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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 _