Study of Small Turbofan Engines Applicable To General-Aviation Aircraft [PDF]

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NASA CR 114630 AVAILABLF TO THE PUBLIC AIRESEARCH 73-210'148

A STUDYOF SMALLTURBOFANENGINES APPLICABLE TO GENERAL-AViATION AIRCRAFT FINAL REPORT by G, L. Merrill,

G. A. Burnett,

C.C.

SEPTEMBER

Prepared AIRESEARCH

under Contract

E);VISION

(JF

Sky Harbor

No.

NAS2-6799

by

COMPANY

THE

GAbriEl7

Airport,

PHOENIX,

et al.

1973

MANUFACTURING A

Alsworth,

f_r.]_?prJRATi0

OF

ARIZONA

N

402 South 36th Street ,_RIZONA

85034

for Ames

Research Center

NATIONAL AERONAUTICS ANDSPACEADMINISTRATION System=

Study

Division

r

CO:_'fEI_TS

IN-RODUCTION

._./" ,.. .... _' _

SY_'_OLS

3 9

PHASE

I

-

PRELIMINARY

Preliminary

ENGI['_E

Aircraft

Airplane

Engine

cycle

Cost

and

Reduction cost

Materials

and

Turbine

STUDIES

13

Analyses

sizin_ an u_

.c; o

W

c;

64

F

1.0

0.10

o.e

o.o8

I

I

0.4

0.04

10 HP EXTRACTED 600_

I

;

I

I

J

1200

a

2000

,

0

,

5O

i

•

100 VELOCITY,

I

Figur_

lS0

2OO

250

30O

TAS - KNOTS

19. Full-Throttle Takeoff Candidate Engine ZA.

and

Climb

Performance,

65

_

0.10 .......

1 tl _"

_.,.

oo. i

0.8 I

z !

0.6

0.4

-

0.__,

7315 M (24,000 FT) ISA 22 N/MIN (5 LB/MIN) BLEED 10 HP EXTRACTED

350(

70_30O(

I

250¢

== F.=

h-

350 KNOTS

200C

; -m

150( 0

0.2

0.4

0,6

0.8

MACH NUMBER

Figure

G6

20.

Full.-Throttle Performance Altitude, Candidate Engine

at Cruise IA.

.0

F 4

9"

"A_

¸

I

1.00

0.100 I

I

7315 M (24,000 FT) 350 KNOTS TAS 22 N/MIN (5 LB/MIN) 10 HP EXTRACTED

0.95

0.095

m

n"r! Z (.9

.J

.J

ISA

0.90

v

--

I

BLEED

I

0.090

(J U.

U. ¢n

p-

I--

L

_°

0.85

0.0,35 [

0.80 -

0.080 1000

ii _

1200 | 250

1400 THRUST

I 300 THRUST

Figure

21.

1600 - N

I 350 -

1800

2000

,) 400

LB

Part-Throttle Performance at Flight Conditions, Candidate

Design-Point Engine IA.

67

I

0.10 i

1.0

!

o.s

°

_, o.oe

I

•

_

OI

305°K (90OF) 610 1524MM(2000 (5000 FT), FT)--__

0.6

oo_ 0.4

_-

\

_._,

STANDARD

DAY

0.04 NO BLEED 10 HP EXTRACTED 6000

1200

-

m 1000 .

ee 3:

_.,

800

_ANDARD

DAY

X

__

30OO 1524 M (5000 FT)_J 306OK (90OF)

-"_

2000

I 0

Figure

6| .I

22.

K (90°F)

5O

100

160

VELOCITY,

TAS - KNOTS

Full-Throttle Takeoff Candidate Engine IC.

and

200

Climb

250

Performance,

3OO

i

10 lip EXTRACTED

350O

I

-==\ I- 2600

\

I-

380 KNOTS

2OOO

16000

Figure

23.

Full-Throttle Candidate

0A 0.6 MACH NUMBER

Engine

Performance XC.

at

Cruise

03

P

1,0

Altitude,

6_

t .

•

,

.

......

1.0

• ......

°

0.100-.-----...Ilmmmlmmmmm

0.95

"

0.091 7316 M (24,000 FT) ISA 360 KNOTS TAS 22 N/MIN (S LB/MIN)BLEED

=, _0.90

0.09q

!

0.85

0.081

0.80 -

0.080 1000

10 HP E_(TRACTED I

\

1200 l 250

1400 THRUST I 300 THRUST

1600 - N

I 360

1800

2O0O

J 400

- LB

t

Figure

70

24.

