The performance of MIL-F-5624A fuels, grades JP-3 and JP-4, was investigated in an axial-flow turbojet engine over a range of altitude conditions of 10,000 to 55,000 feet. Examination of the fuel flow, combustion efficiency, and net thrust specific fuel consumption showed the grade JP-4 fuel to be slightly inferior to the grade JP-3 fuel, although the altitude ignition limits were essentially equal for the two fuels over a range of flight Mach numbers and fuel-inlet temperatures

Braithwaite, Willis M

Renas, Paul E

English

NASA Technical Reports Server

1 
c 
RESEARCH MEMORANDUM 
COMPARISON O F  TURBOJET-ENGINE  ALTITUDE  PERFORMANCE 
CHARACTERISTICS  AND  IGNITION LIMITS WITH 
MIL-F-5624A FUEL, GRADES JP-3 AND JP-4 
B y  W i l l i s  M I  Braithwaite  and Paul E. Renas 
Lewis  Flight Propulsion  Laboratory 
Cleveland, Ohio 
NATIONAL ADVISORY COMMITTEE 
FOR AERONAUTICS 
WASHINGTON 
February 27, 1952 
https://ntrs.nasa.gov/search.jsp?R=19930094391 2020-06-16T22:19:49+00:00Z
. J!XATIORXG ADvIsoKy C O " l T T Z X  IWR AERONAUTICS 
MIL-F-5624A FUEL, G X D B  Jp-3 AmD Jp-4 
By Willis M. Braithwaite and Paul E. Rems 
The altitude performance and ignitian Units of  axial-flaw 
turbojet  engine -re evaluated in an a l t i t ude  test chamber with 
MIL-F-5624A, grade Jp-3 fuel  ( that  specif ied f o r  this engine) and a 
low-volatility grade JT-4 fue l .  The investigation was conducted over 
a range of a l t i t udes  From 10,000 t o  55,000 feet. 
Use of the Jp-4 fuel resulted in 821 increase in spec i f ic  fuel con- 
aumption of 2 t o  5 percent over that obtained with the Jp-3 fuel. This 
increased specifLC fuel consumption resulted from 8 combination of 
reduced combustion efficiency and the lmr heat of c d u s t i o n  of the 
JP-4 fuel. Alti tude ignit ion limits w e r e  found t o  be essent ia l ly  equal 
for the two fuels mer a rasge of flight Mach nuniber and fuel-supply 
temperature. Inspection of the ccanbuetors af%er 6 hours operation with 
JP-4 fue l  revealed no noticeable carbon deposition. 
The need for a f u e l  having charac te r i s t ics   su i tab le   for  je t  air- 
c ra f t  and that could be produced in large quant i t ies  led t o  the devel- 
opment of MIL-Y-5624AJ grade Jp-3 fue l .  It was found, however, that 
dm- rapid climbs t o  high al t i tudes  rapid boilFng of this fuel 
occurred, thus resulting in large losses of fuel,  both as vazor and as 
entrained liquid (f-). Such fuei losses seriously decreased the 
range o r  enduranoe of the aircraft. 
These losses may be reduced by uaing a fuel of lower vo la t i l f t y .  
Frevious investigations (references 1 and 2) ind ica t ed  t h a t  a reduction 
Fn v o l a t i l i t y  of the  JP-3 fuel f r o m  a Reid vapor pressure of about 
6 pounds per square inch t o  1 pound per square inch resulted in no 
loes in altitude performmoe with re spec t   t o  thrust, fuel consumption, 
land a l t i t ude  bluw-out, but decreased altitude i w t i o n  limits. O n  the 
2 __. NACa RM E5lL05 
basis  of these and other data, a comprcanise fuel speoification, 
pressure t o  2 t o  3 pounds per square inch. It was therefore desirable 
t o  ompare  the performance of this  new speoification fuel wlth 
MIL-F-562411, grade 3p-3 in-a full-scale turbojet  engine developed on 
Jp-3 fuel. 
MIL-F-562411, @ d e  Jp-4, ws issued which limits the Reid ~ a p m  
Such a cmparisan of the a l t l t ude  performame character is t ics  and 
al t i tude  igni t ion  l imite  of a current axial-flow engine using Jp-3 and 
JP-4 f'uele was obtahed in an investigation conducted in an a l t i t ude  
test chamber at the mACA Lewis laboratory. &g€ne perfo3laanm data  
were obtalned a t  simulated altitudes of 10,000, 40,000, and 55,000 f e e t  
a t  a simulated flight Maoh-number of 0.6 for both Suels. Altitude 
igni t ion limits for   both f'uels were obtained a t  flight Mach numbers 
of 0.4, 0.6, and 0.8 with fuel-supply temgeratures between 45O and 80° F. 
