Braithwaite, Willis M

Sivo, Joseph N

English

NASA Technical Reports Server

RESEARCH MEMORANDUM 
EVALUATION O F  ALTITUDE-IGNITION  CHARACTERISTICS 
OF THREE FUELS O F  DIFFERENT  VQLATILITY 
IN A TURBOJET ENGINE 
By Wi l l i s  M. Braithwaite and Joseph N. Sivo 
Lewis  Flight  Propulsion  Laboratory 
C leveland, Ohio 
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NACA RM E 5 3 U  
EVALUATION OF ALTITUDE-IGNITION  CHARACTEBISTICS 
OF TEFEE FUELS OF D- VOLATILITY 
By Willis M. Braithwaite and Joseph N. Sivo 
An investigation  was  conducted on a full-scale  turbojet  engine  in- 
stalled  in an PSACA Lewis laboratory  altitude  chamber  to  obtain  the 
altitude-ignition  characteristics  of  three  fuels with different A.S.T.M. 
distillation  curves. Two of the  fuels  had  Reid  vapor  pressure of 2.7 
pounds per  square inch and  the  third  fuel  had a Reid  vapor  pressure of
fuel  having a Reid  vapor  pressure  of 1.7 had a much lmer distillation 
curve  than  either  of  the  others.  The  altitude-ignition  and flame- 
propagation  characteristics of the  fiels  were  carpared  at an ltitude 
of 50,000 feet and a flight Mach number of 0.8. Also, the ignition and 
propagation  lFmits  of m e  of  the  fuels  were determined aver a range of 
altitudes  for  the  particular  engine,  ignition  system,  and  cmbustor-Inlet 
conditions  of  the  investigation. 
c 1.7 pounds  per  square inch. Hmever, at  temperatures  above lOo F, the 
. 
The  relative  quantities  of  fuel  evaporated  at  the-cmbustor-inlet 
sktic pressure  and  temgerature,  as  indicated by the A.S.T.M. distilla- 
tion  curves,  were  found  to  give a better  indication of the  effect  of 
fuel  volatility on altitude  ignition than did  Reid  vapor  pressure. A s  
noted  in  several  previous  investigations,  the  fuel-air  ratio  required 
for ignition  and  propagation  decreased  with  increasing  volatility.  The 
minimum time for both  ignition  and  progagation was obtained with the 
same fuel-air-ratio  setting  over  the  range  of  altitudes  investigated. 
INTRODUCTION 
The use. of turbojet  engines on multiengine  Ugh-altitude  aircraft 
requires  that  these  engines  be  capable  of  starting  at  altitudes of 
50,000 feet.  The  altitude-ignition  limits  depend o  the  energy  per 
mark and the  spark  repetition  rate  produced  by  the igni im system 
+ 
Y (ref 6. 1 and 2), the  over-all  fuel-air  ratio  of  the  mixture  to  be 
2 NACA RM E53Lll 
ignited, and the vo la t i l i t y  of the fuel  (refs. 3 and 4 ) .  R e i &  vapor 
pressure has been used as an indication of fue l   vo la t i l i ty ,   bu t   the   fas t  
r a t e  of climb obtained with current airplanes requires that the Reid 
vapor pressure be l imited  to  a low value t o  prevent loss of fue l  from 
the airplane fuel tanks during climb (ref. 5) .  In an effort  to improve 
the  altitude-ignition  characteristics of fuels  with low Reid vapor pres- 
sures ,  new blends of fue l  have been developed that  have fewer heavy 
components and, thus, a higher over-al l   volat i l i ty  without an appreci- 
able change i n  R e i d  vapor pressure. Using these fuels requires a dif- 
ferent method  of indicating  the  volati l i ty of a fue l  as it affects  ig- K 
r 
nition. The 15-percent  point on the A.S.T.M. d ie t i l l a t i on  curve  has i; 
been used fo r  this purpose (ref .  4 ) .  
