 B A NO. 2 B

ENG SPD TRIM DECR O F F INCR UMAV002_6

10/2/2003 Page 5 of 14 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

LIGHTS - SECONDARY The secondary lighting control switch is a new lighting control on Black Hawk, which controls the legacy cockpit flood lights. It provides the ability to control the cockpit secondary lights at the same time. These are the primary source of backup lighting to the lower console and instrument panel since the glareshield lights are no longer being used.

FORMATION LTS

EXT LTS MODE

5

4

Clockwise rotation of the Secondary Lights control increases the brightness of all cockpit secondary lights. Full counterclockwise rotation dims or turns the lighting off.

CARGO HOOK LT

3 N O R M

2 1

O F F IR

ON

OFF

ANTI COLLISION LIGHTS

UPPER B O T H

DAY

LOWER

BLUE

DIM

STEADY

BRT

FLASH

O F F

O F F

SECONDARY WHITE

POSITION LIGHTS

NIGHT

LIGHTS CPLT MWC

DIM

LTD SW

BRT

DIM

BRT

PLT MWC

DIM

BRT

LIGHTS SECONDARY

OFF

BRT

CABIN DOME LT

UPR CSL

OFF

LWR CSL

BRT

OFF

BRT

INST PNL

OFF

BRT

FIRE DETR TEST OPER 1

WHITE

O F F

LAMPS 2

Cockpit Secondary Light

TEST

BLUE

ENG OVSP TEST A NO. 1 B A NO. 2 B

ENG SPD TRIM DECR O F F INCR UMAV002_6

10/2/2003 Page 6 of 14 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

The upper console panels controlled by this switch are:

LIGHTS - UPR CSL The LIGHTS UPR CSL rotary control varies the brightness of the upper console lucite panels. Clockwise rotation of this rheostat increases brightness of the upper console panels. Full counterclockwise rotation dims or turns the lighting off.

• • • •

Upper Console left, center, and right panels. Engine Control Quadrant left, center and right side. Right and left gunner’s ICS control panels. Troop Commander’s ICS control panel.

One intensity knob will be used for the upper console. It will control all information panels shown.

FORMATION LTS

EXT LTS MODE

5

4

CARGO HOOK LT

3 N O R M

2 1

O F F IR

ON

OFF

ANTI COLLISION LIGHTS

UPPER B O T H

LOWER

BLUE

DIM

STEADY

BRT

FLASH

O F F

O F F

SECONDARY WHITE

POSITION LIGHTS

DAY

NIGHT

LIGHTS CPLT MWC

DIM

LTD SW

BRT

DIM

BRT

PLT MWC

DIM

BRT

LIGHTS SECONDARY

OFF

BRT

CABIN DOME LT

UPR CSL

OFF

LWR CSL

BRT

OFF

BRT

INST PNL

OFF

BRT

FIRE DETR TEST OPER 1

WHITE

O F F

LAMPS 2

TEST

BLUE

ENG OVSP TEST A NO. 1 B A NO. 2 B

ENG SPD TRIM DECR O F F INCR UMAV002_6

10/2/2003 Page 7 of 14 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

LIGHTS - LWR CSL The LIGHTS LWR CSL rotary control varies the brightness of the lower console lucite panels. Clockwise rotation of the LWR CSL control increases brightness of lower console lights. Full counterclockwise rotation dims or turns the lighting off. The lower console panels controlled by this switch are: • • • • •

Copilots Flight Management System (FMS) Copilots ICS Control Panel KY-100 Control Panel Blade Deice Control Panel Blade Deice Test Panel

FORMATION LTS

EXT LTS MODE

5

4

• • • • • • • • • • • • • • • •

Radio Retransmission Control Panel Load Access Panel Emergency Control Panel Mission Systems Control panel Ice Rate Indicator Stabilator/AFCS Control Panel Parking Brake/Tail Wheel Lock Control Panel Select Jettison Control Panel Auxiliary Fuel Management Control Panel Pilot’s FMS Pilot’s ICS Control Panel ARC-186 Control Panel Chaff Dispense Control Panel APR-39 Radar Warning Control Panel ALQ-144 Infrared Counter Measures Control Panel Auxiliary Cabin Heater Control Panel.

CARGO HOOK LT

3 N O R M

2 1

O F F IR

ON

OFF

ANTI COLLISION LIGHTS

UPPER

POSITION LIGHTS

DAY

B O T H LOWER

DIM

STEADY

BRT

FLASH

O F F

O F F NIGHT

LIGHTS

SECONDARY WHITE

CPLT MWC

BLUE

DIM

LTD SW

BRT

DIM

PLT MWC

BRT

DIM

BRT

LIGHTS SECONDARY

OFF

BRT

UPR CSL

OFF

CABIN DOME LT

LWR CSL

BRT

OFF

INST PNL

BRT

OFF

BRT

FIRE DETR TEST OPER 1

WHITE

O F F

LAMPS 2

TEST

BLUE

ENG OVSP TEST A NO. 1 B A NO. 2 B

ENG SPD TRIM DECR O F F INCR UMAV002_6

10/2/2003 Page 8 of 14 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

LIGHTS - INSTR PNL The INSTR PNL rotary lighting control allows adjustment of the brightness of the instrument panel and MFD bezel buttons. The following control panel items are controlled by this control: • • • •

Copilot’s Outboard MFD bezels Copilot’s Inboard MFD bezels Copilot’s FD/DCP Copilot’s Collective Stick Grip

FORMATION LTS

EXT LTS MODE

5

4

• • • • •

Pilot’s Outboard MFD bezels Pilot’s Inboard MFD bezels Pilot’s FD/DCP Pilot’s Collective Stick Grip Digital Clock

The INSTR PNL rotary control will increase the brightness of the instrument panel and MFD bezel buttons when turned clockwise. Counterclockwise rotation dims or turns instrument panel lighting off. The INSTR PNL lighting control also functions to enable dimming of the Pilot and Copilot Master Warning Panels.

CARGO HOOK LT

3 N O R M

2 1

O F F IR

ON

OFF

ANTI COLLISION LIGHTS

UPPER B O T H

DAY

LOWER

BLUE

DIM

STEADY

BRT

FLASH

O F F

O F F

SECONDARY WHITE

POSITION LIGHTS

NIGHT

LIGHTS CPLT MWC

DIM

LTD SW

BRT

DIM

BRT

PLT MWC

DIM

BRT

LIGHTS SECONDARY

OFF

BRT

CABIN DOME LT

UPR CSL

OFF

LWR CSL

BRT

OFF

BRT

INST PNL

OFF

BRT

FIRE DETR TEST OPER 1

WHITE

O F F

LAMPS 2

TEST

BLUE

ENG OVSP TEST A NO. 1 B A NO. 2 B

ENG SPD TRIM DECR O F F INCR UMAV002_6

10/2/2003 Page 9 of 14 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

CABIN DOME LT The Cabin Dome Light control switch allows the use of as well as the selection of the color of the three cabin dome lights. The Cargo Hook Light control switch location has changed slightly from the UH-60A/L. Use of the Cargo Hook Lights control switch is identical to the UH-60A/L. Utility Lights The function and purpose of the UH-60M cockpit utility lights is identical to the UH-60A/L.

FORMATION LTS

EXT LTS MODE

5

4

CARGO HOOK LT

3 N O R M

2 1

O F F IR

ON

OFF

ANTI COLLISION LIGHTS

UPPER B O T H

DAY

LOWER

BLUE

DIM

STEADY

BRT

FLASH

O F F

O F F

SECONDARY WHITE

POSITION LIGHTS

NIGHT

LIGHTS CPLT MWC

DIM

LTD SW

BRT

DIM

BRT

PLT MWC

DIM

BRT

LIGHTS SECONDARY

OFF

BRT

CABIN DOME LT

UPR CSL

OFF

LWR CSL

BRT

OFF

BRT

INST PNL

OFF

BRT

FIRE DETR TEST OPER 1

WHITE

O F F

LAMPS 2

TEST

BLUE

ENG OVSP TEST A NO. 1 B A NO. 2 B

ENG SPD TRIM DECR O F F INCR UMAV002_6

10/2/2003 Page 10 of 14 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

The lights tested by this control are:

LAMPS TEST The LAMPS TEST pushbutton allows all pushbuttons, annunciator lights, and Master Warning Panels (except the Fire Detector Lights) to be illuminated in the bright or dim mode in which they are set to allow verification proper operation and illumination of the lights under test. The Fire Detector lights are tested using the Fire Detector Test Switch. The reversionary panels and annunciators on the FD/DCPs will switch from bright to controlled dimming when the instrument panel lights are turned on.

• • • • • • • • • • • • • •

Copilots FD/DCP Copilot’s Reversionary Switch Panel Pilot’s FD/DCP Pilot’s Reversionary Switch Panel Copilot’s ICS Control Panel Blade Deice Control Panel Blade Deice Test Panel AFCS Control Panel Parking Brake/Tail Wheel Lock Control Panel Chaff Dispense Control Pilot’s ICS Control Panel Left Hand Gunner ICS Control Panel Right Hand Gunner ICS Control Panel Troop Commander ICS Control Panel

10/2/2003 Page 11 of 14 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Exterior Lighting

Anti-Collision Lights UPPER/BOTH/LOWER

Formation Lights

The function and purpose of the UH-60M AntiCollision Lights are identical to the UH-60A/L.

The function and purpose of the UH-60M Formation Lights are identical to the UH-60A/L.

The ANTI COLLISION control switch location has changed slightly from the UH-60A/L. Use of the Cargo Hook Lights control switch is identical to the UH-60A/L.

The FORMATION LT control switch location has changed slightly from the UH-60A/L. The function and use of the FORMATION LT control switch are identical to the UH-60A/L. FORMATION LTS

EXT LTS MODE The External Light mode control switch is a new lighting control introduced by the UH-60M modernization. The EXT LTS control is a two-position toggle switch that allows aircraft external lighting to be visible to the eye or in the infrared spectrum. The NORMAL position enables external lighting systems to use visible lighting fixtures and disables all infrared fixtures. Conversely, the IR position disables all external lighting visible fixtures and enables all infrared fixtures.

3 N O R M

2 1

O F F IR

ANTI COLLISION LIGHTS

UPPER B O T H

LOWER SECONDARY WHITE

BLUE

POSITION LIGHTS

DAY

NIGHT

STEADY

BRT

FLASH

Position Lights

LIGHTS CPLT MWC

DIM

BRT

LTD SW

DIM

BRT

PLT MWC

DIM

BRT

LIGHTS SECONDARY

OFF

BRT

CABIN DOME LT

UPR CSL

OFF

LWR CSL

BRT

OFF

BRT

INST PNL

OFF

BRT

FIRE DETR TEST OPER 1

WHITE

The function and purpose of the UH-60M Cargo Hook Light is identical to the UH60A/L.

DIM O F F

O F F

The function and purpose of the UH-60M Anti-Collision Lights are identical to the UH-60A/L. The ANTI COLLISION control switch location has changed slightly from the UH-60A/L. Use of the Cargo Hook Lights control switch is identical to the UH-60A/L.

ON

OFF

O F F

Cargo Hook Light

CARGO HOOK LT

EXT LTS MODE

5

4

Anti-Collision Lights DAY/OFF/NIGHT

LAMPS

The function and purpose of the UH-60M Position Lights are identical to the UH-60A/L. The POSITION LIGHTS – DIM/OFF/BRT control switch location has changed slightly from the UH-60A/L. The function and use of the POSITION LIGHTS – DIM/OFF/BRT control switch is identical to the UH-60A/L.

2

TEST

BLUE

ENG OVSP TEST A NO. 1 B A NO. 2 B

ENG SPD TRIM DECR

The Cargo Hook control O F switch location has changed F INCR slightly from the UH-60A/L. Use of the Cargo Hook Lights control switch are identical to the UH-60A/L.

UMAV002_6

10/2/2003 Page 12 of 14 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Landing Light – Collective Stick The function and purpose of the UH-60M Landing Light is identical to the UH-60A/L.

Search Light – Collective stick The searchlight is a new dual-element fixture, with 400 watts visible and 200 watt infrared modes. The searchlight control switch on the collective stick is not dimmable. Control switch is push on/off and extends forward, retracts aft, and turns left and right.

Landing Light Control Switch

Search Light Control Switch Landing Light

Collective Stick Grip The Landing Light control switch location has changed slightly from the UH-60A/L. Use of the Landing Light control switch is identical to the UH60A/L.

Search Light Landing and Search Light Locations

10/2/2003 Page 13 of 14 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

NOTES

10/2/2003 Page 14 of 14 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

- Radar Warning Receiver (RWR)/Laser Detection Set (LDS) - Stormscope Weather Mapping System - Digital Map

Section 2-10.1 Flight Display System (FDS) System Description

FDS Interfaces

The Flight Display System (FDS) of the UH-60M helicopter replaces the existing conventional analog FDS, electromechanical/pneumatic gages, vertical display systems, and caution/ advisory panels of the UH-60A/L helicopter. This system allows simultaneous display of primary flight systems, tactical situational awareness, and aircraft systems status from either or both pilot stations. It allows the operator convenient access to controls associated with the primary displays.

Each display receives data directly from various aircraft avionics, primarily via digital data bus. These aircraft avionics/sensors include: • • •

Two Embedded GPS/INS (EGI) Two Air Data Computers (ADCs) One Radar Altimeter (RADALT) Transmitter/Receiver (T/R) Two Data Concentrators Units (DCUs) Two FMS Control Display Units (CDUs) - Each FMS also passes through information from the following navigation radios: - One VHF Omni-directional Ranging (VOR)/Instrument Landing System (ILS) receiver - One Automatic Direction Finder (ADF) receiver

• •

The FDS provides integrated control and display of essential flight and mission information. The FDS is configurable to accommodate both the UH-60M and UH-60M Medical Evacuation (MEDEVAC) helicopter configurations. The primary difference between the two helicopters is the navigation sensors installed. NON SECURE RADIOS WILL NOT BE KEYED

RA D IO C A LL 24432

WHEN USING ANY SECURE RADIO OR THE INTERCOM FOR CLASSIFIED COMMUNICATIONS #1 ENG OUT

#2 ENG OUT

MASTER

#1 ENG OUT

CAUTION PRESS TO RESET

FIRE

#2 ENG OUT

MASTER CAUTION

LOW ROTOR RPM

ON

ON

VID

BRT

160

10

10

3 2 1

10

1 2 3

7 136 6 5 10

N

03

B

3

A RO

28.82in

100

M

55

60

5

50

10 15

45 BK LT

VID

BRT

CON

ETC

40 35

CON

ATT

FMS

ADC

HDG

DCU

REV

REV

REV

REV

NAV SRC

OBS

REF SEL

BRG 1

BRG 2

REF ADJ

LOC CAP

GS CAP

RALT 250

ALTP 2500

ALT 1500

IAS 120

HDG 240

VS 500

SEL

CPLD ****

OFF

0

120

BRT

VID

BRT

40 22 20 00

180

VID

ON

OFF

OFF

BK LT

LOW ROTOR RPM

VID

BRT

VID

BRT OFF

REV

PRESS TO RESET

FIRE

ON

30

20

CTRL

OFF

ON

GA

DECL CAP

HVR P-SYNC

P-SYNC

P-SYNC

P-SYNC

P-SYNC

BK LT

VID

BRT

CON

25

ENGINE IGNITION

NAV SRC

OBS

REF SEL

BRG 1

BRG 2

REF ADJ

LOC CAP

GS CAP

GA

DECL CAP

RALT 250

ALTP 2500

P-SYNC

P-SYNC

ALT 1500

IAS 120

HDG 240

P-SYNC

P-SYNC

VS 500

HVR P-SYNC

BK LT

VID

BRT

CON

CPLD **** ATT

FMS

ADC

HDG

DCU

REV

REV

REV

REV

REV

UMAV001

The FDS provides display features for numerous flight management functions, including: • •

• • •

Primary flight data display Navigation/guidance data display - Deviation display - Map (Flight Plan) display - Hover guidance display Engine/Transmission indication Crew Alerting System (CAS) display Sensor data display - Forward Looking Infrared (FLIR) sensor (MEDEVAC only)

• •

• • • • •

One FM receiver One Flight Control Computer (FCC) (note- left and right side displays interface to independent output FCC channels) One FD/Display Control Panel (FD/DCP) (note- left and right side displays interface to separate FD/DCPs) One Stormscope system RWR/LDS One Improved Data Modem (IDM) Two Data Transfer Systems (DTS)

10/2/2003 Page 1 of 48 UH60M AIRCRAFT1 FIRST FLIGHT TRAINING DATA

•

Three Multifunction Displays (MFDs) (the on-side display and the two crossside displays)

Operator Interface The UH-60M FDS allows operator input to directly affect the display presentation and content via built in display control features (bezel controls) and external interfaces to aircraft controls. The operator control capabilities include: • •

•

Each of the display modes may be selected in any of the display positions. The displays default to particular modes/formats on power up based on the position they are installed in. These defaults are: • •

Outboard displays (1 and 4 positions) default to Primary Flight Display (PFD)Full display mode/sub-mode Inboard displays (2 and 3 positions) default to Engine Indication Crew Alerting System (EICAS) display

Display mode and format selection Display selection and setting controls including: - Display feature selections - Display parameter settings Data source selection

Multifunction Display – Primary Flight Display 10/2/2003 Page 2 of 48 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Flight Display System Components Multifunction Displays (MFDs) The UH-60M FDS consists of four MFDs, installed on the instrument panel from left to right. Each MFD is a 6”x 8” landscape Active Matrix Liquid Crystal Display (AMLCD) and has a resolution of at least 768 vertical and 1024 horizontal color pixels. It also provides a minimum of 262,144 selectable colors. The MFDs perform normal operations up to 20,000 ft above sea level and withstands transport altitude to 50,000 ft above sea level.

The MFD has incorporated an embedded BuiltIn Test (BIT) to determine the health and status of the MFD The BIT includes: •

•

•

Start-Up BIT (SBIT) – Automatically runs to indicate the readiness of the MFD. The SBIT will only run when the aircraft is on the ground. Continuous BIT (CBIT) – Continuously runs in flight or on the ground without interfering with the MFD operation and will report failures as it occurs. Initiated BIT (IBIT) – Determines the status of each Shop Replaceable Unit (SRU) and detects faults to the unit level. The IBIT will only run when the aircraft is on the ground.

The MFDs sends and receives messages via the MIL-STD-1553B data bus and ARINC-429 digital data bus to related navigation and communication systems.

Multifunction Display – EICAS Display 10/2/2003 Page 3 of 48 UH60M AIRCRAFT1 FIRST FLIGHT TRAINING DATA

Flight Director /Display Control Panel (FD/DCP) The FD/DCP, located below each operator’s inboard MFDs, optimizes both visibility and physical access to the panels. It provides independent pilot and copilot control of both FD and display selections/settings. The displays interface with FD/DCP to provide additional pilot selection and setting control capability. The FD/DCP interface provides control of the following selections/settings through the displays: • • • • • • • • •

Navigation Source selection Bearing Data Source selections Selected Course settings Selected Heading setting Time Display selection Barometric Correction setting Low RAD ALT setting High RAD ALT setting FCC selection/settings

Each display interfaces to its on-side FD/DCP, with one of the on-side displays functioning as a “master” controller, performing all of the DCP related selections/setting for that side of the cockpit. The other on side display functions as a “slave,” receiving the selection/settings from the “master” display. The outboard PFD displays are the default display control “masters,” with the inboard displays capable of assuming this function if the outboard displays are failed. A setting performed on either side of the cockpit is common to all displays. If both pilots were to try to adjust selected heading at the same time, whichever pilot started turning his control first would assume control of the setting until that pilot stops adjusting the setting. OBS

REF SEL

BRG 1

BRG 2

REF ADJ

LOC CAP

GS CAP

GA

DECL CAP

RALT 250

ALTP 2500

P-SYNC

P-SYNC

ALT 1500

IAS 120

HDG 240

P-SYNC

P-SYNC

The reversionary switch panels provide a means of manually switching data sources for the pilot and copilot MFDs. During normal operation, the MFD uses its on-side sources for display of information. If one of these sources fails, the pilot would press the Reversionary Switch for the failed source. The MFD would then revert to the alternate source. There are two reversionary switch panels for the UH-60M Blackhawk helicopter that are identical to each other. One reversionary switch panel is located on the Pilot’s instrument panel below the outboard MFD, and the other reversionary switch panel is located on the Copilot’s instrument panel below the outboard MFD. Both reversionary switch panels have five pushbuttons labeled: • • • • •

FD/DCP Control Coordination

NAV SRC

Reversionary Switch Panel

VS 500

ATT HDG ADC FMS DCU

REVERSIONARY SWITCH PANEL

CPLD ****

HVR P-SYNC

UMAV001_5

FD/DCP

10/2/2003 Page 4 of 48 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Flight Display System Controls

•

Display Mode/Format Selection Each of the display modes may be selected in any of the display positions. The displays default to particular modes/formats on power up based on the position they are installed in. These defaults are: • •

Outboard displays (positions 1 and 4) default to PFD-Full display mode/submode Inboard displays (positions 2 and 3) default to EICAS display

The selection of the display modes and submodes is accomplished via bezel buttons located along the bottom of the display that are dedicated to this function.

•

Software Controlled Bezel Control (bottom buttons of the MFD) • • • • • •

MFD Bezel Controls There are a total of 24 controls located (6 on each side), around the MFD display. Each button has a turn on/off functionality assigned. The functionality of each button depends on the MFD display application. The display includes numerous bezel selection/setting controls, which are functional in some or all display modes. The bezel buttons have dedicated functions and have functional labels on the bezel or the button, or be “soft” bezel buttons – where software and a label defining the bezel button function program the function is displayed on the LCD adjacent to the bottom. The bezel controls are defined as the following:

Video contrast setting – The last two buttons on the bottom right side of the MFD. Up arrow increases video display contrast (tactical display only). Down arrow decreases video contrast (tactical display only). Display power on/off control switch – The first button on the top left side of the MFD

Display mode/sub-mode select bezel controls Distance display units selection (nm or km) Navigation map display range selection CAS message window paging controls CAS message window cancel/recall selection Hover velocity source selection

These bezel controls only affect the display on which they reside. Actuation of a bezel control on one display does not cause the other displays to change. Bezel Control Function Enable Many of the bezel button controls are active in certain display modes and/or only with certain display features active. When the “soft” bezel control function is active, the associated bezel button label is displayed, and when the bezel button control function is inactive the associated label is not displayed. For the “hard” bezel buttons, the labels are always displayed on the button and/or bezel.

Dedicated Bezel Controls •

•

Display luminance setting – The last two buttons on the bottom left side of the MFD. Up arrow increases backlight brightness. Down arrow decreases backlight brightness. Video brightness setting – The first two buttons on the top right side of the MFD. Up arrow increases video brightness (tactical display only). Down arrow decreases video brightness (tactical display only).

10/2/2003 Page 5 of 48 UH60M AIRCRAFT1 FIRST FLIGHT TRAINING DATA

Dedicated Bezel Control Buttons

Software Controlled Bezel Buttons

10/2/2003 Page 6 of 48 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

FD/DCP Switch Functions Navigation Source (NAV SRC)/Omnibearing Select (OBS) Control The pilot via a knob on the FD/DCP selects the current Navigation Source (NAV SRC), with the selection applying to the data displayed on that side of the cockpit. Each detent on the knob selects a different navigation source for display on the Horizontal Situation Indicator (HSI). The display receives inputs from the FD/DCP and determines the current navigation source from the sources available. The functions controlled by this NAV SRC control include: • • • •

NAV SRC/OBS NAV SRC

OBS

REF SEL

BRG 1

BRG 2

REF ADJ

LOC CAP

GS CAP

GA

DECL CAP

RALT 250

ALTP 2500

P-SYNC

P-SYNC

ALT 1500

IAS 120

HDG 240

P-SYNC

P-SYNC

VS 500

CPLD ****

HVR P-SYNC

UMAV001_5

BRG 1/BRG 2 FD/DCP FD/DCP Reference Select Control

VOR ILS FMS FM HOMING

The navigation source selects between VOR or ILS based on the frequency the radio is tuned to. There is single FMS navigation source selection between FMS 1 or FMS 2 depending on which side (pilot or copilot) is displayed. Adjusting the OBS is performed during any time of ground and flight operations as long as a navigational source is selected, which displays the selected course. Bearing 1/Bearing 2 (BRG 1/BRG 2) Source Select Control The current bearing 1 and bearing 2 sources are selected by the pilot via knobs on the FD/DCP, with the selection applying to the data displayed on that side of the cockpit. The display receives inputs from the FD/DCP and determines the current bearing source (1 and 2) from the sources available. The functions controlled by this by this BRG 1/BRG 2 control include: • • •

There is single FMS navigation source selection between FMS 1 or FMS 2 depending on which side (pilot or copilot) is displayed.

VOR/ILS FMS ADF

Four of the settings/selections performed by the FD/DCP are performed by a single, multifunction reference set (“Ref Sel”) control. This control has two knobs associated with it, one for selecting the function that is active, the setting/selection and the other for adjusting the parameter. The REF ADJ control adjusts the value of the reference selected by the REF SEL control. The functions controlled by this “Ref Set” control include: • • • •

“BARO” – Barometric Correction setting “RA-L” – Low Radar Altitude setting “RA-H” – High Radar Altitude setting “TIME” – Time Display selection

The reference set functions are sequenced through in the order indicated above when the reference set function select control is turned clockwise, and in the opposite order when the control is turned counterclockwise. REF SEL NAV SRC

OBS

REF SEL

BRG 1

BRG 2

REF ADJ

LOC CAP

GS CAP

GA

DECL CAP

RALT 250

ALTP 2500

P-SYNC

P-SYNC

ALT 1500

IAS 120

HDG 240

P-SYNC

P-SYNC

VS 500

CPLD ****

HVR P-SYNC

UMAV001_5

The VOR/ILS navigation radio only provides a bearing when tuned to a VOR frequency. There is no bearing for ILS frequencies. If an ILS frequency is tuned, the bearing data displayed reflects this condition.

REF ADJ FD/DCP NOTE At this time all other switch functions will not be available for first flight.

10/2/2003 Page 7 of 48 UH60M AIRCRAFT1 FIRST FLIGHT TRAINING DATA

Reversionary Switch Panel Functions ATT REV (Attitude Reversionary) When the ATT REV is pressed to reversionary position, it displays attitude information from the other EGI. ATT1 or ATT2 is annunciated in yellow on the failed side and white on the operating side with a box indicating both sides are using the same source. HDG REV (Heading Reversionary) When the HDG REV is pressed to reversionary position, it displays heading and turn rate information from the other EGI. MAG1 or MAG2, or TRU1 or TRU2 is annunciated in yellow on the failed side and white on the operating side with a box indicating both sides are using the same source.

FMS REV (Flight Management System Reversionary) The FMS REV allows continued use of all displays after a single FMS failure. FMS is annunciated in yellow on the failed side and white on the operating side with a box indicating both sides are using the same source. DCU REV (Data Concentrator Unit Reversionary) The DCU REV allows continued use of all displays after a DCU failure. In such an event, the DCU REV switch on the failed side is placed in the reversionary position. DCU is annunciated in yellow on the failed side and white on the operating side with a box indicating both sides are using the same source.

ADC REV (Air Data Computer Reversionary) When the ADC REV is pressed to reversionary position, it displays air data information from the other EGI. ADC1 or ADC2 is annunciated in yellow on the failed side and white on the operating side with a box indicating both sides are using the same source.

Reversionary Switch Panel

10/2/2003 Page 8 of 48 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Flight Display System Indications Engine Indication, Crew Alerting System (EICAS) The UH-60M EICAS display integrates display of primary engine/transmission data, fuel quantity data, and a CAS, alerting messages. The EICAS display also provides annunciation of engine/propulsion parameter monitoring functions.

EICAS display The primary EICAS page display information consists of three groups of parameters/data: • • •

Engine/Rotor/Transmission data (engine #1 and #2, main rotor, main transmission) Fuel Quantity data Cautions and Advisories

Several characteristics of the primary flight display (PFD) are related to multiple functional elements of the display to include: • •

Data source annunciations Parameter comparison and annunciation

Engine Indication Crew Alerting System (EICAS) Display

10/2/2003 Page 9 of 48 UH60M AIRCRAFT1 FIRST FLIGHT TRAINING DATA

EICAS Data Source Annunciation The source for EICAS engine/rotor/transmission, fuel, and CAS data (Data Concentrator Unit (DCU1 or DCU2)) annunciation operates and appears exactly as the engine data (DCU) source annunciation on the PFD and is located in the same display position.

