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N A S A TECHNICAL
ME M O R A N D U iVI
CONOCXI
t
X
NASA TM X-2630
FACTORS AFFECTING * •*ALTITUDE RELIGHT PERFORMANCEOF A DOUBLE-ANNULAR s
RAM-INDUCTION COMBUSTOR
by Donald F, Schultz and Edward J. Midarz
Lewis Research Center
and
U.S. Army Air Mobility R&D Laboratory
Cleveland, Ohio 44135
N A T I O N A L A E R O N A U T I C S AND SPACE ADMINISTRATION • WASHINGTON, D. C. • SEPTEMBER 1972
https://ntrs.nasa.gov/search.jsp?R=19720024127 2018-07-13T12:01:13+00:00Z
1. Report No.
NASA TM X-26304. Title and Subtitle .FACTORS AFFEC
PERFORMANCE OF A DOUBL
RAM- INDUCTION COMBUSTOI
2. Government Accession No.
:TING ALTITUDE RELIGHTE- ANNULARI
7. Author(s)
Donald F. Schultz and Edward J. Mularz
9. Performing Organization Name and Address
NASA Lewis Research Center and
U.S. Army Air Mobility R&D Laboratory
Cleveland, Ohio 44135
12. Sponsoring Agency Name and Address
National Aeronautics and Space Administration
Washington, D. C. 20546
3. Recipient's Catalog No.
5. Report Date
September 19726. Performing Organization Code
8. Performing Organization Report No.
E-6788
10. Work Unit No.
764-74
11. Contract or Grant No.
13. Type of Report and Period Covered
Technical Memorandum
14. Sponsoring Agency Code
15. Supplementary Notes
16. Abstract
A test program was conducted to evaluate the altitude relight capabilities of a short- length,double-annular, ram-induction combustor which was designed for Mach 3 cruise operation.The use of distorted inlet-air flow profiles was tried to evaluate their effect on the relightperformance. No significant improvement in altitude relight performance was obtainedwith this approach. A study was also made to determine the effects of the reference Machnumber, the fuel temperature, and the fuel volatility (ASTM-A1 against JP-4) on the altituderelight performance. Decreasing the reference Mach number, increasing the fuel temper-ature, and using more volatile fuel all decrease the combustor pressure necessary forrelight.
17. Key Words (Suggested by Author(s) )
Jet engine;' Combustor s (annular); Combus-tors (double annular); Hydrocarbon combus-tion; Altitude relight; Diffuser radial profile
19. Security Classif. (of this report)
Unclassified
18. Distribution Statement
Unclassified - unlimited
20. Security Classif. (of this page) 21. No. o
Unclassified 2f Pages 22. Price"
7 $3.00
* For sale by the National Technical Information Service, Springfield, Virginia 22151
CONTENTSPage
SUMMARY 1
INTRODUCTION 1
TEST FACILITY 2
INSTRUMENTATION 3
CALCULATIONS 5Reference Mach Number 5Units 6
TEST COMBUSTOR 6
Combustor Design 6Combustor Configurations 9
TESTING PROCEDURE 12
RESULTS AND DISCUSSION 12Preliminary Investigation 13Performance Evaluation 14
Effect of variation in combustor reference Mach number 14Effect of variation of inlet-fuel temperature 15Comparison of ASTM-A1 and JP-4 fuels 16
SUMMARY OF RESULTS 16
CONCLUDING REMARKS 17
REFERENCES 18
iii
FACTORS AFFECTING ALTITUDE RELIGHT PERFORMANCE OF
A DOUBLE-ANNULAR RAM-INDUCTION COMBUSTOR
by Donald F. Schultz and Edward J. Mularz*
Lewis Research Center
SUMMARY
A test program was conducted to evaluate the altitude relight capabilities of a short-length, double-annular, ram-induction combustor which was designed for Mach 3 cruiseoperation. The combustor was modified by the placing of airflow distortion plates in thediffuser inlet to redirect the inlet airflow as well as removing air swirlers on the innerannulus while adding blockage to the swirlers in the outer annulus. The intent of thesemodifications was to reduce the airflow through the outer annulus by diverting the air-flow toward the inner annulus. The resultant lower velocity in the outer annulus shouldfacilitate altitude relight since decreasing velocity increases flame stabilization. Inaddition, two sets of fuel nozzles were used to improve relight performance by providingbetter atomization at low fuel flow. However, no significant improvement in altituderelight performance was obtainable with any of these modifications. The relatively highpressure loss of this combustor redistributed the distorted inlet-air profile in such away that very little decrease in reference Mach number occurred in the outer annulus.