" 
I 
In the present investigation, the dietillation curves obtained at 
the couibustor operating pressure were consfdered as an indication of 
the fuel  volat i l i ty .  The investigation was conducted on a full-scale 
turbojet engine in s t a l l ed   i n  an al t i tude chamber a t  the NACA Lewis lab- 
oratory  to  compare the  altitude-ignition and flame-propagation charac- 
t e r i s t i c s  of three fuels having d i f fe ren t   d i s t i l l a t ion  curves; two of 
the fuels  had the same Reid vapor pressure. This investigation w a s  con- 
ducted at a simulated al t i tude of 50,000 fee t  and a f l i g h t  Mach number 
of 0.8. I n  addition, ignition and propagation limits, as defined by 
the  fuel-air-ratio ranges i n  which ignition and propagation could be 
obtained, were determined fo r  one of the  fuels  over a range of a l t i tudes 
a t  a flight Mach number of 0.8. 
d 
APPARATUS 
Engine Instal la t ion 
A J47-GE-.27turbojet engine was in s t a l l ed   i n  a 10-foot a l t i tude 
t e s t  chamber fo r  use in this investigation. This engine has a 12-stage 
axial-flow compressor, 8 tubular conibustors, a single-stage turbine, 
and a nonafterburning ta i l  pipe. The fue l  system was the standard 
equipment supplied with the engine. 
Two comercially available capacitor-discharge ignition systems 
were used during the investigation. Each system f i r ed  two fixed-gap 
spark plugs, one i n  each of  two tubular cambustors 1800 apart. 'Both 
ignit ion systems had spark rates of 1 t o  4 sparks per second and spark 
energies which were 20 t o  30 percent of the stored energy. The igni- 
t ion  system used for   the   fue l  comparison had a stored energy per spark 
of 2.7 joules. The system used for  determining the altitude-ignition 
and propagation characteristics had a stored energy per spark of 1.35 
joules. 
NACA RM E53Lll 3 
Instrumentation - 
The data used f o r  the windmilling-performance and air-flow calcu- 
lations of this investigation were obtained with the instrumentation . l is ted  in   the  fol lowing  table:  
Engine i n l e t  
Compressor 
out le t  
Combustor 
Turbine 
ou t l e t  
Total- 
temperatur 
probes 
12 
12 
pressure o r i f  ices 
  robes 
24 14 stre.am, 6 wall 
12 
" 
2 wall 
2 wall 
15 I 3 wall 
Air flow was calculated from t o t a l  temperature and t o t a l  and s t a t i c  
pressures measured at the engine inlet .  Strip recorders were used t o  
record the variation with t i m e  of engine speed, fue l  flow, fuel- 
manifold pressure, cornbustor static  pressure,  and t o t a l  temperature of 
the gas i n  each combustor. - 
Y 
Fuels 
Analyses of the  three  fuels used in   this   invest igat ion are pre- 
sented i n  table  I, and the A.S.T.M. d i s t i l l a t i o n  curves are presented 
in f igure l(a) . Dis t i l l a t ion  curves f o r  the   fue ls  at a pressure equal 
t o  the combustor s ta t ic   p ressure   a t  an a l t i tude  of 50,000 f e e t  were 
obtained from the A.S.T.M. d i s t i l l a t i o n  curves (which me obtained a t  
sea-level  pressure) by the method of reference 6 and are  presented  in 
figure l (b )  . These three fuels  are, in general, within the specifica- 
t ions f o r  MIL-F-5624A grade JP-4 fuel .  
PROCEDURE 
The engine-inlet  total  pressure and the anibient s ta t ic   pressure 
were s e t  to correspond with NACA standard conditions, and the   in le t  
-&tal temperature was maintained a t  -40' F f o r  a l l  f l ight  conditions.  
After the flight conditions were established, the engine was brought 
t o  i ts  f ree  wlndmilling speed; the result ing combustor-inlet conditions 
a re   l i s t ed   i n   t ab l e  11. The f u e l  flow was then set a t  a given value 
and the igni t ion system was energized. A time limit of 30 seconds was 
allowed f o r  ignit ion t o  occur. Upon successful ignition, as indicated 
by a temperature r i s e   i n  one of the  combustors, an interval  of 30 
* 
4 __ NACA RM E53Lll 
seconds w a s  allowed for  propagation of the flame t o  the remaining com- 
bustors. A t  each a l t i tude  this procedure was repeated Over a range of 
fue l  flows from the lean  to   the  r ich limit as indicated by nonignition 
or unt i l  the  limit of the fuel-meter -capacity-was reached. After each 
ignition above 50,000 feet, the  a l t i tude was increased i n  1000-foot in- 
crements un t i l  the al t i tude limit was reached. 
- 
- 
RESULTS AMD DISCUSSION 
In i t ia t ion  of combustion i n  a turbojet engine requires a suitable 
m i x t u r e  of fue l  vapor and air. This dxture is dependent upon the  fue l  
vo la t i l i t y  and the over-all   fuel-air   ratio as discussed i n  the follow- 
ing  sections. .. 