DATA SOURCE ANNUNCIATION DISPLAY

Engine/Rotor/Transmission Data Display

display by an “engine OT” symbol, which again provides quick recognition of which engine is not operational, reducing the opportunity for confusion, which could result in shutting the wrong engine. Engine/Transmission Icons There are six groups of engine/transmission parameters accomplished by icons. The engine parameters are grouped in two icons per engine. The main transmission parameters are grouped, and rotor speed is grouped with the engine 1 and 2 power turbine speed. The icons are always displayed regardless of whether or not there is any valid data displayed within them. The icon color reflects the condition of the parameters it encloses, with the icon taking on the color of the most severe limit exceeding currently active for those parameters.

The EICAS display of engine/rotor/transmission parameters functionally and physically groups related indicators on the display. This is accomplished by enclosing the related indications within “icons”. The indicators themselves consist of vertical tapes and numeric readouts, with labels annunciating what parameter the indicators represents. Loss of an engine is indicated on the

Engine/Rotor/Transmission Icons

10/2/2003 Page 10 of 48 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Fuel Quantity Data Display

Total Fuel Quantity Display

The primary EICAS display of fuel quantity consists of:

The total aircraft fuel quantity is displayed as a numeric readout. The aircraft total fuel quantity readout is centered within a white box outline, and has a maximum of five digits and a range of 0 to 99999 lbs. The readout includes a label displayed above the numeric fuel quantity value within the readout outline. The label “TOTAL” is white and the readout is normally green. If a fail occurs, the label remains white but the readout is displayed in red text and does not include any quantity for failed tanks.

• • •

Main tank fuel quantity tapes and readouts Left and right auxiliary tank summary readouts Total aircraft fuel quantity readout

Main Fuel Quantity Display Each main tank fuel quantity is represented on the display by a tape and readout. The main fuel quantity display includes several labels – “FUEL” and “MAIN” displayed between the two tapes, and individual labels on each tape (”1” on the left tape and “2” on the right tape). The main tank quantity fuel tapes are scaled linearly from 0 lbs at the bottom of the tape to the maximum main tank quantity, 1200 lbs, at the top of the tape. There is a yellow low fuel quantity caution line displayed at the 172 lb position on each tape. If the fuel quantity readout for either tank is at or below fuel caution value (172 lbs) the readout and tape for that tank will change to yellow. The “MAIN” and “FUEL” labels and associated labels are white and always displayed.

Main Fuel Quantity Display

10/2/2003 Page 11 of 48 UH60M AIRCRAFT1 FIRST FLIGHT TRAINING DATA

Crew Alerting System (CAS) Display The CAS display provides the pilot with immediate indication of a malfunction, a degraded system state, or an abnormal system state, which the pilot either needs to be aware of, or must respond to by taking some action. The CAS function of the EICAS page displays alerting text messages to the pilot indicating the aircraft/system malfunction or condition. Message Display Areas CAS Caution and Advisory messages are presented in seven-display message “windows” distributed around the EICAS display. Of the

seven message windows, one (the lower left) is designated for display of advisory messages only and the other six windows are designated for display of caution messages only. Within each of the windows, messages are displayed in chronological order, with the most recently activated message displayed at the top of the window. The messages displayed lower in the window have been active for progressively longer periods; the further down they appear in the window. The lower left and right message windows have boxes around them because they have unique characteristics that allow them to display more active messages than the message window size allows.

Caution Message Windows

Advisory Message Window

EICAS Caution and Advisory Message Areas

10/2/2003 Page 12 of 48 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Message Characteristics All messages can be up to 18 alphanumeric characters in length (including spaces). The advisory messages are always displayed as white text. The caution messages are displayed as black text on yellow background when they initially are activated, and revert to yellow text when acknowledged by the pilot. The acknowledgement is accomplished by pressing the master caution light on a crew-alerting panel in the cockpit. The master caution light is illuminated any time a new message becomes active and remains active until the pilot acknowledges the message or the new message becomes deactivated causing the message to go away.

content of a page may change during viewing. This can result in some active alerts disappearing or new alerts appearing on the current page as they are moved upward or downward on the alert list. This occurs when a higher priority alert becomes inactive, creating a hole in a lower numbered page or a new alert became active, pushing an older onto a higher numbered page.

Caution Message Overflow Operation Each caution message is assigned one of the Caution display windows as its default location. However, the size of each of these message windows is limited. If the number of messages active in one of the caution message windows exceeds the space available in that window, then messages begin to overflow. In a message overflow condition, the oldest message in the window is removed from that window and displayed in the lower RH caution window, which is designated as the caution message overflow window. All advisory messages are assigned to the lower left message window, so there is no overflow operation associated with the lower left advisory window. Message Paging Operation Both the lower left advisory and lower right caution overflow windows allow paging of messages to handle situations where there are more messages active than can be displayed at one time. If the lower right caution window fills up with messages, and then additional messages become active in that window, then message paging becomes active for that window. The paging of message sequences through a full window of messages at a time, with each press of a page up or down, with the bezel button key. When the first page of messages is displayed, the message paging bezel key labels are displayed. The message “pages” are continuously updated, so the

10/2/2003 Page 13 of 48 UH60M AIRCRAFT1 FIRST FLIGHT TRAINING DATA

Primary Flight Display

•

The UH-60M FDS integrates the function of an Attitude Directional Indicator (ADI), flight instruments, and a navigation display into a combined electronic display format, “The Primary Flight Display” (PFD).

•

PFD Sub-Modes

•

Full Rose HSI (“PFD FULL”) – Traditional HSI with a 360° compass rose ARC Map (“PFD ARC”) – 80° compass arc with sector map displaying flight plan, navigation data and weather radar data Hover – 360 °compass rose, with precision hovering flight data

The PFD includes three sub-modes, which provide different navigation data presentations. These sub-modes are:

Stabilator Indicator

Engine Instrument “Power Pod” 10/2/2003 Page 14 of 48

Airspeed Indicator

Attitude Direction Indicator

Barometric Altitude Indicator

Navigation Data Display Primary Flight Display Layout UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Vertical Speed Indicator

Radar Altitude Indicator

Attitude Direction Indicator (ADI) The electronic attitude direction indicator (ADI) displays the following information: Attitude Display • • • •

Aircraft symbol Artificial horizon Pitch scale Roll attitude pointer and scale

Deviation Display • • • •

Lateral deviation scale (“Standard” scale or “Localizer” scale) Lateral deviation pointer (“Standard” or “Localizer” scale) Vertical deviation scale Vertical deviation pointer (GS or Vertical Navigation (VNAV) pointer)

Pitch Attitude Scale The pitch scale shows the aircraft pitch angle. The pitch scale moves vertically as the aircraft pitches. Pitch angle is read at the aircraft symbol nose (the center dot of the split axis symbol). The pitch scale is 180° (+/- 90°) full scale with 40° shown on the ADI. Red warning chevrons are displayed on the scale at extreme pitch angles. Roll Attitude Pointer and Scale The roll attitude pointer and scale provide an indication of the aircraft bank angle, in addition to the sky/ground indication. The roll attitude scale has markings at 0 (filled triangle), +/-10°, +/-20°, +/-30°, +/-45° (open triangles) and +/60°. The roll pointer moves along the fixed scale to indicate the current roll attitude. Pitch Scale Display Roll Scale Roll Pointer

Rate of turn pointer and scale Comparator warning annunciations (“PIT”/”LOC”/”GS”) Radar Altitude alert annunciation MB annunciation FD display • • • • •

Pitch command bar Roll command bar Collective command FD mode annunciation Active FD annunciation

Attitude display Artificial Horizon and Aircraft Symbol The stationary aircraft symbol in the center of the ADI provides the aircraft reference for display of attitude and related symbology. Attitude is shown by the blue-colored sky and brown-colored ground moving about the aircraft symbol to provide a quick visual cue to current aircraft attitude.

Aircraft Symbol Reference Point And Outer Edges

10/2/2003 Page 15 of 48 UH60M AIRCRAFT1 FIRST FLIGHT TRAINING DATA

Unusual Attitude

Failed Attitude

If the aircraft attitude reaches extreme limits, several scales, indicators, and FD annunciations not related to attitude are removed. The unusual attitude pitch limits are 30° nose up and 20° nose down. The bank limit is 65° left or right. All ADI information is restored when the pitch angle is reduced to between 25° nose up and 125° nose down, and bank angle is reduced to less than 60°. The items removed are:

If valid pitch and roll inputs are not received, the attitude display symbology is removed as well as other symbology that uses attitudes as a reference to include FD command cues. An attitude fail flag is displayed if the input is failed or the input data is not received.

• • • • • • •

FD pitch, roll, and collective commands and active FD and mode annunciations. FD reference indicators (airspeed, altitude, Vertical Speed Indicator (VSI), and radar altitude) Lateral and vertical deviation indicators Turn rate scale/indicator Radar Altitude alert annunciation MB annunciator Associated miscompare/fail flags

Failed Attitude ADI Deviation Display Lateral Deviation pointer and scale

Unusual Attitude Indication

The lateral deviation pointer and scale provide lateral course deviation guidance with the selected navigation source. The lateral deviation scale has two dots on each side of the center index mark. The lateral deviation pointer moves +/- 2.5 dots. The lateral deviation pointer (plus sign) shows whether the course is left or right of the aircraft (center of scale). A “Standard” lateral deviation scale is displayed when the selected navigation source is an FMS source, VOR/ILS receiver tuned to a VOR frequency, VOR/ILS receiver tuned to an ILS frequency flying a backcourse approach, or an FM Homer signal. An “Expanded” lateral deviation scale is displayed when the NAV source is ILS and not flying a backcourse approach. The Localizer

10/2/2003 Page 16 of 48 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

deviation pointer is a diamond shaped symbol, which moves only two deviation dots (one per side) behind the lateral deviation scale (Expanded or Standard) to indicate the deviation. When lateral deviation is invalid, the lateral deviation scale and pointer are removed. The NAV source annunciation is displayed as a fail flag when the lateral deviation input data or its status is failed. There is no flag displayed in place of the lateral deviation symbology.

Standard Deviation Scale/Pointer

Expanded Lateral Deviation Scale/ Localizer Deviation Pointer

Lateral Deviation Display

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Vertical Deviation Pointer and Scale The vertical deviation scale and pointer indicate vertical deviation from the ILS GS or FMS VNAV guidance when those navigation sources are selected. The GS deviation pointer, a diamond sign, displays when NAV source is ILS and the navigation radio is receiving localizer and GS signals. The vertical deviation pointer, a plus sign, also displays when VNAV source is FMS and the FMS is supplying vertical navigation inputs. The vertical deviation pointer indicates that the vertical path is above or below the aircraft (center of scale). The limit of movement for the GS deviation pointer is the same distance from the center of scale as the limit of movement for the VNAV deviation pointer (2.5 dots above or below center of scale). When the vertical deviation input (GS from VOR/ILS or VNAV deviation/scaling from FMS) is invalid, the vertical deviation scale and pointer are removed. When the status of the input is failed, a fail flag is displayed.

VNAV Deviation Pointer/Scale

Vertical Deviation Fail Flags

GS Deviation Pointer/Scale

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Rate-of-Turn indicator The rate-of-turn indicator consists of a white bar and a three-box scale. When the turn rate is either left or right, the marker deflects in the turn direction from the center box. Deflection is linear from zero turn rate to maximum deflection. When the bar is centered between the center and outer scale markings in either direction, it indicates a standard rate turn (3°) in that direction. When the bar moves to a position where it is centered under the outer scale markings, it indicates a 6° per sec turn rate (double a standard rate turn). The bar remains parked in the 6° per sec (left or right direction) when inputs are greater than 6° per sec.

Rate of Turn Indicator

Rate of turn scale

Rate of turn pointer

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Barometric Altitude Display (BAR ALT) Barometric altitude (BAR ALT) display symbology includes the following: • • • • •

BAR ALT dial and pointer BAR ALT rolling digital readout Barometric correction set readout Selected BAR ALT symbology Pre-selected BAR ALT symbology

BAR ALT Dial/Pointer The analog BAR ALT presentation provides information that is not easily interpreted from the rolling digital readout that provides accurate interpretation of current altitude. The white BAR ALT dial represents a 1000-ft altitude window. The dial and pointer alone do not provide complete BAR ALT resolution. The readout is required to determine which 1000-ft window the aircraft is in. The dial includes tick marks at 20-ft increments (with long tick marks at 100-ft increments), numeric labels of tick marks every 100 ft, and a label indicating the units of the displayed parameter (“FEET”). When the BAR ALT goes below –20 ft, the normal dial is replaced with a dial representing negative BAR ALT. The negative dial remains displayed until the BAR ALT rises to 0 ft. The BAR ALT pointer travels around the dial to indicate the current BAR ALT. The pointer travels clockwise for increasing altitude and counterclockwise for decreasing altitude. BAR ALT Rolling Digit Readout The BAR ALT digit readout provides a digital readout with numeric indication of 20 ft. The BAR ALT color is white and includes up to five

digits and a surrounding window, which helps distinguish the readout from the dial and pointer. A “NEG” label is displayed to the left of the readout (within the readout window) when BAR ALT is less than 0 ft. Barometric Correction Set Control/Readout The barometric correction set readout provides the pilot ability to accurately set the barometric correction factor used by the ADC to compute BAR ALT from pressure altitude. The value is displayed and adjusted in inches of mercury (In Hg). Each pilot sets the barometric correction setting from HSI FD/DCP using the FD/DCP reference set control. The settings adjust up for clockwise rotation and down for counterclockwise rotation when the selected function for that knob is barometric correction. The displays compute the barometric correction from the reference set knob turn information received and coordinate with the on-side displays to assure that there is a common setting. The barometric correction setting has a range of 22.00 to 31.00 In Hg, with a resolution of 0.01 Hg. The barometric correction settings default to 29.92 (“standard pressure”) when the reference set knob “sync” button is pushed. The barometric correction setting defaults to the last set value on power up. Selected BAR ALT Symbology The selected BAR ALT symbology provides the pilot a reference altitude indication that is associated with a currently active FD mode. The selected BAR ALT symbology includes a digital readout of the selected altitude value and an indicator that is displayed on the BAR ALT dial. It is enabled by the following circumstances:

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• • • •

BAR ALT dial is displayed FD flag fail is not displayed A valid input from the FD Unusual attitude conditions do not exist

Pre-selected BAR ALT Symbology The pre-selected BAR ALT symbology provides the pilot a second reference altitude indication associated with a currently active flight mode. The pre-selected BAR ALT symbology includes a digital readout of the pre-selected altitude value and an indicator that is displayed on the BAR ALT dial. It is enabled by the following circumstances: • • • •

BAR ALT dial is displayed FD flag fail is not displayed A valid input from the FD Unusual attitude conditions do not exist

Vertical Speed Display Vertical speed display symbology includes the following: • • •

Vertical speed scale, pointer, and tape Vertical speed readout and direction indicator Selected vertical speed symbology

Vertical Speed Indicator (VSI) The analog vertical speed presentation provides information that is not easily interpreted from the readout that provides the accurate interpretation of current altitude. The white vertical speed scale provides a display range of +/-4000 ft per minute (fpm). The scale includes markings at 500 fpm increments throughout its entire range (longer tick marks at 1000 fpm), with numerics at the 1000 fpm graduations. The vertical speed pointer angles up and down the scale to indicate the current vertical speed. If the input exceeds +/-4500 fpm, the pointer remains at that position above/below the scale. The vertical speed tape follows the pointer motion to enhance the pilots’ ability to recognize the direction of pointer deflection and determine information. The tape is blue when the displayed vertical speed is greater than 0 fpm and brown when the displayed vertical speed is less than 0

fpm (drag line is not displayed for vertical speed of 0 fpm). Vertical Speed Readout The vertical speed readout provides a numeric indication of aircraft vertical speed. The vertical speed readout color is white and includes a direction indication, which improves the pilots’ ability to distinguish vertical speed direction when the vertical speed is near 0 fpm and the pointer is nearly horizontal. Selected Vertical Speed Symbology The selected vertical speed symbology provides the pilot with a reference vertical speed indication that is associated with a currently active FD mode. The selected vertical speed symbology includes a digital readout of the selected vertical speed value and an indicator that is displayed on the vertical speed scale. The selected vertical speed symbology is enabled by the following circumstances: • • • •

Vertical speed scale is displayed FD flag fail is not displayed A valid input from the FD Unusual attitude conditions do not exist

Selected VSI The selected vertical speed indicator is positioned at the selected vertical speed value on the vertical speed scale when the selected vertical speed symbology is enabled. The VSI is removed if the selected vertical speed value exceeds the vertical speed range on the scale. Selected Vertical Speed Readout The selected vertical speed readout is displayed above the vertical speed readout, with a direction indication to indicate whether the flight director reference vertical speed is up or down. Display of the selected vertical speed readout is normally under control of the flight director, allowing the flight director to turn the readout on under the following circumstances:

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• •

When the flight director mode is initially activated When the selected vertical speed setting is being changed by the pilot

Radar Altitude Display The Radar Altitude is represented as a round dial analog indicator with a digital readout. The radar altitude dial and pointer provide the analog indication of the radar altitude. The analog presentation provides information that is not as easily interpreted from the digital readout. The white radar altitude dial provides a display range of 0-1000 ft. The dial consists of three separate linear ranges: • • •

0 to 100 ft (10 ft increments) 100 ft to 400 ft (50 ft increments) 400 ft to 1000 ft (100 ft increments)

“LO” and “HI” Radar Altitude Symbology The “LO” radar altitude setting/display provides a pilot selectable reference setting that defines the altitude at which the low radar altitude alert is activated when the aircraft descends. The low radar altitude alert can be used as a monitor when flying an approach or when in a hover, to alert the pilot that the aircraft is passing through the threshold defined by the low radar altitude setting. The “HI” radar altitude symbology provides the pilot a high radar altitude reference that triggers an alert indication if the aircraft climbs through the high altitude setting. The high altitude can be used as a monitor during certain phases of flight to alert the pilot that the aircraft is passing through the threshold defined by the high radar altitude setting. Both the “LO” and “HI” radar altitude values can be set by either pilot from the FD/DCP using the reference set control on the panel to adjust the value up (clockwise) or down (counterclockwise) when the selected function for that knob is high or low radar altitude. The displays compute the altitudes (HI or LO) from the reference set knob turn information received and coordinate with the on-side displays to assure that there is a common setting.

If the scale is currently displayed, it remains displayed until the radar altitude readout value rises above 1050 ft, at which point the scale is removed along with the pointer. The radar altitude pointer travels around the dial to indicate the current radar altitude. The pointer travels clockwise for increasing radar altitude and counterclockwise for decreasing radar altitude. The pointer color is white when both the high and low radar altitude alert conditions are inactive. When either the high or low radar altitude alert is active, the pointer color is yellow. Radar Altitude Readout The radar altitude readout provides a numeric representation of aircraft height above ground with displayed resolution that varies with the radar altitude. The radar altitude readout includes a label “RA” displayed below the readout. The readout displays to four digits, with a range from –20 ft to 1500 ft. The readout and label color is set to match the radar altitude pointer color (either white or yellow).

Selected Radar Altitude Symbology The selected radar altitude symbology provides the pilot a reference altitude indication that is associated with a currently active FD mode. The selected radar altitude symbology includes a digital readout of the selected altitude value and an indicator that is displayed on the radar altitude dial. It is enabled by the following circumstances: • • • •

Radar altitude dial is displayed FD flag fail is not displayed A valid input from the FD Unusual attitude conditions do not exist

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Radar Altitude Alert Annunciation A radar altitude alert annunciation is displayed when either a “low” alert or a “high” alert occurs. The radar altitude passing through either the pilot selected “low” or “high” radar altitude settings activate the alerts. The “low” and “HI” radar altitude alert annunciations are displayed in large yellow text. The logic for determining the alert conditions prevents both alerts from being activated simultaneously. The alert annunciation flashes for the first five seconds it is displayed, then is displayed constantly while the alert condition is active.

RAD ALT “LO” ANNUNCIATION

RAD ALT “HI” ANNUNCIATION

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PFD Navigation Data Display The lower center and lower RH areas of the PFD is the navigation data display area. The data presented in the HSI area varies by PFD sub-mode. The different sub-modes are: • • •

Full Rose Horizontal Situation Indicator (HSI) mode (“FULL”) Arc Map mode (“ARC”) Hover mode (“HOVER”)

The full rose HSI provides a full rose compass presentation with a lateral deviation indication, much like the traditional electromechanical HSI indicator. The electronic HSI provides numerous additional features that enhance the deviation/heading display and provide increased navigation capability in the cockpit. The PFDFull navigation display data includes the following groups of features: • • • • • •

• • •

Heading Display Selected heading display Ground track display Ground speed display Wind display Navigation source data display, including: - Navigation source annunciation - Navigation system operational mode - Navigation station identifier - Selected course/desired track - Course deviation - To/From indication - Distance associated with selected navigation source - Time readout - FMS navigation state annunciations FMS message alert display FMS command heading display Bearing display

Heading Display

• •

Heading reference/source annunciation Directional gyro (DG) annunciation

Aircraft Symbol The aircraft symbol is fixed reference point at the center of the compass rose; around which all symbology oriented relative to heading rotates (including the compass rose. The nose of the aircraft points towards the current aircraft heading (lubber line). The aircraft symbol is always displayed, regardless of the status of the heading input. Compass Rose The full (360°) compass is an azimuth scale providing a graphical indication of aircraft heading by rotating about the aircraft symbol to place the scale markings/position corresponding to the aircraft heading at the lubber line Compass Benchmarks The compass benchmarks provide angular references around the compass rose enhance the pilots’ ability to quickly interpret the position of indicators displayed around the compass rose. The compass benchmarks are fixed in position. They are displayed at 45° intervals around the compass rose, except that no benchmarks are displayed at 0° and 180°. Heading Readout The heading readout provides a numeric indication of the displayed heading value that corresponds to the graphical indication provided by the compass rose. The heading readout range is 001° - 360°. The readout always displays three digits centered in a white box. The bottom edge of the readout box is triangular with the point extending down to the compass rose.

The aircraft heading display provides the pilot with aircraft orientation information necessary for navigation and aircraft control. The heading display symbology includes: • • • •

Aircraft Symbol Compass rose Compass benchmarks Heading readout/readout box/lubber line

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HSI Full Rose Display

HSI Arc Map Display

Hover Heading Display 10/2/2003 Page 25 of 48 UH60M AIRCRAFT1 FIRST FLIGHT TRAINING DATA

Heading Reference/Source Annunciation The heading reference/source annunciation indicates the current heading reference (magnetic north or true north) and indicates the source of the displayed heading. The annunciation is boxed (yellow) if both pilots have the same source selected. Directional Gyro (DG) Annunciation The DG annunciation is displayed below the heading reference/source annunciation when heading data source is operating in the DG mode (not slaved to the associated magnetic flux detector). When displayed, the color of the annunciation matches the color of the heading reference/source annunciation. Heading Fail When the heading input is invalid, the compass rose, compass benchmarks, and heading readout/readout box are removed from the display and the heading fail flag is displayed. If active, the heading reference/source and the DG annunciations are still displayed to provide the pilot with additional information for diagnosing the problem or to select an alternate source.

HEADING FAIL DISPLAY Ground Track Display The aircraft ground track is indicated by the track indicator on the compass rose, which is positioned at the current aircraft ground track value on the compass rose. The ground track indicator is cyan in color and overwrites the compass rose markings, to enhance visibility.

The track indicator is removed from the display when: • • •

The displayed aircraft heading is removed (aircraft heading input is invalid), or The ground track input from the FMS is invalid, or The heading reference is selected to magnetic and the true north referenced ground track input can not be compensated for magnetic variation (the magnetic variation input from the FMS is invalid)

Ground Speed Display The aircraft ground speed is indicated by a digital readout, displayed bottom right side of the airspeed, and above the wind indicator, and the compass rose. The readout includes a label (“GS”) displayed adjacent to the readout numeric The groundspeed readout has a display range of 0 – 999 kts, with a resolution of 1 kt. When the ground speed readout exceeds 999 kts, the displayed readout value is limited to 999 kts.

GROUND SPEED READOUT

GROUND SPEED FAIL

The ground speed readout numeric is replaced by 3 red dashes when the input from the FMS is failed (or not received). The numeric is removed (and nothing displayed) when the ground speed readout is not valid, but is not failed.

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Wind Display

Wind Speed Readout

Wind is represented by a wind speed readout and a wind direction This symbology represents either the current wind direction and magnitude, or a stored value (“remembered” wind), whichever is most appropriate for the conditions The wind symbology is white when the current wind information is displayed and is yellow when “remembered” wind is displayed. The wind symbology is removed from the display when:

The wind speed is indicated by a digital readout, displayed above the wind direction indicator. The readout has a range of 0-255 kts (corresponding to the input range), with a resolution of 1 kt. displayed adjacent to the readout numeric. The readout includes a label (“R”) when the wind symbology represents a “remembered” wind, rather than the current wind The wind speed readout has a display range of 0 – 999 kts, with a resolution of 1 kt.

• • •

•

The displayed aircraft heading is removed (aircraft heading input is invalid), or Either the wind direction or wind speed input from the FMS is invalid, or The heading reference is selected to magnetic and the true north referenced wind direction input can not be compensated for magnetic variation (the magnetic variation input from the FMS is invalid), or The wind display is not removed based on the “remembered” wind is not being displayed, and the wind speed drops to less than 4 kts.

Navigation Source Annunciation The navigation source annunciation identifies the source of the navigation data displayed. There are different annunciations possible for the FMS and VOR/ILS navigation sources, depending on the conditions. When the crossside source is selected (source 1 is cross-side for pilot and source 2 is cross-side for copilot), the annunciation is displayed in yellow text. If the on-side source is selected, then the annunciation is displayed in white.

Wind Direction Pointer The wind direction pointer rotates around the pointer center (spins in place) to indicate the direction the wind is blowing relative to the current aircraft heading. If the wind direction pointer is oriented straight up on the display, the wind is coming from directly behind the aircraft. If the wind direction pointer is 90 degrees clockwise from straight up, the wind direction is directly across the aircraft from left to right. This pointer is centered directly below the wind speed readout.

WIND SPEED READOUT

REMEBERED WIND SPEED

Cross-side Normal NAVIGATION SOURCE ANNUNCIATIONS When both the pilot and copilot have selected the same navigation source, then the annunciation is boxed, to indicate the condition so that the pilots are aware that they are subject to a misleading display on both sides of the cockpit. There is no independent source for monitoring the navigation data. When the lateral deviation data associated with the selected navigation source data is failed, the navigation source annunciation is displayed as a failed flag.

Same Source Failed NAVIGATION SOURCE ANNUNICATION

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When the FMS is the selected navigation source and the inputs indicate that the navigation solution is degraded, the navigation source annunciation is displayed as black text on yellow background.

based on the input received from the FMS (the FMS computes the desired track into a FMS flight plan waypoints or flies a manually entered course via the FMS by the pilot into a waypoint depending on the FMS operational mode. Selected course computation and display is disabled when the navigation source is FM Homer. Selected Course Computation

Degraded NAVIGATION SOURCE ANNUNCIATION

The selected course associated with radio navigation sources is entered using the FD/DCP. The range of the computed selected course is -180° to +179 °, computed in increments with a resolution of 1°.

Active Waypoint Identifier Display

Course Pointer

The active waypoint identifier is an alphanumeric label associated with the active (fly to) waypoint in the flight plan. The waypoint identifier is displayed when the navigation source is FMS and a valid flight plan is received from the FMS. The active waypoint identifier is white in color, up to 12 characters long, and is displayed immediately below the selected navigation source annunciation. If a valid flight plan is not received, or if the active waypoint does not have an identifier, then the waypoint identifier appears as a blank line on the display.

The course pointer rotates around the aircraft symbol reference point to position the head of the pointer at the selected course/desired track value on the compass rose The white pointer is oriented vertically on the display. It has a head and a tail separated by a gap the size of the course deviation bar. When the aircraft is on a desired course, the deviation bar is centered and it “connects” the course pointer head and tail. The course pointer is removed if FMS is the navigation source and the FMS desired track/selected course input is invalid, or if the navigation source is selected to FM Homer. Course Readout

ACTIVE WAYPOINT IDENTIFIER Selected Course Desired Track The selected course symbology indicates the desired course into a radio navigation station or the desired ground track into a waypoint on the FMS flight plan. The selected course associated with the radio navigation sources is computed by the displays from FD/DCP inputs and displayed when one of those sources is selected. When the navigation source is an FMS source, selected course or desired track is displayed

The course readout provides a numeric indication of the current displayed course value. The course includes a white three digit numeric value, with a label to the left of it - either “CRS” for display of selected course (for VOR/ILS and navigation source, or FMS navigation source when the FMS is in To-From mode) or “DTK” for display of desired track (for FMS navigation sources only). The course readout range is 001 to 360°, with 1° resolution (corresponding to the full range of the input/computed course value, 180 to +180°). The readout always displays three digits. The course readout is removed if FMS is the navigation source and the FMS desired track/selected course input is invalid, or if the navigation source is selected to FM Homer. If a FMS is the navigation source and the desired track input is failed, the readout is replaced by 3 red dashes.