A study was also made relating the reference Mach number, the fuel temperature,and the fuel volatility with altitude-relight performance. Significant improvements inaltitude relight performance were obtained by reducing the reference Mach number, byincreasing the fuel temperature from 251 to 425 K (-8° to 305° F), and by using a morevolatile fuel, JP-4 instead of ASTM-A1.
INTRODUCTION
This report presents the results of two efforts. One was an attempt to improve thealtitude relight capability of a double-annular, ram-induction combustor by using distor-ted inlet-air flow profiles. The other shows the effects of reference Mach number, fuel
*U.S. Army Air Mobility R&D Laboratory.
temperature, and fuel volatility on altitude relight performance.Previous testing of a short-length, double-annular, ram-induction combustor
(ref. 1) indicated a serious altitude relight problem. Relight could not be obtained atinlet-air pressures and temperatures below ambient with reference Mach numbers above0. 05. In an effort to improve altitude-relight capability, several modifications to thiscombustor were undertaken in light of its otherwise fine performance.
The previous testing indicated that the altitude-re light performance was very sensi-tive to reference Mach number. Decreasing the reference Mach number improved re-light performance. Because the double-annular combustor has the unique feature thatcombustion can be maintained in either annulus independent of the other, several meth-ods were proposed to improve relight performance by reducing the reference Mach num-ber in the outer annulus while maintaining the same diffuser inlet Mach number. Reduc-tion of velocity in the outer annulus was chosen because the igniters are located there.Hence, the combustor modifications all involved attempts to lower the airflow (i.e., thereference Mach number) in the outer annulus by directing part of the airflow that wouldnormally go to the outer annulus to the inner annulus while holding diffuser inlet condi-tions constant. In application, diffuser bleed (ref. 2) or variable combustor geometryhave been proposed for use on engines to redirect the airflow during altitude relight. Todetermine if airflow redistribution would enhance relight capability, a flow distortiondevice was used to redirect the airflow.
Other factors that may affect relight performance are fuel temperature, fuel vola-tility, and ignitor location. Fuel temperature was varied from 251 to 425 K (-8° to305° F) to evaluate its effect. A few tests using the more volatile JP-4 fuel were madeto compare its performance with ASTM-A1 fuel. Igniter location and energy were notvaried during the test program.
Combustor blowout and relight limits were obtained for several configurations. Inaddition, using the best relight configuration, maximum combustor average temperaturerise was determined as a function of inlet-air temperature, total pressure, and com-bustor reference Mach number. These average temperature rise data are useful indetermining how well the combustor could accelerate an engine.
TEST FACILITY
The altitude-re light capabilities of the combustor were studied in a closed-duct testfacility. A flow path of this facility is shown in figure 1. Airflow rates for combustionfrom 2.3 to 136 kilograms per second (5 to 300 Ibm/sec) at pressures from 1. 7 to103 newtons per square centimeters (2.5 to 150 psia), could be cooled to 260 K (8° F)or heated to 922 K (1200 F) without vitiation before entering the combustor.
Inlet flowcontrol valve
ri'UIxxr/-Heat exchange
;' temperature ri/ when used. 13
/ 333 K( 250° tot
ternAir-measuring 117orifice
I'l <III
r-.
