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RESEARCH MEMORANDUM

ALTITUDE EVALUATION OF SEVERAL AFTERBURNER DESIGN

VARIABLES ON A 347-GE-17, TURBOJET ENGINE

By Willis M. Braithwaite, Curtis L. Walkerand Joseph N. Sivo

Lewis Flight Propulsion LaboratoryCleveland, Ohio

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. NAti6&i~’i?&AtI$%;ORY COMMI I I tEFOR AERONAUTICS

WASHINGTONOctober 23, 1953

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NACA FM E53F1O

-NATIONAL ADVIS(2RYCOMMITTEE FOR

RESEARCH MEMORANDUM

AERONAUTICS

ALTITUDE EVALUATION OF SEW?RAL AI’I’ERBURNERDESIGN V3KM3LES

ON A J47-GE-17 TURBOJET ENGINE

By Willis M. Braithwaite, Curtis L. Walker,

An investigation

and Joseph N.

SUMMARY

was conducted in an

Sivo

NKW altitude chamber toevaluate the ef=ectiveness of turbine-outlet gas-straightening vanesand vortax generators, fuel distribution modifications, and after-burner shell cooling as means of improving afterburner performance.Installation of the turbine-outlet gas-straightening vanes and vortexgenerators resulted in a lower total-pressure loss through the after-burner and an altered air-flow profile at the afterburner inlet.Therefore, it was aecessary to modify the fuel distribution to providea good fuel-air environment at the flame holder. Another result of

* the installation of the turbine-outlet gas-straightening vanes andvortex generators was the reduction of the afterburner shell temper-ature by approximately 100° R. An additional 100° R reduction in the

* afterburner shell temperature was obtainedcorrugated liner.

The best afterburner ccmfiguration ofrated the turbine-outlet gas-straighteningthe ceramic-coated corrugated liner, and a

with a ceramic-coated

this investigation incorpo-vanes and vortex generators,fuel distribution that

provided a good fuel-air environment at the flame holder. This config-uration had higher afterburner combustion efficiencies than the orig-inal configuration, and the altitude limit was in excess of 54,000feet. At low altitudes, the operation of this configuration wac notlimited by afterburner shell temperatures.

INTRODUCTION

Current and proposed military aircraft require thrust in additionto normal engine thrust for take-off, climb, and high speed at highaltitudes. Using an afterburner is one method of meeting these require-

2 NACA RM E53KL0

ments for additional thrust. Accordingly, the NACA is actively con-A

ducting several related afterburner programs.

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The investigation reported herein presents information on designfactors and modifications of the production afterburner for the

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J47-GE-17 turbojet engine designed for medfum temperature operation.——

The present report is concerned only with t&e afterburner performance ._ .=and operating characteristics. Altitude-starting characteristics oftwo of the configurations in this report are discussed in reference 1. _ G

This investigation was conducted with an engine equipped with a 3

variable-area exhaust nozzle and operated ov-era range of simulatedflight conditions in a lo-foot-diameter altitude test chamber at the

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IfACALewis laboratory.n

The initial configuration was developed by the manufacturer from aprevious production afterburner by incorporating design modificationsindicated by investigations reported in references 2 to 5. In thepresent investigation, attention was focused on three primary factorsin order to improve the performance and opeYating limits: (1) turbine-outlet gas whirl, (2) matching of the fuel distribution with the mass-flow distribution of the turbine gases, and (3) afterburner shellcooling. Previous investigations (ref. 6] indicate that turbine-outletgas whirl can have a detrimental effect on the performance character-istics of the afterburner. Accordingly, the effect of turbine-outletgas-straightening vanes and vortex generators on performance wasevaluated in this investigation. Several fuel-spray-bar configurationswere also evaluated in order to provide satisfactory matching of thefuel distribution with the mass-flow distribution. It was desirable toevaluate methods of afterburner shell cooling, since maximum thrust canbe limited by afterburner shell temperatures” ~o methods of cOO1ing.the afterburner shell were investigated: (1] the incorporation of amethod of fuel distribution that provided a lesn fuel-air mixture nearthe outer shell and (2) installation of a ceramic-coated corrugatedliner in the burner section.

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The performance with these modifications incorporated iS presentedover a range of altitudes up to 50,000 feet. Data are presented intabular and graphical form to show the effects on afterburner perform-ance of turbine-outlet gas-straighteningvanes and vortex generators,fuel distribution, and a ceramic-coated corrugated liner and to

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illustrate the effect of variations in flight conditions on one of thebest afterburner configurations.

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NACA lM E53F1O 3

KPPARATUS AKO II?STRU3ENIATION

Installation

The engine was installed in an altitude chamber that is 10 feet indiameter and 60 feet long (fig. 1). A honeycomb installed in thechamber upstream of the test section straightened and smoothed the flowof the inlet air. A forward bulkhead, which incorporated a labyrinthseal around the forward end of the engine, separated the engine-inletair from the exhaust and provided a means of maintaining a pressuredifference across the engine. A 14-inch butterfly valve in the forwardbulkhead provided cooling air for the engine compartment, and a rearbulkhead prevented recirculation of exhaust gases about the engine.The exhaust gas from the jet nozzle was discharged into an efiaust dif-fuser to recover some of the kinetic energy of the jet and thus to ex-tend the capacity of the exhaust system. The combustion in the after-burner was observed through a periscope located directly behind theengine.

Engine

A J47-GE-17 afterburning turbojet engine was used in this inves-tigation. The engine has a static sea-level thrust rating of 5420pounds without afterburning at the rated engine speed of 7950 rpm and aturbine-discharge temperature of 1760° R for an inlet air temperatureof 519° R. At this operating condition, the air flow is 104 pounds persecond. The over-all length of the engine and afterburner is approx-imately 228 inches, and the maximum diameter is 41 inches. The dryweight of the engine and afterburner, including the electronic controland airframe mounted components, is 3553 pounds. The electronic control(described in ref. 1) controls the engine speed by regulating enginefuel flow and controls the turbine-outlet temperature by regulating theexhaust-nozzle area.

Afterburner Assembly

A diagram of the afterburner assembly is shown in figure 2. Thefollowing were common to all configurations: conical diffuser, two-ring V flame holder mounted by struts from the inner body, convergingconical burning section, and variable-area clamshell exhaust nozzle.The conical inner body, mounted from the outer shell by four tubularrods, contained the fuel manifolds and a depressed flame seat in thedownstream end. Fuel was supplied to the afterburner by an air-turbinefuel pump driven by compressor bleed air.

4 NACA RM E53F1O

Modifications to the afterburner, which were incorporated in sixconfigurations, are listed in table I. The fuel-distribution patternsand spray-bar designs are presented in figure 3. A photograph of theturbine-outlet gas-straightening vanes and vortex generators is shownin figure 4. The ceramic-coated corrugated liner is shown installed inthe afterburner section in figure 5. The ceramic used was I?ationalBureau of Standards number A418, primarily composed of chromium oxide.

The six configurations are grouped as to purposes of the modifica-tions. The first group of modifications was selected to show the effectof straightening vanes on engine and afterburner performance. Thisgroup included the manufacturers original configuration A and config-uration B, which was made by incorporating in configurationA straight-ening vanes and vortex generators at the turbine outlet and a 4-inchshorter fuel mixing length (the radial fuel distribution was equivalentto that of configuration A). The shorter fuel mixing length was aresult of moving the spray bars downstream of the inner-cone supportingstruts to allow equidistant circumferential spacing of the bars. --

The second group of configurations was selected to illustrate theeffect of modified fuel distribution. These configurations were geo-metrically similar, all having straighteningvanes, vortex generators}and corrugated liners. Configuration C had a radial fuel distributionequivalent to the original configuration. In configurations D and E,the fuel distribution was modified to compensate for the shift in mass-flow profile caused by the installation of the straighteningvanes andvortex generators. This modification was based on the total-pressureprofile at the diffuser outlet. The spray bar for configurationD had”equally spaced holes of varied diameters, while the bar for configura-tion E had varied spacing of equally sized holes. The bar for config-uration E, designed by the manufacturer, was a uniorifice bar with ametering orifice at the inlet to the bar that permitted the use oflarger spray holes to prevent plugging by foreign material in the fuel.

The third group of modifications was selected to evaluate severalmethods of cooling the outer shell. ConfigurationsA, B, C, and T arecompared. Configuration was the original configuration already de-scribed. Configuration had a fuel distribution that was lean near theouter shell but was “otherwisethe same as configurationA. Configura-tion B differed from configuration in that turbine-outlet gas-straightening vanes and vortex generators were added, and configurationC differed from configuration B only by the addition of the corrugatedliner.

Instrumentation

Engine-inlet air flow was determined by pressure and temperaturemeasurements at the compressor inlet (station-

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1, fig. 6(a)). Instrumen-

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NACA RM E53!3’1O 5

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9

tation measured the engine midframe air bleed, which was subtractedfrom the engine-inlet air flow in order to obtain the afterburner airflow. Turbine-outlet temperature was calculated from engine performance,and turbin-outlet temperature for the electronic control was deter-mined by averaging the engine manufacturer’s eight thermocouples. Theangle of whirl of the gas flow was measured at the diffuser outlet bymeans of a rotatable rake (fig. 6(b)). Twenty-five thermocouplesattached to the afterburner shell measured temperatures at station 8,

located79~ inches do-stream of the t~bine-outlet flange (fig. 6(c)).