A t  a flight Mach number of 0.6, a l t i tude   ign i t ion  limits were  also 
obtained for f'uel-eupply temperatures of about -35O and Oo F.. 
8 
nJ 
P 
5 
Engine 
L 
A modern axial-flow turbojet engine was used for  this investigation. 
Liners designated a8 a smokeless type were installed in the cambustors. 
The ignition system used had a 15,000-volt, 400-cycle output t o  the plugs 
in  two diametrically opposite cambustors. 
Altitude T e s t  Chamber 
The engine w a s  i n s t a l l e d   i n  an altitu&e test chamber 10 f e e t i n  
diameter and 60 feet in  length, as sham in  f igure  1. The front  of 
the engine &ended through the f ront  bulkhead by means of a labyrinth 
seal. This bulkhead separated the inlet and exhaust sections of the 
chamber. The inlet section was connected to the laboratmy canbustion- 
a i r  supply and the engine exhausted in to  a diffuser elbow tha t  w a s  
connected t o  the laboratory exhaust system. A rear  bulkhead was 
located near the exhaust nozzle of the engine t o  prevent recirculation 
of exhaust gas around the engine. The altitude-chamber f u e l  system 
included 8 cooler t o  provide a range of fuel-supply temperatures f rm 
about -40 t o  80' F. 
NACIII RM E5lL05 3 
Wtrumentatian 
Radial survey rakes at  several  circumferential  positions were 
installed at  the W e t  and outlet of each ccanponent of the engine t o  
measure total  temperature and t o t a l  and s ta t ic  pressure. The air 
flow was computed from the temperatures and preseures measured at the 
engine inlet, and the  jet  thrust was cmputed frm this a i r  flaw and 
s ta t ic  a d  t o t a l  pressures measured at  the eawt-nozzle inlet .  The 
fuel flow was measured dlrectly by rotameters in the  fuel-supply line 
and the engine speed, by remote indicathg tachameters. 
P 
The fuels used in  this  investigation were MIL-F-S624A, grades JP-3 
and Jp-4. The specification and an analysis for each of the fuels are 
presented in table I. !The JP-3 fuel is the fuel presently specified 
f o r  the engine used in thie investigation. The JP-4 fuel is me of the 
lower quality  fuels permitted under this specification in that i t e  Reid 
vapor pressure was 2.1 pounds per square inch, and the aromatics con- 
tent was essentially the maximum allowed. p e  final boiling polnt of 
the fuel (561O F) slightly exceeded the 550 F allowed by the specifica- 
tion. The high final boiling point  is believed Wiepificant.  
Perfomnance data m e  obtained at altitudes of 10,000,  40,000, 
and 55,000 feet  f o r  a flight Mach number of 0.6. At each flight 
condition, the engine-*let temperature and pressure were set  at  
values correspanding t o  the stagnation pressures and temperatures In 
flight, based on MCA standard atmospheric conditions. The e a u s t  
section of the chamber was maintained at   the s t a t i c  pressure corres- 
ponding t o  the particular altitude being simulsted. The engine speed 
was varied from rated speed or the maximum engine speed limlted by a 
tall-pipe gas temperature of 1300° F t o  a n  engine speed where the  tai l-  
pipe temperature m e  as luw aa 300'to 500° F. 
The ignition limits were obtained a t   f l igh t  Usch nunibere of 0.4, 
0.6, and 0.8 with the fuel-supply temperature between 450 and 800 F. 
A t  a flight Mach number of 0.6, imit ion limits were also obtained with 
fuel-supply temgeratures of about -35' and Oo F. Because fuel charac- 
ter is t ics  have o n l y ,  a secondary effect cn propagation of flame from 
one combustor t o  another and on engine acceleration, only the  iepition 
phase of an altitude skr t  was investigated. For the ignition studies, 
the engine-fuel manifold was modified t o  supply fuel only t o  the t w o  
combustors containing spark plugs, which reduoed the amount of fuel 
that would accumulate in the engine during an ignition attempt. If the 
combustors containbg the spark plugs i s i t e d ,  propagation t o  a l l  con- 
b u s t ~ ? ~ I  was assumed possible. 