Part-Throttle Point Flight Engine IC.

Performanue Conditions,

at DesignCandidate

f r

i

1.0

0.10_

0,S

0.0:_

.

610 M (2000

-."

FT)

-__s24M(r,ooo_')==.__ __'"_

_""'_

0.4

_,'_-

•

,.._ SEA LEVEL

_

STANDARD

_'_

I

i DAY

NO BLEED 10 HP EXTRACTED

1200 f"-305OK (90OF) FT) 610 M (2000

w,

..SEA

1000 ;_

_'

I.-

•

L_

_

LEVEL

STANDARD

DAY

800

p

6OO

•

2000

b

1

o

_o

'too VELOCITY,

F£gu=e

25,

Full-Th_o_tle Candidate

Engine

Tak_of£ IZA.

leo

_

TAS -

and

,

:so

j

_Oo

KNOT_

Climb

_ ,

Per_ormanoe_

71

t

|

T 1

i

1.0

0.10

o.s

0.08

I i

T

/

/

I --

0.6

- __0.o6/j 0.4 -

0.04

7315 M (24,000 FT) ISA 22 N/IWIN (5 LB/MIN) BLEED 10 HP EXTRACTED

7OO

?

sod

I •':.

360 KNOTS

2OOO I

4OO

,m

1600

".

0

0.2

0.4

0.6

0.8

MACH N_'MBER

_ .

FAgu_e

i_¸ ,

, __

?2

26.

Full-ThrottZe CandAdate

Performance EngAne ZZA.

at

Cx'uAse

AltAtude,

1.0

I &.

0.90

m

0.85

0.090

-

_0.80

-

u_

7315 M (24,000 FT) ISA 350 KNOTS TA& 22 N/MIN (5 LB/MIN) BLEED 10 HP EXTRACTED

0.085

-

_ 0.080 U.

0.75 -

0.075

0.71

°°'°_L

;#

\ i

,_

i

1400

1600

THRUST

THRUST

1800

2000

- N

- LB i

I

Figure

27.

Part-Throttle Point Fli_ht Engine IIA.

Performance Conditions,

at DesignCandidate

.°

?3

r

1.0

-

0.10

C

0.08

0.6 F-

.

LL.

i_.o o._ c._

-___

305°K (90OF)

--STANDARD SEA LEVEL DAY

__ 0.4

0.04

r

t

eooo

NO BLEED 10 HP EXTRACTED

!I

5000 qk

__. 3000

.___._.__

_____._

000

11_4

_

_

M (6000

FT)._.

0 306°K

_;;_

(2000 FT) (90°F)

_

_

"J 200

260

! 396OK (90OF) i

2(_0

b

0 'o

'

60

'

I

100

'

VELOCITY,

TAS - KNOTS

f

)igure

t

28. Full-ThrottleTakeoff Candidate Engine

_r

?4

and IlC.

Climb

Performance,

#

1,0

=, -r,

-

Q

0.10

o.s - _ o.oe r3

,,,J

I

I

0.06

0.6 7315 M (24,000 FT) ISA 22 N/MIN (5 LB/MIN) BLEED 10 HP EXTRACTED 500

--

700000

m.

60O

E zI--

50O .I-

_,

350 KNOTS 2OO0 4OO

16000

0.2

0.4

0.6

0.8

1J)

MACH NUMBER

Figure

29. Fu11-Throt_le Performance Candidate Engine IIC.

at

Cruise

Altltude,

7S

t

f

r

0.90

0,090

-

7315 M (24,000 FT) ISA 0.85 0.085

22 N/MIN

(5 LB/MIN)

BLEED-_

350 KNOTS TAS 10 HP EXTRACTED

=/-.

0.80

-

_

0.080

•

.

I (J

u.

LL

C4 I--

0.75

-

0.075

0.70

0.071 IOO0

lb..

1200

1400

1600

1800

t

THRUST I 260 ,•.

,.

Figure

30.

I 300 THRUST

- N

I 35O - LB

Part-Throttle Performance at Flight Conditions, Candidate

I 400

Design-Point Engine IZC.

2000

r,',

e

1j .ii!