Effect of Fuel  Volati l i ty and Fuel-Air Ratio on Altitude- 
Ignition and Propagation Characteristics 
Ignition was obtained at an al t i tude of 50,000 f ee t  and a f l i gh t  
Mach number of 0.8 i n  the following fuel-air-ratio ranges (f ig .  2) : 
0.012 t o  0.019 for f u e l  52-368, 0.018 to 0.027 fo r  fuel 53-39, and 0.025 
t o  0.038 f o r   f u e l  52-288. Fuel 53-39 had the same Reid vapor pressure 
as fue l  52-288 (2.7 Ib/sq in.), but ignition occurred at leaner fuel- 
a i r  r a t i o s .  However, t he  d i s t i l l a t i on  curves fo r  t he  fue l s  a t  the pres- 
sure level of the conibustor in le t   ( f ig .   l (b)  ) indicate that, a t  a tem- 
perature of approximately 20O0 F, fue l  53-39 would yield more fuel vapor 
per pound of fuel than would fue l  52-288. (The tenrperature of 200' F 
was selected a8 the  approximate temperature the atomized fue l  would 
likely reach for the coufbustor-inlet air temperature of 250° F obtained 
in this investigation.)  Although fue l  52-368 had a lower Reid vapor 
pressure (1.7 lb/sq in .  ) than the other two fuels, it ignited st leaner 
fuel-air ratios. Figure l(b) indicates that  fuel 53-368 would yield 
more fue l  vapor a t  a temperature of 200' F than would the  other two 
fuels.  It is  evident then that the disti l lat ion curves of figure l(b) 
gave a better qualitative indication of t he   fue l   vo la t i l i t y  as it af- 
fects  ignition  than did the Reid vapor pressure. 
d 
The tests indicated that the more vola t i le   fue l  (52-368) required 
more t i m e  for ignit ion than the less volati le fuel (52-288). Because 
of the  limited amount of data on the time for   igni t ion and the  large 
scatter,  a quantitative measure of the difference in ignit ion t i m e  
could not be established. 
The fuel-air-ratio range i n  which flame propagation.with each of 
the three fuels was obtained a t  an a l t i tude  of 50,000 feet i s  shown i n  
figure 3. The range for  each fue l  is similar to, but somewhat  more I 
IC 
M 
3 
NACA RM E53Lll  - 5 
limited than the ignition range. The operable range of  fuel-air  ra t ios  
fo r  propagation i s  shown t o  be a function o f  the   fue l   vo la t i l i ty   as  
indicated  by  the  dist i l lat ion curves at the cmbustor-inlet  conditions. 
The minimum time  required f o r  propagation increased with increas- 
ing   fue l   vo la t i l i ty .  This can be explained by considering the mechan- 
i c s  of the propagation of the flame t o  an adjacent couibustor. As re- 
ported in  reference 1, igni t ion   in   the  first combustor and the accom- 
w panying couibustion cause a temperature rise which r e s u l t s  i n  a pressure 
r rise i n  this  confbustor. The pressure rise causes a pressure differen- 2 t i a l  between the adjacent combustors and a f low of burning  as through 
the cross-fire tube. Since the more vo la t i l e  f u e l  had the leaner fuel- 
a i r  r a t io  a t  i gn i t i on ,  t he  temperature and, consequeniXLy, the pressure 
r i s e  due t o  couibustion would be less than  for a l e s s   vo la t i l e  f u e l .  
Hence, t he   r a t e  of flow of the burning gas through the cross-fire tube 
would be lawer and the time required f o r  the propagation of the  flame 
t o   t h e  second and a l l  other couibustors would be greater. 
Effect of Altitude  Variation on Altitude  Ignition 
and Propagation  Characteristics of Fuel 53-39 
A t  a flight Mach number of 0.8, the range of fue l - a i r   r a t io s   i n  
which ignition was obtained decreased with increasing  alt i tude  (fig.  
4) .  The a l t i tude  limit of ignit ion with a 1.35-joule-per-spark ig- 
ni t ion system a t  a f l i g h t  Mach nuuiber of 0.8 was approximately 54,000 
f e e t   a t  a fue l -a i r  ra t io  of 0.025. Ignition was possible over the 
range of al t i tudes investigated in as little as 2 seconds. 