9 to Y7600° F) Y
^N *-T*A \ J-57 jet engine
J with afterburner
- S
Flow-straighteninplates
Ii!l *
P*
9
Testsection \
\^ Water
dilution
lerature rise when used,to 300 K (210° to 540° R)
A 260 to 3 UK (8° to 100° F),V " up to 136 kg/sec (300 Ib/sec),
r-U (60° to 90° F), 0 to 700 N/cmz
Y (0 to 1000 psig)
AX"CV 79-N/cm2(115-psia)j .S ] steamline
<r-
. ~T . Altitude orr^Jrl » ntmn'nhrriclfu 7 exhaustExhaustcontrolvalve
Figure 1. - Schematic of test facility combustion air and fuel.
Fuel was available for the tests at a wide range of temperatures. For desired fueltemperatures below ambient down to 251 K (-8° F), the fuel was supplied from a con-verted nitrogen trailer which was chilled off site.
For fuel temperatures from ambient to 425 K (305° F), the fuel was passed througha steam heated heat exchanger. Further description of the test facility is given in ref-erence 3.
INSTRUMENTATION
Figure 2 shows the axial location of the instrumentation stations and the placementof the airflow distortion plates. Four, five-point pitot-static tube rakes at station 3.5were used to measure the airflow profile at the diffuser inlet.
Combustor airflow passage instrumentation consisted of two rakes in each of thethree passages to measure airflow distribution between the outer, inner, and centerflow passages (fig. 3). Each rake consisted of three total-pressure tubes and a static-pressure tube. The rakes were located at the entrance to each flow passage.
Combustor- outlet total temperature and pressure were measured at 6 incrementsaround the exit circumference. At each 6° increment, five temperature and pressurepoints were measured across the annulus. In addition, 17 randomly spaced chromel-
Punched plate-,
1.430 or 0.968(0.563) or (0.381)
Support backing plate
Punched plate: 29 percent openarea; 0.061 cm (0.024 in.) thick;0.0838 cm (0.033 in.) holes—''
Mr Exit probe shields
r Exit probe,
Figure 2. - Axial location of combustor instrumentation. (All dimensions are in cm (in.).)
Typical scoop(End view)
Airflow passages:
OuterCenter^ \Inner -v\ \
80.8 71.1(31.8) (28.0)diam diam
69.9 89.9(27.5) (35.4)diam diam
CD-11294-28
Figure 3. - Cross section of double-annular ram induction combustor. (All dimensions are in cm (in.).)
alumel thermocouples, five near the outer wall, were used to determine when ignition orblowout had occurred.
CALCULATIONS
Reference Mach Number
The reference Mach number was computed from the total airflow, air density at thediffuser inlet, and reference area. The actual combustor reference area is 0. 428
osquare meter (662.8 in. ), the minimum cross-sectional area into which the combustorwill fit. However, this combustor was designed to operate in an engine that had a com-
9bustor reference area of 0. 448 square meter (695 in. ). The reference Mach numberwas computed-using the larger area. Therefore, all pressures, temperatures, andairflows were consistent with the altitude windmill operation of that engine.
Units
The U.S. Customary system of units was used for primary measurements and cal-culations. Conversion to SI units (System International dfUnites) is done for reportingpurposes only. In making the conversion, consideration is given to implied accuracyand may result in rounding off the values expressed in SI units.
TEST COMBUSTOR
Combustor Design
The combustor is referred to as a double-annular, ram-induction combustor. Thedouble-annular design permits a considerable reduction in combustor length while main-taining an adequate ratio of length-to-annulus height. The ram-induction principle uti-lizes the kinetic energy of the inlet air to provide rapid mixing both in the primary zoneand in the secondary zone. The advantages of this combustor are a shorter combustorlength, a shorter diffuser length, and a reduction in the film-cooling-air requirement.A cross section of the combustor may be seen in figure 3. Figure 4 shows photographsof the combustor.