Total pressures at the a&aust-nozzle inlet were obtained with a water-cooled survey rake (fig. 6(d)), and ambient pressure in the region ofthe exhaust-nozzle outlet was determined by static-pressure probes inthe plane of the e@wust nozzle. Engine and afterburner fuel flowswere measured by calibrated rotameters. The fuel used in this inves-tigation was MIZ-F-56241 grade JP-4.

PR(xlmIRE

The inlet and exhaust conditions for these tests were determinedby the altitude and flight l&ch number according to RNA standardatmosphere; 100-percent ram pressure recovery was assumed. Afterburnerperformance data were obtained over a range of altitudes from 15,000to 50,000 feet at a flight Mach number of 0.6 and a range of flightMach numbers from 0.4 to 1.0 at an altitude of 30,000 feet. For eachflight condition, data were obtained at rated engine speed and turbine-outlet temperature as maintained by the electronic integral control fora range of fuel-air ratios.

The range of fuel-air ratios at each flight condition represents,in general, the practical operating range for the afterburner. Therich operating limit was determined by afterburner shell temperaturelimit (2010° R) or bymaximm nozzle opening. It should be noted that,with the electronic control, exhaust-nozzle area is a function ofexhaust-gas temperature. Therefore, maximm nozzle opening was eitherthe physical limit of the nozzle or the area obtaine&with maximumtemperature for those configurations that reached a point of decreasingtemperature with increasing fuel flow. The lean limit of operationwas indicated by unsteady (oscillatory) combustion in the afterburneror blow-out of the flame. Over the range of conditions investigated(turbine-outletpressures up to 2950 lb/sq ft with fuel-air ratiosup to 0.04), this unsteady combustion at the lean limit was the onlyoscillatory combustion phenomenon encountered. ITosignificant altitudelimits were obtained for these configurations because of limitations ofthe altitude exhaust facilities. For afterburning conditions, the fac-ility was limited to about 54,000 feet, which corresponds to a turbine-outlet total pressure of about 580 pounds per square foot.

6 NACA RM E53F1O

The jet thrust produced by the engine andby a self-balancingnull-type pneumatic thrust

0

afterburner was measuredcell. A sketch of the

linkage from the engineThe symbols and methods

to the thrust cell is presented in ftgure l(b). ?of calculation are presented In the appendix.

RESULTS AND DISCUSSION

The results of this investigation are discussed in relation to the Nthree factors considered for improving the afterburn~ performance; Ethat is, (1) reduction of whirl in the turb~_ne-outletgas, (2) matching

+

of the fuel distribution with the mass-flow distribution, and (3) cool-ing of the afterburner shell. The altituder_perf’ormancedata are pre-sented in tabular form in table II and In graphical form In figures 7to 17. —

Effects of Turbin*Outlet Gas-StraighteningVanes and

Vortex Generators on Afterburner Performance

The performance of an afterburting engine may be detrimentallyaffected by high angle of whirl of the turbine-outlet gases (ref. 6).The addition of a whirl velocity produces high resultant velocities —

that make burning more difficult. Furthermore, a high angle of whirlproduces high total-pressure losses in the burner that are particularly

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noticeable during nonafterburning operation: Whirl losses are lesspronounced during combustion, probably because the high level of tur- dbulence during combustion reduces the whirl component of the flow.

Since whirl angle increases with diffusion (ref. 7), measurementsof the whirl were made at the diffuser outlet to obtain the maximum .-.

angle and the results are presented in figure 7. The method of obtain-ing this curve is the same as used in reference 6. The direction ofwhirl was opposite to turbine rotation, as &ndicated by the negativeangles on figure 7, and was great= than 30 over most of the passage.The installation of the turbine-outlet gas-straighteningvanes andvortex generators, designed by the manufacturer to reduce the whirlangle and to provide a uniform velocity profile, resulted in only 10°whirl in the same counterrotatfonal direction.

Effect on afterburner-inlet temperature and pressure profiles. -installation of the turbine-outlet gas-straighteningvanes and vortexgenerators caused a shift in afterb&ner-inl& temperature profile inrelation to the control thermocouple. After the vanes and vortex gen- .

erators were installed, the average burner-inlet gas temperature wasfrom 40° to 70° R lower for the same indicated control temperature. w—.

NACA RM E53E’1O 7

*This effect can be seen by inspection of data from table II. Thesetting of turbine-outlet temperature with the manufacturer’s instru-mentalion resulted in operatIon at a lower turbine-outlet temperature.and resulting lower nonafterburning thrusts for the configurateionshaving straightening vanes and vortex generators.

Another effect of the straightening vanes and vortex generatorswas a slightly more uniform total-pressure distribution (fig. 8) at thediffuser-outlet (station 6). This trend in total-pressure profile

+ indicates more nearly uniform mass-flow and velocity distribution of:N the combustion gases. However, a comparison of the pressure levels on

figure 8 is invalid because of the different temperature levels.

Effect on internal performance. - A comparison of the total-pressure-loss ratios before and after the straightening vanes andvortex generators were installed is presented in figure 9. In theconfiguration with vanes and vortex generators (configuration B), theturbine-outlet total-pressure measurements were made downstream of thevanes. Therefore, the total-pressure loss through the straighteningvanes and vortex generators was not included in the pressure-loss ratio,defined as the ratio of the total-pressue loss through the burner tothe burner-inlet total pressure. The pressure-loss ratio through thevanes and vortex generators was determined to be approximately 0.025 at30,000 feet and 0.035 at 50,000 feet from the engine pumping character-istics. This value has been added to the data on figure 9 for config-uration B, and the combined pressure loss is shown by the dashed curve.Thus the installation of the straightening vanes and vortex generatorsreduced the over-all pressure-loss ratio by 0.02 to 0.01, the greaterreduction occurring at low fuel-air ratios where the whirl is believedto have been greater.

The afterburner combustion efficiencies before (configuration)and after (configuration B) straightening vanes and vortex generatorswere installed are compared in figure 10(a). For the range of fuel-airratios investigated there was no appreciable difference in combustionefficiency at 30,000 feet, while at 50,000 feet the configuration withvanes and vortex generators had lower combustion efficiency. This 10SSin efficiency with vanes was due to a poor fuel-air distribution. Thefuel distribution was designed for the mass-flow profile that existedbefore the vanes were installed. Following installation of the vanes,this mass-flow profile was modified, but the fuel distribution remainedthe same and resulted in a different fuel-air ratio profile.

The decrease in combustion efficiency for configuration B at 50,000feet resulted in lower exhaust-gas temperatures (fig. 10(b)). The lowerturbine-outle~ temperatures also contributed to the decrease in after-burner exhaust-gas temperature.

Effect on over-all performance. - As previouslyof the straighteningvanes and vortex generators wasafterburner pressure loss, and an increase in thrustHowever, because of the difference in turbine-outlet

NACARME53F1O

*noted, the effectto reduce thewould be expected. .—conditions in this

investigation, the increase in thrust was not realized and the augmented %jet thrust ratio was the only valid basis for thrust comparison. (The 3nonafterburning jet thrust was calculated for each point with theturbine-outlet conditions obtained at that point and with the pressuredrop assumed through a standard tail pipe as explained in the appendix.)The augmented jet thrust ratios are presented in figure 1O(C). At30,000 feet, where combustion efficiency was essentially the same,installation of the straightening vanes and vortex generators resultedin an increased jet thrust ratio of about 0.03. At 50,000 feet, how-

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ever, the reduction in combustion efficiency countered the reductionin afterburner pressure loss, and an increase of only 0.01 was observedin jet thrust ratio.

Net thrust specific fuel consumption, presented in figure 10(d),is affected by the tail-pipe pressure loss, -tyrbine-outletgas temper-ature, and combustion efficiency. At 30,000 feet, combustion efficienc~-was unaffected by the turbine-outlet gas-straighteningvanes, and thereduced tail-pipe pressure loss should have resulted in decreased specificfuel consumption. However, the lack of tiprovement illustrates the factthat the control should be rescheduled to give the same engine operating “~conditions after the temperature profile sh\ft due to the installationof the straightening vanes. Also, the shtft_in mas$-flow profile due t-othe addition of the vanes required an altera~ion to the fuel distributionto prevent a decrease in combustion efficiency at 50,000 feet. Theresults of modifying the fuel distribution after straightening vanesand vortex generators were installed are discussed in the followingsection.

Effect of Fuel Distribution on Performance

The installation of the turbine-outlet straighteningvanes andvortex generators resulted in a loss in combustion efficiency thatwas attributed to a poor fuel-air distribution. Before alteration ofthe fuel system, a ceramic-coated corrugated liner was installed, theeffects of which will be discussed later. Configurations C, D, and Ediffered only in fuel distribution.

A comparison of the combustion efficiency of configuration C,which had a fuel distribution equivalent to the original configuration

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NACA RM E53F1O 9

b A, and configuration D, which had a fuel distribution designed to matchthe mass-flow profile, is presented in figure U.(a). The peak effi-CieIICy Of COIIf@UatiOn D occurred at higher fuel-air ratios than it

. did for configuration C. Also, configuration D had about 0.11 highercombustion efficiency than configuration C at 50,000 feet. Forexample, at 50,000 feet, combustion efficiency for configuration Dhad a peak value of 0.82 at a fuel-air ratio of 0.037 and was about0.70 at a fuel-air ratio of 0.06.