,_•

o.8 o.o8

•

_ I-

,

1524 M (5000 FT)-----_ 305OK (90OF)

_

\

,

o.,,

-

_

0.4

__

0.04

_

ALEVEL STANDARD

l

DAY

NO BLEED 10 HP EXTRACTED

I

7OOO

1400

1200

=, I

1ooo == I8OO

3OO(

308OK (00OF) 20O0 0

50

100 VELOCITY,

150

200

260

300

TAS - KNOTS

ruZ1-Throttle Takeoff and Candidate Engine IIC/gBPR.

Climb

Performance,

77

t

1.0

-

0.10

=. o.o I 0.6

o.

f

' _ o.oe

0.4

J

0.04

0.02 7315 M (24,000 FT) ISA 22 N/MIN (5 LB/MIN) BLEED 10 HP EXTRACTED 4OOO 8OO b"

_,,oo

350 KNOTS

_J 00

mtm

L

1000 0

0.2

!).4

0.6

0.8

1.0

MACH NUMBER

Figure

?0

32.

Full-Throttle Performance Candidate Engine IIc/gBPR.

at

Cruise

Altitude,

j+

r

7"

2:

0.90

0.090

+

7315 M (24,000 FT) ISA 360 KNOTS TAS - 22 N/MIN (5 LB/MIN) BLEED 10 HP EXTRACTED

0.85 .J _C 3: !:

m .4

0.085 0.80

I

v I

0.080

\

0 LL

_r

0.75

0.075

0.70

0.070

-

!

IOO0

1200

1400

1600

THRUST

l 250

THRUST

Figure

- N

l, 300

1800

I 350

J 400

- LB

Part-Throttle Performance Point Flight Conditions, Engine IIC/9BPR.

at DeeignCandidate

20O0

D)IISN 'ISrlI:IH •

:)_ls/9"l/g'l

.,

JL:Ol =ll03dS I

"JSlll:lH

I

I :DI:IIO3dS

l

8"I/I:IH/81

'0:IS1

II0

mmmmm

r I,

Engine

IC Core Other

En@ine :#

compressor components

- single-stage centrifugal (modi fled) - same descriptions and as IA.

- TPE331 derivations

IIA Core compressor - 5-stage axial - ATF3 Core turbine - single-stage axial - TSE36-10 Other components - same descriptions and derivations as IA

z

Engine

IIC Core

- single-stage centrifugal - TSE36-10 (modified) Core turbine - single-stage axial - TSE36-10 Other components - same descriptions and derivations as IA Engine

compressor

IIC/9BPR Fan - single-stage axial - TFE731-2 (modified) Core compressor - single-stage centrifugal - TSE36-10 (modified) Core turbine - single-stage axial - TSE36-10 Other components - same descriptions and derivations as IA

The derived components were modified substantially in mechanical design. However, care was taken to avoid affecting aerodynamic performance. The candidate engine components are considered to be conventional and reflect the current state-of-the-art in their projected performance. Acoustic anal_sis results and attenuation treatmene. - The results of the acoustic'analysis of the five candidate engines are presented in Table VIII. As indicated, the 85-PNdB goal could not be achieved with full attenuating treatment of the inlet and both walls of the bypass duct. Substantial redesign would be r&quired in all five engines to meet the 85-PNdE goal. Figure 35 was prepared to illustrate the severity of the 152-m (500-foot)/85-PNdB goal versus present and proposed regulations at a 0.25-nauticalmile measuring point. The five candidate engine noise levels in thls figure are slngle, unattenuated, bare-englne predictions, with no allowance made for airplane shielding, flight speed effects, or tone correction. Because these points fall substantially below the proposed 1980 regulation, the engin_ redesign required to reach the 85-PNdB goal was not attempted.

;

$I

j

_k

t

i

TABLE

VIII.

ACOUSTIC ENGINES ENGINES].

ANALYSIS RESULTS [152 M (500 FT)

Unattenuated Noise (PNdB)

k k

OF FIVE SIDELINE

Bypass Duct Treatment for 95 PNdB

|

Engine

Forward

RearWard

Weight kg (ib)

Length om (in)

CANDIDATE NOISE - TWO

Noise with Full Inlet and Bypass Duct Treatment

(PNdB)

IA

94.5

97.0

3.86 (8.5)

30.5 (12)

IC

94.5

99.0

7.26

(16.0)

61.0

(24)

90.0

95.5

98.0

5.22

(11.5)

43.2

(17

89.0

94.5

97.5

5.67

(12.5)

40.6

(16)

88.5

95.0

99.0

5.44

(12.0)

30.5

(12)