The alt i tude  ignit icm limit was d e t e m e d  by slmulatin@; a partio- 
u l w  a l t i t u d e  and flight speed in  the  a l t i tude  chamber. When the engine 
ro tor  speed stabil ized, the igni t ion circui t  was energized and the 
t h r o t t l e  advanced. If ignition did not occur, the t h r o t t l e  WEW closed 
and slowly advanced once more. This process of repeatedly varying the 
fuel flow was continued until ignition occurred or a time limit of 8 
45 seconds had elapsed. Upon completion of two successive starts, the z 
a l t i t ude  was increased 2500 f e e t  and the procedure repeated. men an 
a l t i t ude  was reached where ignit ion could not be obtained after several 
attempts of 45-second duration, the alt i tude wa8 lmm t o  that at 
which igni t ion has prevdously occurred. If ignition w&s again obtained, 
this a l t i t ude  was, by definit ion,  the a l t i t ude   i@t ian  limit for  the 
engine. 
The ef fec ts  of different fuel types on turbojet-e-e performanoe- 
will be apparent on variables including fuel oaneumption, canbustar- 
outlet-temperature profile, and engine s tabi l i ty  character is t ics .  
Iaasmuch as the dissimilarity between the fuels is not p a t ,  the 
combustor-outlet--temperature-profile effect- was not considered. "he 
' engine exhibited stable operation over the range of engine speed& for 
the  flight conditions investigated. Consequently, fuel flow, ocdms- 
t ion  efficiency, and spec i f i c   fue l  consumption are the only performance 
variables discussed. The data  obtained in this hve f l t i ga t im  a ~ e  pre- 
sented in table  II. 
c 
Performance Variables 
The cambustion eff ic iencies  obtained with the JP-3 and JP-4 fue ls  
at- a eimuLated flight Mach number of 0.6 and altitudes of 10,000, 
40,000, and 55,000 f e e t  are presented In figure 2. The combustian . 
effioiency presented herein is defined as the r a t i o  of the enthalpy 
rise &CZ-OEB the combustor to   the  enthalpy &vaiLable frm cmplete com- 
bustion of the fuel (reference 3 ) .  The JP-3 fuel gave consistently 
higher  effioiencies at all three alt i tudes  than  did the JF-4 fuel. 
A t  peak efficiencies, this difference did not exceed 2 percent; a t  
off-peak conditions, the largest. differenae in combustion eff'ioiency 
was about 4 percent. 
The lmr  cambustion efficiency with the JP-4 fuel would require 
8 higher  fuel flow with t h i s  fuel than xith JP-3 fuel;  in addition,  the 
5 
n 
c 
. 
lower heating value of the JP-4 fuel is less than that of the JP-3, 
requfring a st i l l  higher fuel flow. The observed fuel flow is shown 
in figure 3. At. an altitude of 10,000 feet, the increase in corrected 
fuel flow of the JP-4 fuel over that of the JP-3 fuel was slightly less 
than the &crease fndioated for the higher altitudes. A t  a correoted 
engine speed of 7500 rpm (approximste speed f o r  peak ccanbustion efficiency) 
and altitudes of 40,000 and 55,000 feet,  the corrected fuel flow was 
about 3 percent s e a t e r  f o r  the grade JP-4 fuel than for the JP-3 fuel. 
The Increased fuel f l o w  required when JF-4 fuel was ueed resulted 
In a correspondingly higher corrected net tbrmt specific  fuel con- 
sumption, as aham in  figure 4. The corrected net thrust specific fuel 
consumptima with JT-4 fuel mre  2 t o  5 percent higher than with the 
JT-3 fuel. These data inaimbe no significant trend d t h  increashg 
al%itude . 
Altitude Ignitfan 
The effects of flight Mach number and fuel temperature on altitude 
ignition limits for the two fuels are ccangsred in figure 5. On the 
basis of previous investigations (for example, reference 2),  the more 
volatile Jp-3 fuel might be expected t o  allar ignition at  slightly 
higher altitudes than the less volatile JP-4 fuel. Nevertheless, as 
shown i n  figure 5(a), the  altftude  ignition limits of the two fuels 
differed by lese than the  altitude increment of 2500 feet ueed in 
determining the ignition limits. The altitude ignition limits pre- 
sented in figure 5(b) for a range of fue1-inle-b temperatures at a 
flight Mach number of 0.6 were alBo essentially equal for the two 
fuels. A slight decrease in altitude i a t i o n  limit occurred for 
both fuels as the fuel-inlet temperature was reduced from approXhBtely 
60° t o  -35O F. 