The exhaust-gas temp=ature (fig. n(b)) and net thrust ratio(fig. 11(c)) were higher for configuration D above a fuel-air ratioof 0.035 at 30,000 and 40,000 feet and higher at all fuel-air ratiosinvestigated at 50,000 feet. Net thrust specific fuel consumption -(fig. n(d) ) was lower for configuration D above a fuel-air ratio ofQ.035 at 30,000 and 40,000 feet and lower at all fuel-air ratios inves-tigated at 50,000 feet. The redesigned fuel distribution provided animprovement in combustion efficiency at fuel-air ratios in excess of0.035 and, hence, in over-all afterburner performance.

.Comparison of configurations D andE (figs. n(a) to (c)) shows

that the use of equal-diameter spray holes with various spacing(configurationE) resulted in about the same performance as withvarious-diameter spray holes with equal spacing (configuration D).Furthermore, the use of a metering orifice at the inlet to the spraybar did not affect performance.

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Afterburnm Shell Coolingw

A maximum temperature limit of 2010° R for thewas imposed by the structural strenmh requirements

afterburner shellof the afterburner.

The shall tem~erature of an afterbu&er is a function of the exhaust-gas temperature and the burner pressure (gas density). It was foundthat with a rich fuel-air mixture the original configuration waslimited at 30,000 feet and below by maxiurumshell temperature. !There-fore, to permit higher exhaust-gas temperatures at low altitudes,methods of cooling the afterburner shell were investi~ted.

Reduction in turbine-outlet whirl. - A reduction in the temperatureof the afterburner shell of from 80” to 120° R resulted from theinstallation of the straightening vanes and vortex generators (fig. 12)at an es&aust-gas temperature of 3100° R and altitudes of 30,000 and50,000 feet. The larger reduction occurred at the lower altitude,where shell cooling was more critical. The data further indicatethat the reduction would have been greater at higher fuel-air ratios.This decrease in shell temperature with the installation of strai@ten-ing vanes was due to two factors. The deorease in whirl reduced the

10 NACA RM E55F1O

tendency for the fuel to centrifuge toward the outer shell and also 4probably increased the thickness of the boundary layer at the outer

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shell..

Fuel distribution moved away from shell. - A previous investigation(ref. 5) showed that shell temperatures could be lowered by modifyingthe fuel distribution to reduce the amount of burning near the shell.Configuration F incorporates such a modification, as shown by a com-parison of figures 3(a) and (e). This modification reduced the shell

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temperature approximately 30° at 30,000 feet (fig. 12). However, itj

lowered the altitude limit from over 54,000 feet for configuration Ato 50,000 feet for configuration F at a fliqht Mach number of 0.6. .

Ceramic-coated corrugated liner. - Configuration C incorporated aceramic-coated corrugated liner installed in the afterburner in additionto the straightening vanes and vortex gen~ators. Over the range ofexhaust-gas temperatures investigated, the combination of straighteningvanes and liner reduced the temperature 200° to 300° R. For configura-tion C an increase in exhaust-gas temperature from 2800° to 3300° Rat 30,000 feet (fig. 12(b)) resulted in only 50° R increase in shelltemperature. The combination of straighteningvanes and liner was themost effective means of reducing shell temperatures investigated andwas equivalent to the use of approximately 8 percent of engine air flowfor cooling air in a conventional convective cooling system.

The coating used on the liner was a Bureau of Standards numberA418 ceramic. After approximately 33 hours of afterburning, thecoating and liner showed no evidence of deterioration.

A comparison of the shell temperatures for the three configurations(A, B, and C) over a range of turbine-outlet pressures with an e*ust-gas temperature of 3100° R is presented in figure 13. The temperaturefor the original configuration increased with increasing pressure,while that for configuration C (liner plus straighteningvanes andvortex generators) was essentially constant. This figure illustratesthe greater effectiveness of the vanes and liner at low altitudes(high burner pressures), where shell temperatures were found to becritical on the original configuration.

Performance of Final Configuration

Performance data are presented for configuration D over a rangeof altitudes from 15,000 to 50,000 feet at a flQht Mach number of0.6, and for flight Mach numbers of 0.4 and 0.8 at an altitude of 30,000

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NACA RM E5SF1O 11

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feet. This configuration was used for evaluation because it had per-formance and operating and cooling characteristics equivalent orsuperior to any of the other configurations.

Effect of altitude and flight Mach nsmber variations. - For agiven afterburner fuel-air ratio, increasing the altitude at a constantflight Mach number tended to lower the combustion efficiency, exhaust-gas temperature, and augmented net thrust ratio (fig. 14). The peakcombustion efficiency deoreased from 96 percent for an afterburnerfuel-air ratio of 0.027 at 15,000 feet to 82 percent for a fuel-airratio of 0.037 at 50,000 feet. The corresponding e-ust-gas temper-atures were 296@ and 2880° R. For a constant fuel-air ratioof 0.03, the temperature decreased from 2960° R at 15,000 feet to2710° R at 50,000 feet, while the combustion efficiency decreased from96 to 76 percent. For a wide-open eihaust nozzle, the ekhaust-gastemperature decreased from 3550° R at 30,000 feet at a fuel-air ratioof 0.056 (high temperature data were not obtained at 15,000 ft) to3380° R at 50,0~ feet at a fuel-air ratio of 0.061. This decreaseresulted in a decrease in augmented net thrust ratio from 1.6 to 1.52.

The net thrust specific fuel consumption decreased as the altitudeincreased up to analtftude between 30,000 and 40,000 feet. A furtherincrease in altitude resulted inan increase in specific fuel consump-tion. This decrease was due to the increasing engine cycle efficiencywith the decreasing compressor-inlet temperature as the altitudeincreased until the tropopause (35,000 ft) was reached, which more thancompensated for the decrease in combustion efficiency. Above thetropopause, the temperature remained constant, and the decreasingcombustion efficiency caused an increase in specific fuel consumption.

The effect of varying flight Mch number at an altitude of 30,000feet is presented in figure 15. A decrease in flight lhch numberaffected the afterburner combustion efficiency (fig. 15(a)) and theexhaust-gas temperature (fig. 15(b)) in the same manner as did anincrease in altitude; that is, combustion efficiency and exhaust-gastemperature decreased with decreasing Mach number for a given after-burner fuel-air ratio. However, since the range of flight Mach num-bers was small (0.8 to 0.4), the effect was slight. The augmentednet thrust ratio (fig. 15(c)) decreased with decreasing Mach number,but the net thrust specific fuel consumption (fig. 15(d)) increased asa result of the decreasing combustion efficiency.

Afterburner combustion efficiency as a function of turbine-outlettotal pressure is presented in figure 16 for afterburner fuel-airratios of 0.025, 0.035, 0.045, and 0.055. At the low fuel-air ratio,the efficiency decreased rapidly with decreasing pressure, but decreas-ing pressure &ffected the

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efficiency less at the higher fuel-air ratios.

12 NACA RM E55J’1O

Thrust Generalization

A method of generalizing jet thrust for a turbojet engine has beendeveloped in reference 8. The basis of this method is that the jetthrust is a function of the flow through a nozzle and may, therefore, bedescribed in terms of the nozzle-inlet total pressure, the ambient staticpressure, and the throat area of the nozzle. If the experimentallydetermined jet thrust is ad~usted for nozzle area variation, it willgenaalize when plotted against (1.25P9 - po). This method of general-ization provides a calibration that may be used in determining turbojetengine jet thrust in a flight installation for take-off and flight con-ditions.

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✎

This method of generalizing jet thrust has been used for the dataof configuration D (fig. 17). The factor used to adjust the jet thrustfor area variations of the variable-area nozzle, which is presented infigure 17(b), is the ratio of Jet thrust to jet thrust for maximumnozzle area as a function of the ratio of actual nozzle area to maximumnozzle area. Therefore, the jet thrust presented in figure 17(a) isthe jet thrust obtainable with the maximum nozzle area.

The generalized thrust curve maybe used to obtain the Jet thrustfor a given flight condition and afterburner fuel-air ratio. The valuesof the exhaust-nozzle-inlettotal pressure and ambient static pressure —may be obtained from the flight conditions and the engine characteris-tics. With these data, the jet thrust for maximum nozzle area can be

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obtained. The exhaust-nozzlearea for a given afterburner fuel-airratio can be obtained from figure 17(c), and from this the ratio ofjet thrust at the given exhaust-nozzle area to the jet thrust at max-

*

imum nozzle area can be obtained. A numerical example of this cal-culation is given in the appendix.

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C!OIVCLUDINGRXMARE3

An investigationof the J47-GE-17 afterburning engine showed thatthe installation of turbine-outlet gas-straighteningvanes and vortexgenerators reduced the counterrotationalwhirl by approximately 20°.This change modified the burner-inlet (turbine-outlet)total-tempera-ture profile so that the calculated average turbine-outlet gas temper-ature was from 40° to 70° R lower for the same indicated control tem-perature. This decrease in turbine-outlet temperature resulted in

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lower thrust for the same afterburner fuel-air ratio. However, thetotal-pressure-loss ratio was reduced by 0.01 to 0.02, and the augmentedjet thrust.ratiowas increased by 0.01 to 0.03. Because of

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a change in total-pressure profile and thus a change in masa-flow profile with the same fuel distribution, the fuel-air distribution

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NACA RM E53F1O 13

at the flame holder was not as suitable with the vanes and vortex gen-erators, and the combustion efficiency decreased.