Propagation of the flame from the first t o   t h e   l a s t  combustor w a s  
obtained i n  approximately the same range of fuel-air   ra t ios   as  w a s  ig- 
ni t ion f o r  a l t i tudes up t o  50,000 f ee t  ( f ig .  5). The minimum t i m e  f o r  
propagation increased with increasing altitude. This increase in  t i m e  
was from less than 1 second a t  45,000 f e e t   t o  about 5 seconds a t  
50,000 feet ;  no propagation occurred a t  54,000 f ee t .  A t  50,000 fee t ,  
the minimum time f o r  propagation occurred a t  a fue l -a i r   ra t io  of 
approximately 0.025. The minimum time f o r  both igni t ion and propaga- 
t i on  occurred a t   t h e  same fue l - a i r   r a t io   a t  the conditions of this 
investigation. 
CONCLUDING REMARE 
a A t  the conibustor-inlet  conditions of this  investigation,  the  rela- 
t ive  quant i t ies  of fue l  evaporated at  the  operating  pressure and tem- 
perature of the combustor, as indicated by the A.S.T.M. curves corrected 
t o   t h e  conibustor pressure, were found t o  provide a better  indication 
of the  effect  of f u e l  v o l a t i l i t y  on al t i tude  igni t ion and propagation 
- 
6 - NACA FM E53Lll 
than  Reid  vapor  pressure.  The  fuel-air  ratios  required  for  ignition 
and  propagation  decreased  with  increasing  fuel  volatility; for one
fuel,  both  ignition  and  propagation  were  obtained  at  the  same  fuel- 
air  ratio.  Due  to  scatter of the  data,  no  reliable  trend of time  for 
ignition with fuel volatility was discernible. However, the data did 0 
indicate a trend of increasing  minimum  times for propagation  with  in- 
creasing  fuel  volatility. 
The  fuel-air-ratio  range  over  which  ignition and propagation  were 
obtained  decreased as the  altitude  increased. The altitude-ignition 
limit with a 1.35-joule-per-spark ignition system at a flight Mach M 
nuniber of 0.8 was approximately 54,000 feet  at a fuel-air  ratio f 
0.025. 
P- 
Ln 
rl 
With  the  proper  selection  of  fuel  and a 2.7-joule-per-sppk  ig- 
nition  system,  ignition  could  be  established  in  all  combustors  of  the 
engine  investigated  in as little as 2 seconds at an  altitude of 50,000 
feet  and a flight  Mach  number  of 0.8.
Lewis  Flight  Propulsion  Laboratory 
National  Advisory  Committee  for  Aeronautics 
Cleveland, Ohfo, Deceniber  7,  1953 
1. Armstrong,  John  C.,  and  Wilsted, H. D.: Investigation of Several 
Techniques for Improving  Altitude-Starting  Limits  of  Turbojet 
Engines. mACA RplI E52103,  1952. 
2.  Wilsted,  H.'D.,  and Armstrong, J. C.:  Preliminary  Results  of 
Turbojet-mine Altitude-Starting  Investigation. NACA RM E5lH30, 
1951. 
3. Wilsted, H. D., and Armstrong, J. C. : Effect  of.  Fuel  Volatility 
on Altitude  Starting  Limits of a Turbojet  Engine. NACA RM E O G 1 0 ,  
1950. 
4.  Foster,  Hampton H., and  Straight,  David M.: Effect of Fuel  Volatil- 
ity  Characteristics  on  Ignition-Energy  Requirements  in a Turbojet 
Combustor. NACA RM ES2J21, 1953. 
5. Braithwaite,  Willie M., and  Renas,  Paul E.: Comparison  of  Turbojet- 
Engine  Altitude  Performance  Characteristics  and  Ignition  Limits 
with MIL-F-5624A Fuel,  Grades JP-3 and Jp-4. NACA RM E51L05, 1952. 
6. Hougen, 0. A . ,  and  Watson, K. M. : Industrial  Chemical  Calculations. 
Second ed., John Wiley & Sons, Inc., 1936, p,  73, 
NACA RM E53Lll - 7 
8 
i 
TABLE 1. - FUEL ANALYSES 
NACA f u e l  number 
Gravity 
OAPI 
Specific 
In i t ia l   bo i l ing   po in t  
Percentage evaporated 
A.S .T .M. d i s t i l l a t i o n  D86-46, ?F 
5 
10 
20 
30 
40 
50 
60 
70 
80 
90 
95 
Final  boil ing  paint 
Residue, percent 
Loss, percent 
-40° F 
LOO0 F 
S i l i ca   ge l  
Viscosity, centistokes 
Aromatics, total   percent  by volume 
Reid vapor pressure, lb/sq in .  