The combustor has 64 fuel nozzles, 32 in each annulus. Two sizes of fuel nozzleswere used to better tailor fuel atomization for high- and low-fuel flow operation require-ments. Figure 5 shows the fuel flow against fuel nozzle differential pressure for thetwo sizes of fuel nozzles used. Simplex fuel nozzles were used in this test program.
The combustor had two igniters located in the outer annulus, 180° apart. Eachigniter was supplied by a 20-joule capacitance discharge power supply. A cross-firetube located in a center air passage scoop adjacent to one of the igniters was used toassist ignition of the inner annulus. Figure 4(a) shows this crossfire tube. A more de-tailed description of this combustor and the ram-induction concept may be found in ref-erence 1.
Inner diameterheadplate airdeflector
Crossfire holein center sec-ondary scoop
C-72-851
primary scoops-^
Inner secondary scoops-.
Inner annulus
Headplate air deflectors
Outer primary scoops-J \
Outer secondary scoops
Outer annulus-1
Center primary scoops
^-Center secondary scoopsLFuel nozzle -
swirler locationsC-69-2929
(a) Closeup view.
Figure 4. - Double-annular ram-induction combustor.
Inner exit transition liner-^
Diffuser air-flow spreaders-y
-Diffuserstrutlocations
C-68-3611
(b) Side view (outer transition liner removed).Figure4. -Concluded.
.6
4000 r—
e 2000f .2
o1—100 200 300 400
Fuel nozzle differential pressure, N/crrr500
200 400 600Fuel nozzle differential pressure, psid
800
Figure 5. - Total fuel flow for 64 nozzles (32 on each annuli) as function of pres-sure drop across fuel nozzles.
8
Combustor Configurations
Previous tests (ref. 1) have shown that reducing the overall combustor referenceMach number usually improves the altitude-re light performance of a combustor. How-ever, it may not be practical to reduce the overall reference Mach number of an engine(combustor) in flight. Since the double-annular combustor has two annuli for combus-tion, diverting airflow from the outer annulus to the inner annulus should reduce theeffective reference Mach number in the outer annulus while maintaining the same overallreference Mach number. This reduction in effective reference Mach number in theouter annulus would then permit improvements in altitude relight performance.
One way to experimentally change the airflow distribution between the annuli is toinstall an airflow deflector upstream of the diffuser. The purpose of the deflector wasto mechanically distort the airflow toward the inner annulus. A moderately hub peakedprofile was obtained with a punched plate installed perpendicular to and mounted on theouter wall of the combustor housing. The location of this plate is shown in figure 2(a).A photograph of a plate segment is shown in figure 6.
p-Outer diameterI plate segment
C-71-1994
Figure 6. - Inlet airflow distortion plate.
In addition to airflow distortions in the inlet, the open-hole areas of the two com-bustor annuli were also changed in one case.
The five most significant combustor configuration designs are summarized intable I, and are described below:
Configuration 1 (fig. 7(a)) was essentially the original combustor design of refer-ence 1. The inlet-air profile was flat, the low-flow fuel nozzles with radial air swirlerswere used, as well as the original transition liners.
Configuration 2 (fig. 7(b)) was the same as configuration 1 except that the inlet-air
9
, ,-7-Fuel supply tubes
r Igniter
)-Exit/ transition
Airflow ;/spreaders-'
//• *-»«\.^•Fuel nozzles \
and swirlers V|nner shroud
(a) Configuration 1.
r ModeratelyIn let air / hub peakedprofile / —(fig. 8) /
"T"I
^-Punched plate
(b) Configuration 2 and 3.
Igniter-'
Fuel nozzlesO Disconnected from fuel supply• Connected to fuel supply
-Igniter
(c) Circumferential view of configuration 3representing fuel nozzle arrangement.
i-ModeratelyInlet air / hub peakedprofile /
No fuel supplied to i.d.