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Course Deviation The course deviation is displayed on the PFD by a course deviation scale and a deviation bar. It displays a “fly-to” cue to the pilot, indicating the direction to fly to capture the desired course into the waypoint/station. The course deviation scale symbol appears the same as the ADI lateral deviation scale, with two dots on either side of the scale. The course deviation bar is a magenta segment and is the same length of the gap in the deviation pointer. The bar moves laterally along the scale to indicate the current deviation using the “fly-to” convention. When the deviation input is invalid, the course deviation scale and bar are removed. There is no flag displayed in place of the course deviation symbols. To/From Indication The To/From indication provides a cue to the pilot indicating whether the displayed course deviation should be interpreted as deviation from the selected course/desired track for flight into a station/waypoint, or flight from the station/waypoint. The indicator is a white triangle pointing toward the station or waypoint the deviation is based on. The To/From indicator is only displayed when the navigation source is either VOR or FMS.

COURSE DEVIATION DISPLAY Bearing Source Annunciation The bearing source annunciations identify the sources of the displayed bearing 1 and bearing 2 data. The bearing source annunciations label is displayed following the annunciation, which is a miniature version of the head of the corresponding bearing pointer. The bearing pointers have different shapes, so these labels have correspondingly different shapes. This label, along with the color-coding, reinforces the

relationship between the bearing source annunciation (and associated distance readout) and the bearing pointer (located on the compass rose). The labeling and color-coding are provided to eliminate misinterpretation of the bearing information (source, pointer, and distance) caused by the data being physically separated on the display. Bearing Distance Annunciation The display of distance associated with each of the bearing sources is provided by numeric readouts with a label following it indicating what units are being displayed (“NM” for nautical miles and “KM” for kilometers) displayed immediately below the corresponding bearing source annunciation. If the distance input associated with a selected bearing source is failed, the corresponding readout is replaced by three red dashes.

DISTANCE ANNUNCIATION

DISTANCE FAILED Bearing Pointers The bearing 1 and 2 pointers rotate around the aircraft symbol reference point to position the head of each pointer at the compass rose position corresponding to the bearing to the associated selected source. The symbols for the bearing 1 and 2 pointers are different in addition to the bearing 1 and 2 color coding to assist in differentiating them on the compass rose.

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Message Alert Display The FMS displays status messages related to aircraft health and/or FMS navigation performance on the FMS. When a message is activated, the FMS provides an output to put an alert in the pilots’ primary field of view, prompting the pilot to look at the FMS. This is displayed regardless of what the navigation source selected is – unlike some other FMS related annunciations. The message alert is displayed as “MSG” in yellow text. When activated by the FMS, the alert flashes for the first 5 seconds displayed, then remains continuously on while the FMS indicates it is active. Once the pilot has read and cleared the message on the FMS, the FMS deactivates the alert and it is removed from the display.

The navigation map display symbology is referenced to the aircraft symbol reference point, which represents the current aircraft position. The symbology is oriented relative to the aircraft symbol based on waypoint symbol positions, range and bearing to radio navigation stations or hover mark points, and displayed heading.

ARC NAV MAP DISPLAY

MSG ALERT DISPLAY

ARC Navigation Map Display The PFD-Arc navigation map display provides lateral navigation guidance via a graphical presentation of FMS flight plan/radio navigation station information. It is proportionally scaled on the PFD display based on the navigation map display range selection made by the pilot. The range selection allows the pilot to expand navigation map/flight plan display to reduce extraneous cutter and focus on the map/flight plan segment being flown. The FMS flight plan is always displayed (if valid) when the PFD-Arc format is selected (regardless of what the selected navigation source is). Additional radio navigation station position/bearing symbology is presented when the selected navigation source is either Tacan or VOR. The ADI lateral and vertical deviation guidance remains based on the selected navigation source. This combination of navigation data displayed in PFD-Arc sub-mode enhances the pilot situation awareness during transition between FMS flight plan and another navigation source.

The navigation map symbology rotates and translates as the aircraft position and orientation (heading) change, with the aircraft symbol remaining stationary. The navigation map symbology is restricted to the area bounded by the dashed yellow line on the sides and bottom, and the white compass arc on the top. Any annunciations, which appear in this display area are either outlined in black or have black background fills behind them to assure their visibility above the map symbology.

ARC MAP DISPLAY FMS Flight Plan The active FMS flight plan is represented by a series of map symbols connected by arcs/vectors, which graphically show the pilot

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entered flight plan to which the FMS is currently navigating/providing guidance. The display receives the FMS flight plan as a series of waypoint records, each of which has the following associated parameters: • • • • •

Position (latitude and longitude) Waypoint (symbol) type Waypoint identifier (alphanumeric “name” for the waypoint) On-route or off-route waypoint type Flight plan interconnecting “leg” characteristic (straight vector (default), arc, or no connection (i.e. gap in flight plan))

A complete FMS flight plan may contain up to 127 waypoints (waypoint records). The display processes each waypoint record as it is received to: • •

•

Validate the record (correct labels, correct checksum, etc..) Decode the pertinent informatio n from the data stream (position, waypoint type, waypoint identifier, etc..) Sort the waypoints into active flight plan waypoints

The PFD computes display position for each waypoint based on aircraft present position, aircraft heading, and pilot selected display range, once an entire set of flight plan data has been received and validated. The flight plan including off-route waypoints is then displayed by rendering each flight plan waypoint, with the interconnecting flight plan “legs” generated as the successive waypoints are “drawn”. The display only renders those waypoints and associated flight plan legs that fall within the flight plan display area. If a flight plan waypoint falls outside the display area, any flight plan legs

associated with it are drawn to the edge of the flight plan display area and then “clipped” to not extend beyond this area. Since the display allows up to 127 flight plan waypoints, transmission of the flight plan data provides one waypoint every 100 milliseconds. Therefore, the updating position of each waypoint in the flight plan is based on the present ground speed and ground track of the aircraft. Flight Plan Waypoints The flight plan waypoints are those points, which specifies the on route waypoints the FMS is navigating to. Off-route waypoints are points from the navigation database or manually entered by the pilot, which are in the vicinity of the current active flight plan and provide additional situation awareness for the pilot. The different types of navigation map symbols are displayed as flight plan waypoints for onroute and offroute. The active waypoint the FMS is currently steering to is displayed as a magenta symbol. All other on-route flight plan waypoints are shown as white waypoints. The off-route waypoints are displayed as cyan map symbols and the identifier associated with a point are displayed to the lower right of the symbol. When the FMS indicates an impending flight plan leg change and the aircraft is within a specific time interval of “arriving” at the active waypoint, the active waypoint flashes. This corresponds to the flashing of the To/From pointer on the PFD-Full display. When the active waypoint sequences, the new active waypoint turns magenta and the old active waypoint turns white.

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Flight Plan Legs The lines connecting the active flight plan waypoints form the flight plan “legs.” The flight plan legs are shown as solid magenta lines on the navigation map except that the leg into the active waypoint may vary in appearance depending on flight plan mode. The lines may be straight vectors, or circular arcs based on the flight plan leg information provided by the FMS with each waypoint. The “active” flight plan leg varies with FMS flight plan modes. The flight plan mode specifies the way the FMS provides guidance into the current active waypoint. For the normal To-To Steering mode, the FMS provides guidance into the active waypoint exactly as the flight plan was entered. The line to the active waypoint is displayed as a solid line with the previous and active waypoints connected. The second flight plan mode is the Direct-To mode. The pilot may select a waypoint, which may or may not be the current active waypoint and instruct the FMS to fly directly to that waypoint and displays the active waypoint with a short-long dash. The FMS then computes a desired track directly into that waypoint from the aircraft position at the time the mode was activated, and provides guidance to that desired track. In this case, the previous waypoint (from waypoint), is no longer a part of the flight plan, and is not represented on the display. The third flight plan mode is To-From, in which the FMS provides guidance along a “selected course” into the active waypoint that is manually entered by the pilot and displayed with a shortlong dash. It is much like flying a radial into a VOR station. The active flight plan leg is displayed differently when the mode is not the “normal” To-To Steering mode, to keep the pilot aware of the fact that the FMS navigation/guidance is not being provided in the normal manner. The characteristics of the display of the flight plan for the three different modes Under Automatic flight plan waypoint sequencing, the FMS transitions from guidance to the current active waypoint to the next active waypoint once the “capture” is met.

For Manual flight plan waypoint sequencing, the FMS continues to provide guidance along the desired track/course into the active waypoint continuing past the waypoint until the pilot manually selects to sequence to the next waypoint in the flight plan (even after the waypoint “capture” is met. The “normal” FMS guidance mode is a flight plan mode of To-To, with automatic waypoint sequencing. In this condition, the FMS provides guidance to the current desired track into the next (active) waypoint, automatically sequencing to provide guidance (per the flight plan desired track) to the next waypoint when the current waypoint has been “captured.” Flight Plan Fail

The display receives the flight plan data from the FMS and performs the checks listed below to determine if the flight plan data is valid. If a transmission is received with any invalid condition, the display ignores the remainder of the data in the current flight plan transmission and waits for the start of a new flight plan transmission. If three transmission intervals pass without a valid flight plan being receive, the display declares a flight plan fail condition and activates the flight plan fail annunciation. Flight plan data is considered to be invalid when any of the following occur in the data received from the FMS: • Any waypoint record label has an invalid status. except the To/From waypoint label may have a status of NCD and still be considered valid • The number of waypoint records received for the flight plan does not match the number the FMS indicates are in the flight plan

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•

The checksum received in a waypoint record does not match the calculated value

When the FMS flight plan input is invalid, the flight plan is removed and a “FLT PLAN FAIL” annunciation is displayed.

Navigation Map Range Display The navigation map scaling and the “size” of the navigation map display area is determined by the pilot selected navigation map range The range is displayed by the inner range arc and associated numeric range annunciation, and the outer range arc by the inner edge of the compass arc and associate d numeric range annunciat ion. The numeric range annunciat ion displayed on the inner range arc is ½ the selected navigation map display range. The numeric range annunciation displayed on the outer range arc is the full selected navigation map display range. The range arcs and numeric range annunciations are white. Bezel buttons in the lower right corner of the display controls the navigation map range. The map range selection is independent on each display allowing a pilot to select different range scales on the inboard and outboard displays to see different levels of detail. For each display, the map range selection is shared between the PFD-Arc map and the other navigation map display formats. When the pilot changes the range scale on one map format, that range is the active display range when another map format is selected on that display. On power up, the navigation map range defaults to the 10nm range selection.

Navigation Map Display The UH-60M FDS navigation display mode provides an expanded version of the navigation information included in the PFD-Arc map display sub-mode. It includes two sub-modes with unique flight plan / navigation data presentations, including a full compass rose, heading up navigation map display, and a north up, decluttered “plan” map display. The full rose navigation display (ND-Full) provides a heading up orientation of the FMS flight plan and other selected navigation data, along with the compass rose and other navigatio n data, which provides increase d situation awarene ss. The ND-Full aircraft heading display symbolo gy appeara nce and operatio n are the same as the PFD Full except that the ND-Full compass rose is significantly larger and therefore positioned differently on the display. The ND-Full selected heading appearance and operation are the same as the PFD-Full selected heading characteristics except for the following indications: •

•

The selected heading indicator is never “out-of-view” near the bottom of the compass rose (like on the PFD-Full compass rose) The selected heading readout is only displayed when the selected heading is changing, not when it is more than 150° from the displayed heading value.

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PFD Navigation Map Display

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Map Display

ARC Map Display

The ND-Full navigation map display (including flight plan display, radio navigation station position/bearing display, hover mark point display, and range operation/display) is identical in appearance and operation to the PFD-Arc navigation map display, except as follows:

The PFD ARC map provides an expanded compass arc presentation with a flight plan/map presentation of navigation data. In many cases the appearance and function of the display is identical to the features as the full rose HSI except for the following:

• •

•

•

The navigation map display area encompasses the entire area within the compass rose The range symbology includes an inner range ring (no explicit outer range ring), and there is only a single (inner) range annunciation displaying ½ the selected navigation map range. Otherwise, the range symbology is the same as the PFD-Arc display. Range control is disabled when the CAS “pop-up” message window is displayed and multiple “pages” of CAS messages are active (this is described in section Error! Reference source not found.). The flight plan fail annunciation (“FLT PLAN”FAIL”) is positioned below the compass rose (rather than below the aircraft symbol as on the PFD-Arc).

•

• • •

•

The compass ARC format is an 80° section of the full compass rose that rotates centered on the PFD ARC aircraft symbol. There are no compass benchmarks in the ARC mode The position of the aircraft symbol is different than on the Full compass. The selected heading indicator rotates around the PFD ARC aircraft symbol position and is removed once it is more than 40° from the displayed aircraft heading and redisplayed when it is 39°. The ground track indicator is removed when it is more than 40° from the displayed aircraft heading and redisplayed when it is 39°.

Full Rose HSI The PFD full rose HSI provides a full rose compass presentation with a lateral deviation indication much like the traditional elctromechanical HSI indicator that the UH60A/L utilized. The digital HSI provides numerous features that enhance the deviation/heading display and provide increased navigation capability in the cockpit.

PFD ARC MODE

Hover Display The PFD Hover mode provides a full compass rose with a navigation data display primarily consisting of a graphic and numerical presentation of the hover point and of the aircraft velocity and acceleration. In many cases the appearance and function of the display is identical to the features as the full rose HSI or ARC display except for the following: •

Compass rose appearance is different from the PFD Full compass rose

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• •

There are no benchmarks in the hover mode Ground speed data displayed is the source selected hover velocity rather than the selected FMS velocity

The annunciation source text color is white when the primary source is selected on-side and yellow when the secondary source is selected on-side. A yellow box outlines the annunciation when both pilots are selected to the same source.

Flight Director Display The display of Flight Director (FD) symbology is normally enabled when the associated inputs are valid and a FD fail condition does not exist. The FD includes: • PFD HOVER MODE (TYPICAL)

Data Source Annunciations Display annunciations are provided to indicate the selected source for: • • • •

Attitude data Air Data Engine/CAS data Heading data

When each pilot has HSI primary source selected, the source annunciation is not displayed, as it would unnecessarily clutter the display to indicate a “normal” state. In the case where a pilot has selected as alternate source for a data group or both pilots have the same source selected (creating the potential for misleading data affecting both pilots), the source annunciation for that data group is displayed. The source annunciations are: • • •

“ATT x” for attitude “ADC x” for air data “DCU x” for engine/CAS (DCU) data

Where the “x” is either “1” or “2” depends on what source is selected. SOURCE ANNUNCIATIONS

• • •

Steering cues - Pitch, roll, and collective command cues Mode annunciations Coupled FD annunciation Flight parameter reference indications (for selected/pre-selected “fly to” values)

Pitch Command Bar The FD pitch command bar is a horizontal, magenta bar that provides a steering cue to capture desired aircraft pitch angle. The pitch command bar is displayed when enabled by the FD with a valid command input. When a pitch command input of 10° is received, the pitch command bar is positioned at a distance equivalent to 10° on the pitch tape above the aircraft symbol reference point. The command bar will “park” at the 20° position (up or down) when a pitch command input from the FD exceeds +/- 20°. When the pitch axis steering is not enabled or the command input is not valid, the pitch command bar is removed. Roll Command Bar The FD roll command bar is a vertical, magenta bar that provides a steering cue to capture desired aircraft roll angle. The FD with a valid command input enables the roll command bar. The scaling of the roll command bar motion is such that for a 0° command the command bar is positioned at the aircraft symbol reference point, and for a 45° command, the command bar is positioned at the outer edge (left or right) of the aircraft symbol. The roll command bar motion is limited to the 45° position (left or right) when a roll command input is received from the FD, which exceeds 45°. When roll axis steering is not enabled or the command input is invalid, the roll command bar is removed.

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Collective Command Cue and Reference/Scale The FD collective command cue is a magenta symbol shaped like a collective grip that provides a steering cue to collective position, which provides the desired vertical flight profile. The collective command cue and reference/scale are displayed when enabled by the FD with a valid command input. The scaling of the command bar motion is such that for a 0° command, the command bar is centered in the reference symbol outline, and the maximum travel up or down is when the collective symbol reaches the white triangular collective scale boundaries. Collective Command Reference/Scale

Roll Command Bar

Pitch Command Bar Collective Command Cue FD Annunciations The FD annunciations are displayed along the upper edge of the display. These annunciations include the active FD annunciation and FD mode annunciations. The green colored annunciations indicate an engaged (coupled) condition.

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Marker Beacon (MB) Annunciation

PFD Engine Power Pod

An MB annunciation is activated when the VOR/ILS receiver detects that the aircraft is in the vicinity of the outer, middle, or inner MB. The two-letter designation and color distinguish between the outer, middle, and inner marker annunciations are as follows:

The display of engine data on the PFD includes the information required for normal performance monitoring and flight operations, as well as quick recognition/diagnostics of abnormal engine conditions. This allows the pilots increased flexibility to use cockpit MFDs for display of mission/situational awareness data, rather than dedicating one or more displays full time to an EICAS format. This also serves to reduce pilot scan and minimize display format switching. The PFD engine data display is located primarily in the lower left quadrant of the display. It displays an integrated rotor/turbine (NR/NP1/NP2) tape and readout, torque (Q1 and Q2) tapes and readout, abbreviated turbine gas temperature (TGT1 and TGT2) tapes and readouts, gas generator turbine (NG1 and NG2) readouts and fuel tape and readout. The DCU source for engine data display is indicated in the upper right of the engine data display. Under normal configuration, with each pilot using HSI primary on side DCU source, this readout is not present. The data source used matches the EICAS source selection.

• • •

Outer Marker – “OM” – BLUE Middle Marker – “MM” – YELLOW Inner Marker – “IM” – WHITE

The annunciations flashes for the first five seconds it is displayed, and then is displayed until the VOR/ILS receiver no longer detects a marker signal. The annunciation also flashes for five seconds if it transitions directly from one state to another. The MB annunciation is disabled if the active marker input is not received from the VOR/ILS receiver.

MB Annunciation

ENGINE POWER POD

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Airspeed Display

Vne Indication

Airspeed symbology includes the following:

The Vne indication is a red arc displayed behind the airspeed dial markings, extending from the not to exceed airspeed (Vne) to the maximum value represented on the dial. Under normal conditions, Vne is 193 kts and the red arc extends from the 193 kts position to the 201 kt mark on the dial.

• • • •

Airspeed dial and pointer Vne indication (not to exceed airspeed) Airspeed rolling digit readout Selected Airspeed indicator/readout

Indicated Airspeed (IAS) Dial/Pointer The airspeed dial and pointer provide an “analog” indication of airspeed. The analog airspeed presentation provides rate and trend information that is not as easily interpreted from the rolling digit readout. The dial provides a full scale range of 0 to 200 knots (kts), and includes tick marks at 5° increments (alternating in length between 5 and 10 kt increments), numeric labels of tick marks every 20 kts, and a label indicating the units of the displayed parameter. The airspeed scaling is compressed at the lower end of the dial, in the region between 0 and 40 kts where the airspeed data is less accurate and subject to sudden, rapid fluctuations. The airspeed dial is displayed when the airspeed input is not failed. The airspeed pointer (white) travels around the dial indicating the current speed and when the IAS is below the displayed Vne value. The pointer color turns red when the airspeed is at or above the Vne value. The pointer will remain red until the airspeed displayed drops at least 1 kt below the Vne value. The airspeed pointer travels just beyond the end of the dial for an airspeed input of 201 kts. If the airspeed exceeds 201 kts, the pointer will remain at 201 kt position. If the airspeed is below 0 kt the pointer will remain at the bottom of the dial.

IAS Rolling Digital Readout The airspeed rolling digital readout provides a digital readout with resolution actually finer than 1 kt. The readout includes 3 digits (1 kt, 10 kt, and 100 kt digits, and a surrounding window, which helps to distinguish the readout from the dial and pointer. Selected Airspeed Symbology The selected airspeed symbology provides the pilot a reference airspeed indication with a digital readout that is associated with a currently active FD mode. The selected airspeed symbology includes a digital readout of the selected airspeed value and an indicator that is displayed on the airspeed dial. The selected airspeed symbology is magenta, consistent with the convention that the color magenta on the display represents a “flyto” indication. The selected airspeed symbology is enabled under the following circumstances: • • • •

Airspeed dial is displayed Flight Director fail flag is not displayed Selected airspeed input from the flight director is valid Unusual attitude conditions do not exist

Selected Airspeed Indicator The selected airspeed indicator is positioned at the selected airspeed value on the airspeed dial when the selected airspeed symbology is enabled. The airspeed indicator is removed if the selected airspeed value exceeds the airspeed range on the scale.

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data from the other side of the cockpit. This provides monitoring of both the inputs from the external sensors and monitoring of some of the processing performed internal to the display. When ”miscompare” is detected, a corresponding miscompare annunciation is displayed on the PFD. The miscompare annunciations on the PFD are displayed as a yellow parameter label (“PIT”, “ROL”, “HDG”,…) with a double headed arrow above it. The annunciation remains displayed until the miscomparison condition ceases.

Selected Airspeed Readout The selected airspeed readout is displayed above the airspeed dial, labeled with a symbol that matches the symbol used for the selected airspeed indicator. Display of the selected airspeed readout is normally under control of the flight director, allowing the flight director to turn the readout on under the following circumstances: • •

When the flight director mode is initially activated When the selected airspeed setting is being changed by the pilot

Individual parameter comparisons are disabled when a cross side parameter is not valid, or not available for comparison. A white miscompare annunciation is displayed when the associated comparison is disabled (with the exception of the Glide slope (GS) and localizer comparisons, which have no “disabled” annunciations). It is not automatically assumed that both pilots will have ILS selected as navigation source.

PFD Miscompares The displays perform a comparison of primary flight parameters to help protect against undetected display of misleading information. This function looks for differences between onside source selected data and source selected

Airspeed

Localizer

Pitch

Heading

Roll

Radar Altitude

Barometric Altitude

Velocity

Glideslope

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Data Source Annunciations Display annunciations are provided to indicate the selected source for: • • • •

Attitude data Air Data Engine/CAS data Heading data

When each pilot has HSI primary source selected, the source annunciation is not displayed, as it would unnecessarily clutter the display to indicate a “normal” state. In the case where a pilot has selected as alternate source for a data group or both pilots have the same source selected (creating the potential for misleading data affecting both pilots), the source annunciation for that data group is displayed. The source annunciations are: • • •

“ATT x” for attitude “ADC x” for air data “DCU x” for engine/CAS (DCU) data

Where the “x” is either “1” or “2” depends on what source is selected.

Data Source Annunciations

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DISPLAYS STATUS PAGE

Flight Management System Status And Test Indications STATUS The main status screen displays the overall status of each functional group. The screen for each functional group, in turn, displays the overall status of each subsystem associated with the functional group. Pressing SK-6 for Displays accesses the status page for all four MFDs, both FMSs, and both Display Control Panels (DCPs).

MFD STATUS PAGE TEST

MAIN STATUS PAGE Status for MFD 1 can be selected from this page by pressing SK 1. Status for the MFD will then be displayed. All faults and failures will be displayed under the MFD if a fault or failure should occur.

The Test Menu is accessed by the function key and provides the ability to conduct and review tests of all systems and subsystems or line replaceable units (LRU’s) that can conduct an Initiated Built In Test (IBIT). Faults are provided as 16-character text description of the fault(s) that are active. Pressing SK-6 for Displays accesses the test page for all four MFDs, both FMSs.

MAIN TEST PAGE

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Test for MFD 1 can be selected from this page by pressing SK 1 for MFD 1. An Initiated Built-In Test (IBIT) can be initiated for the MFD. Upon completion of the test, MFD 1 IBIT screen will display the overall subsystem test results only. All faults and failures will be displayed under the MFD 1 if a fault or failure should occur.

Flight Display System Operation The FDS provides integrated control and display of essential flight and mission information. The PFDs will display information from inputs via ARINC-429and MIL-STD-1553B data bus through Electronic Function Interface System (EFIS) provided by DCUs, FMSs, and FCCs. When turning the four MFDs to the on position the MFDs warms up in approximately five seconds the two inboard MFDs defaults to EICAS and the two outboard MFDs defaults to PFD. A Start-up Built-In Test (SBIT) runs automatically when the MFDs are powered up and will only run when the aircraft is on the ground (Weight-On-Wheels (WOW)).

DISPLAY TEST PAGE

Simultaneously pressing the BRT and CON switches and releasing, the MFD displays the DISPLAY STATUS screen. An Initiated Built-In Test (IBIT) is performed on this screen by pressing R1bezel key for approximately two seconds until the IBIT confirmation message is displayed at the bottom of the screen and is written in black over a solid yellow background. The IBIT sequence is completed within two minutes and the MFD returns to its normal screen. The IBIT only runs when the aircraft is on the ground (WOW).

MFD TEST PAGE

MFD DISPLAY STATUS SCREEN Continuous BIT (CBIT) runs continuously, in flight or on the ground, without interfering with the operation of the MFDs and report failures as they occur.

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Section 2-10.2

Pitot/Static System Components

Flight Instruments Pitot/Static System System Description The UH-60A/L Pitot/Static System has been modified to accommodate the UH-60M Air Data Computer System (ADC) and the Electronic Standby Instrument System (ESIS). The Air Data Computer System is a stand-alone system suitable for use as the “primary air data computer” in flight critical applications. The ADC replaces the current UH-60A/L Blackhawk air data and air speed transducers. The two electrically heated pitot tubes with static ports are connected to the pilot and copilot’s air data system and to the Electronic Standby Indicator System (ESIS). In addition, airspeed data is sensed for operation of stabilator, flight path stabilization, and flight director. The Barometric Altimeter, Air Speed indicator, and the Vertical Speed Indicator are now displayed on the Primary Flight Display (PFD) via ARINC-429 data bus.

The Air Data Computer System (ADC) is comprised of the following components: • Two ADC • Two Outside Air Temperature probes • Two Accelerometers • Two Pitot/Static heater switches

Air Data Computer The ADC measures the physical parameters of impact and static pressure from the current Blackhawk pitot static system. The ADC accepts outside air temperature (OAT) from external aircraft OAT probes and barometric correction. It computes and output various air data information for use by the aircraft’s automatic flight control system, flight management system and primary flight displays. The ADCs are located in the broom closets behind and outboard of the pilot and copilot seats.

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Outside Air Temperature (OAT) Probes The OAT probes are located external to the nose of the aircraft just outside of the center windshield.

Accelerometers The accelerometers are located inside the cabin overhead area just forward of the L/H cargo door.

Pitot/Static System Controls The Pitot/Static heater switches have been relocated on the UH-60M overhead console center panel. Previously located on the outboard mid portion of the right panel, the switches are now located on the upper portion of the right panel. A second switch has been added to the Pitot/Static system in the event one fails. The Pitot/Static heater switch purpose and functionality remains unchanged.

WINDSHIELD WIPER PARK

OFF

LOW

VENT BLOWER

HEATER MED O F F

O F F

HI

ON

ON

ENG NO. 1 O F F

PITOT LEFT

ON

O F F ON

ON

WINDSHIELD COPILOT CTR

ANTI-ICE PILOT

O F F

O F F

O F F ON

ON

HI

HEAT RIGHT

O F F

O F F ON

ACCELEROMETERS

ANTI-ICE NO. 2

OFF

ON

FUEL IND

TEST

UMAV002_5

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Flight Management System Status And Test Indications FMS can check the status of the ADCs. By selecting Fixed Function key STS on the FMS accesses the MAIN STATUS PAGE. Pressing (SK) 1 from the MAIN STATUS PAGE accesses the EGI ADC status page. From the EGI ADC Status page, pressing SK 2 or SK 7 accesses status for pilot ADC and copilot ADC. This menu selection will allow a user to view the status of the ADCs. All faults and failures will be displayed under the ADCs if a fault or failure should occur. ADC STATUS PAGE These faults are as follows:

MAIN STATUS PAGE

LVL1 SYSTEM LVL2 ADS OAT IN DIGITAL BOARD ANALOG BOARD POWER SUPPLY SELF TEST STATUS PARAMETER LOCK BARO CORR INPUT ACCEL INPUT CNFG/SEC PARITY CONFIG ERROR P/SSEC DISABLED MAINT MODE

EGI ADC STATUS PAGE

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Air Vehicle (Dynamics Segment)

PITOT STATIC

S

P

ADC

NO.1 DC PRI BUS

2

ADC No.1

28 VDC

NO.1

J1 WOW

S 2

J1 WOW

J1

ANALOG ±15 VDC

J1

TEMP

J1

OAT 1 P/N: A9506-01-A-1.85

P

ADC No.2

28 VDC

NO.2

TEMP

Accelerometer 1 P/N: 65610-03031-105

PITOT STATIC

ADC

NO.2 DC PRI BUS

J1

P/N: 70600-01821-101 J1

Configuration

Air Vehicle (Dynamics Segment)

ANALOG ±15 VDC

Accelerometer 2 P/N: 65610-03031-105

P/N: 70600-01821-101 J1

Configuration

OAT 2 P/N: A9506-01-A-1.85

Air Data System Block Diagram

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NOTES

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SECTION 2-10.3 ELECTRONIC STANDBY INSTRUMENT SYSTEM (ESIS) System Description The Electronic Standby Instrument System (ESIS) is used as a backup to the aircraft’s primary flight displays. The ESIS supplies the pilot and copilot with a visual display of pitch and roll attitude, airspeed, altitude, vertical speed, heading, and validity indications. When operating in stand-alone configuration, the ESIS will operate for a minimum of 30 minutes using the Emergency Power Supply.