13gadjusting the fuel distributionto the.mass-flow profile thatresulted from the installation of the straightening vanes and vortexgenerators, the loss in combustion efficiency was overcome. The peakvalue at 50,000 feet altitude was approximately 82 percent at a fuel-air ratio of 0.037 and ~S about 70 percent at a fuel-air ratio of 0.06.The combustion efficiency, the eximust-gas temperature, and the aug-mented net thrust ratio were higher at higher fuel-air ratios, while netthrust specific fuel consumption was lower for the modified fuel distri-bution. The final fuel distribution was achieved both with spray barshaving equally spaced various-diameter holes and uniorifice spray barshaving variously spaced equal-diameter holes.

Afterburner shell temperatures were reduced approximately 100° Rby the installation of the straightening vanes and vortex generatorsfor an exhaust-gas temperature of 3100° R at 30,000 feet altitude. Anadditional reduction of 100° R in shell temperature was achieved by theinstallation of a ceramic-coated corru~ted liner. The liner was usedfor approximately 33 hours of afterburner operation with no noticeabledet~ioration or change in appearance.

Except at the lean limit of combustion, no oscillatory combustionwas encountered during this investigation, which covered burner pressuresup to 2950 pounds per square foot with fuel-air ratios up to 0.04.

Lewis Flight Propulsion LaboratoryRationalAdvisory Co?mnitteefor Aeronautics

Cleveland, Ohio, May 18, 1953

14 IIACARM E53F1O

APEENDIX - CALCULATIONS

symbols

The followimg synbols are used in this report:

oross-sectionalarea, sg,ft

velocity coefficient, ratio of actual jet velocity to effec-tive jet velocity

jet thrust, lb

calculated nonafterburning

net thrust, lb

oalcul.atednonafterburning

fuel-air ratio

Jet thrust, lb

net t@ust, lb

acceleration due to gravity, 32.2 ft/sec2

enthalpy, Btu/lb

lower heating value of fuel, Btu/lb

Mach number

total pressure, lb/sq ft abs

total pressure at exhaust-nozzle survey station in standardengine tail pipe, lb/sq ft abs

static pressure, lb/sg.ft abs

gas constant, 53.4 ft-lb/(lt)(OR)

total temperature, %

velocity, ft/sec

effective velocity, ft/sec

air flow, lb/see

——

*

.

NACA RM E53F1O 15

* Wa,c compressor leakage air flow, lb/9ec

Wf fuel flow, lb/hr.

Wf/Fn specific fuel consumption based on total fuel flow and netthrust,(lb/lrr)/lbthrust

N ‘g gas flow, lb/9ec@3

Y ratio of specific heats for gases

v combustion eff~ciency

Subscripts:

a

ab

e

f

. g

m.

n

nc

s

T

tc

tp

o

1

3

5

air

afterburner

engine

fuel

gas

maximum exhaust-nozzle

exhaust-nozzle

nozzle cooling

labyrinth seal

total

%urline cooling

turbine pump

outlet, vena contracta

free-stream conditions

engine inlet

compressor outlet at engine combustor inlet

turbine outlet or tail-pipe diffuser inlet

16 NACA RM E53F1O

9

10

exhaust-nozzle tilet

exhaust-nozzle outlet

Methods of Calculations

Flight Mach num%er and airspeed. - Flight Mach number and equiva-lent airspeed were calculated from engine-inlet total preesure and tem-perature and free-stream static pressure with complete total-pressureram recovery assumed:

1 r y.-l -I l— “.L

and

L%(f)“

.

Air flOW. - Air flow was determined from pressure and temperaturemeasurements obtained in the engine-inlet annulus by the followingequaticm:

r27~g‘a,l ‘“0”g8p111~

.71-1 -

~

()

P1

g-1

. .

where the 0.98 accounts for the 0.02 leakage between measuring stationand compressor inlet. Air flow at the com@essor outlet (statIon 3)was obtained by deducting the compressor leakage, turbine and nozzlecooling-air flows, and compressor bleed air used to drive the turbtie.fuel piimp:

Wa,3 =Wa,l - Wa,c

Gas flow. - Afterburner gas

%,9 =wa,3 +

.

.

.

.

- Wa,tc - Wa,nc - ‘a)tp

flow is

‘f,e ‘Wf,ab3600 + %,tc

—- —_

.

.

NACA RM E53F1O

Fuel-air ratio. - The engine fuel-air ratioing equation:

‘f efe s~

3600 Wa,3

17

is given by the follow-

The afterlwz’nerfuel-air ratio used herein is deftied as the weightflow of fuel in~ected intodivided by the weight flowcombtiing air flow, enginecombustion efficiency, theratio is obtained:

the afterburner plus the unburned engine fuelof unburned air entering the afterburner. Byfuel fluw, afterburner fuel flow, and engtiefolluwing equation for afterburner fuel-air

(1 - qe)vf,e + ‘f,abfab =

~ewf,e3600(Wa,3+Wa,tc) -

0.0675

where 0.0675 is the stoichiometric fuel-air ratio for the engine fuel.

The total fuel-air ratio for the engine and afterburner is

‘f,e + ‘f,abf= ~600(Wa,3 +Wa,tc)

.

Engtie combustor efficiency. - Engine combustor efficiency is theratio of the enthalpy rise through the engtie divided by the product ofengine fuel flow and the lower heating value of the fuel:

of ref. 9).

Engine combustion efficiency was obtained from a correlation ofthe engine with a standard tail pipe. For this investigation T5 was

then obtained by solving the above equatim for ~,5.

Afterburner combustion efficiency. - Afterburner combustion effi-ciency was obtained by dividing the enthalpy rtse through the afterburnerby the product of the afterburner fuel flow md lower heating value ofthe fuel:

(ha,lo - ha,l) +f~lo - ~efehc. qab = fhc - ~efehc

18 NACA RM E53F1O

Augmented thrust. - The jet thrust of the combined engine and.

afterburner was detezmdned from the thrust-system measurements by theequation .

where Fd is equal to the thrust-system scale reading adjusted for the

pressure difference on the lhds connecting the thrust bed in the testchamber and the measuring cell outside the test chamber, and the last

‘em‘e”o(+:~:vlis the difference between the pressure forces

on the bellmouth and the momentum at the bell.mouthoutlet.

The augmentedstream momentum of

net thrust was obtained by subtracting the f&ee-the inlet air from the jet thrust:

‘a,lFn = Fj -0.98g ‘0

Nonafterburning jet thrust. - The jet thrust for the nonaf%erburningengine was calculated for each afterbumtig point with measured turbine-outlet total pressure, engine gas flow, calculated turbine-outlet totaltemperature, and an assumed tall-pipe pressure drop. Experhnental datafrom the nonafterburning engine indicated that the total-pressure lossthrough a standard tail pipe between stations 5 and 9 was approxhately0.04 at rated engine speed, that is, P9 = 0.96P5. The nozzle velocity

coefficient for the nonafterburning engine was assumed to be 0.99,while that for the afterburning engine was assumed to be 0.97. Thejet thrust equation used was

where (vef/+EF) iS ~ effective velocity parameter derived in

reference 10; and the nonburmlng net thrust was

Exhaust-gas total temperature. - The total temperature of theexhaust gas was calculated from the jet thrust and conditims existingat the exhaust nozzle:

—

.

—

-.

-. —

.

NACA FM E53F1O 19

.

.

Nco3

()‘efwhere —

-is the effective velocity parameter.

Calculated jet thrust from generalized Jet thrust. - It has beenshown (ref. 8) that the jet thrust of a turbojet engine maybe relatedto the nozzle-inlet total pressure, ambient static pressure, and ties.This relation is

Fj =A10(l-2~g - PO]

This eq.uationmaybe represented by two plots as figures 17(a) and (b).

For an assumed flight condition of 40,000 feet altitude and aflight Mach num%er of 0.6, the pressure values are found from table IIas P9 = 1009 pounds per square foot and p. = 392 pounds per square

foot. Then

1.25Pg - PO = 1.25(1009) - 392 = 869.5 lb/sq f%.

Erom figure 17(a), the jet thrust for an open nozzle (AIO.m =,3.5 S~ ft),which would require an afterburner fuel-air ratio of 0.0575, is 2790pounds.

If an afterlnwner fuel-air of 0.040 is desired,

Alo = 3.17 sqft (from fig. 17(c))

and

%0 3.17—=—= o.906‘lO,m 3.50

With the area ratio, the thrust correction ratio may be found fromfigure 17(b) to be

‘~/Fj,m = 0.912-

Therefore, the jet thrust for this fuel-air ratio is

F~ = Fj,m @j/Fj,In) = 2790(0.912) = 2544 lb

REmmENms .

20 NACA RM E53F1O

1. Jansen, Rnmert T., and Harvey, Ray W., E&.: Transient Data ObtainedDuring Altitude Starts of the J47-GE-17 Afterburner. NACA m

-,

E52J17, 1952.