87.5

IIA IIC

iIC/ 9BPR

%

In order to meet _:he 95-PNdB during takeoff, only the rearward suppressed. This can be achieved bypass ducts. metal) honeycomb tion requirements the attenuators honeycomb IIC, and

12

value at propagation by lining

500

feet sideline noise needs to be the outer wall of the

Low-cost attenuators, using per£orated metal (holeabsorbers, were designed to meet the noise reducfor each engine. The w_Jghts and lengths of are given in Table VIII. The thicknesses of the

absorbers are 9.5 mm 30.5 mm (1.2 in.) for

(0.375 engine

in.) for IIC/gBPR.

engines

IA

through

t

r

d_

s

g_'Nd3 "13A3"1 351ON O:ZAlj3tl3d

::lAIJ.O:l:ld3

83

!

t.

Engine

Configuration

and

Weight

Analysis

Enuine and component confi urations.T _ ... i, . _ he basic layouts of the zlve canalaaue engines and their dimensions are presented An Fi_,res 36 through 40. The layouts were prepared in accordance with the engine design principles established in Phase I. The flow paths represent the aerodynamic components and the cycles determined in the candidate engine definitions conducted in Phase II. The configurations shown are supported by a considerable amount of preliminary mechanical design and stress analysis. The layout_ also represent th,_ materials and manufacturing methods that we_e final

4

selected as basal.lines for this study. The layouts sufficiently comprehensive co provide a basis for and cost analyses of the five candidate engines.

/

IIC

One fan engines.

was designed eud sized This fan has slightly

engine, because of each engine cycle.

\

are judged to credible weioht

for use on the different match

the small difference in The nominal 100-percent

IA, IC, points

fan airflow speed for

be

IIA, and for each

requi-_•ed by this 1.48-

pressure ratio fan is 17,000 rpm. A separate fan design and size, producing 1.3 pressure ratio at ii,ii0 rpm, was required for the IIC/9BPR engine. Both fan baseline designs used Ti6AI-4V titanium blades, with part-span d_mpers, and aspect ratios of approximately 2.5. The fan bla_es are attached to high-strength, low-alloy steel hubs and have integral stub shafts. On all engines, the fa_, outlet starer was divided between the inner and outer flow paths. The inner starer was located close behind the bypass splitter. The outer starer was positioned about 51 mm '2 inches) downstream of the fan, in an effort to achieve a low-noise design. Both starer vane rows are integral with a steel weldment, attached to the engine walls,

front frame, and include and th_ nose section of

the the

inner bypass

and outer splitter.

fan

flow-path

The front frame of all engines contains the thrust bearings for all rotors, the radial accessory drive shaft, and, _n the IIC/9BPR engine, the fan-drlvlng reduction gear. In all engines, the frame is a shell-mold aluminum casting with six integral struts. The accessory gearbox is mounted on the bottom of the frame, with the radial drive shaft, and with the front bearing eavit F ell drain passing through the bottom strut, into the qearbox. The five candidate pressor configurations. O

engines

differ

primarily

in

their

core

com-

Engine IA has a four-stage axial c_pressor, providing a 4.27 pressure ratio at 37,280 rpm. Each starer stage im an integral-wheel-type investment _ast!ng of 17-4PH steel alloy. The individual stages, _g,_ther with front and rear shafts, are Joined by electron-beam welds. The cantilevered, strip-stock stainless steel tstator vanes

._+

........

_

+'

+.

•

+I

'I

.

',

•

i + _

I +

+..

•

+++'_ '

\

++

,,

",

+ +++.

: ....

~-,

I

•!+

,+

• .I + , ,++r"_'.+' +_

• .+o

_+.,r:._.

!

+

++

+. +-

• -+-!

•

I ,f

+v o!

,

!

BASIC

DIAMETER

LENGTH

I_ROH

53.34 INLET

CM

FLANGE

(21.0 TO

_

IN.)

PLANE

OR

PRI_,RY

JET

I I

'IqZMARY

JET

NOZZLE

127.76

CM

(50.3

IN.

)

rA_re

36.

(:and Ldate IngLne ConfLg_,tLon XA.

es/ee

_

"

''"

"

"

_"

,'i,

f

\

\ /

v

ID_.IIZC D'_ MIIII_ o

_

riwnl

/'.1_1

62.23 ZUI,I_

r.JI (24.S

/'Z,_IWII

'_

ZB.) PEJUIII 0/'

I_NlkltY

,:rrL,

......J_

° *

112,77

QI

(47,0 Zm,)

07/80

f

I

.!