ESIS GH-3001 The ESIS GH-3001 unit is a self contained solid state instrument requiring only input power to provide attitude and slip/skid information. Information is displayed on a color Active Matrix Liquid Crystal Display (AMLCD). The GH-3001 has the capability to display the following settings and information: • • • •

Air data information, which is provide by the external ADC-3000 Heading information Attitude information Barometric (BARO) Settings

Components ESIS is comprised of the following major components: • • • • •

Electronic Standby Attitude Instrument – GH-3001 (instrument panel) Detachable Configuration Module – DCM-3000 (ESIS) Air Data Computer (ADC) – ADC3000 (RH Broom Closet) Magnetometer – MAG-3000 (Tail Cone) Emergency Power Supply – PS855C (2.5 ampere hour (ah) Jet Battery Pack) (Nose Shelf)

ESIS GH-3001 INDICATOR The ESIS indicator has the ability to adjust and initiate various displays via a Menu. The Menu will appear at the bottom part of the screen. To access the menu, press the “M” button located below the display screen. NOTE A 3 ATI mounting clamp is provide with the ESIS kit and has a specific requirement in orientation of the clamp. If the clamp installation is incorrect, damage to the indicator may result during operation due to excessive heat build up.

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Detachable Configuration Module DCM-3000 The DCM-3000 is a solid-state device that retains software and hardware configuration information for specific aircraft. If the GH3001 were to be removed, the DCM-3000 will remain with the aircraft via a chain that is attached to the GH-3001 wire harness.

DCM-3000

Magnetometer The Magnetometer uses three axis magnetic sensors that senses magnetic fields and converts the signals digitally and is transmitted via RS-422 data bus to the ESIS. The signal is then used with the pitch and roll attitude of the indicator to compute magnetic heading of the aircraft.

MAGNETOMETER MAG-3000

Air Data Computer (ADC) 3000 The ADC accepts total pressure and static pressure inputs and provides altitude, Indicated Airspeed (IAS), and vertical speed via ARINC-429 data bus and displays the information on the indicator. The ADC must be configured along with the DCM-3000 or “Configuration Required” will be displayed during self-test time period of the indicator.

Emergency Power Supply The Emergency Power Supply is used as a backup system in the event that the UH-60M helicopter has a complete systems fail. The ESIS will operate for a minimum of 30 minutes with the use of the emergency power supply.

ADC-3000

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EMERGENCY POWER SUPPLY

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ADC-3000

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Controls A switch on the center overhead switch panel controls the ESIS. This switch is used to ensure that the battery pack is fully charged when in the test position. When the switch is in the armed position, the ESIS will be activated in normal flight conditions. In the event of an aircraft failure, the ESIS will be armed and ready for flight.

When placing the Standby Instrument Battery “STBY INSTR BATT” switch on the center overhead console, in the “TEST” position, a signal will be sent to the Emergency Battery Pack to indicate the status of the battery. When the switch is in the “OFF” position, a Crew Alerting System (CAS) message via the Data Concentrators (DCUs) will appear on the Multifunction Display (MFD). Once the switch is placed in the “ARMED” position, the CAS message will disappear.

ESIS Indications BARO Settings Indications STBY INST SWITCH There is an adjustment knob, located in the lower right corner of the ESIS indicator that is used when in menu mode to scroll up and down the menu as well as pressed to select from the menu list. This knob is also used to adjust the BARO settings when in the normal mode.

Indications

The ESIS indicates BARO settings by adjusting the BARO knob by rotating the knob clockwise or counter-clockwise. To indicate a standard BARO pressure press the knob and “STD” will appear in the BARO settings window. The ESIS will indicate invalidities and will display a loss of information from associated sources. Invalid information will appear as a RED “X”. It will appear individually or in several screens depending what information is lost.

Center Overhead Switch Panel There are two capsules “BATT LOW” or “BATT GOOD” on the center overhead console that indicates whether the Emergency Battery Pack is low or fully charged.

INVALID DISPLAY

BATTERY GOOD CAPSULE

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Extended Maneuver Indications Extended Maneuver indication “EXT MANUV” occurs in the lower left corner after a long duration attitude event is detected for a period exceeding six minutes at roll angle greater than 7° from level or if the indicator is within 8° from a magnetic heading. The “EXT MANUV” disappears when one of the following occurs: • •

Normal erection and heading alignment are restored. The attitude failure indication is displayed.

The ESIS’s Attitude Display indications are similar to that of the Primary Flight Display (PFD):

AIR DATA INDICATIONS Heading Indications Heading indications is displayed on a sliding tape and includes tick marks representing degrees, heading index, and A DG (directional gyro) indicator. The heading tape moves right or left to keep the current heading under the heading index line. The viewable heading tape will display a span of 65° left to right. The DG indicator will be displayed when information from the magnetometer becomes unavailable. The DG indicator is an indication that the free heading mode is activated.

ATTITUDE DISPLAY Air Data Indications Air data indications include altitude, air speed, and vertical speed. The indications are shown in either a scrolling tape or digital readout depending on the information. The ADC provides the indicator with an altitude range of –1,000 to +55,000 ft.

HEADING INDICATION

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System Operation The ESIS is a stand-alone system that is used as a backup system in the event that the UH-60M helicopter has a complete systems fail. The ESIS will operate for a minimum of 30 minutes with the use of the emergency power supply. With aircraft power off, the emergency power supply can be tested by placing the STBY BATT SW on the overhead switch panel to TEST position or by pressing the TEST button on the front panel of the power supply. Depressing either switch, places a load across the battery output unit under test for five seconds. When the battery voltage under load is sufficient, the LED on the front panel of the power supply and “BATT GOOD” capsule will illuminate. If the capsule or the LED extinguishes before the test switch is released, the emergency power supply must be charged or retested. During normal operation, the aircraft 28vdc bus supplies the necessary charging voltage to the emergency power supply. In the event of an aircraft electrical power failure the emergency power supply will automatically take over and supply DC power to the ESIS. Upon power–up and the STBY BATT SW on the overhead switch panel placed in the ARMED position, the ESIS starts an automatic process of self-diagnostics prior to normal operation. The display screen will be blank for approximately 15 seconds while the tests are performed.

If no system failures are detected the ESIS will then display ATT FAIL with a message ALIGNMENT IN PROGRESS and a timer/counter below the aircraft symbol. Sensor alignment mode will reach normal operation mode with three minutes after power is on. The ESIS will not enter the operational mode if the sensor alignment is not successful. A message, “ALIGNMENT FAILURE”, will appear on the screen. An alignment must be performed when the aircraft is stationary or in a straight and level non-accelerated flight. While the ESIS is operating normally, the system will continue to perform diagnostic self-tests, to assure the pilots of accurate information. Normal Indicator Operation During normal operation the adjustment knob on the indicator is used for BARO corrections. Rotating the knob clockwise will increase BARO settings counter-clockwise will decrease the BARO settings ranging from 22.00 In HG (inches of mercury) to 31.00 In HG or (745 HPA (Hectopascals) or MB (Millibars) to 1050 HPA or MB) in increments of 0.01 In HG (1HPA). Pressing the BARO knob the pilot can select between HPA, MB, or In HG settings. The BARO correction will not operate in the menu mode.

If no failures are detected during self test mode the identification screen will appear. The aircraft type, operation counter, and software ident numbers will appear on this page. Should the ESIS detect a fail, the Identification screen will show either a clearly stated error message or an error code.

BARO SETTINGS

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Menu Mode Operation By pushing the Menu “M” access button below the display screen, activates the menu mode display. During menu mode operation the adjustment knob is used to scroll the menus and initiate the highlighted menus items. The highlighted items will activate after pressing the knob. Menu items that are followed by three dots have sub-menus that will appear when selected. The sub-menus will give additional instructions for adjusting functions and values. Menu access will end when a setting is initiated or by pressing the menu access button. The menu will automatically end after 15-20 seconds of inactivity.

MENU ACCESS

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FCC #1

FCC #2

CVR/FDR

Data Concentrator Unit (DCU) No.1

Data Concentrator Unit (DCU) No.2

DCU Cross Communication

EFIS J-BOX #1

FMS #1

EFIS J-BOX #2

20 20 10 10

20 20 10 10

20 20 10 10

20 20 10 10

10 10 20 20

10 10 20 20

10 10 20 20

10 10 20 20

T

T

T

T

MFD #1

MFD #2

MFD #3

FMS #2

MFD #4

ESIS BLOCK DIAGRAM

10/2/2003 Page 9 of 10 PRELIMINARY REVIEW COPY

10/2/2003 Page 10 of 10 PRELIMINARY REVIEW COPY

Section 2-11 Active Vibration Control System (AVCS) System Description The UH-60M uses an Active Vibration Control System (AVCS) to address 4/rev fuselage vibration and a main rotor hub-mounted bifilar to suppress vibration resulting from in-plane rotor loads. The AVCS takes the place of the passive vibration absorber system that was used in the UH-60A/L. The AVCS reduces helicopter cockpit and cabin vibrations for enhanced pilot/crew comfort. This is accomplished by mechanically generating additional airframe vibratory loads outof-phase with the main rotor induced 4/rev vibrations such that the resulting cockpit/cabin vibration is reduced. The AVCS offers several benefits over the vibration suites found on the UH-60A/L. It weighs approximately 65 lbs less than the absorber suite it replaces in the UH-60L, while achieving comparable vibration levels at nominal rotor speed. Additionally, the AVCS maintains consistent vibration performance with varying

rotor speed (i.e. 98 -105% Nr), whereas the passive absorber suites are optimized for 100% Nr. Unlike the passive systems, the AVCS adapts to changing flight conditions and aircraft configurations such as weight and center of gravity variations resulting from fuel burn off or mission equipment installation. The AVCS is able to provide these benefits at a lower weight because it is a distributed system. The passive absorbers can only respond to and attenuate vibration at the locations where they are mounted. The AVCS however, can generate arbitrary force levels at any of its actuators to affect vibration attenuation at other locations. The UH-60M AVCS consists of an Active Vibration Control Computer (AVCC) that processes vibration sensor signals and provides command signals to the AVCS Electronic Unit (EU). The EU in turn converts AVCC data into analog drive commands and sends the data to the appropriate force generators (FGs). Three FGs generate forces at the appropriate frequency and phase that cancel cockpit and cabin vibrations based on the EU drive commands.

10/2/2003 Page 1 of 8 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Vibration Control Actuation System (VCAS)

AVCS Components Active Vibration Control Computer (AVCC) The AVCC is located in the lower avionics bay of the cockpit. The AVCC performs the input/output (I/O) and control processing functions of the AVCS. The AVCC uses an adaptive optimal control algorithm to process aircraft vibration data generated by cockpit and cabin motion sensors. The resultant AVCC output represents the frequency, amplitude, and phase of forces required by the force generators to minimize airframe vibrations. The AVCC command signals are transmitted via an ARINC-429 data bus to the Electronics Unit (EU).

The Vibration Control Actuator System (VCAS) harnesses the centrifugal loads produced by counter rotating wheels to produce tightly controlled 4/rev airframe vibratory loads. The VCAS is comprised of an electronic unit (EU) and up to three pairs of force generators, each pair of which operates in coordination to generate a sinusoidal force with controllable amplitude and phase at a prescribed frequency. Electronics Unit (EU) The EU is located in the lower avionics bay of the cockpit. The EU converts the AVCC digital force commands into analog motor drive signals used to control the speed and phasing of the Mechanical Units (MUs) to generate commanded forces. The EU provides control of the vibration control system three force generators. The portion of the EU responsible for controlling one FG is defined, as a “channel” thus the UH-60M EU is a three channel EU. The EU also transmits VCAS status and Built-In-Test (BIT) information to the system controller for maintenance actions.

VCAS AVC Computer

FWD

Antenna Well – Looking Up 10/2/2003 Page 2 of 8 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Force Generators (FG) The three FGs are installed in the same locations that the passive absorbers were located in UH60A/L. An FG capable of generating 1000 lbf is located in the forward cabin overhead, and FGs capable of generating 450 lbf are located in the left stub-wing and in the nose beneath the nose avionics compartment. Each FG is consists of two MUs that utilize counter rotating eccentric masses rotating at 4/rev to generate forces at the appropriate frequency and phase. The FG forces then cancel the 4/rev vibratory response at the cockpit and cabin accelerometer locations.

Each MU contains two counter rotating concentric wheels, each with a mass eccentricity. Turning the wheels at the blade passage frequency causes an MU to generate a sinusoidal force with constant harmonic amplitude. When operated in concert with a second MU and tightly controlling the phasing relative to the first MU, a force of arbitrary amplitude (up to two times the single MU force output) and phasing can be produced. The illustration below relates the operation of a pair of mechanical units, or force generator (FG), with an example in which the position of the mass eccentricities for MU #2 lags that of MU #1 by 90°.

0 lb.

0 Deg.

500 lb. 0 lb. 500 lb. - 500 lb.

0 lb.

Mechanical Unit No. 1

0 lb.

- 500 lb.

90 Deg.

180 Deg.

270 Deg.

Mechanical Unit No. 2 10/2/2003 Page 3 of 8

UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Vibration Control System Sensors

Cockpit/Cabin Accelerometers

The AVCS uses inputs from both the rotor speed sensor and accelerometers installed in the cockpit and cabin areas. The AVCS sensors provide signals to the AVCC that are representative of the amplitude and phase of aircraft vibration at the blade passage frequency.

A total of ten accelerometers are located in the cockpit and cabin areas. Four vertical and one lateral accelerometer are mounted behind cabin soundproofing panels in forward and aft cabin areas. The cockpit sensors also include four vertical and one lateral accelerometer. They are located in the pilot and copilot seatwells as well as on the cockpit deck forward of the pilot’s and copilots yaw pedals. The accelerometers sense motion, or the acceleration generated by airframe vibrations and converts the sensed motion into a corresponding electrical signal. In the case of the UH-60M active vibration control system, the sensors detect inertial acceleration that enable the AVCC to generate counteracting vibrations that reduce detected vibrations to acceptable or comfortable levels for the crew.

Rotor Speed (Nr) Sensor The Nr sensor, mounted on the right Main transmission accessory module provides the AVCC a tachometer signal. The rotor speed indication is used to calculate rotor system vibrations.

VV

V

V

Lt

V

V

V

F

V

Cabin Nos

F

F

E

V

L V

V

L

LL V

V

V

Cockpit and Cabin Accelerometer Locations

10/2/2003 Page 4 of 8 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

AVC

V

V

AVCS Controls The VIB CONT switch is located on the overhead console center panel. The two position switch allows the AVCS to be selected to the ON or OFF state. The switch is located at the aft end of the center panel.

CARGO HOOK

VIB CONT NO. 1 DC ESNTL BUS NO. 1 ENG

LIGHTS

IFF

AUX CB

CABIN ICS

NO. 1 ESNTL

5

5

5

35

4

2

FIRE DET

SEC

SENSE

JTSN OUTBD

NO.1 STAB

SAS

ECP

CARGO HOOK

NO. 1 VHF

HOIST CABLE

ESSS JTSN

5

2

5

5

5

7

1 2

PWR

PNL SPLY

EMER RLSE

BOOST

FM

7

7

SHEAR

1 2

1 2

INBD

EMERG REL NORM

EXT LTS MODE

5

4

CARGO HOOK LT

BATT

O F F

N O R M

1

O F F

ANTI COLLISION LIGHTS

B O T H LOWER

DIM

BRT

CABIN DOME LT

OFF

BRT

ON

APU TEST

VHF AM

AUX CB

VOR/ ILS

NO. 2 ESNTL

PLT ICS

5

5

35

2

2

2

PNL SPLY

MSTR WARN

PRI

PLT CDU

NO. 2 ADC

BACKUP HYD

5

2

5

DIM

BRT

O F F

PLT MWC

DIM

ENG

WINDSHIELD WIPER PARK

OFF

LOW

O F F

BRT

DCP

TEST LT

FUEL BOOST PUMP NO. 1

BRT

7

INBD

1 2

ENG NO. 1

BATT LOW

OIL HOT

ACC LOW

APU FAIL

APU ON

O F F

BATT

MED O F F

ANTI-ICE NO. 2

ON

O F F

O F F ON

OFF

PITOT LEFT

HI

O F F ON

ON

O F F

NO. 2

WINDSHIELD COPILOT CTR

ANTI-ICE PILOT

O F F

O F F

FUEL IND

ON O F F ON

ON

TEST

O F F

O F F ARM

ON

ON

R O E F S F E T

1

WHITE

ON

R O E F S F E T

NO. 2 TEST

R O E F S F E T

ON

APU TEST

ON

EXT PWR

TEST

ENG OVSP TEST A NO. 1 B A NO. 2 B

O F F INCR

ENG

O F F

O F F

O F F

DECR

BOOST PUMP

CONT

RESET

ENG SPD TRIM

AIR SCE HI/STRT

APU

LAMPS 2

BLUE

NO. 1 TEST

BRT

FIRE DETR TEST OPER

O F F

ARMED

GENERATORS NO. 2 TEST

NO. 1 TEST

STBY INST TEST

HEAT RIGHT

INST PNL

OFF

ALL

SHORT

ON

3

PRI

FUEL BOOST PUMP O F F

ARM SAFE

O P E N

HEATER

O F F

HI

BATT GOOD

ON

APU APU FIRE

ICS ICU

NO. 2 DCU

25

VENT BLOWER

APU

EMER RLSE

ON

OFF

1 2

CONTR CKPT

ON

AIR SCE HI/STRT

ON

SENSE

PLT MFD

PLT FD/

7

CONTR

FLASH

LWR CSL

BRT

R O E F S F E T

O F F ON

RESERVE

STEADY

LTD SW

UPR CSL

NO. 2 TEST

NO. 2 EGI

2

ON

LIGHTS

OFF

O F F

ON

SECONDARY

R O E F S F E T

BOOST PUMP

CONT

FIRE EXTGH

LIGHTS

BRT

ON

MAIN

NIGHT

DIM

NO. 1 TEST

APU

RESET

O F F

CPLT MWC

ON

EXT PWR

O F F BLUE

ON

POSITION LIGHTS

DAY

O F F

SECONDARY WHITE

O F F

O F F

R O E F S F E T

PLT

ON

OFF

UPPER

NO. 2 TEST

ON

IR

ARMED

GENERATORS

NO. 1 TEST

O F F

3 2

ARM SAFE

ALL

SHORT

ON

STBY INST TEST

ARM

FORMATION LTS

CONTR CKPT

O P E N

O F F

O F F

NO. 2 DC ESNTL BUS

CARGO HOOK

VIB CONT

ESSS

EMERG REL NORM

ON

ON

O F F ON

APU

UMAV002_1

FIRE EXTGH RESERVE O F F MAIN

FUEL BOOST PUMP O F F

NO. 1

APU APU FIRE

ON

EMER RLSE TEST LT

BATT GOOD BATT LOW

OIL HOT

ACC LOW

APU FAIL

APU ON

FUEL BOOST PUMP O F F

NO. 2

ON

Upper Console Center Panel

UMAV002_4

10/2/2003 Page 5 of 8 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Active Vibration Control System Indications Caution/Advisory Indications The AVCS Caution/Advisory Indications are as follows: AVCS INOP - Advisory Indicates that AVCS is inoperative.

Flight Management System Indications FMS Status Messages The FMS allows review status of AVCS Line Replaceable Units (LRU’s) that report continuousBuilt In Test (CBIT) and Power-up Built In Test (PBIT) fault messages. Selecting the STS fixed function key to access the MAIN STATUS page on the FMS can check the status of the AVCS. By selecting soft key (SK) 4 from the MAIN STATUS PAGE accesses the AIR VEHICLE status page. Selecting SK 5 accesses air vehicle equipment status for AVCS. This menu selection allows the operator to view the status of the AVCS. All faults and failures will be displayed under the AVCS if a fault or failure should occur.

AVCS STATUS PAGE

AIR VEHICLE STATUS PAGE

10/2/2003 Page 6 of 8 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

FMS Test Initiation/Indications Using the Test Menu and associated Active Vibration Control (AVC) test functionality, users are able to monitor AVCS fault status as well as initiate and review overall test status of the AVCS system line replaceable units (LRU’s) conducted by an Initiated Built In Test (IBIT). The Test Menu is accessed by the function key and provides the ability to conduct and review tests of AVCS line replaceable units (LRU’s) that can conduct an Initiated Built In Test (IBIT). By selecting soft key (SK) 4 from the MAIN TEST PAGE accesses the AIR VEHICLE Test page. Selecting SK 4 accesses air vehicle equipment test for AVCS This menu selection allows the operator to initiate an IBIT of the chip detectors. All faults and failures will be displayed under the AVC if a fault or failure should occur. AVC TEST PAGE

AIR VEHICLE TEST PAGE

10/2/2003 Page 7 of 8 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

System Operation The AVCC determines the frequency, amplitude, and phase of forces required to the FGs to produce the desired vibration environment in the aircraft. The AVCC then processes and calculates the required FG commands and transmits them digitally via a data bus to the EU. The EU then converts the digital FG commands to analog motor driven signals to the FGs. The forces that the FGs generate cancel the 4/rev vibratory loads produced by the rotor system. Each FG uses its own inertia to produce the commanded force at its attachment point in the fuselage, as opposed to generating the force by exerting equal and opposite contact points between two points in the fuselage.

AVCS Block Diagram

10/2/2003 Page 8 of 8 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Section 3-1 Flight Management System (FMS) System Description The FMS and Emergency Control Panel (ECP) is a stand-alone system for use as the primary operator interface for the Avionics Management System (AMS) for the UH-60M program. The FMS and ECP provide the pilot and copilot with the primary operator interface to the aircraft AMS. The FMS provides the aircrew with the ability of data entry, a central point of control and display for secure and nonsecure communication equipment, radio navigation equipment, and the Embedded Global Positioning System/Inertial Navigation Unit (EGI) navigation system. Each FMS is capable of being either the MIL-STD1553B Bus Controller (BC) or a remote terminal (RT) on the MIL-STD-1553B bus. At any one time, only one FMS is required to perform the BC task. The second FMS becomes the Backup Bus Controller (BBC). The FMS also interfaces to various aircraft subsystems via MIL-STD1553B data bus, ARINC-429, Ethernet and discrete signals.

Components FMS The primary operator interface for the UH-60M FMS consists of two identical FMS and one ECP. The FMS are located in the lower center console for access and operation by the aircraft’s pilot and copilot. The ECP is also located in the lower center console in order to provide both the pilot and copilot access to the panel. The FMS allows primary control functions (select preset, select manual frequency, and select plain/cipher) for all radios from a single, top-level display screen. Secondary controls such as squelch, TR or TR+G, ECCM/FH setup and ERF, slew tuning, and saving manual frequencies as new presets are done on lower level display screens.

FMS CONTROL PANEL

ECP The ECP provides primary emergency control functionality to the pilot and copilot. The ECP has the capability to select which FMS will operate as BC, provide zeroize signals to specific aircraft subsystems, and provide UHF Guard select and Identification Friend or Foe (IFF) Emergency/Normal control. The ECP interfaces with the FMS and avionics subsystems via discrete signals.

E M E R G

Z E R O I Z E

CDU-1

CDU-2

GUARD

A U T O

N O R M

IFF EMERG N O R M

+

EMERGENCY CONTROL PANEL

UMAV003_3 EMERGENCY CONTROL PANEL

10/2/2003 Page 1 of 20 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Copilot Outbd MFD

Copilot Inbd MFD

Pilot Inbd MFD

Pilot Outbd MFD

Coll Control

ACRONYM LIST

Coll Control

1

2

3

ESH – Ethernet Switching Hub CDU – Control Display Unit DTU – Data Transfer Unit EGI – Embedded GPS Inertial Unit ICU – Interface Control Unit FDR/CVR – Flight Data Recorder/ Cockpit Voice Recorder MFD – Multi-Function Display

4

Slew Control

Slew Control (MFD Display Control)

DTU#1

Conformal Antenna Above Tail Driveshaft

DTU#2

ESH #1 Copilot CDU

EGI #1 H-764GU

LF/ADF AN/ARN-149

ESH #2 Pilot CDU

Emergency Control Panel (Zeroize, UHF Guard Rcvr)

VHF Blade Antenna on Bottom of Fuselage

VOR/ILS AN/ARN-147

Conformal Antenna Aft of the Upperdeck 2

1

EGI #2 H-764GU

Collective Radio Control (Radio Page Control) VHF Whip Antenna in Tail Pylon

FM Homing Antennas

UHF Flush Antenna on Tub/Belly

VHF-FM1 AN/ARC-201D COMM 1

VHF-FM1 AMP Low Pass Filter

FDR/CVR

FM Homing Steering Outputs to MFD’s

ARING-429

MIL-STD-1553B

ETHERNET

ANALOG & PTT

AUDIO

UHF-AM AN/ARC-164 COMM 2

RS-485

KY-58

VHF Bent Blade Antenna on Tailcone

*

VHF-AM AN/ARC-186 COMM 3

VHF Tail Pylon Cover Antenna Low Pass Filter VHF-FM2 AN/ARC-201D COMM 4

HF Towel Bar Antenna on Left Side of Tailcone HF AMP HF AN/ARC-220 COMM 5

IFF AN/APX-118

KY-100

Digital ICS ICU

* UH-60M ONLY RF DIGITAL INTERCONNECTS & PTT

Low Profile Control Audio Panels (LPCAP’s)

FMS BLOCK DIAGRAM 10/2/2003 Page 2 of 20 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

STS

CDU

STS

D

CDU

D

IDNT

Z E R O I Z E

E M E R G

CDU-1

EMER

A U T O

CDU-2

IDNT

IFF EMERG N O R M

N O R M

+

EMERGENCY CONTROL PANEL TAIL SERVO

BACKUP HYD PUMP

NORMAL

M I S C

BACKUP

1

2 N

3

A

B

C

D

E

F

4 W

5

6 E

G

H

I

J

K

L

7

8 S

9

M

N

O

P

Q

R

BRT DIM

PRV

S W I T C H

HYD LEAK TEST

ON

A U T O

N O R M

ON

OFF

OFF SERVO OFF 1ST STG

2 N

3

A

B

D

E

4 W

5

6 E

G

H

I

J

K

L

7

8 S

9

M

N

O

P

Q

R

.

0

+/-

COM

NAV

EGI

FPN

DAT

CLC

FIX

CLR

XPD

INI

ZRO

STS

TST

PPS

1

BRT DIM

NORM

+

2ND STG PRV

C

F

NXT

NXT

.