2. Jobnscn, LaVern A., and Meyer, C=l L.: Altitude ~erfo~ceCharacteristics of Turbojet-Engine Tail-l?ipeBurner withVariable-&ea Exhaust Nozzle Using Several Fuel Systems andFlame Holders. NACARME50F28, 1950. E

E3. J~sen, Emmert T., and mo~, H. CWI: Altitjude perfo~ce

Characteristics of Tail-l?ipeBurner with Variable-Area ExhaustNozzle. NACARME50E29, 1950.

4. Flemtig, W. A., Conrad, E. William, and Young, A. W.: ExperimentalInvestigation of Tall-Pipe-BurnerDesign Variables. NACA RME50K22, 1951.

5. Conrad, E. William, and Jansen, Emnert T.: Effects of IntenalConfQuration on Afterbuzmer Shell Temperatures. NACA EM E51107,1952.

6. Braithwaite, Willis M., Renas, Paul E., and Jansen, Emme& T.:Altitude Investigation of Three Flame-Holder and Fuel-SystemsConfigurations in a Short Converging Afterburner on a TurbojetEngine.

.NACA FM E52G29, 1952.

7. Schwartz, IraR.: Investigations of an Annular Diffuser-Fan Com-bination Handling Rotattig Flow. NACA RM L9B28, 1949.

8. Hesse, W. J.: A Simple Gross Thrust Meter Installation Suitable F&fidicatlng Turbojet Engine Gross Thrust in Flight. Tech. Rep.No. 2-52, Test Pilot T&aintig Div., Naval Air Test Station,Apr. 3, 1952.

9. Turner, L. Richard, and Bogart, Donald: Constant-Pressure CombustionCharts Including Effects of Diluent Addition. NACA Rep. 937, 1949.(SupersedesNACATN’s 1086 md 1655.)

10. Turner, L. Rlohard, Addie, Albert N., and Zfmmerman, Richard H.:Charts for the Analysis of One-Dimensional Steady CompressibleFlow . NACA TN 1419, 1948.

configuration] IWmber of

fuel-sprqbara

A 19

B 20

c 20

D 20

E 20

TF 19

I

k6Z 1

!CKE!LEI. - MODITICATIONSIN03~IM!lTD

Irifloeaper

Bgtray bti

10

10

10

18

16

10

Kixlllglength

h.

26

22

22

22

22

26

5WW-Wlocatlon(fig. 2(8):

1

2

2

2

2

1

MelaiStii-butkm(fig. 3;

(a)

(b)

(b)

(c)

(d)

(e)

,

Wdlfioat Ions

Manufacturer’s original oon-

fIglrmtkm

Fuel distribution equivalenttto oonf~tim A; turbine-outlet gas-~traightm~vaues and vortex generatmwadded

Cem.mlo-ooa-ted Corrugatedliner added to configom-tjmn B

kiodlfledfuel distribution;other .mriables same as

Oonfl.goratlonc; finalCmm.guration

Unioriflce qmay bar with dM -tributim of oonffguratlm D;

other variables same as con-figuration C

Fuel diatkributlonaway flmmshell; no turbine-outletatralghtening vanes, vortexgenerator~, or ooatedllner

NACA RM E53F1O22

.TABLE m, - ALTITUDE

Alt;~de , F$llt Fre:g~am En#g- ?3n1313~-Tu#fi- Tdrblna - Mantu%c - 3n2in*- 12n# 4St8rburn8r To-l 211el-Afterburnm.outlet turOr ,s inlet ru.1 flew, airf ratio, fuel-air

number, prnwure, total t0t21 total total t:lx:g - air flow, flow,% P(JJ

‘f,ab, ratio,pr.nmure, temper. prmnurr, tK!pera-

“Wa,l’ %,*B E ‘abplt

ature, P5 ●tluw, oontrol

:;, %* tempWa- ~ ~ i-mS(Ift abn lb lb

ture, 8*Osq rt abn *

hrml Ct abs

T5,mI

%

:

15,01m 0.589,600.8C!2.5s1.599

30,002 0;:::

.s83

.3s1

.38s

.381

50,000 0.580.575.578.Sas.594.504

s,mo 0.813.793.798.7S2.7W

-.802

7040 0.04S25992 .04125*7 .03464672 .03684011 .03ss

530s O:g;:43613622 .c4eo2980 .04132408 .03721974 .0340

4e93 0.04864311 .04723621 .04453S80 .0412.2888 .03812350 .034t

5853 0.04984674 .04C94112 .0418SL=1 .0373.?209 .0s332194 .0s8

Crlglnal

1210119711s812Q71199

%863840630631

6320406356386286s3

:3

8306S46376W

15301526152615281527

.%’700697697697

767801796800796797

r72959658959956962

547648546560549

423425425427427426

,441441441433440441

4n4894874W4664’96

2678.2688298126782961

1614160.91611159018061604

160716031798182317951792

212420742020

20722093

17761783176117721771

1803lao31203179317981804

178417791785177a17841770

17e617711777176917761769

175517451760174s17W

177017701770177017e51770

176517601780.176017651780

176a17301750175017451745

72.3072.1212.7612.2472.08

36.9s38.9839.0438.75;::;;

43.2JI43.7643.E144.0043.5343.3’9

31.155U.615!2.71Y3.7e50.5730,64

0.0374.0321.0288.0249.Gr’15

0.0446,0463.0314.0309,0250.0205

0,0420.0s92.03:1.0306.0266.021h

0.0463.0379.0319.0283.0224.0171

30,000 0.9.9s.9s3.6’93,998.99s.992

6396338s36316326s4

1188116611881188lIW1188

497496497494494494

252425232556234423482538

1789177518071779179317.66

178a17701785177517751775

60.2120.7530.4460.6380.3060.80 L

4571 o::Y~43714052 .0370SEW .0337

.05422207 .0318

0.0303.0286,0268.0239.0216.0184

+“40,W0 0.579

.381,581.570.387.370.569

50,CKI0 0.643.620..643.650.600

40137397398393395395

2392422s923924E.

W3499499496495493492

313318316312312

425425425425425426425

4384334543419421

1149115411541144113011391143

710710711724720

183118s118051631163116U1837

1638162216221818102s

17801720177517W178017201700

17701770176517701763

27.7727.6728.1027.5327.47

m18.8717.1317.2417.3217.33 ~- &

4473 0.08603723 .05693284 .03262734 ,04e42312 ,04411641 ,0406155a .0366

2807 0.06782810 .08192261 .057616-21 .06161.272 .0482

0 .08s3.0E33,0470,0404,0343.0292.023S

0.0670.0622.053.8.04:7.0401

00nfiwntion B, 8tralghtwiln&”-”

.—

.

.

NACA RM E53F1O 23

.

ZT!RFORMANCERATAmust-*m~- =Xmust-AfteFb-mp~ed Aye;ted Net

nc=zl*- nor.zle tmWPer6AW=#ted Net treat 3n61n0 Afterburner HAxl*n.xl-Aftevtmmm

tkm.lst, apeolffo ambus- Ocdllsticm presmUw-

lnlet tdd arma , tit, thrust Fn, thrust rlml om- tkm ei’fici*ncs, burner 10BS

total taer- A1O * W% F,, ratio , lb Cullptlc’n, efflci-FZ’ (wf,.4wf,ab)

‘ab ah*ll ratio,

pressur*, ature ,Wq rt lb r~~

UIop ,?9, ~lo :C

teKUm&- P5-P*

ne

lb %‘n %4 —P5

gq ft abn

.

N

;

Cm&

3131

fig2574

3362S27431S329822870233a z8246 1.297

8106 1.2626013 1.23557a7 1.2Q65591 1.167

5552 1.3U1.317

3417 1.2243249 1.248s123 1.1813037 1.144

0.1034.10L8.1013.0S68.09C19k

267026.94287928882S77

1427142S14?,5141914441443

1s101206lea1853lels1622

S.042.922.892.822.72

3.22S.lo2.982.922.7S2.6S

1.7631.68s1.6171.54s1.4s3

1.8251..9161.7391.~631.4251.407

4776 1.427 2.412 0.s20 0.883 - 1s204617 1.379 2.264 .990 .916 ::;:4514 1.339 2.191 .930 .91s4314 1.297 2.109 .92a .2Q7 1KX34104 1.242 2.056 .920 .266 17P.

Silo 1.409 2.5m 0.990 0.763 2022x141 1.364 2.s22 .220 .647 19s12955 1.344 2.132 .92a .912 120s2826 1.226 1.992 .920 .961 17292663 1.219 1.s8 .990 .2C7 1222257a 1.174 1.eQ5 .990 .874 1549

0.1159.1120.1093.1076.1CQ9.m4

3Yd3030229927612544

0. low.1043.lLm7.1031.1014.0964

S.073.012.992.922.622.62

1.8141.7771.ee71.e301.54e1.430

4130 1.3264Q57 1.397s94e 1.2763918 1.2423772 1.2133Sd7 1.1s3

3341 I.43e 2.2e2 o.92a o.a74 19343278 1.410 2.21e .eeo .ae9 lam31ee 1.3ee ?.139 .s20 ..9e5 17e23152 1.S51 2.035 .eeo .eee lWO2974 1.287 1.964 .9eo .897 1750

1.219 1.978 .990 .947 1663

r1294125ele70le7s18721200

324835.9e2e23270124W23e6

S.lo3.012.es$8~

2.34

O:;J

.Iole

.099!.

.oee:.