!&e increased aramatic,content and decreased vola%ility of the 
JP-4 fuel would be expected t o  increase the carbon depositim. Beam- 
ination of one of the ccanbustor liners revealed no notioeable carbon 
formation after 6 hours operation with thfs fuel; however, most of 
t h i s  operation was at high-altitude conditions x b r e  carbon deposition 
is minimized. 
A comparison of the performance of MIL-F-5624A grade JP-3 and a 
low-volatility MIL-F-5624A grade Jp-4 f u e l  In a  turbojet engine indi- 
cated that the canbustion efficiency with the JP-4 fuel was 2 t o  4 per- 
cent lower than with the JP-3 fuel a t  each of the  three  altitudes 
6 - NACA IN E5lLO5 
investigated, 10,000, 40,000, and 55,000 feet. AB a result of this 
lower combustion efficiency with the JP-4  fusl, together with approx- 
imately a. l-percent lower heating value, the  engine net thrust apecFfic 
fuel caneumption was 2 t o  5 peroent higher with this fuel than with the 
JP-3 fuel. 
The a l t i tude-  igni t ion limits -re essentially equal for the two 
fuels over a range of flight Mach numberer fram 0.4 t o  0.8. Sfmilarly, 
the a l t i tude   ign i t ion  limits of  the two fuels -re eesent ia l ly  eqwl 
over a range of fuel-inlet temperatures f r a m  approximately 60' t o  -35' F. 2 2 
Af'ter 6 hours operation with the JP-4 fuel, which haa a re l a t ive ly  
high aromatic content, exmination of one coldbustor revealed no 
noticeable oarbon formstion. 
Lewis Flight Propulsion Laboratorg 
Cleveland, Ohio 
National Advisory Camittee for  Aernnsutios 
1. Wilsted,-H. D., and Amnstrong, J. C.: Comparison of P e r f o m c e  of 
AN-F-58 and AN-F-32 Fuels in 533-A-23 Turbojet ErtgFne. 
NACA FM E m 4 ,  m 9 .  
2. Wilsted, E. D., and Armfltrang, J. C. : EPfect of Fuel Vola t i l i ty  an 
Altitude Starting Umits of a Turbojet Engine. rrmclA FM E50G10, 
1950. 
. 
3. Turner, L. Richard,  and Lord, Albert M.: Thermdpxmic C h a r t s  f o r  
the Corngutation of Conibulstion and Mixture Temperatures at Canstant 
Pressure. HACA TN 1086, 1946. 
NACA RM E5lLO5 7 
c 
A.S.T.M. d i e t i l l a t i o n ,  
(9) 
Initial boiling po in t  
Percentage  evaporated 
5 
10 
20 
30 
40 
50 
60 
70 
80 
90 
95 
Final boiling point 
Residue, (peroent) 
LOSS, (percent) 
Etreez- point,  9 
Aramaticer, (percent by 
volume) 
A.S.T.M. D-875-46T 
Silica gel 
Bromine number 
Reid vapor pressure, 
Bydrogen-carbon r a t i o  
( W S ¶  in.) 
H08t O f  odUEtiOll ,  
(B.t;u/lb ) 
Wavity,  ( OAPI) 
A i r -  jet residue, 
(mg/100 d) 
Sulfur, (peroant by 
-%w) 
hil ine PO&*, 9 
Viscos i ty  at 70° F, 
(cen t idokee )  
"" 1 117 
"" 
"" 
156 
205 
178 
226 
" 246 
"" 267 
"" 292 
"" 
363 -1-1 
322 
452 
1.5 =x.) 1.0 
1.5 ma3.) 1.0 
-76 (-0) Belm -76 
"" 
"" 
400 (Inin.) 
600 (=.I 
415 
481 
"" 
1 
10 
9 
0.5 
5 t o  7 
0.172 "" 
5.8 
57.6 45 t o  63 
18,680  18,400 (min.) 