0

COM

NAV

EGI

FPN

DAT

CLC

FIX

CLR

XPD

INI

ZRO

STS

TST

PPS

S

+/-

U

T

V

W

EGI 1 Y

X

M S N

SPC

Z

S Y S

ENT

RAD ALT

ON

IFF

TEST

OFF

OFF

S

U

T

V

W

EGI 2 X

ON

N O R M

O N

OFF

M4 HOLD

Y

SPC

Z

ENT

PRESS TO TEST

1

2

3

4

ICS

VOX

HOT MIC

CALL

1

2

3

4

5

B

A

ICU

TX ICS

PVT

STABILATOR

O F F

2 BU

3 7

5

6

ON AUTO FLIGHT CONTROL TRIM SAS 2

SAS 1

ON

ON

2

3

4

VOX

HOT MIC

CALL

1

2

3

4

5

ICU

R E S E T

2

V O L

V H R

S Q D I S

T O N E

12

EMER AM FM MAN

FAILURE RESET

TEST

PROGRESS

L

O F F

UNLKD

PRESET

50

DF

5

TR

ARM

CHAFF 3 0

3 0

R I P F P I L R E E

PARKING BRAKE

LOCK

5

OFF

TAIL WHEEL

M

ICS VOX

4 6

8

PRE

CPTR 2

DISP CONT

MODE MA AUTO T

AL NU

D E I C E

TES T IN

3

5

LOAD

POWER ON

MVOL 1

BU 7

FLARE

B L A D E

B

A

TX ICS

RMT

ON

CPTR 1

ON

PVT

FPS

ON

SAS/BOOST

1

ICS

AUTO CONTROL

TEST

DOWN

ICS VOX

4

RMT

CONTROL

MAN SLEW UP

MVOL 1

ARM MAN

PGRM

SAFE

BLADE DEICE TEST NORM SYNC1

MA IN

TAIL

RTR

RTR

S T O R E S JE TT IS O N

EMER JETT ALL

OUTBD

BOTH

OAT EOT

CHAFF DISPENSE

JETT INBD

SYNC2

BOTH

ON OFF

L

L OFF

ALL

RADIO RETRANSMISSION FM 1/FM 2

I R C M

FM 2/UHF

FM 1/VHF

FM 2/VHF

FM1/UHF

L OUTBD XFER/ REFUEL

VHF/UHF OFF

LOAD

DEVICE SELECT 1

2

10 9

3

8

4 7

5 6

RCPT

A U X F U E L

L INBD XFER/ REFUEL

SELF

PWR

R

R

PWR

R INBD XFER/ REFUEL

R OUTBD XFER/ REFUEL

CLOSE

CLOSE

CLOSE

CLOSE

XFER/ REFUEL

AUX CABIN HEATER XFER AUTO

CLOSE

REFUEL PRESS TO

CUT

HOIST POWER ON

DOWN

OFF NO. 2 BATT BUS

NO. 1 BATT BUS

NO. 2 ENG

NO. 2 ENG

APU

NO. 1 ENG

APU CONTR

FUEL PRIME

NO. 2 CEFS

5

2

5

5

5

CONTR

FIRE DET

WARN LIGHTS

INST

BOOST

NO. 1 EGI

NO. 1 BATT

BUS &

NO. 1 CEFS

5

2

5

5

SEC

BUS CONTR

IFF

FIRE

NO. 1 DCU

1

5

7.5

5

5

5

5

7.5

2

BKUP

EXTGH

SEC

SEC

INST

SEC

EXTGH

SEC

BUS CONTR

5

B U S

HTR INOP

+

OFF

PRESET VALVES

PILOT OVERRIDE UP

HTR ON

RESET ON

O F F

CABLE

APU GEN

AUDIO

OFF

INT

MANUAL

B A T T

DESCRM ON OFF

TEST

ON

CONV WARN

CDU

NO. 1 BATT UTIL BUS

APU CONTR

B A T T B U S

NO. 2 EGI

FIRE

UTIL BUS

2

5

5

WARN LIGHTS

FIRE DET

CKPT

NO. 2 DCU

NO. 2 BATT

NO. 2 BATT UTIL BUS

UMAV003_1

UH-60M CENTER CONSOLE

10/2/2003 Page 3 of 20 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

viewing. The exception to independent operation is that system annunciations can be displayed simultaneously. The keyboard is divided into the following types of keys.

FMS Controls FMS Display Screen The FMS display screen is a flat panel display using Thin Film Electro Luminescence (TFEL) technology and provides a 3” x 5” display area. The display also provides 20 lines of 21 characters using the standard character font. These lines are organized as follows: •

Lines 1-4 are the “header” body providing essential data Line 5 is the status line providing alerts Lines 6 thru 19 are the main screen portions of the display Line 20 is the scratch pad.

• • •

3” 2

HEADER

3

• • • •

Access lower level pages Toggle between modes of a function, if the scratchpad is blank Select the field via the SK, enter the data and press ENT to place data into field Delete data in the scratchpad by pressing CLR

SKs only perform a function when there is an arrow beside the key. Inward arrows are for functions performed on the screen. Outward arrows take the operator to a new screen. SK arrows are removed if the function is no longer available.

4

STATUS LINE

5 6 7

MAIN SCREEN

8 9

XXXX_21CHARACTERS_XXX

5”

12 13 14 15 16 17 18 19 20

Ten “SKs” whose functions depend on the current display screen presentation border the display. These SKs are used to:

SK Arrows

1

[

Soft Keys (SK) (Line select keys)

__SCRATCHPAD__

]

Up/Down Arrows Whenever the up/down arrow appears in the lower right of the screen, next to the scratchpad, this signifies that the PRV/NXT key (or collective switch) is operational for paging functions. These include stepping through preset or NAVpoint records or stepping through flight plan or quick-review pages. In addition, PRV/NXT key allows the operator to navigate between (1/X) pages. Display Pages

The display is bordered by 10 “soft-keys, which have access to additional FMS pages or select a parameter for editing, and perform a switch function. Control of the systems is provided by two FMS mounted on the lower center console. The FMS provide simultaneous and independent operation. Simultaneous means that one FMS may display system status, while the other is editing a flight plan. Independent means that one FMS may edit a display that the other is

The FMS displays all data on a series of screens or pages. Pressing a fixed function key (FFK) accesses the top or main page. Fixed Function Keys (FFKs) The FFKs are dedicated keys on the lower portion of the FMS panel. Pressing these keys will cause the first page of a series to be displayed regardless of what page was

10/2/2003 Page 4 of 20 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

previously displayed. The functions provided by these keys are: • • • • • • • • • • • • • • •

COM: Radio Communications control NAV: Radio Navigation control EGI: Embedded GPS Inertial system control FPN: Type of steering sequence and flight plan selection DAT: Mission data base data entry, editing, and review CLC: Calculate functions (range/bearing, time, fuel, hover) FIX: Position update and “on the fly” waypoint storage CLR: Clears FMS alerts on FMS status line and clears scratchpad entries. XPD: Transponder control and display INI: System initialization, mission cartridge load, and scramble functions ZRO: Zeroize functions STS: System status display TST: Preflight and test functions PPS: Present position displays ENT: Enters value placed in scratchpad for the selected item

Entries and Scratchpad A box drawn around a parameter indicates that it is selected for editing. The operator can step through editable fields on a page by pressing the ENT key, or by pressing the SK associated with the parameter. SKs associated with editable fields are normally shown with inward arrows pointing to the field label. Boxed Parameters and Editing Pressing the ENT moves the edit box around the screen among the editable fields. As long as the scratch pad is empty, pressing ENT will not change the parameters when moving among fields. Also, when a soft-key next to an editable field is pressed, the box moves to select that field. A boxed parameter such as “ACTIVE” on the flight plan page signifies that is not editable and will be ignored since the flight plan is active. Inverse Video Inverse video is used to draw the operator’s attention. Inverse video is used to:

• • •

Highlight soft-key switch settings Highlight alert conditions on the status line Highlight new alert conditions (as opposed to previously acknowledged conditions)

If an error in entry is detected, “ERROR” will be displayed in inverse video on the scratchpad line for two seconds followed by a description of the expected entry: • • • • •

“EXPECT N OR S” “EXPECT 0-3” “EXPECT LTR” “RANGE ERROR” “INVALID WPT”

During entries CLR clears the sratchpad and terminates the edit. ENT accepts the entry. PRV deletes the last keystroke and NXT pulls the characters down from the current value. For decimal entries, the editor function of the FMS aligns the entry based on the decimal point position. If no decimal is entered it is assumed that one is there after the last character is keyed. For special format entries such as lat/long, zone/coord, or date/time, the editor fills unentered positions with blanks or zeros as appropriate. Action Confirmation Inadvertent soft-key actuation could adversely affect the mission performance such as deletion of data or changing the navigation solution, the operator may be required to confirm the action. Upon initial actuation of the soft-key, the function label is over written by the text “CONFIRM” in inverse video. Navigation to a different page before conformation resets the soft-key. Asterisks (*****) Asterisks or stars (****) instead of numbers, signify that the value is out of range, invalid or otherwise undefined. A range value that is beyond 999.9 displays as ***** and a cross track error with an invalid “TO” waypoint also displays as *****

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Annunciators

ECP Controls

There are two incandescent annunciators at the top right of the FMS:

Zeroize

•

•

STS: A change in system status has occurred FMS: This illuminates when Built-In-Test (BIT) detects a failure within the FMS.

Latitude/Longitude (Lat/Long) and Military Grid Reference System (MGRS) All positions are presented with the LAT/LONG format: • •

LAT N XX XX.XX LONG W XXX XX.XX

Values are in degrees, minutes, and hundredths of minutes. Entry range for latitude is from S to N 89 59.99. For Longitude it is W to E 179 59.99. All these field s can be changed to MGRS coordinates display if desired by the operator. Most pages displaying LAT/LONG positions also show a SK (normally SK-6) labeled MGRS. Pressing this key changes the label to L/L and changes the position display to MGRS zone/coordinate as follows: • •

ZONE NNA AA COORD XXXX: XXXX

The zones values are two numbers an a letter (designating the UTM zone) plus two letters specifying the grid. The next set of numbers is the grid coordinates. Once the MGRS are selected, all position information on the FMS is in MGRS format until the L/L key is pressed. The FMS always store and processes position information based on the WGS-84. The FMS locally computes and corrects the displays, plus converts entered local datum positions to WGS84 before storing. This local datum correction affects the presentation of both LAT/LONG and MGRS positions. The local datum is selected from either FMS via the FFK INI.

The zeroize switch is a toggle switch with a red guard. This switch initiates a direct zeroize function of all nonvolatile, secure, and sensitive information storage. This switch zeroizes data in: • • • • • • •

Data Transfer System Radios KY-58 and KY-100 IFF KIT-1A computer GPS KEYS IDM encryption data Mission data

FMS-1/AUTO/FMS-2 This switch is normally left in the center (AUTO) position. In this case, the FMS units automatically determine which FMS will be BC of the MIL-STD-1553B bus and the ARINC-429 bus. The FMS1 position forces BC assignment to the copilot’s unit while FMS-2 forces assignment to the pilot’s unit. EMERG This switch should be left in the center position (there is no down position), except when both FMS units have failed or are deactivated. When active (GUARD), this switch sets the radios to their perspective emergency frequencies. EMERGENCY IFF The “IFF EMERG” switch on the ECP should be left in the center position. The up position selects emergency IFF codes and is equivalent to the emergency switch on the APX-100 control panel. The down, momentary position, is equivalent of the “refuel hold” function.

Indications FMS Header Display The top five lines of the display indications are totally independent from whatever display screen is active on the bottom 15 lines. The

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header provides continuous indications of essential information in one of two formats: • •

The Communications Header The Navigation Header

One FMS defaults to the communications header while the other FMS displays the navigation header. Either operator can change header selections by using HDR command on the status page or on the present position page. The communication header provides continuous status of the currently selected transmit radio for both the pilot and copilot. The FMS units’ monitor the ICS transmit selection and provide the following information:

• • •

• •

•

System time Time-to-go for the current leg Planned altitude for current leg (either Planned Enroute Altitude (PEA) or Planned Ground Clearance (PGC) altitude, Above Ground Level (AGL) or mean sea level (MSL) (only displayed if one was entered in the flight plan definition) Ground speed for current leg Waypoint type (displays “WPT” for regular waypoint, waypoint number, and the first five characters of the waypoint name) The system status display line (line five of the header)

COMMUNICATION HEADER • • • • • • • •

System time Current navigation mode Selected radio band and mode Selected radio secure voice status Selected radio frequency, net, or channel Radio call sign Transmit active indication by either the PLT or CPT label changing from normal video to inverse during transmission. The system status display line (line five of the header).

COM HEADER INDICATIONS

NAV HEADER INDICATIONS FMS Status Lines The FMS status lines provide an indication BC status as well as 15-character alert line for the display of system and subsystem alerts. The BC status “BC” in small text displays if the FMS is the BC. The contents of the alert field are the same for either operator and provide the same data independent of the screen being displayed. The Multiple Alert Symbol (↓) is displayed if more than one alert can be displayed. Each alert has a priority associated with it. The higher priority alerts are placed on top of the alert list relative to the lower priority alerts. The alerts are in order of priority from the highest to the lowest. The highest priority will be displayed in the status line.

NAVIGATION HEADER • • • • •

Range to “TO” waypoint Bearing to “TO” waypoint Next leg course (for turn anticipation) Wind direction and Velocity Current Navigation Mode

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Reset Alerts can be removed by pressing the CLR FFK with an empty scratchpad on either FMS or it can be removed if the condition that provoked the alert no longer exists. Some alerts are defined as non-clearable in which case actuating the CLR FFK with an empty scratchpad and an alert that is still active will be moved to the lowest priority. All other non-clearable alerts will reset its priority and place it in the appropriate order of priorities.

MAIN STATUS PAGE

FMS STATUS LINES

Status The STS FFK on the FMS accesses the Status Menu. It provides the ability to review status of all subsystems or line replaceable units (LRU’s) that report continuous built in test (CBIT) and Power up built in test (PBIT) fault messages. Faults are provided as 16-character text description of the fault or faults that are active. The main status screen displays the overall status of each functional group. The screen for each functional group, in turn, displays the overall status of each subsystem associated with the functional group. Pressing SK-6 for Displays accesses the status page for all four Multifunction Displays (MFDs), both FMS, and Both Display Control Panels (DCPs).

DISPLAYS STATUS PAGE The FMS status page displays specific faults, interface faults, and subsystems overall status. Pressing SK-6 and then SK-3 or SK-8 displays the FMS-1 or FMS-2 status screen layout: Title Lines For each system or subsystem, the title line contains the name of the system or subsystem and number where “X” is the number of the applicable system (i.e. FMS-X for FMS-1 and 2). The title line also contains the displayed page number according to the following: •

As the number of active faults exceeds the number of available fields, additional pages are provided to accommodate the remaining faults with same formatting as the initial page.

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• • •

•

“(X / Y)” in the title line is displayed to the right of the status page, where “X” is the number of the currently displayed page and “Y” is the number of pages total. The PREV and NEXT FFKs allow the operator to scroll through the status pages associated with the selected system or subsystem. Upon screen entry the first page is displayed.

DATA BUS INTERFACE Depending on which system or subsystem, each screen displays status of interface between the subsystem and the FMS. In some cases, the interface is between the subsystem and an intermediate subsystem. Line 7 MIL-STD-1553 Data Bus displays the following states: • • • • •

“GO”- both buses are active “NGO-A”-bus A has failed. Subsystem rolls-up as degraded “DEG”. “NGO-B”-bus B has failed. Subsystem rolls-up as degraded “DEG”. “NGO-T”- 1553 bus communication does not exist between the unit and the FMS. “NGO-T” rolls up to a NGO status for the subsystem.

Line 8: ARINC data bus status displays the following states: • •

“GO”- ARINC data bus is active “NGO”- a bus fault exists. Subsystem is NGO.

Line 9: ENET (Ethernet) bus status displays the following states: • •

“GO”- no faults exist for the Ethernet bus “NGO”- a bus fault exists. Subsystem is NGO.

FMS STATUS PAGE Lines 10-19 for each system or subsystem screen has a dedicated left justified column of 10, 16-character fields dedicated to the display of active faults. If no faults are active, the fault area is blank. Unacknowledged (new) faults are displayed above acknowledged faults within the list. Within the acknowledged/unacknowledged fault groups, all faults are listed in the order in which they were received with the most recent on top of the list. The fault that is displayed remains in the failed or degraded state for 2 seconds. Faults that are no longer active are removed from the list. FAULT ACKNOWLEDGEMENT Unacknowledged (new) faults are displayed in inverse video. This includes the interface faults. The displayed faults transition to acknowledge (normal video) once the pilot has actuated the ACKnowledge select key. Unacknowledged faults on additional pages are not acknowledged. New faults that have occurred less than 1 second prior to the actuation of the ACKnowledge select key remains in inverse video (unacknowledged) to minimize acknowledgement of unobserved faults. Status and Fault Rollup Overall status of a subsystem is a rollup of the individual subsystem faults. This includes the unacknowledged state of any lower level screen. Activation of a new fault activates “MSG” alert on the MFD as well as the “STS” light on the FMS.

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Status of each subsystem rolls up to the overall status of its functional group. Upon report by a subsystem of a “NGO” or “DEG” condition, the check status light on the FMS is lighted and the “MSG” alert on the MFD is displayed. Main status indications consist of: • •

• • • •

“GO”- indicating all subsystem’s associated with the particular functional group are good (no faults), “DEG”- indicating a condition in which at least one subsystem associated with the particular functional group is degraded. AFCS and AVC supply overall status of DEG “NGO”- indicating a condition in which at least one subsystem associated with the particular functional group is NGO “TST” – a subsystem is in test mode (IBIT). “OFF” – all subsystems have alert reporting set to OFF “---“ – Status for system is unknown

Test The Test Menu is accessed by the TST FFK, which provides the ability to conduct and review tests of all systems and subsystems or Line Replaceable Units (LRU’s) that can conduct an Initiated Built-In-Test (IBIT). Faults are provided as 16-character text description of the fault(s) that are active. Pressing SK-6 for Displays accesses the test page for all four MFDs, both FMSs, and both DCPs.

ALERT REPORTING Alert reporting activates the MSG alert on the MFD as well as the STS light on the FMS. Actuating SK-7 toggles alert reporting “ON”/“OFF”. When in the “OFF” state, the fault rollup, MSG alert, and STS alert become inactive.

MAIN TEST PAGE

FAULT COUNT DISPLAY AND CONTROL On line 14 of each subsystem screen, the number of transitions from “GO” to “NGO” or “DEG” displayed. Actuating of SK-8 resets the fail count to 0 for the corresponding subsystem. Actuating SK-10 on the main status screen resets the fail count for every subsystem to 0. This does not affect the display of active faults. SOFTWARE VERSION OR HARDWARE ID Where applicable, software version numbers are right justified on line 8 and, as needed, on line 9. In some cases software ID can be accessed via SK-6.

DISPLAY TEST PAGE TITLE LINE For each subsystem the title line contains the name of the subsystem and number where “X” is the number of the applicable system (i.e. FMS-X

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for FMS-1 and 2). The title line also contains the displayed page number according to the following: •

•

• •

As the number of active faults exceeds the number of available fields, additional pages are provided to accommodate the remaining faults with same formatting as the initial page. “(X / Y)” in the title line is displayed to the right of the test page, where “X” is the number of the currently displayed page and “Y” is the number of pages total. The PREV and NEXT FFKs allow the operator to scroll through the test pages associated with the selected subsystem. Upon screen entry the first page "(1 of Y)" is displayed.

SCREEN ENTRY Upon screen entry the last known test results are displayed with the any faults displayed in the fault window. If no test has been conducted, “---“ is displayed as the overall LRU test results indicating that a test has not been performed and Test Message Line are cleared. TEST INITIATION

SUBSYSTEM TEST AND FAULT ROLLUP The test state is a rollup of the individual subsystem faults. A “CHECK TEST” alert is displayed on the FMS status line upon report by a subsystem of a “NGO” or “DEG” state. The system’s test has the following states: • •

Upon screen entry, the main page displays the overall test of each functional group. Main test indications consist of: o “GO”- indicating at least one subsystem associated with the particular functional group is good (no faults), o “DEG”- indicating a condition in which at least one subsystem associated with the particular functional group is degraded (only applies to AFCS and AVC), or o “NGO”- indicating a condition in which at least one subsystem associated with the particular functional group is NGO. o “TST” – a subsystem is in test mode (IBIT). o "---" – No Tests within a functional group have been performed or results unknown.

Upon initiation of the test, the “TEST” label is displayed in inverse video and the fault list is cleared. Upon completion of the test, the label returns to normal video. If operator attempts to initiate a test without Weight On Wheels (WOW), the test becomes inhibited and “GROUND ONLY” is displayed in the Test Message Line. Some test; however, allow testing in flight by design. Lines 10-19 on each system and subsystem screen has a dedicated left justified column of 10, 16-character fields dedicated to the display of active faults. If no faults are active, the fault area is blank. All faults are listed in the order in which they were received with the most recent on top of the list. The faults are displayed until the next test is initiated. Each test screen section provides the fault list associated with that particular subsystem.

TEST PAGE The main test screen displays the overall test status of each functional group. The screen for each functional group, in turn, displays the overall test status of each subsystem associated

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with the functional group. The subsystem level displays specific faults and subsystems overall test status. Pressing SK-6 and then SK-3 or SK8 displays the FMS-1 or FMS-2 test screen layout:

KEYBOARD TEST SK-3 accesses DISCREET INPUT, which provides a single screen view of the FMS discreet interface data. FMS TEST PAGE Pressing SK-2 accesses KEYBOARD TEST screen, which permits the verification of operating each key on the FMS keyboard. All keys are shown in this screen with the exception of the BRT and DIM. Actuating each key on the FMS blanks its corresponding representation on the screen. Actuating each soft key blanks its adjacent number that is displayed. Normal operation of the fixed function keys is suspended during testing of this screen. Once the test is completed and all the indications have been blanked pressing SK-10 RTN is then the last and final key test and returns to the FMS Test screen. FMS DISCREET PAGE SK-7 accesses the initial FMS DISPLAY TEST screen, which provides a lamp test of the FMS annunciators and provides access to a subsequent series of display test patterns which are sequenced using the and keys. This test must be stepped through to completion prior to exiting.

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All waypoints and steering calculations are done within FMS. Based upon EGI position, the FMS provides: • • • • DISPLAY TEST PAGE Entering this page, the FMS starts an annunciator lamp test sequence that lights in order the annunciators for between 1 and 2 seconds each. The sequence is: "ADV", "STS", "MSG", "FMS", and then, turn off. At the end of the sequence, the original lamp settings are restored. The FMS lamp remains on during sequence of the first three lamps if an FMS BIT fault has been detected. The indication "NOW IN PROGRESS" changes to "COMPLETE" when the lamp test sequence completes.

System Operation When power is first applied to the system, the FMS units perform an internal startup BIT. The power-up BIT is completed within 60 seconds after power is applied. The immediate indication of operation will be an “on” then “off” (lamp test) transition of the two FMS annunciators. The FMS then displays a graphic UH-60M plus the software version for up to 10 seconds. The display then switches to the initialization screen. If startup occurs properly, FMS -1 (copilot FMS) becomes the BC. The FMS starts up displaying the time and date from the GPS clock if available and the last position, which was stored in the FMS nonvolatile memory.

Waypoint to waypoint or Direct-toWaypoint for either of two flight plans consisting of up to 98 waypoints. The ability to construct flight plans from the database of 99 navigation points or 25 store-points. The ability to capture 25 nav-points “on the fly” based on over-flying the position or on range/bearing sensors Calculation functions along the active flight plan

Both the inertial and GPS determine their position independently. The GPS determines its position from the GPS satellite constellation. The pilots can manually exclude GPS or inertial data from the solution resulting in an inertial only or GPS only. The FMS monitors the pure GPS, pure inertial, and blended output. Flight plans Pressing FFK FPN on the FMS accesses the flight plan screen. The flight plan screen provides: • • • •

Selections of the flight plan used to steer the aircraft Editing of the flight plans Selections of steering options Manual stepping through flight plan legs

Three flight plans and three search patterns are provided within the flight plan screen: • • • • •

Direct to Flight Plan 1 and 2 Creeping Line Expanding Square Sector Search

Note that none of these engages without a valid present position.

Once the EGI is operational and in “NAV”, “NAV READY”, or ”DEG NAV RDY” mode, the pilots can then initiate steering.

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operator wishes to remove data from either the MFD or flight instruments pertaining to flight plans no longer being followed. Flight Plans 1 and 2 The two flight plans that are provided are made up of up to 98 waypoints each. The waypoints can be placed in any order and repeated if required. Waypoint "numbers" shown here refer to the reference number in the waypoint database. Waypoints are sequenced as part of "legs" in the order shown - progressing down the display screen from top to bottom.

FLIGHT PLAN SELECTION PAGE AUTO/MAN (SK-1) selection determines whether the sequence from one leg to the next is automatic or manual. In manual mode, arriving at the "TO" waypoint causes the header to display "WPT" on the status line and flashes the "WPT" label on the navigation header. Pressing SK-7 NXT sequences to the next leg. In the automatic selection, the "WPT" warning appears for only a second prior to sequencing to the next leg. If the last leg of a flight plan is reached, then the system continues to navigate to the last waypoint.

Pressing SK-4 or SK-5 accesses the flight plan control and review screen. This page shows the current "from" and "to" points, plus the next 8 waypoints in the sequence. When a leg is finished, the "from" waypoint disappears and all the waypoints move up one position on the display. The PRV/NXT key is used to show groups of waypoints following or preceding the displayed group of ten waypoints

If SK-1 is selected to AUTO, it is switched to MAN when sequencing to a next leg. The EARLY/OVER (SK-2) selection determines the capture criteria for the current "TO" waypoint and determines when automatic leg switching occurs. OVER (or overfly) selects the waypoint that must be passed within a 4.0 nm range. Early captures determine turns of less than 120°; a rate-one turn results in a rollout exactly on the next leg course. For turns above 120°, the calculation is still done based on 120° and a teardrop turn will result. When there is no "next" leg, Overfly is enabled. The solution uses actual track rather than desired track in calculating turn anticipation. This allows for low-level tactical flight where the final track to a waypoint may not be along the desired track. The STOP (SK-6) selection disengages the active flight plan or search pattern, resulting in a no computed-data state. This is used when the

FPLN EDIT PAGE When EDIT is pressed from either flight plan, the screen goes to a screen that identically mirrors this screen for waypoint content. Again the PRV/NXT key is used to get to other waypoints in the sequence beyond the ten being displayed. The "SEQ↑" and "SEQ↓" SKs forces the flight plan to re-fly the previous leg or advance to the

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next leg. Pressing either of these keys will change the sequencing mode to "manual" rather than automatic. A "Skip immediately to" function is provided with the "GO TO" key. By entering a flight plan sequence number in the field, or by pressing the left-side SKs to specify one of the displayed points, a "go to" point is specified and appears in the field associated with SK-6. A flight plan may be edited, even if it is active. Editing allows insertion, deletion, or modification of the waypoint sequence by specifying either name or number. If the flight plan is full then “FULL” is displayed in place of “INSERT” upon selection.

b) An offset range and bearing from a specified waypoint or storepoint c) A manually entered LAT/LONG or MGRS coordinate. In the case of (b) or (c), the flight plan manager creates a phantom waypoint in memory for steering purposes. For fast reference, the last-captured, mission data base storepoint is listed for fast selection. This allows for a rapid set-up of a fly-back-to maneuver. ENGAGING AND RESUMING FLIGHTPLANS Pressing SK-5 engages or resumes any flight plan or search pattern. The flight plan manager function in the FMSs then marks the current aircraft position, creating a phantom waypoint in memory. It then uses this as the "from" waypoint in creating a leg that goes to the first start or the resume point of a flight plan. Engage re-marks and re-starts the steering if the ENGAGE mode is already active (inverse video). In order to re-start a flight plan from the beginning once it has been partially navigated, the "GO TO" function of the flight plan control screen must be used.

FPLN EDIT PAGE Pressing SK-6, the flight plan does a "direct-to" leg from the present position mark point to the specified "go-to" point, and then continues sequencing from there. Direct-To Flight Plan The direct-to flight plan allows flight directly from the current position to a specified position. The specified position can be either a waypoint, a range/bearing offset from a waypoint, or a coordinated position entry. The direct-to flight plan, when engaged, marks the current aircraft position and then provides steering from the mark-position to the specified waypoint. The direct flight plan screen provides the ability to define the destination point as: a) Any mission data base waypoint or storepoint

DIRECT–TO FLIGHT PLAN PAGE If the waypoint number points to an "empty" waypoint record then an error message "WPT NOT DEFINED" is displayed in the scratchpad and no edit is performed. If a waypoint number changes, the waypoint corresponding name is displayed.

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Changing a waypoint name or number, the offset range and True bearing are set to zero and the position is set to match the waypoint location. Magnetic bearing is computed. Whenever a non-zero offset range is entered, the lat/long is updated to display the resulting position. For steering purposes, the waypoint is now considered a phantom waypoint and has the computer-generated number "--" and an asterisk "*" is prefixed to the waypoint name. When a position edit occurs, the waypoint is now considered a phantom waypoint for steering purposes, and has the computer-generated number "--" and name "MAN-LL". Entering an offset bearing is as a magnetic bearing, it must be converted to true prior to use in the range/bearing-to-lat/long calculation. The displays of magnetic and true bearing remain consistent with the operator entry. If a magnetic bearing is entered then the corresponding true bearing is computed while if a true bearing is entered then the corresponding magnetic bearing is computed. Both FMSs display the same bearing information. The conversion to true is performed each time a magnetic bearing entry occurs. In order to determine and apply the magnetic-variance at the waypoint, the magnetic-variance prediction calculation is performed specifically for the base waypoint position

FPLN LEG DATA PAGE

Position Fix Pressing the FFK FIX on the FMS can access position Fix. These screens enable two types of position fixing: • •

Update – Updating the Doppler sensor from a known good position Store – Storing a position as a waypoint relative to the aircraft position.