.09;92

0.0979.09el.0939.09$1.oe39.0914

o.12e8.121s.U70.1110.u4e.1089.LC27

1 .e371.7441.6451.5271.402lxie

5137 1.3454932 1.3054201 1.2304601 1.2124s70 1.1s64s11 1.130

2277227S2319230223072308

275e23e32a372708263e24m

:::g2.752.222.652.5e

1.2s31.2131.5701.5211.4721.377

I.eee1.6061.7521.e241.6461.3.621.415

1.2021.8721.e571.7291.7W

+

3oe7 1.27360S2 1.25e.eoIa 1.2395662 1.2156791 1.1933522 1.153

2730 1.3es26ee 1.3272644 1.SQ2254s 1.275261a 1.2522408 1.21122ee 1.154

34723we31es310130182261

349e341133as31443104

3.353.223.103.042.952.852.72

8.3s3.313.223.10S.04

e12e12els6326S1 L

16e7 1.3e41224 1.%71224 1.33e1644 1.2871607 1.271

ue4 1.d97 2.949 o.9e7 0.224 18171S52 1.471 2.234 .9e7 .727 lam1348 1.423 2.38b .esa .780 1781131e 1.3e7 2.387 .eea .776 17171229 1 .38s 2.261 .2e.9 .e29 16e8

0.1377.1s81.13s1.1?70.1?44

1 v-ma and vortex sm-mratorsaddd

u3e3 3s30 3.26 I.e2e 427S 1.4111601

345e 1.3e3 2.094 0.93U 0.879 17ee 0.07763270 3.12 1.e7a 4203 1.3e6

181135ee

32221.327 2.e23 .920

3.12.726

1.841 4152171e .0716

1.3e61606 3152

3330 1.303 2.427 .920 .7e53.08 1.209 4073

17181.s48 3269 1.477 2.245

.0628

lele 2290 2.es.e90 .e66 1702 .0423

1.E47 33e7 1.27e ?L156 1.3e1 2.CZ?7 .e90 .% 16:,9 .Om:,

1002 3.23 I.82e 2s56 1.368Ices %%

214e 1.EC3 2.774 o.ee7 o.ec4 1661 o.07e23.17 1..617 2e32 1.354

1013 2ee61.421 2.399

3.03.987 .7e6 1650 .0737

1.e93 2530 1.297lo2e 26e9

%% 1.40s2.87

2.081 .e671.517 2ses

.8941.223

lffi9 .oe94lee5 1.301 1.948 .98e .a4e 1629 .oe22

6Q7 Sloa 3.24 1.760 1.311ml

1.33629.9e

127a 1.4e4 2.6M, o.9e4 o.621 16Q0 o..xen3.12 1.ee2 1372

e=1.2S7

26131239

2.97 1.4921.410 2.418 .966 .730

1466 I.al13J30 .0234

227 25711152 1.290 2.225

2.62.9e3 .eei 7J74

1.46e 1460 1.197 1123.0754

1.273 2.003 .9ee .776 1L3S .oe92

r●tmigh

eeel~ses27112717ee742e772es92722

3228 3.12 l.ale e454 1.350 4936 1.512 2.571 O:;S o.a253106 :.:: 1.765 6524 1.327

15214elo

o.06a91.4W

3M6 1.7162.452 .643 1s33

1.3oa 4705.0622

1.U22ee7

?.3242:91

.920 .873 15W .06451.e34 6oa3 1.271

273S4s41

2.821.39e 2.220 .990

1.558 3esl.e7.2

1.25e1640 .0621

26704313 1.34e

2.m2.129 ,eao

1.474 5e2e.eee

1.2Q01519 .0634

23224102 1.22e 2.009

2.58.a20 .eo2 L616 .05e4

1.331 3333 1.13-221S3

332e I.m 1.941 .960 .790 1494 .0s:02.49 1.223 5107 I.oel 320e 1.133 1.93n .e90 .633 142a .0s22

24 NACA RME53F1o

.

TABLE II* - Concluded..

15,0W

So,ooo

30,000

50>000

40,aoo

50,000

17s117X31749.;;::

17s9

176817681766176717701767

160117921758180812D6

178517651765176517051785

1765176S1770172517651765

176517651765176S1760

43.6243.4643.8943.8143.7345.74

27.7927.7727.6127.EO27.7127.97

17.0517.0716.“7417.0817.07

con:

1863 4431 0.06441665 3784 .05801’649 .OWIla32 %’; .043418s7 1607 .os5ala72 loea .0353

m

Lfterburnorl-ml-air

ratio ,

‘lb

mraticac,

CI:aa;

.043s,os3a.0233.013s

0.0631.0G31.0433.0514.0187

Canfixumt Ion D. mdlfied furl dimtrlbut lm

0.614 lla5

I

1527 63a I 2aa7 I 1737.61s ma 1s52 3.33

1780 73.sa 43a7 6021 0.0422 0.04162!347 1738

.614 lla71720

1531 53774.14 4464 Sam

2920 1741.0s9s .0301

17an u.67 4434 4612 .0S5J ,0241

con fimmaticm B. mdlC18d fml

40,000 0.5ea 3aa Soo 421 1092 1740 1730 2a.ol 182s s94a o.05a.3 0.0555.alo 38a 423 106S 1742.aol

1745 22.lasa4 50s

1817 3eao .05s7423 1087 1742 1745 28.15 1625

.0s14

.567 405 sol 4223sa4 .052s

1022 1742.0470

174s.573 401

g.: 1s23 3L26 .030sW1 422 1021 1741

.oa4a

.5241745 1825 2776

5a7 600 423 1020 2aI15.0465

1742.0327

1745 la2s 22a5 .0415 .0320

X!,000 O.6S2 237 310 432 a59 :;g,a23 243

1770 17.0-2S16

1171 2717 0.0649431 668

0.06381780 17.37

.626 2S6 3071151

4332275

660 1773.0s8’6

1770.0526

.62416.90

2S6U42 2030 .0535

507 432 6al 17a4,0462

1766 16.aa 1137 1488 .0441 .0352

30,000

40,MW

Conflzuratlon F, original with

o.sa2 834 604 443 1785 1763,690

1760 43.87am

2S76 64957aa 441

o.oall175a 1774

o.05a7

.6911753 45.44 2a7a

332 aoo5a55

440 1770.osao

1780.0516

.s921760 4a.a7 2all

6=4955

7a6 44s.Osla

1757 :;:;.0401

.590175s

6s345.23 2268 4s11 .047s

aol 438 1774 1735.osaa

.3S0 63643.74 2211

8043222

4s6 177a.Oa

177a.0224

1755 43.S2 2926 2243 .0533 .0204 -

0.520 395 m 417 1149 la12 1780 2a.07 1920 4073 0.081s o .05a7.593 5a4 Wo 415 l14a lalo.664

176Q 2a.oa 1980394

3475 ,05524aE 414 1140 1797

.05021770 26.01

.593lasa

3912634 .04aa

4a6 426 l13a 1601.0423

.W617ao

SW27.57

4951826 2406

429.04s5

1130 1616 1775 27 .S7.035s

laos 1901 .0s2a.590 3a5 600 414 11s1 lam 1775

.028222.10 1986 1466 ,0s46

.

..-

NACA BM E53I?1O 25

.

ALTZKJDE PI!31’ORMANCEDATA.

FnE

I 266227S22735

S406

Is.12

2657 2.95.?726 2.82

1.s621.7011.S67

6662 I 1.39.?6501 1.2945686 1.2*1

5127

I1.577

4781 1.4314446 1.s53

2.420 IO:;g2.1562.035 .990

0.274.954.648

161216a21669

0 :O&

.06s4

F1409141414211464

15911594162s16021625 --t-

3557 3.=!5236 3.222616 3.012s23 2.62

3546 S.503356 3.s13142 S.ls262S 2.2224!36 2.72

2.0s41.6S91.6661.3s9

2.0321.s311.2011.6141.407

+

3669 1.4243470 1.3533273 l,a7s2994 1.153

4416 1.4s44274 1.39140S6 1.3405799 1.261.mm 2..174

5474 1.447ML? 1.36s4953 1.s43

1.279M 1.216

---1S165299127652494

35873447321229.972754

1.6221.4s71.X51.1s9

1.5641.65s1.4271 .s571-239

1.6751.s991.5221.4191.S25 *

2.672 o.9m2.252 .9202.010 .9W1.8*6 .93-3

2.56S 0.990Q.381 .62o2.1S6 .s901.961 .9901.77a .9sa

2.619 0.9902.290 .96a2.166 .9601.964 .9M1.875 .92a

0.841.093.627.741

1616157015361469

17s916271560MM1420

1772

R1526

0.0286.0794.0700.0579

0.0256.0207.0726.0643.0669

o:%

.0707

.00ss

.0383

0.8S6.862.914.943.206

0 :80;

.914

.9&7

.201rlam18751=419101921 T3:68 S.60

3546 5.2231a S.lo2870 2.922614 2.75

2.0531.9241.72s1.6421.495 141964019

366936963S76

lm3lm7lmb10151027

3407 S.w3306 3.293300 3.1828s3 Z.sa2s93 2.75

1.SU1.6231.7611.62.61.359

2773 1.4m2722 1.s752563 1 .ml2461 1.2s8

l.lm.