"" 1.0 
20.0  5.0 
0.4 (max.) 0.1 
"" 122 .o 
"" 1.0 
I " 
"F-5624A 
Jp-4 
3pecif ioat ions 
"" 
"" 
250 (=.I 
"" 
"" 
"" 
"" 
"" 
"" 
"" 
"" 
"" 
550 (msr.) 
1.5 (ms1.1 
' 1.5 (mx.) 
-76 (=a) 
25.0 (u. ) 
30.0 (a.) 
2 t o  3 
"" 
18,400 (e. ) 
40 t o  58 
"" 
20.0 (Inax.) 
0.4 (max.) 
"" 
"" 
148 
188 
218 
255 
288 
519 
349 
378 
409 
441 
475 
499 
561 
1.1 
1.0 
b l a ~  -76 
"" 
25 
8.0 
2.1 
0.160 
18,500 
.46.9 
ll 
15 
0.1 
114.1 
1.1 
9 
Iiun A l t i -  
- 
1 
2 
3 
4 
5 
6 
7 
8 
9 
10 
ll 
12 
13 
14 
15 - 
- 
16 
17 
19 
20 
21 
22 
23 
24 
25 
26 
27 
28 
29 
30 
l a  
- 
Maoh 
Duniber 
10,000 0.601E 
.611E 
.590'i 
.595? 
.596'i 
40,000 0.6152 
.6252 
.593€ 
.6052 
.618E 
55,000 0.720€ 
.6286 
.6336 
. 6 4 4 7  
.6204 
1 1 I 
Grade JP-3 
7643 
6829 
6067 
5563 
4932 
8000 
7586 
6827 
6072 
5562 
8023 
7531 
6848 
6421 , 7201 
3.5126 
5569 .5800 
6072 .5€M 6453 .5896 
7078 .6038 
7389 
D. 6103 7998 
.62l2 7592 
,6196 7064 
.6228 6457 
.6l59 5562 
3.6059 7587 
.6344 7062 
.6196 6827 
.a55 6072 
.6173 6450 
1458 
1457 
1447 
1455 
1462 
387.4 
386.3 
395.7 
389 04 
385.4 
179.6 
189.2 
186.5 
187.3 
188.0 
1862 
520 1860 
52 2 1849 
52 0 1863 
520 1845 
519 
500.1 417 
502.8 417 
502.3 418 
498.7 416 
498.9 417 
253.8 
428 246.9 
426 
420 243.7 
422 247.7 
424 244.4 
. .  . . . .  
e a d e  JP-4 
1454 
1855 1477 
1855 1472 
1846 1459 
1864 1457 
1859 
387.8 
501.1 386.3 
498.7 
245.2 185.3 
243.3 187.8 
245.6 187.3 
242.7 189.4 
500.9 387.8 
502.8  387.1 
500.3 386.2 
187.9 243 -0 
521 
523 
526 
524 
525 
414 
413 
414 
416 
418 
422 
422 
413 
425 
421 
3399 
2 627 
2 l l 5  
1882 
1725 
lll3 
1062 
913.2 
699.2 
574.7 
592.4 
525.2 
458.2 
402.9 
493.8 
1671, 
43.48 1029 
53.06  1096 
64'33 1343 
75.98 
957 35.66 
1749 24.06 
1604 23.84 
1363 22.33 
1089 18.79 
957 15.92 
1876 11.51 
1715 11.29 
1510 10.53 
1295 9.52 
1555 11.21 
3133 
2828 
2303 
2114 
1887 
1127 
1069 
971.2 
811.5 
583.9 
525.8 
479.1 
449.8 
337.6 
389.5 
1554 
1425 
1202 
1105 
102 6 
1766 
1608 
143 7 
1230 
961 
1720 
1554 
1474 
1204 
1301 
72.72 
67.71 
57.40 
52.30 
43.39 
23.89 
23.88 
23.16 
20.92 
15.81 
11.23 
10.94 
10.61- 
8.76 
9.68 
2E WACA RM E5lLo5 
c 
9 
I i_ Net  Thruet spe- Cor- Cor- Cor- Corrected thrust oifio fuel reoted rected rected thrust C ~ U S -  tion effic- iency sir flow (k) 
Graae JP-3 
7643 86.34 5059 1.345 0.979 1 
6822 73.86 3005 1.436 .974 2 
6061 60.33 1680 1.880 -970 3 
5546 49.91 1222 2.926 .971 4 
4927 40.61 864 10.209 .957 5 
8925 91.27 8017 1.477 0.944 6 
8463 89.94 6878 1.407 0953 7 
7608 84.41 5009 1.355 .958 8 
6782 71.37 2986 1.489 .949 9 