Note: At this time the Updates for position fixing is not available for first flight.

Flight Plan Leg Data The Flight Plan Data provides means to enter or modify the planned altitude for the leg. Upon screen entry, the FMS performs a NAV calculation between the “TO and “FROM” waypoints to compute a circle distance and a desired track for the leg.

FIX MENU PAGE

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Updates Updates are used to correct for position drifts, which build up over time in a sensor.

Pressing SK-9 STORE allows the operator the option to immediately add waypoint attributes such as name or vertical definition immediately following a storepoint capture.

Stores Store actions create "store points" in the mission database, which can be later used for navigation. These are also known as "Targets of Opportunity". In this case the aircraft DGNS position is used as the reference to create one of 25 possible store points, which are designated as "A" through "Z" less "O". Overflying the point and striking MARK can create a storepoint. Pressing SK-6 accesses the Store FLYOVER screen. The Flyover screen allows for the creation of a storepoint by marking an overflown position. SK-5 marks and holds the current position. MARK appears in inverse video when a position is marked. Subsequent selection releases position and allows display to update from the aircraft present position.

STORE-POINT DATA PAGE Present Position Present position provides a summary screen of relevant navigation information. These screens, along with the COMM screen, are the screens that are usually left on the display while in normal flight. Pressing the FFK PPS accesses the present position screens. Pressing SK-5 PPOS COMP accesses the PPOS COMP 1/3, which provides the current aircraft present position, EGI 1 and EGI 2 blended positions. The position differences between the aircraft present position and the EGI 1 blended positions are displayed along with the 95% position error for each EGI blended solution.

FLYOVER STORE PAGE

PRESENT POSITION PAGE

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FMS CONTROL PANEL

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NOTES

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In addition, it provides audio signal interfaces between various other avionics system devices such as navigation radios, aircraft survivability equipment, cockpit voice recorder (CVR), and aural warning generators, etc.

Section 3-2.1 Communication Systems Digital Intercommunication System (ICS) Description

The Digital ICS also has the capability to receive direct aural warnings and internally generate aural warnings to be triggered by discrete input signals and to be heard at pre-defined crew stations. Certain aural warnings may be made available to aircraft pilots only and may not be made to other crewmembers.

The ICS provides the means of secure interphone and radio communication for aircraft crewmembers and ground crewmembers. Each crewmember has the ability to select and volume control all interphone channels and radios at a control panel independent of all other control panels.

The Digital ICS provides a MIL-STD-1553B digital interface to the aircraft 1553B data bus for the communication of status and control messages.

The Digital ICS allows secure transmission and reception of voice and data information over aircraft communication radios.

CPLT Cyclic

PLT Cyclic

CPLT HDST

PLT HDST

Audio CPLT PTT Foot SW

Audio

CPLT Jack

PLT PTT Foot SW

PLT Jack

PLT CTRL PNL

CPLT CTRL PNL ICS

ICS

CVR Audio

Radio

Radio

CVR Audio FMS-1

FMS-2 CVR/FDR ICU

CVR Audio MIL-STD-1553B

Analo g RS48

ICS Station (No. 1 and No. 2) 10/2/2003 Page 1 of 10 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

LHG HDST

LHS EXT MAINT HDST

RHG HDST

RHS EXT MAINT HDST

Audio

Audio

Audio

Audio

LHG Jack & PTT SW

LHS EXT MAINT Jack & PTT SW

RHG Jack & PTT SW

RHS EXT MAINT Jack & PTT SW

LHG CTRL PNL LHG PTT Foot SW

RHG CTRL PNL

Audio

Audio

Radio PTT

ICS PTT

Avionics Relay Unit

WOW

RHG PTT Foot SW

Audio

Audio

Radio PTT

ICS PTT

ICU Analog RS-485

UH-60M Digital ICS (Station No. 3 and No. 4)

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The ICU provides the following major capabilities:

Components The Digital ICS is a centralized system consisting of one Interface Control Unit (ICU), five control panels, cyclic, and collective sticks. These control panels are located in the following crewmember stations: • • • • •

Pilot Copilot Right Hand (RH) Gunner Station Left Hand (LH) Gunner Station Troop Command Station

ICU The ICU is the primary method of interfacing the aircraft analog audio subsystems for the Digital ICS. The ICU has the capability to receive discrete inputs, trigger internally generated noncontrolled aural warnings, and send these warnings digitally to the desired control panels. In the event of a fault that would prohibit digital communication, the ICU will be required to output a discrete signal notifying of a system fault.

• • • •

• • • • • •

Call Function Voice Operated Xmit Vox Interphone Communications External Communications - Transmission on radios - Radio retransmission - Monitoring of selected navigation aids and other audio assets - Tone generator - Secure voice communications Interface with existing headsets Interface with the Maintenance Terminal (MT) 1553 Interface to Mission Computer CVR Preflight Audio Test Central Aural Warning System (CAWS) Built-in Test (BIT)

ICS ICU

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ICS Control Panels

ICS/RADIO

The ICS control panels serve as the interfaces for all crewmembers to communicate with each other and with other audio systems internal and external to the aircraft. Using the control panels, crewmembers are able to select and control audio signals such as interphone, radio communication, radio navigation, aural alerts, and interface with other associated equipment such as CVR and data links. The control panel utilizes one rotary control for the selection of interphone and radio communication modes, a master volume rotary control, and 11 push/pull rotary controls for on/off and volume adjustment of individual audio sources. The primary operational modes consist of PVT, ICS, COM 17, RMT, and BU.

This 2-detent trigger switch provides the pilots with push-to-talk radio control for either the internal communication system or external radio communication. The first detent position activates the ICS radio while the second, fully depressed position activates the external radio. Additionally, a foot-activated switch on the inboard side of each pilot station will activate transmission on the Intercommunication system (ICS) or the external radio as selected on the ICS panel.

ICS CONTROL PANEL

Cyclic Sticks The cyclic stick assembly function and use is slightly different from the UH-60A/L. The UH60M cyclic stick assembly replaces the UH60A/L cyclic with one adapted from Naval Hawk aircraft. The two major changes made to the cyclic sticks are: the stick grip and the number and location of switches located on and about the grip.

Collective Sticks RADIO SEL The radio select switch on each collective stick grip provides four signals to the to the on-side FMS. • EXT NEXT - provides the same function as the NXT (next) buttons on the FMS keyboard, allowing cycling through a selected list on the FMC display. • COMM/STEP DN - allows the list of available radios to be cycled through in descending order. • EXT PREV - provides the same function as the PRV (previous) button on the FMS keyboard, allowing cycling through a selected list on the FMS display. • COMM/STEP UP - allows the list of available radios to be cycled through in ascending order. HOOK EMER REL

LDG

LT EXT PUSH ON OFF

HUD

CRSR SLEW PUSH-SELECT

PG

UP

DCLT

EXTD RA SE D L UP

BRT L

DIM

R

PG DN

RETR DN

CARGO REL

C/F DISP

TR IM DN L R UP

WEPS REL

TRIM REL

RMT SBY

TRIM

UMFC001_6

UMFC001_7

COLLECTIVE STICK GRIP

CTCLIC STICK GRIP

10/2/2003 Page 5 of 10 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Digital ICS Controls

BU (Backup)

PVT (Private)

In the event that digital communications is lost, BU mode allows for both the pilot and copilot the capability to bypass the Digital ICS digital bus to allow for a hardwired connection to a communication radio. The BU mode provides the pilot and copilot capability to receive and transmit analog headset, microphone, and control signals independently to a dedicated communications radio and not be affected by a loss of digital communications.

The Digital ICS has the ability to provide up to two private nets via embedded programmable software. 1. PVT 1 Implementation • This method of accessing PV mode is available when the operator selects PV and then actuates the Radio PTT key. Access to the private network is enabled and all stations that have selected PV will hear the operator’s audio. Actuating the ICS PTT key input at any station will cause audio to be heard at all stations when any one or all stations are in PV mode.

BU mode is also provided immediately after power is first applied to the Digital ICS regardless of the TX control position to allow analog communication until the Digital ICS initialization is complete. After initialization is completed, the BU mode will default to ICS mode.

ICS ICS mode is available to all ICS stations and is used only for interphone communication with all crewmembers. This mode is not available to external headset cords when the aircraft is in flight. COM 1-7 (Communication 1-7) All aircraft communication radio will be selected by the COM 1-7 positions on the TX control and only one radio can be selected for transmission at any given time when provided a Radio PTT. RMT (Remote) The RMT (Remote) mode allows the pilot and/or copilot to toggle through all communication radios and presets channels via the thumb control switches installed on the pilot and copilot collective sticks. The Digital ICS also provides a method for accepting sequential step up and down control inputs from the aircraft BC for the remote selection of all seven communications radios. The inputs shall be available by both discrete input lines and by the MIL-STD-1553B data bus.

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MVOL (Master Volume Control) The MVOL control has the capability to raise or lower all individually controlled volumes at the same time. The MVOL control does not have any effect on uncontrolled aural alerts that may be generated by external sources or internally by the Digital ICS. The Digital ICS has the capability to store the last MVOL to be recalled during system power up. Individual Volume Control Knobs • • • • • • •

COM1: VHF-FM NO.1 COM2: UHF-AM COM3: VHF-AM COM4: VHF-FM NO.2 COM5: HF COM6: PLS (HH-60M Only) COM7: Spare (SATCOM Future Growth)

• • • •

MON1: VOR/ILS (& TACAN on HH-60M) MON2: MB MON3: LF/ADF MON4: RADALT

4. HOT MIC • This mode provides for interphone communication at all times without requiring the operator to key ICS PTT. HOT MIC is available for both ICS and PVT. 5. CALL (Call Override) • Pressing and holding the CALL pushbutton selects this mode. CALL mode overrides all other interphone communications at all stations, except for warnings and alerts. It overrides all interphone transmissions and shall be heard at all stations at full volume regardless of any control settings. This mode remains active for as long as the operator presses the call pushbutton. ICS CONTROL PANEL

Pushbutton Mode Control The pushbutton mode control is used to select one of four different modes relating to the operator interphone audio. Mode control utilizes illuminated pushbuttons with embedded NVIS Green B LED indicators. Any selection of the pushbutton mode control should not affect or override the means by which radio communications is accomplished. These four pushbutton modes are: 2. ICS • This mode requires that an operator perform either hand or foot ICS PTT keying in order to communicate by interphone. 3. VOX (Voice Operated) • This mode provides for hands off and foot off voice activated keying of the ICS interphone

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Pressing SK-8 accesses the status page for ICS.

Indications The four LED pushbutton indicators illuminate indicating that the specific pushbutton mode is operational. Momentarily pressing the specific pushbutton will cause the LED indicator on the pushbutton to flash, indicating that the operator can adjust the volume on the MVOL control to the desired level. The Individual volume controls have Night Vision Goggles (NVG) compatible white reference dot that illuminates during selection of the control and extinguish during deselection of the control.

Flight Management System (FMS) Status and Test The FMS provides status and test via MILSTD-1553B data bus to the COM system. STATUS Pressing the STS button on the FMS and then pressing soft key (SK)-3 will access the status page for the COM.

COM STATUS PAGE The status page provides the pilots with the software version for the ICU and the onside ICS control panel. It also provides the status of each control panel and will detect a fault if any of the control panels or if the ICU has failed. The FMS Bus Controller (BC) provides a Digital ICS BIT status via MIL-STD-1553B data bus.

MAIN STATUS PAGE When the Status page appears, an overall status of the COM will be displayed. ICS STATUS

10/2/2003 Page 8 of 10 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

TEST

System Operation

Pressing the TST button on the FMS and then pressing SK-3 will access the test for the Communication Suite. When the Test page appears, an overall status of the radios will be displayed. Test for ICS can be selected from this page by pressing SK-8 for ICS. An Initiated Built-In Test (IBIT) can be initiated for the ICS. Upon completion of the test, all faults and failures will be displayed under the ICS if a fault or failure should occur.

The Digital ICS completes its system bootup in less than 15 seconds after first application of the ICU and control panel input power. Boot-up time begins as soon as aircraft power is applied to the ICS. During boot-up, the crewmembers can communicate by transmitting and receiving using the designated emergency communications radio. Warm-up time is not required for the Digital ICS to operate normally when aircraft power is applied. Upon power up, a Start-up BIT (SBIT) will run automatically to indicate that the Digital ICS system is up and running. The SBIT, as well as the IBIT, only runs when the aircraft is on the ground in Weight-On-Wheels (WOW) status. The SBIT will not re-run after Digital ICS input power interrupts of 0 vdc for less than 100 milliseconds. A Continuous BIT (CBIT) will run continuously, in flight as well as on the ground, without interfering with the Digital ICS operation, and will report failures as they occur.

COMM TEST PAGE

Single Point Failures The Digital ICS has no single point failure that causes significant degradation to the operation of the system. Single point failures include loss of power or damage to any of the individual components, damage to a bus connector, or damage to a bus connection between components. In the event that the Digital ICS is no longer available, it will have hardwired analog connections between all control panels to provide uninterrupted interphone communication. The ICU will provide four hardwired communication/navigation emergency radio connections.

ICS TEST PAGE

The Digital ICS will detect and implement these features automatically during the loss of digital communications. A reversionary feature is selectable by the pilot and/or copilot on their respective control panels.

10/2/2003 Page 9 of 10 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

NOTES

10/2/2003 Page 10 of 10 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Section 3-2.2 AN/ARC-164(V) UHF AM RADIO System Description The AN/ARC-164 UHF AM radio system on the UH-60A/L aircraft is being replaced with AN/ARC-164(V) UHF AM radio system on the UH-60M aircraft. A UHF AM radio control panel located in the lower center console controlled the UHF AM communications on the UH-60A/L. Now the UHF AM radio is controlled by the Flight Management System (FMS) via a MILSTD-1553B data bus. The AN/ARC-164(V) UHF AM radio set provides line-of-sight, two-way AM voice communications in the frequency range of 225 to 399.975 MHz, selectable in 25 kHz increments. It operates in guard frequency of 243.000 MHz. This radio also operates in a frequency-hopping antijamming (HAVEQUICK) mode. The set is configured for use with a KY-58 processor to provide narrow and wide band encrypted voice communication. The UHF AM radio set is fully integrated into the avionics FMS and is controlled by the pilots through the FMS via MIL-STD-1553B data bus.

Components The AN/ARC-164(V) UHF AM radio set is comprised of the following components: • • •

Receiver Transmitter (R/T) UHF Antenna TSEC KY-58

UHF AM R/T The R/T, located in the cockpit nose tunnel, receives and transmits both clear and secure voice communications. Besides operating in normal mode, it is also capable of operating in the Havequick modes. The transmitter has an output capacity of 10 watts.

UHF AM R/T

UHF Antenna (ANT) The antenna, mounted in the same location as the UH-60A/L, (underneath transition section), transmits and receives radio frequencies (rf) energy.

TSEC KY-58 The TSEC KY-58 for the UHF AM radio on the UH-60M is located on the top shelf of the nose avionics compartment. It provides the radio with wide and narrow band secure speech capabilities.

Controls The AN/ARC-164(V) UHF AM radio set is controlled by the FMS in the UH-60M helicopter. The Communication Summary (COM SUM) screen is accessible by a single keystroke on the FMS, or by the first left or right operation of the collective stick switch. The screen appears with the radios each displaying: • • • • •

Band (the radio number is included when the band is not unique to one radio) Preset number (1 through 20) Frequency/Channel or Net Secure status Call sign, station name, or any eightcharacter sequence chosen

Pressing soft key (SK)-2/SK-10 allows you to set and tune the radio (UHF-AM).

10/2/2003 Page 1 of 11 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

VHF/AM #2 COMM

IRCM

#2 GPS

VHF/AM #1 COMM

VHF/AM COMM

UPPER IFF

#1 GPS

VHF HOMING (BOTH SIDES)

TROOP CMDR ANT

RADAR WARN (BOTH SIDES)

GLIDESLOPE

RADAR WARN (BOTH SIDES)

AVR-2A LDS SENSOR

AVR-2A LDS SENSOR

HF COMM

LF/ADF

VOR/LOC

LOWER IFF

RADAR WARN BLADE ANT RADAR ALTIMETER

STORMSCOPE MARKER BEACON

AVR-2A LDS SENSOR

UHF/AM COMM

UH-60M ANTENNA ARRANGEMENT

UH-60M Antenna Arrangement

10/2/2003 Page 2 of 11 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

UMHV008

UHF AM COM SETTING PAGE

NOTE There are no secured communications available at this time for first flight.

UHF AM COMM SUMMARY PAGE

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KY-58

Nose Avionics Compartment

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UHF AM R/T COCKPIT NOSE TUNNEL

10/2/2003 Page 5 of 11 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Indications

FMS Status and Test

The indications for the UHF AM radio will be displayed in the header section of the FMS. The header will always reflect the tuned status of the radio. Once the radio is tuned to its channel/frequency and call station, it will also be displayed on the Primary Flight Display (PFD).

STATUS Pressing the STS button on the FMS and then pressing SK-3 will access the status page for the Communication Suite. When the Status page appears, an overall status of the radios will be displayed. Status for UHF AM radio can be selected from this page by pressing SK-7. Status for the UHF AM radio will then be displayed. There are no faults associated with the UHF AM radio.

HEADER DISPLAY On the lower left (copilot) and lower right (pilot) parts of the PFD, the selected/tuned communication radio data will be displayed. COM STATUS PAGE

COMMUNICATION DATA DISPLAY

UHF AM STATUS PAGE

10/2/2003 Page 6 of 11 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

TEST

System Operation

Pressing the TST button on the FMS and then pressing SK-3 will access the test for the Communication Suite. When the Test page appears, an overall status of the radios will be displayed. Test for UHF AM radio can be selected from this page by pressing SK-7. An Initiated Built-In Test (IBIT) can be initiated for the UHF AM radio. Upon completion of the test, UHF AM radio IBIT screen will display the overall subsystem test results only. “SYSTEM FAIL” will be displayed should a fail occur.

Operation and communication control of the UHF AM radio is provided by the FMS via MILSTD-1553B data bus. UHF AM data will be displayed on the PFD via UHF AM outputs. The main transceiver operates in any one of the 7,000 channels, spaced 0.025 MHz units in the 225.000 to 399.975 MHz UHF military bands. The guard receiver is fixed tuned to 243.000 MHz (emergency frequency). The radio is capable of voice data communications in both clear and secure, in normal and HQ modes of operation. When the HQ mode is active, an electronic counter-countermeasures ECCM synthesizer module provides jam resistance and makes direction finding difficult. Operation of the UHF AM radio is accessed through the FMS. The pilot, via the FMS test page to ascertain the system status, can initiate a Built-In Test (BIT). After performing BIT, the radio reinitializes in its power-up state. The radio must then be tuned to its frequency/channel station and have audible side tone before proceeding

COM TEST PAGE

Control and selection of UHF AM (COM 2) radio is through the FMS via MIL-STD-1553B data bus. Depressing the COM function key on the FMS keypad or the RAD SEL switch on the collective stick grip brings you to the primary COM SUM page. SK arrows will appear in inverse video when the corresponding operator’s ICS transmit select switch is set to that radio. The arrow (either inverse or plain) changes to an asterisk for up to 5 seconds after the radio is tuned indicating that the tune is in progress. If the ICS select position is changed while on this screen, both the inverse video indication and edit box will change to the new selection.

Manual Tuning UHF AM RADIO TEST PAGE

To manually tune the selected radio to a desired frequency, the operator presses the SK next to the radio a second time or presses to move the edit box to the frequency field and enter the new frequency on the scratchpad line. Once a manual entry has been made, the preset number for the radio is set to “0” and the call sign is blanked.

10/2/2003 Page 7 of 11 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

“LAST” Function

HVQK Operation

For the edit-selected radio, the status is automatically saved whenever the radio is detuned from a manually entered frequency. Pressing SK-5 will restore the last manually tune frequency, while saving the current frequency (preset or manual). A second press of the SK swaps back to the saved frequency.

When changing between UHF AM and HVQK operation, (any waveform), the radio is always restored to the frequency/net preset which was in use when the radio was last in that band. If first time since power on, then the default frequency or net for that band will be used.

The “LAST” function saves one setting per band of operation.

Changing Band For bands with frequency format displays, manual tuning to a new frequency outside the current band automatically changes the radio to that band. Otherwise, bands can be changed via the “BAND” screen. Whenever the operator commands a radio to switch from one band to another, the radio is restored to the last frequency/channel/net and preset that was current when the band was last in use.

PT/CT (Plain Talk/Cipher Talk)

HVQK I, II, and NON-NATO, selections are actually sub-modes (waveform selections) of HVQK operation, and are normally associated with HVQK presets. Preset tuning will change the submode from preset to preset without having to invoke this screen. Invoking this screen will change the selected waveform for the current operating HVQK Net, regardless of what is specified in the preset. Changes in waveform via this screen do not change the stored preset value. Upon returning from Emergency Comm, switching back to HVQK mode, the ARC-164 radio always goes back to HVQK I mode no matter what mode it was in before (II or NONNATO).

The PT/CT legend will only appear next to this function key if a crypto is configured for the editselected radio. When configured, the key will toggle between PT and CT and the selected PT or CT will appear in inverse video. Also, SEC, in inverse video, will appear after the radio frequency when the radio is in cipher. If it is in plan, MHz or no text will be displayed after the frequency.

UHF AM SETTING

UHF AM SUMMARY

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The selected radio set remains in the retransmit mode until there is no rf input to the set. Retransmit control is then removed and both radio sets operate in the receive mode.

RADIO RETRANSMISSION FM 1/FM 2 FM 1/VHF

FM 2/UHF FM 2/VHF

FM1/UHF

VHF/UHF OFF

UMAV003_11 HVQK SETTINGS

RETRANSMISSION CONTROL PANEL

Retransmission In retransmission operation, the set is used with No. 2 VHF FM and UHF AM radio sets as a radio relay system to transmit audio between remote terminal stations. Both radio sets are in the receive modes with the function selector switch on the retransmission control panel placed to one of the three radio sets. Reception of an rf signal sends a retransmit control signal from the radio set main receiver, through the radio retransmission control panel, to key the transmitter of the selected radio set. The audio output of the receiving radio set is also applied through the retransmission control panel to the transmitter of the selected radio set for retransmission.

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RCV AUDIO

D-ICS

No.2 DC PRI 28 Vdc, 7.5A

No.2 DC PRI 28 Vdc, 2A

XMIT AUDIO RCV AUDIO

ICU

XMIT AUDIO

XMIT X-AUDIO

COMM 2 UHF - AM

CREW PTT

AVIONICS RELAY PANEL

RCV AUDIO

A/D PTT3 A/D PTT3

XMIT AUDIO

KY – 58 TSECON241800-1

MD-1359/A

RT-1614/ARC164(V)

ARC-201 (FM2) RXMT Control RXMT Signals

Mount MT-4838/ARC164(V)

RXMT Control

Retransmit Panel

RXMT Signals

X-MODE CNTL

DPTT

DEFAULT FREQ = 225 MHZ EMERG FREQ = 243 (AM)

CNTL SIGNALS

1553B Bus A

Copilot FMS

Pilot FMS

DIG DATA_3 TX DIG DATA_3 RCV DIG DATA CLK

RT ADDR = 16H 1553B Bus B

DPTT

Loc: Transition

No. 2 EGI H-764GU

AN/ARC-164

X- MODE GD

CTOIP

CREW PTT SENSE

IDM

TUB

HAVE QUICK TOD

CREW PTT RCV X-AUDIO

UHF

X-MODE CNTL RCV X-AUDIO

Loc: Nose

Location : Transition

Emer/Guard

STATUS

Load Access Panel

CONTROL Zeroize CTOIP SEL

Data FILL

UHF-AM Block Diagram

10/2/2003 Page 10 of 11 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Emergency Control Panel

NOTES

10/2/2003 Page 11 of 11 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Section 3-2.3 AN/ARC-186(V) VHF AM Radio System Description The AN/ARC-186(V) VHF AM radio set provides transmission, reception, and retransmission of amplitude modulated with installation of other associated equipment. The transceiver has a tunable main receiver and transmitter, which operates on any one of 1,469 AM discrete channels, each spaced at 25 kHz apart within the frequency range of 116.00 through 151.675 MHz.

V O L

V H R

S Q D I S

T O N E

12

EMER AM FM MAN

5

PRESET LOAD

PRE

50

8

5

DF TR OFF

UMAV003_24

BACK-UP VHF AM CONTROL PANEL

RT-1300B/ARC-186 (V) The RT-1300B/ARC-186 (V) is a two-band VHF AM radio set that is controlled via MILSTD-1553B data bus. The radio has a 10watt transmitter, 20 preset channels, remote capability, single antenna, dual band operation capability, transmit and receive on the general frequencies.

The VHF AM radio set is fully integrated into the avionics Flight Management System (FMS) and is controlled by the pilots through the FMS via MIL-STD-1553B data bus. In addition, the VHF AM radio has a separate control panel for communication on initial power up and can be used as a backup radio in the event of the FMS system failure.

Components The AN/ARC-186(V) VHF AM radio set is comprised of the following components: • • • •

Control panel – Lower Center Console Receiver Transmitter (R/T) (RT1300) – Left Hand (LH) Seatwell Band pass filter – LH Seatwell Antenna – Tail Cone On Top LH Side

VHF AM R/T

Backup Control Panel The C-10604A (V) (6)/ARC-186 (V) control panel provides the capability to manually select any channel of the 4.080 channels available with the capability of storing 20 preset channels in memory. The control panel provides a key for momentary tone transmission and also a means of disabling squelch. The control also provides audio volume control.

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Forward

VHF AM R/T COPILOT SEATWELL

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Controls The COM SUM screen is accessible by a single keystroke on the FMS, or by the first left or right operation of the collective stick switch. The screen appears with the radios each displaying: • • • • •

Band (the radio number is included when the band is not unique to one radio) Preset number (1 through 20) Frequency/Channel or Net Secure status Call sign, station name, or any eight-character sequence chosen

Pressing soft key (SK)-3/SK-10 allows you to set and tune the radio (VHF AM).

VHF AM SETTING PAGE

COM SUMMARY PAGE

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Indications The indications for the VHF AM radio will be displayed in the header section of the FMS. The header will always reflect the tuned status of the radio. Once the radio is tuned to its channel/frequency and call station, it will also be displayed on the Primary Flight Display (PFD).

VHF AM STATUS PAGE

HEADER DISPLAY On the lower left (copilot) and lower right (pilot) parts of the PFD, the selected/tuned communication radio data will be displayed.

COMMUNICATION DATA DISPLAY

FMS Status and Test STATUS

VHF AM STATUS

Pressing the STS button on the FMS and then pressing SK-3 will access the status page for the Communication Suite. When the Status page appears, an overall status of the radios will be displayed. Status for VHF AM radio can be selected from this page by pressing SK-2. Status for this radio will then be displayed. All faults and failures will be displayed under the VHF AM if a fault or failure should occur.

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TEST

System Operation

Pressing the TST button on the FMS and then pressing SK-3 will access the test for the Communication Suite. When the Test page appears, an overall status of the radios will be displayed. Test for VHF AM Radio can be selected from this page by pressing SK-2. An Initiated Built-In Test (IBIT) can be initiated for the VHF AM Radio. Upon completion of the test, VHF AM Radio IBIT screen will display the overall subsystem test results only. All faults and failures will be displayed under the VHF AM if a fault or failure should occur.

The pilot, via the FMS test page to ascertain the system status can initiate a Built-In Test (BIT). After performing BIT, the radio reinitializes in its power up state. The radio must then be tuned to its frequency/channel station and have audible side tone before proceeding. Control and selection of VHF AM (Com 3) radio is through the FMS via MIL-STD1553B data bus. Depressing the COM function key on the FMS keypad or the RAD SEL switch on the collective stick grip brings you to the primary COM SUM page. SK arrows will appear in inverse video when the corresponding operator’s Intercommunication System (ICS) transmit select switch is set to that radio. The arrow, (either inverse or plain) changes to an asterisk for up to 5 seconds after the radio is tuned indicating that the tune is in progress. If the ICS select position is changed while on this screen, both the inverse video indication and edit box will change to the new selection.

Manual Tuning

COM TEST PAGE

To manually tune the selected radio to a desired frequency, the operator presses the SK next to the radio a second time or presses to move the edit box to the frequency field and enter the new frequency on the scratchpad line. Once a manual entry has been made, the preset number for the radio is set to “0” and the call sign is blanked.

“LAST” Function For the edit-selected radio, the status is automatically saved whenever the radio is de-tuned from a manually entered frequency. Pressing the SK-5 will restore the last manually tuned frequency, while saving the current frequency (preset or manual). A second press of the SK swaps back to the saved frequency. VHF AM TEST

The “LAST” function saves one setting per band of operation.