,^Fuel supply tubes
L Punched plate-Fuel nozzles with
50 percentblocked swirlers
(d) Configuration 4.
^Increased open|area transition'liner
r-Strongly hubinlet air / peakedprofile(fig. 8)
^-Punched plate
Fuel supply tubes
/ sI/
*- Fuel nozzlesand swirlers
(e) Configuration 6.
Figure?. - Cross section of double-annular ram induction combustor.
10
profile was moderately hub peaked, as shown in figure 8.Configuration 3 (fig. 7(b)) was exactly the same as configuration 2 except that fuel
was not supplied to the two 90 sectors of fuel nozzles, which were located between thetwo opposed igniters. This fuel nozzle arrangement is detailed in figure 7(c). By usingonly half as many fuel nozzles it was thought that the increased fuel nozzle differentialpressure for the same overall fuel-air ratio could improve relight performance by im-proving the fuel spray quality.
Configuration 4 (fig. 7(d)) had the same hub peaked inlet air profile as configura-tions 2 and 3. In addition, the low-flow fuel nozzles were replaced with midrange noz-zles on the outer annulus, and the nozzles were removed altogether from the innerannulus. The swirlers were also removed from the inner annulus. The fuel nozzle sizechange was necessary because 32 fuel nozzles would be carrying the flow previouslyhandled by 64. On the outer annulus the radial swirlers were blocked 50 percent byrunning metal ribbon around the inlet of the vanes. Finally, the inner transition linerwas replaced with one that had 230 percent greater open area. These modificationswere made in an attempt to greatly increase the airflow to the inner annulus.
Configuration 5 (fig. 7(e)) was essentially the same as configuration 2 except that awider punch plate was used in order to achieve a larger hub peaked distorted inlet-airprofile. The strongly hub peaked profile is shown in figure 8. A more distorted flow
Configuration Hub shape1
2 and 3
100 rr
£ 60
40
20
Hub
FlatModerately hub
peakedModeratelyStrong hub peaked
-.8 -.6 -.4 0 .2 .4
Inlet air velocity parameter,local
av
.6
Figure 8. - Oiffuser inlet annulus height versus inlet velocity parameter. Aver-age velocity, 116 meters per second (381 ft/sec). (See table I and fig. 8.)
11
was tried in order to force maximum airflow to the inner annulus. The width of thispunched plate was 1.425 centimeters (0. 562 in.),
TESTING PROCEDURE
The data were obtained in the following sequence. First, the combustor was oper-ated at a condition that insured ignition. Upon ignition the combustor fuel flow was ad-justed to give the highest possible temperature rise up to 650 K (1170 R) differentialtemperature with airflow, inlet temperatures, and pressure remaining constant.
If a 650 K (1170° R) temperature rise was reached before it appeared that the max-imum temperature rise had been obtained at that test condition, the airflow and pressurewere reduced at constant reference Mach number and inlet-air temperature. At the newpressure level, the fuel-air ratio was again ranged to determine the maximum temper-ature rise. Data were taken at successively lower pressures at constant inlet-air tem-perature, reference Mach number, and fuel-air ratio giving the highest temperaturerise until combustor blowout occurred. This fuel-air ratio is defined as the optimumfuel-air ratio. Then ignition was attempted at the conditions existing at blowout. Thefuel flow was ranged about the fuel flow rate at blowout. If ignition was not achieved atthat condition, the pressure was increased at constant reference Mach number and inlet-air temperature until ignition was obtained. This whole procedure was then repeated,but with a change in one of the basic parameters of inlet-air temperature, referenceMach number, or fuel temperature.
This procedure differs somewhat from that used in reference 1. Reference 1 useda single chromel-alumel thermocouple located near the outer diameter wall to determinewhen ignition and blowout had occurred. The new procedure required a temperature riseindication on one of 17 randomly spaced chromel-alumel thermocouples for ignition andloss of temperature rise indication on all 17 thermocouples for blowout.