Hydrogen-cubon ratioa 
Heat o f  couibustionb, Btu/lb 
Aniline point, OF 
?Determined i n  combustion furnace. 
52 -288 
50.8 
0.776 
13 9 
224 
253 
291 
311 
324 
333 
347 
363 
382 
413 
442 
486 
1.2 
0.7 
4.27 
0.95 
10.1 
2.7 
0.168 
-8,675 
136 .% 
i2-368 
62.6 
0.729 
190 
201 
203 
208 
211 
2 14 
217 
220 
223 
229 
239 
252 
332 
1.2 
0.8 
1.63 
0.59 
11.8 
1.7 
0.172 
8,850 
138 .O 
53-39 
54.2 
0.762 
142 
195 
218 
239 
258 
269 
285 
300 
319 
339 
373 
415 
464 
1.4 
0.6 
2.57 
0.82 
9.7 
2.7 
0.172 
B ,700 
131.7 
bDetermined by aniline-gravity  constant. 
8 NACA RM E53Lll  
TABLE 11. - WINDPllILLING C0M)ITIONS 
Altitude, Combustor- Combustor- Engine- Windmilling Free wind- 
ft i n l e t  s t a t i c  i n l e t  tern- inlet air afr flow, milling 
engine 
lb/sq f% ab6 OF ature, speed, 
preesure, perature, temper- Lb/sec 
rpm OF 
30,000 
260 250 -40 4.4 2 630 55,000 
270 250 -40 4.6 2590 54,000 
325 250 -40 5.5 2560 50,000 
420 250 -40 7 .O 2560 45,000 
8 65  250 -40 13.9 2560 
. 
NACA RM E53Lll  - 
500 
460 
420 
380 
8 
.% 
!? 340 
4J 0 
g 
2 
% s 
300 
rI 
rI 
rl 
260 
220 
180 
140 
Percentage evaporated 
(a) A.S.T.M. d i s t i l l a t i on  curves. 
Figure 1. - Dist i l la t ion curves for three fuel blends. 
9 
10 - NACA RM E53Lll  
0 20 40 Ed-  80 100 
Percentage evaporated 
(b )  D i s t i l l a t i on  curves for  pressure  of 325 pounds per 
square foot (couitmstor-inlet static pressure at 
50,000 f t ) .  
Figure 1. - Concluded. D i s t i l l a t i o n  curves f o r  three f u e l  blends. 
. 
- 5  
M 
NACA RM E53Lll 
7 
L 11 
0 .( 31 
Ignition No ig- KACA f u e l  Reid vapor 
nition  pressure 
lb/sq in.  
1L = o  52-288 2.7 1 
su 52-368 1.7 + n  53-39 2.7 
.02 .03 .04 
Engine fue l -a i r   ra t io  
Figure 2. - Effect of fue l -a i r   ra t io  on ignit ion of three f u e h  with 2.7- 
joule-spark ignition system. Altitude, 50,OOO feet; f l i g h t  Mach umber, 
0.8. -
12 m NACA RM E53Lll 
32 
24 
u 
9) m 
% 
a! 
r l  
16 
u 
4 
. 
0 -01 . .. .02 .03 .04 
Engine fue l -a i r   ra t io  . .  " -- . . " 
Figure 3. - Effect of fue l -a i r  ra t io  on propagation with three fuels .  A l t i -  - r  
tude, 50,000 fee t ;   f l igh t  Mach number, 0.8. . .  
NACA RM E53Lll  C! 13 
Y 
0 
0 .01 .02 .03 .04 .OS 
Engine fuel-air  ratio 
Figure 4. - Effect of altitude  on  ignition of HACA f u e l  53-39 w-lth 1.35- 
j ou le  spark ignition system. Flight Mach nuaiber, 0.8. 
14 r_ NACA RM E53L11 
0 
0 
0 
.P 
32 
28 
24 
20 
16 
12 
0 
4 
0 
I g n i t i o n  No ig- AltltuBe, 
nition ft 
Tai led  eyuitmls i nd ica t e  no 
propagation 
Engine f u e l - a i r  r a t i o  
Figure 5. - Effect  of altitude on propagation i n  NACA fuel 53-39. 
Flight Mach number, 0.8. 
c 
. 
NACA-Langley - 5-10-51 - 556 



Evaluation of Altitude-Ignition Characteristics of Three Fuels of Different Volatility in a Turbojet Engine

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