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Changing Band

Retransmission

For bands with frequency format displays, manual tuning to a new frequency outside the current band automatically changes the radio to that band. Otherwise, bands can be changed via the “BAND” screen. Whenever the operator commands a radio to switch from one band to another, the radio is restored to the last frequency/channel/net and preset that was current when the band was last in use.

In retransmission operation, the set is used with No. 1 VHF FM, or No. 2 VHF FM, or UHF AM radio sets as a radio relay system to transmit audio between remote terminal stations. Both radio sets are in the receive modes with the function selector switch on the retransmission control panel placed to one of the three radio sets. Reception of a radio frequency (rf) signal sends a retransmit control signal from the radio set main receiver, through the radio retransmission control panel, to key the transmitter of the selected radio set. The audio output of the receiving radio set is also applied through the retransmission control panel to the transmitter of the selected radio set for retransmission. The selected radio set remains in the retransmit mode until there is no rf input to the set. Retransmit control is then removed and both radio sets operate in the receive mode.

RADIO RETRANSMISSION FM 1/FM 2

VHF AM SETTINGS

FM 1/VHF

FM 2/VHF

FM1/UHF

Backup Control Operation When the backup control panel on/off switch is placed to the on position, it sends a ground signal to the avionics relay panel. The relay then switches over enabling the backup control panel and disabling the pilot and copilot FMS and sends a ground signal to the R/T to take control.

FM 2/UHF

VHF/UHF OFF

UMAV003_11

RETRANSMISSION CONTROL PANEL

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Rcv Audio

D-ICS ICU

No.2 DC ESNTL BUS 28 VDC, 5A

Tail Cone

COMM 3 VHF - AM

CREW PTT

RG-142 Control Signals Digital Data TX

IDM MD-1395/A

Blade Antenna AM

Xmit Audio

Digital Data RX A/D PTT4

CREW PTT SENSE

BP FILTER Loc: cplt seatwell RXMT Control

Transceiver RT-1300B/ARC186(V)

RXMT Signals RXMT Signals RXMT Control

Retransmit Panel

Mount MT-6051/ARC 186(V)

A/D PTT4

AVIONICS RELAY PANEL

CREW PTT TAKE CONTROL

RETRAN KEY BLANKING PWR ON/OFF

DEFAULT FREQ = 118 MHZ EMERG FREQ = 121.5 MHZ

Copilot FMS

Pilot FMS

CTRL ENB/DISA RT ADDR. = 04H

Control Panel C-10604A(V)6/ ARC-186(V)

CLOCK Loc: Cplt Seatwell DATA TONE KEY

FM BAND INDICATOR

BACKUP ON/OFF

AN/ARC-186(V) VHF AM BLOCK DIAGRAM

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NOTES

AN/ARC-186(V) VHF AM BLOCK DIAGRAM

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Section 3-2.4 AN/APX-118 IFF System Description The AN/APX-100 IFF Transponder in the UH-60 A/L aircraft is being replaced by the AN/APX-118 in the UH-60M aircraft. The AN/APX –118 has embedded Mode 4 crypto and Mode S capabilities. The AN/APX-118 IFF transponder provides identification, altitude, and surveillance reporting in response to challenges from airborne, ground-based, and surface interrogators.

data link. Mode S coding supports error detection in both directions, as well as error correction on the reply link, thus providing reliable data on aircraft identification, altitude, and other transmitted information.

Components The AN/APX-118 IFF System is comprised of the following components: • • •

Required Capabilities. The purpose of the Global Air Traffic Management (GATM) program is to preserve operational readiness and Army aviation access to global aviation routes into the 21st century by equipping Army aircraft to meet the emerging requirements of the worldwide air navigation system. The International Civil Aviation Organization (ICAO), Federal Aviation Administration (FAA), civil aviation authorities (CAA), and other regional and national planning bodies are developing requirements for the future air traffic environment and plan to implement new air traffic management architecture to relieve the tremendous strain on the current air traffic control (ATC) system. If aircraft are not equipped with the appropriate new technologies, they will not be able to operate in airspaces where new separation standards and ATM procedures are implemented by civil aviation authorities, and will therefore be excluded from those airspaces. Mode S The Mode S interrogator provides surveillance of all transponder-equipped aircraft (Mode 3/A and Mode S) within lineof-sight coverage. The Mode S interrogators can provide surveillance of a large number of aircraft with better accuracy and reliability than current Secondary Surveillance Radar (SSR) systems. Mode S also provides ground-to-air, air-to-ground, and air-to-air

Transponder – Located in aft trans Upper IFF Antenna – Top aft of the transition area Lower IFF Antenna – Bottom of tail cone just aft of UHF Ant

The IFF system is interfaced with the following system components: • • • • • • • •

Avionics Relay Panel – Weight-onWheels (WOW) Audio Junction Box - (Audio out for Mode 4) Data Concentrators - (IFF caution and STBY ADVSY) FMS 1553B Data Bus Air Data Computers - (Altitude Data through ARINC-429) Mission System Control Panel (Mode 4 Hold) Central Suppression Unit (CSU) Emergency Control Panel (ECP) (Mode 4 Zeroize)

IFF Transponder The IFF Transponder is a space diversity transponder, which receives rf interrogations from the upper and lower antennas. The reply to the interrogation is transmitted on the antenna, which receive the stronger interrogation signal. The transponder provides all the capabilities of modes 1, 2, 3/A, C, and 4. In addition, the transponder incorporates Mode S capability.

Antennas The upper and lower antennas can be selected between Top, Bottom, or Diverse (both) from the FMS. When the antenna

10/2/2003 Page 1 of 8 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

selection is set to TOP or BOTTOM it will remain at that state for 10 seconds, then the transponder will change back to DIVERSE. On power up the antenna selection will default to DIVERSE.

Emergency IFF The “IFF EMERG” switch on the ECP should be left in the center position. The up position selects emergency IFF codes and is equivalent to the emergency switch on the APX-100 control panel The down, momentary position, is equivalent of the “refuel hold” function. This switch provides controls redundant to those already provided on the APX-100 control panel but they are provided on the ECP.

E M E R G

Z E R O I Z E

CDU-1

CDU-2

GUARD

A U T O

Controls The AN/APX-118 IFF Transponder is fully integrated with the FMS and is controlled by the FMS via the MIL-STD-1553B data bus. The Transponder can be accessed by a single keystroke “XPD” on the FMS. The screen appears with Transponder Control. The screen provides Transponder Mode control and antenna selection. The screen also provides Mode Code status as well as the capability to change the codes. By pressing SK-10 will allow you to access the Transponder Test screen.

IFF EMERG N O R M

N O R M

+

EMERGENCY CONTROL PANEL

UMAV003_3

ECP

TRANSPONDER CONTROL PAGE When selecting the XPD function key the selected mode will default to M1 accesses this page. Enclosing the mode on three sides, in a “bucket”, indicates the selected mode.

FMS CONTROL PANEL

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AN/APX-118 IFF Transponder

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Indications

Status

The indications for the APX-118 Transponder Radio will be displayed in the header section of the FMS. Once the radio is tuned to its channel/frequency and call station, it will also be displayed on the Primary Flight Display (PFD).

Pressing the STS button on the FMS and then pressing SK 3 will access the status page for the Communication Suite. When the Status page appears an overall status of the radios will be displayed. Status for APX118 radio can be selected from this page by pressing SK 4. Status for this radio will then be displayed. All faults and failures will be displayed under the IFF-XPDR if a fault or failure should occur.

HEADER DISPLAY The audio setting will provide an Alert indicator when a Mode 4 interrogation is received and an Audio indicator when the transponder does not reply to a Mode 4 interrogation. The Alert indicator will be displayed on the FMS Status Line.

COMM STATUS PAGE

Transponder Indications The transponder provides Light Emitting Diode (LED) fault indicator lights: •

•

•

WRA GO LED (green) – Steady illumination indicates the IFF transponder is fully operational. The LED will flash to indicate a fault. LOW BTRY LED (amber) – Illumination of this LED indicates that the internal backup battery voltage is low and that the batteries should be replaced. This LED can be activated only with primary power applied to the transponder. KEY LOADED LED (green) – Flashes momentarily upon completion of a successful keyfill, regardless of whether primary power is applied to the transponder. It will also indicate that both A and B crypto codes have been loaded.

IFF XPDR STATUS PAGE

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Test

System Operation

Pressing the TST button on the FMS and then pressing SK 3 will access the test for the Communication Suite. When the page appears an overall status of the radios will be displayed. Test for APX-118 Radio can be selected from this page by pressing SK 4. An Initiated Built-In Test (IBIT) can be initiated for the APX-118 Radio. Upon completion of the test, APX-118 Radio IBIT screen will display the overall subsystem test results only. All faults and failures will be displayed under the IFF TPDR if a fault or failure should occur.

The transponder has an INT/BIT pushbutton switch to perform a full Initiated-Built-In-Test (IBIT). Activating this switch will invoke a full IBIT. A malfunction causes the LED indicator on the transponder to flash. The pilot via the FMS test page to the system status can also initiate a BIT. After performing BIT the radio reinitializes in its power up state. When power is applied the following modes on the FMS will be in their default state: • Mode 4 Reply indicator will default to “Audio” • Mode 4 will default on the “OFF” state. • Mode 4 Code will default to “A” These functions are not required for first flight. Interrogations

COMM TEST PAGE

The transponder receives radio frequency (rf) interrogation signals from a challenging station. The signals are received by the top and bottom antennas and routed to the transponder. The receiver sections are tuned to an incoming signal frequency of 1030 MHz. The transponder recognizes the interrogation mode of challenging station. Reply The transponder transmits a reply message in the same mode as the interrogation from the challenging station. The transmitter sections are tuned to a 1090 MHz frequency. If the interrogation is for a mode 1,2,3/A, or C reply and if the respective M-1, M-2, M-3/A or M-C control mode is selected on the FMS is placed to “ON” the appropriate reply is automatically transmitted. The identification numbers required for Mode 1,2,3/A are obtained from the Mode selection on the FMS.

IFF XPDR TEST PAGE

Mode C reply requires altitude data, which obtained from the Air Data Computers (ADCs) to the Primary Flight Displays (PFDs) via ARINC-429 data bus.

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Mode 4 interrogations are answered with a reply from the transponder if the Mode 4 is selected on the FMS. When a Mode 4 reply is transmitted, a visual indication is available on the FMS and an audio signal may be selected and supplied to the pilot and copilot ICS stations. An IFF legend will appear on the MFDs Engine Indicating Crew Alerting System (EICAS) when there is an error or malfunction in the Mode 4 reply operation.

Zeroize The IFF encryption codes will automatically zeroize on removal of equipment. The Mode codes will be automatically zeroized on orderly system shutdown unless the refuel hold function is selected.

EMCON (Emission Control) The IFF transponder can be placed in an EMCON mode by removing a ground from the transponder interface connector or by selecting the STBY mode on the FMS. Removing the ground overrides mode control settings from the FMS. It also causes a STANDBY caution message to be displayed on the EICAS. SK-8 XPDR MODE on the FMS will toggle between Standby and Normal when commanded. SK-8 arrow will not be displayed when in EMERGENCY MODE. On power up transponder mode will default to Standby. Emergency Operation Emergency transponder operation is enabled when it is selected on the ECP. Line two of SK-8 on the FMS will display EMERGENCY (in inverse video). Pressing this switch will not change the transponder mode. If this switch is pressed while the transponder is in the Emergency state, the status message “EMERGENCY MODE SET” will be displayed in the Status line of the FMS Comm Header. While the transponder Mode is set to Emergency, the Mode 1,2,3/A, C, 4 and S states will be set to “ON” and mode 3/A code will be set to 7700. All of these settings will not be editable while the transponder is in Emergency Mode. When the APX-118 transponder transitions out of Emergency mode it will reset the ON/OFF states foe Modes 1, 2, 3/A, C, 4, and S to their previous states. It will also return Mode 3/A code to its previous state.

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ICS ICU

Mode 4 Audio Out

MCP Pressure Altitude Data ARINC 429

LED Return

Air Data Cmptr

MFD

Clock Data Fill Panel

Switch Ref Key Fill Ref

Zeroize Zeroize Codes Emergency Cntl Pnl

Emergency

Mode 4 Caution Lt

Data Data Request

APX-118 IFF Common Xpnder (CXP) P/N 1008939G-5 Mount MT-4811 (APX-100) or MT-7221

MFD

DCU #1

DCU #2

MFD

Standby Advisory MFD

MCP Emergency Mode IFF Control Page

Normal

1553B Bus A

Weight on Wheels Gnd=Aircraft on Ground

Copilot FMS

Pilot FMS

Top Ant

Relay Pnl 1553B Bus B Mission Systems Cntl Pnl

Retain Crypto Codes

M4 Hold

AN/APX-118 IFF BLOCK DIAGRAM

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Bottom Ant

Section 3-3.1

Components

AN/ARN-147 (V) VHF Omnirange/Instrument Landing System (VOR/ILS) System Description

VOR/ILS Receiver The VOR/ILS receiver, located in the cockpit nose avionics compartment “tunnel”, contains three discrete antenna connections that feed three receivers (VOR/LOC, GS, MB).

The AN/ARN-123(V) VOR/ILS navigation system on the UH-60A/L aircraft is being replaced with AN/ARN-147(V) VOR/ILS navigation system on the UH-60M aircraft. A VOR/ILS control panel located in the lower center console controlled the VOR/ILS system on the UH-60A/L. Now the VOR/ILS is controlled by the Flight Management System (FMS) via a MIL-STD-1553B data bus. The AN/ARN-147(V) VOR/ILS navigation system is source selected on the Flight Director/Display Control Panel (FD/DCP) via ARINC-429 and displayed on the Primary Flight Display (PFD). The AN/ARN-147(V) VOR/ILS navigation system provides a continuous indication of magnetic bearing and slant range to a selected station. It incorporates three distinct receivers: VOR/Localizer (LOC), Glide Slope (GS), and Marker Beacon (MB). Each receiver operates independently and a failure in one will not affect the performance of the others. The VOR and LOC provide magnetic bearing and course deviation as well as audio identification information to the pilots. The GS receiver determines the deviation from the GS to a selected ILS station and displays it on the Primary Flight Display (PFD). The MB receiver picks audio signals from the MBs along the approach and transmits them to the pilots’ headsets.

VOR/ILS RECEIVER It operates in four separate modes: • • • •

VOR – 160 channels ranging from 108.00 to 117.95 MHz at 50 kHz intervals LOC – 40 channels ranging from 108.10 to 111.95 MHz at 50 kHz intervals GS – 40 channels ranging from 329.15 to 335.00 at 150 kHz intervals MB – fixed at 75.00 MHz

The GS receiver is activated with the corresponding frequency when the VOR/LOC receiver is tuned to a localizer frequency. The VOR/ILS navigation radio only provides a bearing when tuned to VOR frequency. There is no bearing for ILS frequencies. If an ILS frequency is tuned, the bearing data displayed on the PFD will reflect this condition.

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Antennas The VOR antennas (ANT) as well as the location on the UH-60M helicopter differ from the antennas on the UH-60A/L. The VOR antennas are now tubular-type antennas located further aft (one on each side) of the tail cone.

The FMS allows you to store 20 presets when tuning the VOR/ILS to a station. (There are no presets available for first flight.)

VOR/ILS ANT The MB and the GS ANTs and the locations of the ANT are the same as the ANT locations on the UH-60A/L.

Controls

FMS

FMS Navigation The AN/ARN-147(V) VOR/ILS navigation system is fully integrated with the FMS and is controlled by the FMS via the MIL-STD-1553B data bus. The VOR navigation can be accessed by a single keystroke “NAV” on the FMS. The screen appears with Radio Navigation, each displaying VOR/ILS/MB and ADF/ANT information. Pressing soft key (SK)-1/SK-5 brings you to “RNAV SETTINGS” page and will allow you to set VOR/ILS/MB navigation aids.

RADIO NAVIGATION PAGE

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BEARING POINTERS

RADIO NAVIGATION SETTINGS

FD/DCP The FD/DCP controls the navigation sources as well as the sources from the Automatic Flight Control System (AFCS)

The PFD detects the direction of rotation of the outer knob to adjust the No. 1 bearing pointer. It also detects the direction of rotation of the inner knob to adjust the No. 2 bearing pointer. The No. 1 or the No. 2 bearing pointer can be used as the VOR/ILS navigation source or as an FMS navigation source.

There are two knobs on the FD/DCP that control the VOR/ILS: •

NAV Source Select/Omni Bearing Select (OBS) Bearing 1 and Bearing 2 Pointer adjust.

•

The PFD detects the direction of rotation of the outer knob (Reference NAV Source Select) to select a navigation source from the following list: • • •

ILS VOR Back Course (BC)

The PFD synchronizes OBS to the current bearing to the selected source when push-tosync (OBS inner knob) is pressed.

OBS

REF SEL

BRG 1

BRG 2

REF ADJ

Indications Once the VOR/ILS is tuned to a frequency, the frequency will be displayed via Electronic Flight Instrument System (EFIS) junction box and ARINC-429 bus to the PFD. The PFD will indicate VOR or ILS depending on the tuned frequency. When the navigation source is selected on the FD/DCP, the information will be displayed on the PFD.

NAV/OBS SEL NAV SRC

BEARING POINTERS

LOC CAP

GS CAP

GA

DECL CAP

RALT 250

ALTP 2500

P-SYNC

P-SYNC

ALT 1500

IAS 120

HDG 240

P-SYNC

P-SYNC

VS 500

CPLD ****

HVR P-SYNC

UMAV001_5

BRG 1/BRG 2 FD/DCP NAVIGATION SOURCE SELECTED

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FMS Status and Test

TEST

The FMS provides status and test via MIL-STD1553B data bus from the VOR/ILS.

Pressing the TST button on the FMS and then pressing softkey (SK)-2 will access the test for the VOR/ILS. When the Test page appears, an overall status of the VOR/ILS will be displayed. Test for VOR/ILS can be selected from this page. An Initiated Built-In Test (IBIT) can be initiated for the VOR/ILS. Upon completion of the test, VOR/ILS IBIT screen will display the overall subsystem test results only. No fault data will be provided.

STATUS Pressing the STS button on the FMS and then pressing softkey (SK)-2 will access the status page for the VOR/ILS. When the Status page appears, an overall status of the VOR/ILS will be displayed. Status for VOR/ILS can be selected from this page. All faults and failures will be displayed under the VOR/ILS if a fault or failure should occur.

RADIO NAVIGATION TEST PAGE RADIO NAVIGATION STATUS PAGE

VOR/ILS TEST PAGE

VOR/ILS STATUS PAGE

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System Operation Operation and navigation control of the VOR/ILS is provided by the FMS via MIL-STD-1553B data bus. VOR/ILS data will be displayed on the PFD via VOR/ILS outputs. The VOR system is capable of operating from 108.00 to 117.95 MHz. Course information is presented by the deviation pointer and the No. 1 or No. 2 bearing pointer. ILS makes up the combination of the GS and LOC capabilities. The MB portion of the receiver visually indicates on the PFD and aural signals are heard over the pilot’s headset when passing over a transmitting MB. The desired type of operation is selected by tuning the receiver to the corresponding frequency of that operation. ILS operation is selected by tuning to the odd tenth MHz frequencies from 108.0 to 111.95 MHz. VOR operation is selected by tuning from 108.0 to 117.95 MHz. The three receiver sections perform the intended functions independently. Degraded performance within any one of the major sections will not affect performance of the others.

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COCKPIT NOSE TUNNEL

AN/ARN-147(V) VOR/ILS RECEIVER

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No.2 VOR/LOC Antenna 70600-05002-101

No.1 VOR/LOC Antenna 70600-05002-101

LOC1 RF

MB RF

AN/ARN-147(V) VOR/ILS Receiver

LOC2 RF

GS Antenna 70600-01043-101

RF

RF

28 28 VDC VDC

VOR/ILS 2 NO.2 DC ESNTL BUS

J1

R-2594/ARN-147(V) (CPN:622-6376-020) VOR/LOC AUDIO

Mount MT-6313/ARN (CPN: 622 -6380-001)

GS RF

Digital ICS ICU

Either FMS Provides Control and Display:

MB AUDIO

Localizer Phasing Coupler M55339/17-00274

• 20 Presets, Preset Names • Tuning 108.00 to 117.95 in steps of .05 • HI/LO Sense Select

Rotor Mod Select J4

REC VOR/ILS RF

J3 J2

RT Address 16 J5

Bus A

Bus A

Bus B

Bus B

ARINC-429: VOR Absolute Bearing LOC Dev GS Dev EFIS J-Box #2

EFIS J-Box #1

#1 MFD

#2 MFD

#3 MFD

#4 MFD

VOR/ILS Block Diagram

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NOTES

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Section 3-3.2 AN/ARN-149 (V) Low Frequency/Automatic Direction Finding (LF/ADF) System Description The AN/ARN-149(V) LF/ADF navigation system on the UH-60 A/L aircraft is being replaced with AN/ARN-149(V) LF/ADF navigation system on the UH-60M aircraft. An LF/ADF control panel located in the lower center console controlled the LF/ADF system on the UH-60A/L. Now the LF/ADF system is controlled by the Flight Management System (FMS) via a MIL-STD1553B data bus. The AN/ARN-149(V) LF/ADF navigation system is source selected on the Flight Director/Display Control Panel (FD/DCP) via ARINC-429 and displayed on the Primary Flight Display (PFD).

Components The LF/ADF system is comprised of the following components: • •

LF/ADF Receiver Antenna

LF/ADF Receiver The AN/ARN-149 (V) LF/ADF navigation system provides a continuous indication of magnetic bearing and slant range to a selected station. It can operate in conjunction with a suitably equipped aircraft, ship, or shore installation. The system operates in a frequency range of 100 to 2199.5 kHz, in 0.05 kHz increments.

ADF RECEIVER

Antenna The Antenna system is a single combination antenna containing both loop and sense elements. The signal from one loop element and the signal from the other loop element are modulated and combined, phase shifted 90°, and amplified. The resulting loop signal is summed with the sense antenna signal and sent to the ADF radio for visual and tone execution.

Controls FMS Navigation The AN/ARN-149(V) LF/ADF navigation set is fully integrated with the FMS and is controlled by the FMS via the MIL-STD-1553B data bus. The LF/ADF navigation can be accessed by a single keystroke “NAV” on the FMS. The screen appears with Radio Navigation, each displaying VOR/ILS/MB and ADF/ANT information. Pressing softkey (SK)-2/SK-5 brings you to “RNAV SETTINGS” page and will allow you to set LF/ADF navigation aids. The FMS allows you to store 20 presets when tuning the LF/ADF to a station. (There are no presets available at this time for first flight.)

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Indications The FD/DCP synchronizes Omni Bearing Select (OBS) to the current bearing to the selected source when push-to-sync (OBS) inner knob) is pressed. NAV/OBS SEL NAV SRC

OBS

REF SEL

BRG 1

BRG 2

REF ADJ

LOC CAP

GS CAP

GA

DECL CAP

RALT 250

ALTP 2500

P-SYNC

P-SYNC

ALT 1500

IAS 120

HDG 240

P-SYNC

P-SYNC

VS 500

CPLD ****

HVR P-SYNC

RADIO NAVIGATION

UMAV001_5

BRG 1/BRG 2 FD/DCP The FD/DCP detects the direction of rotation of the outer knob to adjust the No. 1 bearing pointer. It also detects the direction of rotation of the inner knob to adjust the No. 2 bearing pointer. The No. 1 or the No. 2 bearing pointer can be used as the LF/ADF Navigation source or as an FMS navigation source.

ADF SETTINGS

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FMS Status and Test

TEST

The FMS provides status and test indications via MIL-STD-1553B data bus from the LF/ADF.

Pressing the TST button on the FMS and then pressing SK-6 will access the test for the LF/ADF. When the Test page appears, an overall status of the LF/ADF will be displayed. Test for LF/ADF can be selected from this page. An Initiated Built-In Test (IBIT) can be initiated for the LF/ADF. Upon completion of the test, ADF IBIT screen will display the overall subsystem test results only. No fault data will be provided

STATUS Pressing the STS button on the FMS and then pressing softkey 6 will access the status page for the LF/ADF. When the Status page appears an overall status of the LF/ADF will be displayed. Status for LF/ADF can be selected from this page. All faults and failures will be displayed under the LF/ADF if a fault or failure should occur.

RADIO NAVIGATION TEST PAGE RADIO NAV STATUS PAGE

ADF TEST PAGE ADF STATUS PAGE

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System Operations The ADF system has two functional modes of operation: ANT and ADF. The ANT mode functions as an audio receiver providing only an audio output of the received signal. The ADF mode functions as both audio receiver and an ADF, providing a relative bearing-to-station signal to the No. 1 or No. 2 bearing pointer and an aural output. A TONE sub-mode of operation can be selected in either ANT or ADF mode, providing a 1000 Hz aural output to identify keyed carrier wave (CW) signals. Course deviations are displayed on the MFDs and audio tones are received in the headsets. The LF/ADF operating mode, frequency selection, and tone sub-modes are controlled from either FMS that interfaces with the LF/ADF through the 1553B data bus.

LF/ADF Antenna CPN: 622-6820-003

J1

AN/ARN-149 LF/ADF Receiver

RF PWR Modulating Discretes

RT Address 18

LF/ADF AUDIO

ADF

2

R-2382/ARN-149(V)1 (CPN: 622-6812-002) J1

Digital ICS ICU

28 VDC

Either FMS Provides Control and Display:

NO.1 DC PRI BUS

• 20 Presets, Preset Names • Tuning 100.0 to 2999.5 in steps of 0.5 • ADF or ANT Modes • NORM or TONE Modes

Quadrature Error Correction

Mount MT-6583/ARN-149(V)1

Configuration

(CPN: 622-7210-001)

J2

Bus A

Bus A

J3

Bus B

Bus B

ARINC-429: ADF Bearing

ARINC-429: ADF Bearing

EFIS J-Box #2

EFIS J-Box #1

#1 MFD

#2 MFD

LF/ADF System Block Diagram 10/2/2003 Page 4 of 6 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

#3 MFD

#4 MFD

AN/ARN-149(V) ADF RECEIVER

Cockpit Nose Avionics Compartment “Tunnel”

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NOTES

10/2/2003 Page 6 of 6 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Components

Section 3-3.3 AN/APN-209J Radar Altimeter (RAD ALT) System Description The Radar Altimeter (RADALT) system provides instantaneous indication of actual terrain clearance height. It measures and displays altitude above ground level to the pilot and copilot. The Radar Altitude is represented as a round dial analog indicator and a digital readout on the Primary Flight Displays (PFDs). The RADALT dial and pointer provide the analog indication of the radar altitude. The analog presentation provides information that is not as easily interpreted from the digital readout. The white radar altitude dial provides a display range of 0-1000 ft. The dial consists of three separate linear ranges: • • •

0 to 100 ft (10-ft increments) 100 ft to 400 ft (50-ft increments) 400 ft to 1000 ft (100-ft increments)

If the scale is currently displayed, it remains displayed until the radar altitude readout value rises above 1050 ft, at which point the scale is removed.

The RADALT system is comprised of the following components: • •

RADALT R/T – Right Hand (RH) Seatwell Two Antennas – Cockpit area bottom of the aircraft on both sides

The RADALT system is interfaced with the following system components: • • • •

Mission System Control Panel (MSCP) – On/Off Switch and SelfTest Avionics Relay Panel – Advanced Flight Control Computer (AFCC) Hover Hold AFCC – RADALT Hold Engaged Electronic Flight Instrument System (EFIS) Junction Boxes – Reliability signal to the AFCC

Receiver Transmitter (R/T) The AN/APN-209-J, located in the RH seatwell, looks and acts the same as the RADALT in the UH-60A/L aircraft. The analog adjustment screw on the R/T is used for adjusting the pointer and digital readouts on the indicator shown on the PFD. The digital adjustment screw is not used.

Antennas The RADALT antennas operate in the same manner as the UH-60A/L aircraft. They are also located in the same location as the UH60A/L. One antenna is used for the transmitter and the other one is used for the receiver.

Controls RADALT ANALOG DIAL

Mission System Control Panel (MSCP) The MSCP contains switch functions, which control mission systems as well as controls for navigation, altitude, and identification.

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The RADALT switch turns the RADALT system on and off as well as tests the system. The TEST position tests the system by sending a signal for altitude of 1000 ft. If the system is working properly, the corresponding altitude responds accordingly.

M S N S Y S

EGI 1

RAD ALT

ON

TEST

IFF

ON

N O R M

O N OFF

EGI 2

OFF

OFF

M4 HOLD

UMAV003_5

MSCP

Flight Director/Display Control Panel FD/DCP Both the “LO” and “HI” radar altitude values are set by either pilot from the FD/DCP using the reference set control on the panel to adjust the value up (clockwise) or down (counter clockwise) when the selected function for that knob is high or low radar altitude. The displays compute the altitudes (HI or LO) from the reference set knob turn information received and coordinates with the on-side PFDs to assure that there is a common setting.