RESULTS AND DISCUSSION
This altitude relight program was conducted in two phases. The first phase was apreliminary investigation, and the second phase was a performance evaluation. Thepreliminary investigation screened many combustor configurations to determine the bestone for the performance evaluation.
All test data presented in this report are tabulated in table EL
12
Preliminary Investigation
To begin this investigation, a single test was made to determine the altitude relightpotential of each combustor configuration. Experience has shown that the blowout pointat ambient inlet-air temperature and moderately high reference Mach number (0. 075) isa good basis for comparing different configurations of the same basic combustor design.
Figure 9 shows the effect of inlet total temperature on the combustor pressure atblowout. A dashed line has been drawn through the configuration 1, 3, and 5 points toillustrate that all these points likely represent the same blowout curve.
Table HI shows how the airflow from the various inlet airflow profiles of figure 8was distributed when entering the three main combustor airflow passages as shown infigure 3.
121—
2 10 —
8
CNJ
Eo
— ^
sf5CD 6
Q.
S
"c
4
— Configuration
D 0 1D 2O 3
\ A • A 4_ < * Q 0 5
\^^^ ^Composite blowout limit
^*~ —i r ~o.
- 2 8 — 7 5
61—250 300 350
Inlet total temperature, K400
100Inlet total temperature, °F
200
Figure 9. - Comparison of combustor blowout performance for configurationstested. Combustor reference Mach number, 0.075. (See table I.)
As can be seen in table in, with this relatively high-pressure-loss combustor (5.8percent at a diffuser inlet Mach number of 0.25), the airflow redistributed itself in amanner that nearly nullified the effects of the peaked profiles. This was true even withthe strongly hub peaked inlet profile of configuration 5. Therefore, these efforts to im-prove altitude relight performance were not successful. As configurations 1, 3, and 5exhibit similar performance, configuration 5 was selected for the performance evalua-tion.
13
Performance Evaluation
The performance evaluation was conducted on configuration 5 to determine the ef-fects of combustor reference Mach number, fuel temperature, and fuel volatility on alti-tude relight performance. Subsequent figures showing relight limit, blowout limit, andmaximum temperature rise will be presented.
Effect of variation in combustor reference Mach number. - Figure 10(a) shows theeffect of inlet-total temperature on minimum inlet-total pressure for relight at three dif-ferent combustor reference Mach numbers. The curves indicate the lowest inlet-totalpressure at any given inlet-air temperature at which relight can be obtained at the refer-ence Mach number indicated. Thus, at a reference Mach 0. 05 and an inlet-air tempera-
14
12
10
10
&&
B 6
js&
O
ReferenceMach number
O 0.025D .05O .075
250 300 350 400 450 250Inlet total temperature, K
300 350 400 450
100 200 300 0Inlet total temperature, °f
100 200 300
(al Altitude relight limit. (bl Blowout limit.
Figure 10. - Combustor performance showing variation with combustor reference Mach number. ASTM-A1 fuel at optimum fuel-air ratioand ambient temperature.
ture of 300 K (80 F), combustor ignition should occur at all inlet-total pressures above4.1 newtons per square centimeter (6.0 psia). As expected, as reference Mach numberdecreases, altitude relight performance improves.
Figure 10(b) shows the effect of inlet-air temperature on maximum inlet-total pres-sure for blowout at two different reference Mach numbers. The curves indicate the limit-ing inlet-total pressure at any given inlet-air temperature at which blowout is likely tooccur for a given reference Mach number. Thus, at a reference Mach 0. 05 and an inlet-air temperature of 300 K (80° F), combustor blowout would likely occur at any inlet-totalpressure below 3.3 newtons per square centimeter (4.8 psia). Above this pressure the
14
combustor should remain lit. Comparing figures 10(a) and (b) shows that combustor blow-out occurs at about 0. 6 to 1. 6 newtons per square centimeter (0. 9 to 2. 3 psi) below theprobable combustor ignition point. On a few occasions ignition was obtained at or verynear the blowout point; however, the combustor was only lit in the areas adjacent to theigniters. This resulted in very little overall temperature rise, and blowout occurredwhen the igniters were turned off.