\

_IZC

DT.ANBTB.q_ 53.59

r.,BJlQTH -I'1101( ZNL,IT

CH

FI, ANGI

(21.1 70

IN.) P_ib'E

OP PRZI4_Ry

_'E7

1

5

! ! .__.a

I_MAR¥ J2T

NOZZLE

127.5

Ot

(50,2

IN.) FJgt_e

38.

Cand_dat;e IngJ.ne Con_lgu=stion XXA.

f

.+ 4

I >

l"'


a

propulsion-system performance of sized to match at the end of

specified

takeoff

module provides data relative to the the engine. The propulsion-system is the cruise drag and a rate of climb reclimb. Program options are included , distance

can

be

met

and

the

climb

re-

quirements of FAR Part 25 satisfied. Engine diameter and weight were calculated as a function of engine front-face design-point Mach number, the hub-tip diameter ratio, and the specific weight.

of

The engine

engine length

nacelle dimensions to diameter. The

this study was 2.45. This ratio existing turbofan installations. nacelle drag in terms of engine for the study.

The

unscaled

from

the ratio

ratios for

was established from a survey of The program option to account for net-thrust decrement was selected

generalized-engine-performance-data

program was provided in terms fuel flow, and airflow. The either corrected rotor speed inlet temperature. Resizing basis of specific ery was accounted

were determined length-to-diameter

input

to

the

of corrected values of net thrust, engine power setting was uhe ratio or turbine exhaust to the compressor of the engine was completed on the

thrust and airflow. for directly.

Inlet

total

pressure

of

recov-

A weight and balance analysis was completed on the airplane after the configuration geometry was defined and the engine size and weight were calculated. Weights for the various airplane subsystems were estimated from trend equations based on the correlation included in Reference i0 adjusted for the general-aviation class of airplane weights,

airplane. weight, and the

Available fuel which was computed input gross weight

was determined by summarizing and p_yload.

specifying tip tanks is included in the program. required, the synthesis program will automatically aerodynamics, engine size, and airplane structural plane weight and balance summary outputs provided are shown in Figures 50 and 51.

The

airplane

mission

module

provides

from the empty the subsystem An option for If tip tanks are recompute the weight. Airby the synthesis

computations

of

the

air-

plane performance during taxi, takeoff, climb, cruise, and landing. Options are available in this module for calculating engine-out and accelerate/stop distance, best rate of climb, best cruise liftto-drag ratio, and additional airplane and engine operating characteristics. The effects of gear and flap retraction and ground effect are acco_unted for during the takeoff segment. In the climb segment, speed is restricted to 463 km/hr (250 knots) equivalent airspeed or less at altitudes below 3048 m (I0,000 ft). Fuel reserve inputs accounted for range definition iterates within

a

are acaounted for in the cruise segment. during the climb and cruise segment. When is required, a program option Is utilized

on the specific

airplane gross tolerance of

weight until the required

the calculated range.

Pe.nge is a specific that range

is

J,19

:

F

VOIVE • ULT, LF

333e KTS • 5,70

PROPULSION

VMO • MAN, LF

•

283, KTS 3.80

MMO • GUST LF

(WEP) (WPEI) IWFSS) (WPROP) IWP)

632, 209, 36, O. 875,

STRUCTURES GROUP WING HOR. TAIL VERTe TAIL FUSELAGE LANO|NG GEAR PRIMARy ENG, SECTION TIP TANKS GROUP wEIGHT INCe TOTAL STRUCoGROUP WTe

(WW) (WHT) (WVT) (WB) (WLG) (WPES) (WTIP) (OELWSTI (wSY)

655, 830 66, 666, 279. 170, AS, O, 17630

FLIGHT CONTROLS GRQP COCPIT CONTROLS FIXED WING CONTROLS SAS GROUP WEIGHT INCe TOTAL CONTROL WTe

(WCC) (WCFw) (WSAS) (OELWFC) (WFC)

21, 63. O. O, 83.

WTe OF FIXEO

(dFE)

FIXED

EQUIPMENT

EMPTY USEFUL

OPERATING

661o

IWE) LOAO

WEIGHT

33630

(WFUL)

EMPTY

PAYLOAO

S370

lOwED

GROSS wEIGHT

Figure

50.