The “LO” radar altitude setting/display provides a pilot selectable reference setting that defines the altitude at which the low radar altitude alert is activated when the aircraft descends. The low radar altitude alert is used as a monitor when flying an approach or when in a hover, to alert the pilot that the aircraft is passing through the threshold defined by the low radar altitude setting. The “HI” radar altitude symbology provides the pilot a high radar altitude reference that triggers an alert indication if the aircraft climbs through the high altitude setting. The high altitude is used as a monitor during certain phases of flight to alert the pilot that the aircraft is passing through the threshold defined by the high radar altitude setting. The digital Intercommunication System (ICS) provides a (one shot) “Low Altitude” tone warning to all ICS operator stations. It will not provide a tone warning for the “High Altitude”.

The pilot’s FD/DCP work in conjunction with the pilot’s inboard and outboard PFDs and the copilot’s FD/DCP work in conjunction with the copilot’s inboard and outboard PFDs. LOW RADALT Alert

REF SET KNOB NAV SRC

OBS

REF SEL

BRG 1

BRG 2

REF ADJ

LOC CAP

GS CAP

GA

DECL CAP

RALT 250

ALTP 2500

ALT 1500

P-SYNC

P-SYNC

IAS 120

HDG 240

P-SYNC

P-SYNC

VS 500

CPLD ****

HVR P-SYNC

UMAV001_5

REF ADJUST KNOB FD/DCP

Indications “LO” and “HI” Radar Altitude Symbology

HIGH RADALT Alert

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Radar Altimeter Receiver/Transmitter Right Hand Seat Well

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Status and Test STATUS Status for the RADALT Alert warning tone is available from the Digital ICS Status page via the Flight Management System (FMS).

altitude states based on a comparison of the altitude settings of their respective DCPs and the measured Radar Altimeter altitude. MFD 4 sends an ARINC-429 LO and HI message to FMS 2 for all four MFD LO and HI altitude states. MFD 1 sends an ARINC429 LO and HI message to FMS 1 for all four MFD LO and HI altitude states. Bus Controller (FMS 1 or FMS 2) performs Weight-On-Wheels (WOW) transition logic (e.g. Cold Start and WOW don’t assert LO or HI altitude.) Bus Controller (FMS 1 or FMS 2) sends the Digital ICS a MIL-STD-1553 LO and HI message to trigger the pilot or copilot LO tone warning. Then the digital ICS provides all the ICS stations a momentary (one shot) “Low Altitude” tone warning. The digital ICS does not provide a “High Altitude” tone warning.

The status page provides the pilots with an ALERT ON/OFF indication. It also provides the status of each control panel and will detect a fault if any of the control panels or if the ICS Control Unit (ICU) has failed. TEST Test for the RADALT is provided by the MSCP. When the momentary RADALT TEST switch on the MSCP is placed to the TEST and Hold position the RADALT inhibits a test. On the Multifunction Displays (MFDs) the RADALT pointer and digital readout indicates between 900 and 1100 ft (1000 ft ±100 ft). Once the test switch is released the pointer indicates between 23 ft and 35 ft (30 ft ±7ft) and the digital readout indicates between 25 ft and 35 ft (30 ft ±5 ft).

FD RADALT The Radar Altitude Hold function maintains the aircraft at a reference altitude above ground level (AGL). The radar Altitude Hold Mode is designed to work over water or very flat terrain. The Radar Altitude Hold Mode engagement is only done when the radar altimeter signal to the FD is valid and when it is within the normal operating range of the altimeter.

System Operation The MSCP ON/OFF switch controls power to the RADALT system. Once power is applied, the RADALT system provides all indications to the MFDs. The copilot’s DCP 1 sends a signal via ARINC-429 LO and HI altitude settings to MFD 1 and MFD 2 and the pilot’s DCP 2 sends a signal via ARINC429 LO and HI Altitude settings to MFD 3 and MFD 4. All MFDs set the “LO” and “HI”

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LH ANT RCV RF

RH ANT XMIT RF

RAD ALT CB PWR

28 VDC MSCP

APN-209J RAD ALT

Test/On/Off Aux Analog Alt

FCC #1

Rad Alt Test Inhibit (Note 1)

FCC #2

Rad Alt Test Inhibit (Note 1)

Reliability Aux Analog Alt

Avionics Relay Unit

Rad Alt Test Inhibit

Reliability Reliability Rmt Analog Alt

EFIS J-Box #1

Reliability Rmt Analog Alt Reliability Rmt Analog Alt MFD #1

Reliability Rmt Analog Alt

Suppression Pulse

EFIS J-Box #2

CSU

20 20 10 10

20 20 10 10

10 10 20 20

10 10 20 20

T

T

20 20 10 10

20 20 10 10

10 10 20 20

10 10 20 20

T

T

MFD #2

Reliability Rmt Analog Alt Reliability Rmt Analog Alt MFD #3

MFD #4

Radar Altimeter Block Diagram

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NOTES

10/2/2003 Page 6 of 6 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Section 3-3.4 Embedded Global Positioning/Inertial Navigation Unit (EGI) System Description

If either EGI fails, it will cause a loss of attitude and heading reference displayed on the on-side PFD.

The Embedded Global Positioning/Inertial Navigation Unit (EGI) is replacing the Directional Gyro, pilot and copilot Vertical Gyros, compass control panel, and compass flux valve to include the standard Doppler System of the UH-60 A/L. The primary function of the EGI is to compute attitude, heading, present position, velocity, and turn rate data. The EGI is a self-contained system consisting of an inertial navigation system and an embedded global positioning receiver. The EGI provides information for other aircraft systems to support navigation, time distribution, flight control, and primary flight displays. Navigation and attitude information provided includes position, acceleration, velocity, true and magnetic heading, digital attitude (roll and pitch), attitude rates, and time data. These signals provide navigational information to the pilots through the MFDs and provide reference to control attitude and heading of the aircraft through the Automatic Flight Control System (AFCS).

EGI Components EGI EGI 1 located in the forward nose deck, normally provides attitude and heading information to the copilot PFD and EGI 2 also located in the forward nose deck, normally provides attitude and heading information to the pilot PFD. In the event of both EGIs failing, it will cause a loss of attitude and heading reference to the AFCS resulting in degraded operation.

EGI

Reversionary Switch Panel Functions ATT REV (Attitude Reversionary) When the ATT REV is pressed to reversionary position, it displays attitude information from the other EGI. ATT1 or ATT2 is annunciated in yellow on the failed side and white on the operating side with a box indicating both sides are using the same source. HDG REV (Heading Reversionary) When the HDG REV is pressed to reversionary position, it displays heading and turn rate information from the other EGI. MAG1 or MAG2, or TRU1 or TRU2 is annunciated in yellow on the failed side and white on the operating side with a box indicating both sides are using the same source.

REVERSIONARY SWITCH PANEL

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EGI Controls Mission System Control Panel (MSCP) The MSCP contains switch functions, which control mission systems as well as controls for navigation, altitude and identification. There are two switches associated with the EGI system. EGI 1 turns the copilot’s EGI on and EGI 2 turns the pilot’s EGI on. The EGIs must be shut down before A/C power is shut down in order to perform an orderly shutdown mission data and a Built in Test (BIT) is saved in an orderly shutdown.

M S N S Y S

EGI 1

RAD ALT

ON

TEST

O N OFF

IFF

ON

N O R M

OFF

EGI 2

M4 HOLD

OFF

UMAV003_5

MSCP

Flight Management System (FMS) The FMS provides the ability to manually select the aircraft navigation solution and control of the navigation arbitration mode. All waypoint and steering calculations are done within the FMS.

UH-60M FMS

10/2/2003 Page 2 of 11 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

COCKPIT FWD NOSE

H-764 EGI 1 AND EGI 2

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Initialization Initialization of the EGIs via the FMS, provide entry and display of initialization position, system date and time, UTC date and time, and magnetic variation. It displays local datum and the key load status for each GPS receiver and provides access to the LOCAL DATUM, FILL PORT CONTROL, GPS KEYS and DTS screens. Each EGI independently initializes with system position, date, time, and alignment type based on the weight on wheel state and platform operation selection. The EGI alignment time and mode is displayed for each EGI. It provides control of platform operation, navigation arbitration mode and the automatic navigation mode selection for each EGI. The system initialization position is saved separately from the aircraft position and is displayed whenever the FMS screen is reentered. The date and time of the initialization position is used for RGI initialization with the exception of 1. System initialization position, date and time that are equivalent to the initial software load values and have not been manually entered or updated by the GPS. The Initialization is accessed via the FMS by pressing the fixed function key INI on the keypad.

INI 2 OF 2 PAGE There are modes that the EGIs transitions through via the FMS when EGI 1 or EGI 2 switch on the Mission Control Panel is placed in the on position and when the EGI alignment is in process. These modes are: “OFF” - EGI 1553 communication is not active, “ON” - EGI 1553 communication is active, “STANDBY” - EGI is in standby mode, “ORIENT” - EGI is in orient mode, “FAIL” - EGI is failed, “TEST” - EGI is in IBIT, “ATT” - EGI attitude is valid and not aligning, “AHRS” - EGI attitude is valid and heading is valid and not aligning, “ALIGN” - EGI is aligning, “D-NAV RDY” - EGI has reached degraded Nav ready performance and is fine aligning, “NAV RDY” - EGI alignment complete and in Nav Ready mode, “DEG NAV” - EGI is in degraded NAV mode, “NAV” which indicates that EGI1 is in NAV mode.

INI 1 OF 2 PAGE

10/2/2003 Page 4 of 11 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Indications

FMS Header Indications

All indications for the EGIs are available within the FMS then sent via the data bus system to the Primary Flight Displays (PFDs). The EGI system provides the following via the FMS:

The navigation header on the FMS displays EGI 1 and EGI 2 navigation information of the following solutions:

•

• • •

An aircraft position output selectable as - Pure GPS - Blended GPS/Inertial Solution - Pure Inertial Inertial Velocities (Vx/Vy/Vz and Vh/Vd/Vv EGI status and GPS RAIM condition Primary heading and attitude information

• • • •

IG1/IG2 – Indicates blended solution with GPS aiding I-1/-2 – Indicates blended solution without GPS aiding GP1/GP2 – Indicates pure GPS solution IN1/IN2 – Indicates pure Inertial solution

It also provides access to the following FMS screens: • • •

INS DATA GPS DATA NAV CONFIGURATION

The EGI system can be accessed by a single keystroke “EGI” on the FMS. The screen appears with Navigation, displaying EGI system information. By pressing any one of the softkeys will allow you to manually select settings and controls as well as being able to access INS, GPS data, and NAV configuration pages.

NAV HEADER DISPLAY The Navigation Mode display transitions to inverse video if the selected navigation mode is not usable for navigation. INS DATA The INS DATA provides Protection of Blended Solution (POBS) control for each EGI. Each EGI reports the inertial attitude and heading, heading reference, magnetic variation, wind range, wind bearing and Horizontal Uncertainty Limit (HUL) and displays it on the FMS as well as the PFDs. The time since last GPS update to the EGI blended solutions is also displayed. POBS

EGI NAVIGATION PAGE

The POBS provides a method to protect the integrity of the blended INS/GPS solution by preventing the malfunctioning of GPS satellites from corrupting the Blended navigation solution. The POBS detection occurs when one of the following is met: • •

GPS measurements exceed expected fault free performance. Fewer than five satellites are used in the GPS solution.

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POBS detection can be disabled to allow the GPS aiding to the Blended solution when it is desirable to process less then five satellites. With FMS power on, the POBS Control defaults to “ON” when the aircraft is on the ground and the last saved selection when the aircraft is in flight.

GPS Summary The GPS summary page provides satellite information for both GPS 1 and GPS 2 when selected. It also provides indication whether the satellites are used in the navigation solution. There are 12 channels available for use with the navigation solution. When a specific satellite is used for the navigation solution the CH line is displayed in inverse video.

INS DATA DISPLAY GPS SUMMARY SV PAGE

GPS DATA The GPS data provides Selective Availability (SA) mode control for each GPS receiver. It displays the receiver mode, solution state and solution performance of each GPS receiver and also provides access to the satellite summary screen for each GPS receiver. With power on the GPS SA mode defaults to “MIX” with the aircraft on the ground and the last saved selection is set.

SV – Satellite Vehicle Number ranges from 0 to 32. FQ – Satellite frequency carrier are between “L1” (Link One) or “L2” (Link Two). CODE – Psuedorange Codes are “C/A” Code, “P” Code, or “Y” Code. STATE – Channel State ranges from 1 to 5 and 7. C/N – Carrier to noise ratio of the GPS ranges from 01 to 63 dB Hz.

FMS Status And Test STATUS

GPS DATA PAGE

FMS can check the status of the EGIs. By selecting Fixed Function key STS on the FMS accesses the MAIN STATUS PAGE. Pressing (SK) 1 from the MAIN STATUS PAGE accesses the EGI ADC status page. From the EGI ADC Status page, pressing SK 1 or SK 6 accesses status for EGI 1 and EGI 2. This menu selection will allow a user to view the status of the EGIs. All faults and failures will be displayed under the

10/2/2003 Page 6 of 11 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

EGIs if a fault or failure should occur. These faults are as follows:

NO NAV DATA U GYRO V GYRO W GYRO R ACCEL S ACCEL T ACCEL POWER SUPPLY INERTIAL CALIB SYS PROCESSOR TEST

MIAN STATUS PAGE

The Test Menu is accessed by the function key and provides the ability to conduct and review tests of all systems and subsystems or line replaceable units (LRU’s) that can conduct an Initiated Built In Test (IBIT). By selecting soft key (SK) 1 from the MAIN TEST PAGE accesses EGI 1 test page. Selecting SK 6 accesses EGI 2 test page. This menu selection allows the operator to initiate an IBIT of the EGIs All faults and failures will be displayed under the EGI if a fault or failure should occur. .

EGI ADC STATUS PAGE

MAIN TEST PAGE

EGI STATUS PAGE BATTERY FAIL NO EGR DATA EGR STATUS FAIL EGR CHECKSUM INERTIAL REF

EGI 1 AND 2 TEST PAGE

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is the number the of the EGI to be tested 2. The test is not initiated until the EGI power is set to “ON”. Therefore, if the Screen is exited the test is terminated.

EGI TEST PAGE

•

The operator sets EGI-X power to “ON” 1. Once EGI-X power is set to “ON”, “TEST IN PROGRESS” shall be displayed in the Test Message Line 2. The EGI initiates the fault detection part of the test.

•

The EGI performs an alignment upon completion of the Fault detection. Upon completion of the alignment, The EGI reports any faults to the FMS.

•

System Operation

EGI Initialization

While an EGI is in test, at least one EGI is powered on and available for navigation. If SK-1 is pressed and the EGI that is not being tested is OFF, the test is inhibited and “PWR ON EGI 1 or EGI 2 ” is in the Test Message Line where Y is the number the of the EGI not to be tested. If SK-1 is pressed and the EGI not to be tested is in TST the test is inhibited and “EGI-Y IN TEST is displayed in the Test Message Line where Y is the number of the EGI not to be tested. There is a sequence of events that need to be accomplished in conducting the EGI test: •

•

The operator sets EGI-X power to “OFF”. 1. If SK-1 is pressed and the EGI to be tested is ON the test is inhibited and “PWR OFF EGI-X” is displayed in the Test Message Line where X is the number the of the EGI to be tested, and “READY FOR TEST” is displayed in the Test Message Line. The Test select key is pressed. 1. Upon selection of TEST, “PWR ON EGI-X” shall be displayed in the Test Message Line where X

INITIALIZATION 1 PAGE On power up the FMS displays the initialization page. The latitude and longitude values along with the Greenich Mean time (GMT) must then be entered and accepted. When placing the EGI 1 and EGI 2 switch to the on position on the Mission Control Panel (MSCP), within several seconds the FMS transitions through the following: • • •

OFF STANDBY ORIENT

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Within 45 seconds the following occurs: • •

On the FMS EGI 1 MODE transition to: ATT On the CPLT INBD and OTBD MFDs ATTITUDE Ball is shown with valid pitch and roll

ATTITUDE DISPLAY Monitoring the alignment in progress, with several seconds the following is displayed on the FMS: • •

EGI 1 or 2 MODE ALIGN INIT 1 or 2 is displayed without the arrow on the FMS

After approximately four minutes the FMS indicates: • •

EGI 1 or 2 MODE NAV The heading display and the three digit heading value are still present on the MFDs and the red HDG flag is no longer displayed.

INITIALIZATION 2 PAGE For first flight, PLATFORM (SK-4) is defaulted to LAND and NAV ARB MODE (SK-9) is defaulted to AUTO when weight-on-wheels (WOW). EGI 1 and 2 AUTO NAV is also defaulted to ON when power is applied.

Within the next two minutes the attitude display is still present and the red HDG flag is replaced by a three digit heading value.

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No. 1 MFD +5 VDC Supplied by EGI 1

No. 2 MFD

EGI 2 Tx CH4

EGI 1 Tx CH4

EGI 1 Tx CH4

No. 3 MFD

EGI 2 Tx CH4

EGI 2 Tx CH5

EGI 1 Tx CH5

No. 1 EFIS Junction Box EGI 2 Tx CH1

EGI 1 Tx CH4

Chelton GPS L1/L2 Active Antenna 2 Part No. 20-13K4-11-UH60M

No. 4 MFD

+5 VDC Supplied by EGI 2

EGI 2 Tx CH5

EGI 1 Tx CH5

No. 2 EFIS Junction Box

EGI 2 Tx CH4

EGI 1 Tx CH5

EGI 2 Tx CH5 EGI 1 Tx CH2

Load Access Panel

EGI 2 Tx CH2

EGI 1 Tx CH1

NO. 1 AFCC

No. 1 EGI+429 H-764GU

Crypto I/F

Honeywell Part No. 34209950CN13

ARC-220 HF-Radio PTTI - HAVE QUICK I/F

No. 1 ARC-201D SINCGARS Emergency Control Panel S1

Zeroize Switch

Bus A

5

No. 1 Pri. 28 Vdc

5

No. 1 Bat. Util.

EGI 1 On/Off

No. 2 EGI+429 H-764GU

No. 2 Pri. 28 Vdc 5 No. 2 Bat. Util. 5

S4

Bus A

Bus B

No. 1 ADC

No. 1 CDU

Bus A

No. 2 CDU

Bus B

Bus A

CDU Rx CH2

Bus B

Crypto I/F

Honeywell Part No. 34209950CN13

EGI 2 On/Off

Mission Systems Control Panel S1

CDU Rx CH2

NO. 2 AFCC

EGI 2 Tx CH6 (Spare)

EGI 1 Tx CH6 (Spare)

No. 2 ADC

NAV/COM Redundant MIL-STD-1553 Bus

EGI Block Diagram

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Load Access Panel ARC-164 UHF/AM

Bus B

PTTI - HAVE QUICK I/F

No. 2 ARC-201D SINCGARS Emergency Control Panel S1

Zeroize Switch

NOTES

10/2/2003 Page 11 of 11 UH60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

Section 4 End of Course Administration Course Critique It is requested that Sikorsky Aircraft Training Customer take a moment and let us know how we did. Please complete the course critique form provided and let us know what you thought of the training and training materials. If you felt that something could be added to make the data more relavent to your job performance, please include those observations as well. The Sikorsky UH-60M Program thanks you for your participation in the first UH-60M Training delivered.

3/9/2004 Page 1 of 2 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

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Section 4 End of Course Administration Course Critique It is requested that Sikorsky Aircraft Training Customer take a moment and let us know how we did. Please complete the course critique form provided and let us know what you thought of the training and training materials. If you felt that something could be added to make the data more relavent to your job performance, please include those observations as well. The Sikorsky UH-60M Program thanks you for your participation in the first UH-60M Training delivered.

3/9/2004 Page 1 of 2 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

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Glossary ACRONYMS/ABBREVIATIONS A ADC ADF AFCC AFCS AGL ALE ALT ALTP AME AMLCD AMS ANT APU AS ATC ATM AVC AVCC AVCS

AIR DATA COMPUTER AUTOMATIC DIRECTION FINDER AUTOMATIC (ADVANCED) FLIGHT CONTROL COMPUTER AUTOMATIC FLIGHT CONTROL SYSTEM ABOVE GROUND LEVEL AUTOMATIC LINK ESTABLISHMENT ALTITUDE ALTITUDE PRESELECT AMPLITUDE MODULATION EQUIVALENT ACTIVE MATRIX LIQUID CONTROL DISPLAY AVIONICS MANAGEMENT SYSTEM ANTENNA AUXILIARY POWER UNIT AIRSPEED HOLD AIR TRAFFIC CONTROL AIRCREW TRAINING MANUAL ACTIVE VIBRATION CONTROL ACTIVE VIBRATION CONTROL COMPUTER ACTIVE VIBRATION CONTROL SYSTEM B

BAR ALT BBC BC BIT

BAROMETRIC ALTITUDE BACKUP BUS CONTROLLER BUS CONTROLLER BACK COURSE BUILT-IN TEST C

°C CAA CALL CAS CAWS CBIT CEFS CERFS CDU CIP COM SUM CPU CSU CT CVR CW

CELSIUS CIVIL AVIATION AUTHORITIES CALL OVERIDE CREW ALERTING SYSTEM CENTRAL AURAL WARNING SYSTEM CONTINUOUS BUILT-IN TEST CRASHWORTHY EXTENDED RANGE FUEL SYSTEM CRASHWORTHY EXTENDED RANGE FUEL SYSTEM CENTRAL DISPLAY UNIT CONTROL DISPLAY UNIT COMPONENT IMPROVEMENT PROGRAM COMMUNICATION SUMMARY CENTRAL PROCESSING UNIT CENTRAL SUPPRESSION UNIT CIPHER TALK COCKPIT VOICE RECORDER CONTINUOUS WAVE

3/9/2004 Page 1 of 6 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

D DAFCS DCEL DCP DCU DDFCC DEC DECU DSAS DTS

DIGITAL AUTOMATIC FLIGHT CONTROL SYSTEM DECELERATION DISPLAY CONTROL PANEL DATA CONCENTRATOR UNIT DUAL DIGITAL FLIGHT CONTROL COMPUTER DIGITAL ENGINE CONTROL DIGITAL ENGINE CONTROL UNIT DIGITAL STABILITY AUGMENTATION SYSTEM DATA TRANSFER SYSTEM E

ECCM ECP EFIS EGI EICAS ELT EMC EMCON EME EMG EMI EMP EMV EP ESD ESIS ESSS EU EW

ELECTRONIC COUNTER-COUNTERMEASURES EMERGENCY CONTROL PANEL ELECTRONIC FLIGHT INSTRUMENT SYSTEM EMBEDDED GPS/INERTIAL NAVIGATOR ENGINE INSTRUMENT CAUTION/ADVISORY SYSTEM EMERGENCY LOCATOR TRANSMITTER ELECTROMAGNETIC COMPATIBILITY EMISSION CONTROL ELECTROMAGNETIC ENVIRONMENT EMERGENCY ELECTROMAGNETIC INTERFERENCE ELECTROMAGNETIC PULSE ELECTROMAGNETIC VULNERABILITY ELECTRONIC PROTECTION ELECTROSTATIC DISCHARGE ELECTRONIC STANDBY INSTRUMENT SYSTEM EXTERNAL STORES SUPPORT SYSTEM ELECTRONIC UNIT ELECTRONIC WARFARE F

°F FAA FCC FD FDS FFK FG FH FLIR FMS fpm FPS

FAHRENHEIT FEDERAL AVIATION ADMINISTRATION FLIGHT CONTROL COMPUTER FLIGHT DIRECTOR FLIGHT DISPLAY SYSTEM FIXED FUNCTION KEY FORCE GENERATOR FREQUENCY HOPPING FORWARD LOOKING INFRARED FLIGHT MANAGEMENT SYSTEM FEET PER MINUTE FLIGHT PATH STABILIZATION G

GA GATM GPS GS

GO AROUND GLOBAL AIR TRAFFIC MANAGEMENT GLOBAL POSITIONING SYSTEM GLIDE SLOPE

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H HDG HERF HERO HERP HF HIRSS HMU HIS HSM HUD HVQK/HQ HVR

HEADING HAZARDS OF ELECTROMAGNETIC RADIATION TO FUEL HAZARDS OF ELECTROMAGNETIC RADIATION TO ORDNANCE HAZARDS OF ELECTROMAGNETIC RADIATION TO PERSONNEL HIGH FREQUENCY HOVER INFRARED SUPPRESSOR SYSTEM HYDROMECHANICAL UNIT HORIZONTAL SITUATION INDICATOR HIGH-SPEED MACHINING HEADS-UP DISPLAY HAVEQUICK HOVER I

IAS IBIT ICAO ICS ICU IDM IFF IGB IHIRSS ILS I/O

INDICATED AIRSPEED INITIATED BUILT-IN TEST INTERNATIONAL CIVIL AVIATION ADMINISTRATION INTERCOMMUNICATION SYSTEM INTERFACE CONTROL UNIT IMPROVED DATA MODEM IDENTIFICATION FRIEND OR FOE INTERMEDIATE GEAR BOX IMPROVED HOVER INFRARED SUPPRESSOR SYSTEM INSTRUMENT LANDING SYSTEM INPUT/OUTPUT K

KIAS kts

KNOTS INDICATED AIRSPEED KNOTS L

LDS LED LF LH LNAV LOC LRU LSB

LASER DETECTION SET LIGHT EMITTING DIODE LOW FREQUENCY LEFT HAND LONG RANGE NAVIGATION LOCALIZER LINE REPLACEABLE UNIT LOWER SIDE BAND M

MB MEDEVAC MFD MGB MSCP MSL MT MTOP MTOT

MARKER BEACON MEDICAL EVACUATION MULTIFUNCTION DISPLAY MAIN GEAR BOX MISSION SYSTEM CONTROL PANEL MEAN SEA LEVEL MAINTENANCE TERMINAL MAIN TRANSMISSION OIL PRESSURE MAIN TRANSMISSION OIL TEMPERATURE

3/9/2004 Page 3 of 6 UH-60M AIRCRAFT 1 FIRST FLIGHT TRAINING DATA

MU MVOL

MECHANICAL UNIT MASTER VOLUME CONTROL N

NAV SRC NVG

NAVIGATION SOURCE NIGHT VISION GOGGLE O

O&S OAT OBS

OPERATIONAL AND SUPPORT OUTSIDE AIR TEMPERATURE OMNIBEARING SELECT P

PBIT PEA PFD PGC PSI PT PTF PTT PV

POWER-UP BUILT-IN TEST PLANNED ENROUTE ALTITUDE PRIMARY FLIGHT DISPLAY PLANNED GROUND CLEARANCE POUNDS PER SQUARE INCH PLAIN TALK PEAK TEMPERATURE FACTOR PUSH-TO-TALK PRIVATE R

RADALT RALT RAM RCU Rf/RF RH RMT R/T RT RWR

RADAR ALTIMETER RADAR ALTITUDE RELIABILITY, AVAILABILITY, AND MAINTAINABILITY REMOTE CONTROL UNIT RADIO FREQUENCY RIGHT HAND REMOTE RECEIVER TRANSMITTER REMOTE TERMINAL RADAR WARNING RECEIVER S

SARSAT SAS SBIT SDC SINCGARS SK SRU SSR

SEARCH AND RESCUE SATELLITE-AIDED TRACKING STABILITY AUGMENTATION SYSTEM START-UP BUILT-IN TEST SIGNAL DATA CONCENTRATOR SINGLE CHANNEL GROUND AND AIRBORNE RADIO SYSTEM SOFT KEY SHOP REPLACEABLE UNIT SECONDARY SURVEILLANCE RADAR T

TACAN TBC TGB TGT

TACTICAL AIR NAVIGATION THERMAL BARRIER COATING TAIL GEAR BOX TURBINE GAS TEMPERATURE

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T/R

TRANSMITTER/RECEIVER U

UHF USB

ULTRA HIGH FREQUENCY UPPER SIDE BAND V

VCAS VHLD VIDS VLEA VNAV Vne VOR VOX VS VSI

VIBRATION CONTROL ACTUATION SYSTEM VELOCITY HOLD VERTICAL INSTRUMENT DISPLAY SYSTEM VARIABLE LOAD ENERGY ABSORBER VERTICAL NAVIGATION VELOCITY NOT TO EXCEED VHF OMNI-DIRECTIONAL RANGE VOICE OPERATED VERTICAL SPEED VERTICAL SPEED INDICATOR W

WOW

WEIGHT-ON-WHEELS

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