6205 60.52 2124 2.045 0898 10 
8855 86.97 8782 1.618 0.857 11 
8293 87.85 7530 1.524 0906 12 
7576 82.41 5728 1.457 -936 13 
7121 73.32 4272 1.542 0888 14 
8005 87.54 6513 1.436 -924 15 
Grade n-4 
4452 
2623 
1481 
1071 
760 
1698 
1465 
1067 
630 
4A9 
954 
798 
598 
451 
675 
33ll 
1824 
787 
365 
74 
1282 
1162 
8 78 
473 
245 
651 
576 
454 
324 
522 
1.345 
1.437 
1.882 
2.935 
10.219 
1.324 
1.262 
1.216 
1.333 
1.833 
I. 466 
1.384 
1.317 
I. 391 
1.292 
1.380 
1.424 
1.643 
1.938 
3.133 
1.366 
1.297 
1.248 
1.272 
1.872 
1.412 
I. 350 
1.329 
1.568 
1.395 
- 
16 
17 
18 
19 
20 
2 1  
22 
23 
24 
25 
26 
27 
28 
29 
30 - 
3868 
3108 
1939 
1517 
llo1 
1774 
1525 
1238 
869 
472 
826 
669 
591 
367 
459 
2803 
2183 
1181 
782 
352 
12 98 
1176 
991 
683 
252 
585 
495 
445 
234 
239 
7374 
7050 
6415 
6042 
5536 
8955 
8510 
7 910 
7212 
6198 
8414 
7832 
7653 
6710 
7161 
82.94 
77.17 
66.19 
59.95 
49.79 
90 -52 
89.96 
87.50 
78.83 
59.94 
88.28 
84.97 
82.31 
68.37 
75.95 
4394 1.377 
3514 1.418 
2210 1.633 
1721 1.928 
1249 3.114 
8426 1.530 
7219 1.454 
5863 1.398 
4085 1.421 
2222 2.086 
7886 1.565 
6390 1.497 
5761 1.489 
3500 1.732 
4439 1.548 
0.968 
.911 
.967 
956 .944 
0.922 . 935 . 995 
,945 
861 
0.891 
.923 
.939 
.so1 
.so2 
R 1 
. ,  I I. . . ., 
I ' ": ! 
.. . . . 
3E 
c 
11 
n 
. 
100 
90 
V 
(a) Altitude, 10,ooO feet .  
f 
d 
(b) Altitude, 40,000 fee t .  
4000 5000 6000 7000 moo moo0 10,Ooo 
Corrected engine speed, rpm 
(c) A l t i t u d e ,  55,000 feet .  
Figure 2. - Compariean of combustion efficiencies far MIL-F-562411, grades JP-3 
and n-4 fuels in full-ecale engine. Flight Msoh nuniber, 0.6. 
12 
" 
h 
4E 
F 
NACA RM E5lID5 - 13 
2.0 
1.6 
1.2 
2.4 
2.0 
1.6 
1.2 
(b) Altitude, 40,000 feet .  - 
2.0 
1.6 n - 
-Fs7 1.2 
4000 5000 6000 7000 8000 9000 
Corrected englne speed, rpm 
(c) u t i tude ,  55,000 feet. 
Figure 4. - Comparison of corrected net thrust specific fuel consumptions for 
MLF-5624A, sades 3p-3 and a - 4  fuels in full-scale engine. Flight Mach 
nuniber, 0.6. 
1 4  NACA RM E5lLO5 
40,000 0 Bo ignition 
g 30,000 
. 
a 
2 
R 
5 
4 20,000 
10,000 
0 .2 .4 .6 .8 1.0 
Mach- wmber 
(a) Effect of flight Mach number. Fuel-SuPPly . 
temperatures, 45' to 80' F. 
40, OOO 
a: 30,000 
$ 
* -P 
5 
4 20,000 
10,000 
-40 -20 0 20 40 60 80 
Fuel-eupply  temperature, OP 
(b) Wfeot of' fuel-eupply  temperature. Flight  Maoh number, 0.6. 
c 
SECURITY I N F O R M A T I O N  
.I 
i 


Comparison of turbojet-engine altitude performance characteristics and ignition limits with MIL-F-5624A fuel, grades JP-3 and JP-4

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