2266 1.5s72216 1.m3=76 1.43319S2 1.3481754 1.2Q2

2.84U 0.9872.4S4 .9872.262 .9672.041 .9871.920 .907

0.7W.2.17.833.878.709

17S61520186015261423

0.0915.oa70.0622.0742.0647

L611612614613618618 --L

S324 3.466539 3.573229 5..263132 5.162814 2.902694 2.S9

1.9321.6861.81s1.7931 .s921.s25 w-u L

2.7aa 0.9672.719 .9672.564 .9672.354 .9652.24s .9672.002 .967

0.714.n7.745.614.736.788

1202lzm1693157115431513

O:;mo:

.09%

.0-s64

.0611

.0745

Efuel distrlbutlm may fr.m

1564 3362 S.281561 S587 3.281576 3231 3.161573 32CS s .041s93 2834 2.921606 2322 2.72

I@Ja 3.2=51014 :% 3.19lm2 3207 3.171012 2970 2.981010 2779 2.891031 2418 2.7S +

1.207 41691.002 U261.815 40s01.3n3 S9781.6m 37991.418 3561

1..ss4 27191.a12 26921.725 26311.649 24911.530 23831.341 2294

1.s581.8441.7911.76S1.8541.560

1.3391.S271.3061.2651.2051.123

3S89 I.mz 2.785 0.990 0.720 16123374

0.11391.&9& 2.520 .990 .816 1949

5286.1126

1.442 2.3.99 .s20 .6223162

y;: .10961.412 2.256 .260 .91s

3m4.1047

1.316 2.044 .990 .886 17222723

.2.0201.212 1.056 .s20 .878 mm .0961

2221 1.450. 2.716 0.987 0.7012193

1s3s 0.12Z?71.433 2.47P .987 .761 12Q4

2141.1176

1.40s 2.275 .987 .a63 18321993

.12111.s40 2.157 .s87 .242 1716

1894 1.27S.1OW

2.ma .s87 .8S6 16611796

.10621.163 1.909 .967 .716 1610 .1043.

-.

N03

zg

x=

C2-4 back 2937. 1

stat

(b) Sohematlo diagram of engine imstal.lationshowing thrust~bg6ywkm d Instmmentatlon stations.

Figure 1. - Concltidl. Turbojet engfne equipped with Mterbwner ad titalled h

e.ltltude dwnber.

NIn

1/1 I----G”+

I It’/%a.8r.tarI.Ocatilm

z(~~mB, C, D, d E) II

4 .5-t-n;

WA blcokd area

9.2a m. in.——_

(a] Flam-holds’ ad PIpeJ-hr I.Ocatloln.—.

(II) m

.—amAmlda dutalh

, L

LC6Z

NACA RM E53F1O

Afterhrner

(a) Spray pattern and spray hnr used fn conf@?.’atitxiA.I

.

.(b)Spraypattern arid spray ‘tar used in Cxdlguratlm.q B end C. -“

.M’ter-

burner

shell

Figure 3. - Fuel-distrlbutim pattirns and spray-bar designs.

NACA RM E53F1O

(c)Spraypatternand spraytar umed in configurationD.

.

.

Diffuser so1,Shell> %? A

? I ? ?o 0 0 0 0 0

(+ 25 “ 26,*?3 “ la!’

, 3? z E F>.

L “A

7 “

(d) Spraypatternand eprajbarused h ocmflguratlonE.=&$=

.

.

-----FiEUIW3. - CentInued. Fuel-ciistribu~ionpatternsand spray-bardesigns.

, 2937 , ,

Afterburner

fuel inlet

,,

Diffuser

“’”e ‘+$ “’’’”:4-A

~11●

16

(e)

F@re 3. -

Spray pettem and spray bar used In conf’lgu.ratkm F. ww-

Ccmcluded. l%e~-difrtributim pitterm and sgmay-bar designs.

ter-

umer

Flbell

32 NACA RM E53F1O

-4. - Turbine-outlet@m-s&ai@ten@3 P6, y-e= g~et~s~ @ ~t~manifoldsfar configurationsB, C, D, and E, as viewed h downstreamdireotion.

.

.

.

4 CZ-5 2937

Figure 5. - Cemmlc-cmted ~ted

alreotion.limsr installed In afterburner for cani’iguatim C, D, WI E, as viewed in upstream

0/m

NACA RM E53F1O

●ox

Static-pressuretube(wallorrake)Total-pressuretubeThermocouple

.—

——

.

I(a)EnSineInlet,station1,21 Inches upstream ofleading edge of compre8mor-inletguide vanen.

o00

(b) Diffuseroutlet,station6, 3713Inohemdome tream of turb lne -outlet La..

—

.- —

(c)AfterburnerouterBhell, ntatlonS,74 Inohemdommtreamof turbine-outlet flange.

(d) Nozzle Inlet, mtatlon 9,&2m-mtreain of turbine-outleql%~s

FIwe 6. - Instrumentation,

NACA RM E53I?1O 35

.

o

-20

-40

ConfiguratkmA PercentratedConfigurationB(withoutvanes) enginespeed (with-es)

● 100 095.2 ❑

I

Maxtiumangleof rotation-1

(a) Verylngenglne speed;altitude,30,000feet.

Conf@urationAA1.tltude, ConfigurationB

20,(wlthoutvanes) f% (withvanes)

A 20,000● 30,000

40,000 ~50,000

0 I

oA

* 3

‘1

-20+A

m

g\Maximum angle of rotation-

k for rake

1 I-40 ~E

o 2 4 6 8 10Passagedepth, in.

(b)Vaqingaltitude; 100-peroentrated enginespeed.

IH.gure7. - Effect of enginespeedand altitudeon turbine-outletgaswhirl at flightMach number of 0.6.

.

.0

1800

=@=Presme Configuration ‘

Static Total

-+- ~ A (without vaneE) ,

‘~- ~ B (with vanes)

r-li-ig

ho-50

I

4! ,, !J-. I \t

4‘~\ \

..

1I\

0 2 4 6 8 10

Paasage dapth, in.

Flg.nw 8. - Effect of turbine-outlet straightening vanes and vortex generatoron diffuser-outlet total-pressure profile. Altituda, 30,000 feet; flight

Mach number, 0.6.

I!, 1,!,

NACA RM E53F1O

.

37

.16

.12

kmI d-’

.08

.04

Configuration

[

O A withoutvanes).A B tith vanes)

Net reduction press ure

loss due b vanes

— —- --c) -- --

_- -- - “7 Pressureloss throughturbine-outlet vanes

--- --

L ~4

~ ~ -

d

(a) Altitude,30,W feet.

INet reductionpressure

.12 loss due to vanes

>*”* “

_-

J --- --

.08-- EYessureloss through

turbine-outletvanes_- -- ~ -

-- --+~ -

~ ~~ ~

-.04-

0 .01 .02 .03 .04 .05 . Oe .07

Afterburnerfuel-airratio, fab

(b) Altitude,30,000 feet.

Figure9 - Effect of turbi”ne-outletgae-straighteningvanes,vurtex generators,andcorrugatedliner on afterburnerpressureloss. Flight Mach number,0.6.

38 NACA RM E53I?1O

ConfiguratIon Average calculatedturbine-outlet gas temperature,%

o A (withoutvanes) 1825A B (with vanes) 1754

f’tl

501000 c)

A

Configuration Average calculatedturbine-outlet gas temperature,‘R

o A (withoutvanes) 1782A B (with vanes) 1745

I

30,000

.02 .03 .04 .05 .06 .07

.

AfterburnerfUf31-atiratio;”fab —

(a) Afterburnercombustionefficiency(equivalentradialfueldistribution). “ —-

F@ure 10. - Effectof turbine-outletstraighteningvanesand vortexgener-ators. FlightMach number,0.6.

~. -.

, , b

,?937,

Cmflguration Avemga calculated _outlet gas temperature, %

O A (tithout vanes) 1825

.02 .03 .04 .05 .06 ,(

conflfJlmstlon

l-m -

Avers@ calonlated toxbine -

eutlet gas temperature

PA (without vanee )

AB (with mmes) 174s

A/ -

/&Altitude, 30,000 f%

P. ~/

/

4 ~

d

I

1 .02 .03 .04 .0S ,0: .07Afterburner fuel-alr ratio, fab

(b) =uet-gaa total temperature (equivalent radial

m 10. - Cmtinlld, ~~rector turbine-outlet wtraighteming vaneE and

number, O.6,

fuel diWrlbution) .

vortex generators. Flight Mach wco

40 NACA RM E53T10

.

1.4

Altitude,

1.3 - ft50,000

1.2

Configuration _.-’n!

Average calculatedturbine-

-El outlet gas temperature,‘R\?%V A (without -es) 1825

1.1, 2 B (@th vanes)0’$24 1.sm

E++%@al 1.4

;

a A

2

/30,mo

1.3 - /

d

Configuration1.2

A (withoutvanes) 1782B (withvanes) 1745

I

=%=1.1

.02 .03 .04 .05 .06 ,07

1754

/

Average calculatedturbine-outlet gas temperate, ‘R

Afterburner fuel-air ratio, ‘ab

(c)Au@ented jet tkrust ratio (equivalentradial fueldletrlbution).

Figure 10. - Continued. Effect of turbine-outletstraighteningvanes andvortexgenerator. FlightMaoh number, 0.6.

.

.

NACA RM E53I’1O 41

2.5

.