600o

IWFA)

19310

(WG)

6230,

Summary

(INC,

CREW OF 1)

3900,

(WPL)

FUEL

120

,TS3 2e61

GROiJP

PRIMARY ENGINES PRIMARY ENGINE INSTLe FUEL SYSTEM PROPULSOH WEIGHT TOTAL PROPtGROUP WTe

WEIGHT

•

Output

for

(PAX•

S,)

(WFW•

7940)

Weights.

(WFTPm

6660)

)

F

%

WING LOCATION INFO, FUSELAGE LENGTH WING 1/4C LOCoON CoLe• NAC I/4C LOCATION MAC OISToFRON CoLe WING C,GoLOCATION TiP TANKS CoGeLOCAT[ AIRCRAFT

CeGo

Figure

32013 13099 13089 6047 160_S 13099

• • • •

LOCATION

¢eOoLOCATION OF CoOoOF REHAINING

The operating obtained

•

:

13,81

PNOPULSIONo

WEIGHT

51.

H-TAIL H-TAIL HoTAIL H-TAIL V'TAIL V'TAIL

•

Summary

FT,

OR

V_L, ARM CoG,LOCATION MAC FROM CoLe LOCAT ON VERTe• VOLe ARM ¢*GeLOCATION

0230

by

estimating

Output

equipment costs, and the sales, and manufacturer's

the

16,02 30,21 30S6 025

a •

17001 31020

MAC

EOoBgJ 100601

for

Airplane

economics module calculates airplane cost for a converged airplane design from weights, size, engine, mission,

detemnined

OF

• • •

labor

hours,

appropriate and dealer's

Balance.

first based etc.

material

markups profit.

for

cost and total on the results First cost is costs,

purchased

overhead,

tooling

Total operating cost is broken into variable and fixed costs. Variable costs consist of fuel and oil cost, inspection and maintenance cost, and reserve for overhaul cost. Variable costs are calculated in terms of dollars per hour of operation. Fixed cost is independent of annual utilization rate and is computed on an annual basis. This includes storage, insurance, depreciation, crew salary, and taxes.

The operating

variable costs

per year. ating cost Figures 52

and fixed for annual

Summaries breakdown and 53.

of as

costs were utilization

the airplane determined

in

combined to determine rates from I00 to 800 first the

cost and synthesis

the are

total shown

total hours operin

121

ode ENGINES EMPTY

COST

DATA

NUMBER WETGHTx

CONSUMER

•

2t

3363,

PRICE•

eee TYPEo

7

LRS

306158.

MAX,CRUISE

OOL.

BASIC ADD,

01RECT LAROR ( 40BG,NHRS,) LABOR OVERHEADC|40,PCT) A|RFRAME MATERIALS PURCHASED EQUIP, (ENGINE• 45289e) (PROPe • 0e) (OTHER • _4640e)

153620, 52180, 205799e 29707, 235506e 70652e 306158.

PROFIT(|AePCT)

DEALER-OIST,

Figure

MARKUP(30ePCT)

52.

RANGE•

9Yge

HeN,

FUEL

T8,3

GPH,

Summary

BLOCK RATE•

TOO•

Output

FUEL• 2000e

UTILIZA_|ONINRS/YR) TOTAL OPBeCOSTIDOL/MR)

Figure

122

53.

O,

DOLe

SUB-TOTAL MANUFACTURING DEALER BASIC

|OOe 484,17

PRICE

Basic

Airplane

ISS3e

LBS

BLOCK

ZOO• 2B3,Ig

Output

COST

COST

HOURS/INSPex

FIXED COST STORAGE INSURANCE DEPRECIATION OTHER CREW FAR TAX

TOTAL

Summary

KNOTS

DOLe

COSTs

for

HRSe

VARIABLE COST (DOL/HR) FUEL,OIL 37.05 INSR,,MA|N, 26.20 OVERHAUL RES, 18,95 OTHER O.OG

AZ,_O

3061See

EQUIPMENT

349e

13893, 195130 4995, IIS21B,

ENGtTL_SALES,G-A(3&ePCT) FACTORy

PRICE•

SPEEDo

of

300e _16e19

Total

Cost.

TIMEx IO0,

2,963

HRSe

HRSe

(DOL/YR) 30000 63380 (HULL 2,OPCT) 306160 (8oYRo2OoPCT) O* Oe(OVERHEAO SOePCT) 263e 60197, TOTAL 400, 182,69

Operating

SOO, 162,60

Cost.

800, 131,6S

PHASli

II]

-

LV_Lt, A'i'].ON O}' CAbH)IDATE _