Effect of variation of inlet-fuel temperature. - Figure ll(a) shows the effect ofinlet-air temperature on minimum inlet-total pressure for relight at three different
15
"E
12
10
_ „
I 6
0
Fuel temperature,K <°F)
O 260 (8)D 300(80)O 410 (278)
Fuel temperature,K (°F)
300(80)400(260)
Temperature rise,AT
K (°R)
600 (1080)
C =^500(900)
" —^-300 (540)100(180)
250 300 350 400 450 250 300Inlet air total temperature, K
I I I I I
350 400 450
300 0Inlet air total temperature, °F
100 200 300100 200
(a) Altitude relight limit.
Figure 11. - Combustor performance showing variation with fuel temperature. Combustor reference Mach number, 0.05; optimum fuel-air ratio.
(b) Minimum pressure at which given combustor averagetemperature rise can be obtained.
ASTM-A1 inlet fuel temperatures. As can be seen there is little difference in relightability between 260 and 300 K (8° and 80° F) temperature fuel.
The effect of inlet-air temperature on minimum inlet-total pressure for averagetemperature rises of 500, 300, and 100 K (900°, 540°, and 180° R, respectively) atinlet-fuel temperatures of ambient and 400 K (260° F) are shown in figure ll(b) usingASTM-A1 fuel. Figure ll(b) shows that, as fuel temperature increases, the pressurenecessary to obtain a particular temperature rise decreases. For example, at a 300 K(80° F) inlet-air temperature, a 9. 7-newton-per-square-centimeter (14.1-psia) inlet-air total pressure is necessary to obtain a 300 K (540° R) temperature rise with ambienttemperature fuel, while only 6. 5 newtons per square centimeters (9. 4 psia) is necessaryto obtain this temperature rise with 400 K (260° F) temperature fuel. This is due to the
15
improved fuel atomization resulting from the increase in fuel temperature. However,as inlet-air total temperature increases, a point is reached when there is no longer anyimprovement in temperature rise with increasing fuel temperature. Apparently optimumfuel vaporization is achieved by acquiring heat from the warmer air so no further im-provement can be obtained by having warmer fuel.
Comparison of ASTM-A1 and JP-4 fuels. - JP-4 was selected for this comparisonto determine if a fuel more volatile than ASTM-A1 would make a significant improve-ment in altitude relight performance. Table IV compares various fuel properties ofASTM-A1 and JP-4. The initial boiling point is the most important feature of thesefuels. ASTM-Al's initial boiling point is 433 K (320° F), and JP-4's is only 342 K(154° F). Also, table IV shows that about 60 percent of JP-4 is boiled off at the sametemperature as the initial boiling point of ASTM-A1 fuel.
Figure 12 shows the effect of inlet-air temperature on minimum inlet total pressurefor relight using ASTM-A1 and JP-4 fuels at a combustor reference Mach number of
£ 6
—
D
-0-
150 300
Inlet
1
Type of fuel
D ASTM-A1Q JP4
1350
total temperature, K
1 1
|400
100Inlet total temperature, °F
200
Figure 12. - Altitude relight limit showing variation with type of fuel. Com-bustor reference Mach number, 0.05; optimum fuel-air ratio.
0. 05. Figure 12 indicates that relights are obtainable at lower inlet-air temperatureand pressures with JP-4 fuel than with ASTM-A1 fuel.
SUMMARY OF RESULTS
A test program was conducted to evaluate the factors that might improve the alti-tude relight performance of a double-annular ram-induction combustor. Several meth-ods of distorting inlet-airflow to the inner annulus to reduce the effective reference
16
Mach number in the outer annulus were tried. It was thought that a reduction in outerannulus effective reference velocity would improve altitude relight performance. How-ever, due to the relatively high pressure loss of this combustor it is unlikely that a sig-nificant reduction in outer annulus reference velocity was obtained with any of the con-figurations tested. As a result no significant improvement in blowout was obtained withthese tests.