2.3

2.1

1.9

2.5

2.3

2.1

1.9

Altitude,_ft

50,0CXI

Configuration Average calculatedturbine-outlet gas temperature,‘R

O A (withoutvanes) 1825AB (with vanes) 1754

/

A

/

d

30,003Configuration Average calculatedturbine-

A outlet gas temperature,‘RO A (withoutvanes) 1782A B (withvanes) 1745

.02 .03 .04 .05 .06 .07Afterburnerfuel-airratio, fab

(d)Net thrust specificfuel consumption(equivalentradialfuel distribution).

Figure 10. - Concluded. Effect of turb3ne-outletstraighteningvanes andvortex generators. Flight Mach number, 0.6.

42

1.0

SJ .4F

.8

.6

A

NACA RM E53F1O

u

v ~Altitude,

50,0Q0

mConfiguration Fuel distribution

c OriginalD“ Modified

Q E Modified, with uni-orifi.ce spray bar

40,000

/\Q .

u-0

n

30,000

=$=, 1 1 1 I I I I I I 1

. .

.

N

8?4

.

—

“:01 .02 .03 .04 .05 .06 .07 -.”Afterburner fUel-air ra~iOj fab

(a) Afterburner combustion efficiency. “

Figure 11. - Effect of fuel distribution. Flight Mach number, 0.6.

a 1 ,

Afterburner fuel-air ratio, fab

(b) Ezhaud-gas total temperature.

IMgUre U. - continued. Effeot of fuel distrlbutd.au. FIQht kch number, 0.6.

IF.CN

44 NACA IM E53F1O

1.6

Altitude,

1.450,000

A.-

1.2Configuration Fuel distribution

/n c Original

m o D Modifiedo E Modified, with uni-

1.0orificespraybar

1.64>

; Ill I 40!bQoH-H-t-t“-4rttt2Rll!lllll+i-P

/ /

1.2

c

1.0

1.6 -v

n 3. -u

30,000

1.4 r

A

1.2

T

1.0.01 .02 .03 .C4 .05 .06 .07

Afterburnerfuel.-airratio, fab

(c) Augmented net thrust ratio.

Figure 11. - Continued. Effect of fuel dlstributicm. Flight Mach number, 0.6.

.

.

.

——

.—

.

.

-.

NACA FM E53F1O 45

.

.

3.2

Q

I

2.8 --’.

Altitude,

2.4

&Configuration Fuel distribution

/ ~nc

2.0 /

8

OriginalHtilfied

: Modified,with unl-orificespraybar

uma: t

1.6 -.~~52 -.

$.

s

2t=k

2:0CL 2.4 --—e~c0udU: 2.0 -“””–-—-- -,

0z —.-——.-lualg

1.6 -— “-’----—-Ums: 3.2 -— –- “---“”-

U$ {

2.8 ~— -

..+..—- . / fl

j

/

2.4 ---- -— ----

30,(!00

2.0 ----

~ / ~ y.. .—.-— —

,.:;,..4- .02T

.03 .04 .05 .06 .07Afterburnerfuel-alrratio,fab

(d) Net thrustspecificfuel consumption.

Figure11. - Concluded. Effect of fuel distribution.FllghtMach number,0.6.

46 NACA RM E53F1O

.

2000

1800

1600

1400

1200

Configuration

O A (originalconfiguration)‘A B (with straightening vanes)Q C (straightening vanes and liner)O F (fueldistribution.away from shell)

o/~

0/ /

(

n n 4

Am

.

.—

—

(a) Altitude,50,0C?0feet..

2000

/1800 #s

o

1600

0 ~ -~

T1400

2400 2600 2800 3000 L- 3200 3400 36ti”

Exhaust-gastotal temPerature~TIO~ ‘R

.

(b) Altitude,30,000 feet.

Figure 12. - Effect of turbine-outletgas-straighteningvanes and vortex .

generators,ceramic-coatedCorrugatedliner, and fuel distributionon ‘a-fterburner shell temperature. Flight Mach number,0.6.

, I 9

2(X)O

18CQ

MOO

1403

0

~

n

o 14

mu

Configwation

Ct C (with vanes and line~)

~ ~ ‘

n- ~/

o

‘J

na

) 1600 22C0 26C0 3W0 34CKI

:bine-outlet total pressure, P5, lb/aq ft

Figwe 13. - Effect of turbine-outlet total pressure on afterburner shell tem~ratore

for @s temperature of 31C@ R.

48 NACA RM E53F10

1.0

00

/ *

.8 A //

\

A \A A

.6

.4

(a) Afterburnercombustionefficiency.

36C0

/ g

34CKI o /

3200 /

3030

y /, / Altitu&e, Turbine-outlet Turbine-outletft ~tal pressure, totaltemperature,

P5, lb/sqft T5, ‘R

o 15,003 2950 1739

280CI❑ 30,CO0 172Q 1745

2 :’EO11O3 1754

) 674 1764

1-

Zw

4/

2400 v

-22W

..

-,-..-..

, ..

,.

.

.02 .03 .04 .05 .06- .07 ._=

Afterburnerfusl-airratio,f.gb

(b) Exhaust-gas total temperature.*

Fig-me14.- Effectof-altitudeon perfonne.nceof configura~onD. PlightMach number,0.6.

.—

-

NACA RM E53TIO 49

.

. .g“

1.6

1.4

1.2

2.8

2.4

2.0

1.6

Altitude;ft &bine-outlet Turbine-outlettotal pressure, total temperatureP5, lb/sq ft ~, ‘R

1.5,000 2950 1739: 30,000 1720 1745o 40,(MI 1100 1754A 50,000 674 1764

0

(c) Awnted thrust ratio.

.02 .-03 .04 .05 .06 .07Afterblxner fuel-air ratio, fab

(d) Net thrustspecificfuel consumption.

Figure 14. -.Concluded. Effectof altitudeon performanceof configura-tion D. FlightMach number,0.6.

NACA RME53F1O

1.0

k

.8

c

.6

.4

(a)Afterburnercombustionefficiency,

36000

A @

3400/

c‘

32CKI

3000

Mach Turbine-outlet Turbine-outletnumber totalpressure,totaltemperature,

22C0

o 0.4 ““ 1535 1745”

1746

261XI

2400

2203

.02 .03 .04 .05 .06Afterburnerfuel-airratio,fab

(b)Exhaust-gastotal temperature.

.

—

.

-.

-- .....

—.,

Figure 15. - Effectof flightMachnumberon performance of confi~ration D.Altitude, 30,000 feet. .

a-m!!?

NACA RM E53I?1O

.

51

Mo

$1-

&v

.

.

●

w-l---l““:i~~’~“::~”e’1-lnumber total pressure, total temperature,

o 0.4 1535 1745

1.8

1.6L A .

1.4-

1.2~ /

c

1.0

(c)Augmentednet tbruatratio.

2.8

2.4

2.0

c

1.6 ..02 .03 .04 .05 .06

Afterburnerfuel-airratio,‘ab

(d) Net thrust specifie fuel .onatunption.

Figure15. - Concluded. Effect of flightMach numberon performanceofconfigurationD. Altitude,30,000feet.

mIN

&—-~ Aft&rn& -+

fuel-air ratio,

fab / -X I –l..-L---+--”l--l

.90.035 /

/

.8 /#

0

//

!

.7

. I

600 1000 1400 18~ 22al 2600 3CQ0Turbine-outlet total pressure, P5, lb/sq ft

Figure 16. - Effect of turbine-outlet total pressure on combustion efflc~ency

for varioua afterburner fuel-air ratios for configurations D and E.

,.K6z

!2

II!

NACA RM E53I’1O 53

.

.

b

.

.

7400

7000

6600

6200

5800

3400

3000

2600

2200

1800

1400

Altitude, Machft number

“!

15,000 0.6 P50,000 .4 /30,000 .630,000 .8

d,40,000 .6 (50,000 .6

/

//

L

{

/

‘nk

u- 4600

$

/

@4200z

Bj /x 3800 /z I3

/

/A

//

r

)I

400 60Q 800 1000 1200 14CW 1600 1800 2000 2200 2400Nozzle pressure-drop parameter, 1.25 Pg - Po, lb/acIft

(a) Generalized jet thrust.

Igure 17. - Generalization of afterburnlng jet thrust for configuration D.F

1.COI I

A1O,m73.5 sq rt\

.96\

\

.92\

s.

Q“ \

Wi.

2 .88+dh*

!.84

,,,

.80

.76 I

1.03 .96 .92 .86 .64 .60 .76 .72fiea ratio, A1~A1o,m

(b) Comection fact.or for jet thrust obtained fmm figure 17(a).

F&me 17. - Ccmtinued. km’al.izaticm af afterburning @t thrwt for configw’ation D.

I

, . # *

,, ,,, ,!,, L!2~z,,

NACA RM E53I?1O

.

55

.

.

.

3.6

3.4

3.2

3.0

2.8

2.6

Altitude, Flight Machft number

o 15,000 0.6❑ 30,000 .40 30,000 .6A 30,~0 .8A 40,000 .6

—Maximum exhaust-nozzle area

{

n

L1

.02 .03 .04 .05 .06 .07Afterburner fuel-air ratio, fab

(c) Exhaust-nozzle area for various flight conditions.

Figure 17. - Concluded. Generalization of afterburning jet thrustfor configuration D.

NACA-L@.U@eY- 10-29-5S-325 %