Also investigated were the effects of combustor reference Mach number, fuel tem-perature and fuel volatility on relight performance. The following trends were found:
Decreasing the combustor reference Mach number, increasing the fuel temperature,and using a more volatile fuel all decrease the pressure necessary to obtain relight andincrease the maximum obtainable temperature rise at any air condition. Much greaterpressure decreases are obtainable at ambient temperatures than at higher 365 to 423 K(200° to 300° F) inlet-air temperatures.
CONCLUDING REMARKS
In the preliminary investigation, many attempts were made to improve the altituderelight capability of this relatively high-pressure-loss double-annular combustor bymaking diffuser and combustor modifications. Among those modifications were a mod-erate and severe distortion of the diffuser inlet airflow. It was thought that improve-ments in altitude relight could be made if most of the airflow could be directed to theinner annulus, thus creating a "sheltered" zone in the outer annulus conducive to alti-tude relight. However, it was found that the air always redistributed itself in such amanner that prevented the creation of the desired sheltered zone. This is likely due tothe relatively high combustor total-pressure loss which tended to redistribute the dis-torted inlet-airflow to the more uniform airflow of configuration 1, the original config-uration. Thus no improvement in altitude relight performance was obtainable by thegeometric means attempted. It seems likely that a lower combustor pressure lossmight allow such an airflow redistribution scheme to successfully improve the combus-tor relight limits.
Lewis Research Center,National Aeronautics and Space Administration,
andU.S. Army Air Mobility R&D Laboratory,
Cleveland, Ohio, June 8, 1972,764-74.
17
REFERENCES
1. Perkins, Porter J.; Schultz, Donald F.; and Wear, Jerrold D.: Full-Scale Tests ofa Short Length, Double-Annular Ram-Induction Turbojet Combustor for SupersonicFlight. NASA TN D-6254, 1971.
2. Juhasz, Albert J.; and Holdeman, James D.: Preliminary Investigation of DiffuserWall Bleed to Control Combustor Inlet Airflow Distribution. NASA TN D-6435,1971.
3. Adam, Paul W.; and Norris, James W.: Advanced Jet Engine Combustor TestFacility. NASA TN D-6030, 1970.
18
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24
TABLE in. - COMBUSTOR AIRFLOW PASSAGE DISTRIBUTIONS
Configuration
Number
12 and 3
45
Inlet air velocity profile
FlatModerately hub peakedModerately hub peakedStrongly hub peaked
Percent inletair meas-
ured
83.884.379.593.8
Percent flow in -
Outerpassage
23.619.220.016.0
Centerpassage
49.952.156.055.5
Innerpassage
26.528.724.028.5
TABLE IV. - FUEL PROPERTIES
ASTM-A1 JP-4
Gravity, °APL (D287)ASTM distillation values:
Initial boiling point, K (°F)Temperature at which following percentage
of fuel is evaporated, K (°F):51020
30
4050
60
70
80
9095
Final boiling point, K (°F)Residue, percentLoss, percent
Flash temperature (D56), K (°F)Pour point temperature (D97), K (°F)Viscosity at 239 K (-30° F), m2/sec (cS)Net heat of combustion (D1405), JAg (Btu/lb)
43.1
433 (320)
444 (340)455 (360)
472 (390)
483 (410)
495 (431)
519 (474)533 (500)547 (525)
1.10.9
324223 (-
9.43.27X10 (18 615)
56.05
342 (154)
362 (192)372 (208)
396 (252)
420 (294)
445 (342)
483 (409)
502 (442)518 (471)
1.40.7
43.6X106 (18 794)
NASA-Langley, 1972 28 E-6788 25
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