[american institute of aeronautics and astronautics 36th aiaa/asme/sae/asee joint propulsion...

(c)2000 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization.

&AIAA AOO-36389

AIAA 2000-3119Liquid Methane/Oxygen Injector Study for PotentialFuture Mars Ascent Engines

Huu P. TrinhNASA Marshall Space Flight CenterHuntsville, AL

36th AIAA/ASME/SAE/ASEE Joint PropulsionConference and Exhibit

17-19 July 2000Huntsville, Alabama

For permission to copy or to republish, contact the American Institute of Aeronautics antistronautics,i«m A. lavonHot. Ti^ii HI-IVO «nUo cnnDoctnn VA


AIAA2000-3119

LIQUID METHANE/OXYGEN INJECTOR STUDY FORPOTENTIAL FUTURE MARS ASCENT

Huu P. TrinhNational Aeronautics and Space Administration

Marshall Space Flight Center, Huntsville, Alabama 35812

ABSTRACT

The design of high performance, lightweight injectors fora liquid methane/oxygen system was investigated. In thisstudy, four configurations of unlike and split-tripletimpingement injectors were tested. A total of 40 hot-firings were conducted with chamber pressure andmixture ratio ranging from 343 to 513 psia and 1.2 to3.5, respectively. Results show that combustionefficiencies of the injectors are similar for the testconditions considered. The efficiency was found toincrease with decreasing the mixture ratio. Nocombustion instability was observed. Furthermore, thecombustion ignitability of the subject propellant systemis dependent on initial mixture ratio and chamberpressure.

INTRODUCTION

Mission studies for human exploration of Mars1 havebeen performed at Marshall Space Flight Center(MSFC). These studies indicate that for chemical rocketsonly a cryogenic propulsion system would provide highenough performance to be considered for a Mars ascentvehicle. Although the mission is possible with Earth-supplied propellants for this vehicle, utilization of in-situpropellants is highly attractive. This option wouldsignificantly reduce the overall mass of launch vehicles.Consequently, the cost of the mission would be greatlyreduced because the number and size of the Earth launchvehicle(s) needed for the mission would decrease.NASA/Johnson Space Center has initiated severalconcept studies2 of in-situ propellant production plants.Liquid oxygen (LOX) is the primary candidate for an in-situ oxidizer. In-situ fuel candidates include methane(CH»), ethylene (C2H4), and methanol (CH3OH).

Copyright © 2000 by the American Institute of Aeronautics andAstronautics, Inc. No copyright is asserted in the United States underTitle 17, U.S. Code. The U.S. Government has a royalty-free licenseto exercise all rights under the copyright claimed herein forGovernmental Purposes. All other rights are reserved by thecopyright owner.

MSFC initiated a technology development program for acryogenic propulsion system for the Mars humanexploration mission. Along with investigations oflightweight materials for propulsion structures,propellant tanks, and main chamber, this program hasalso been setup to evaluate propellant injection conceptsfor the LOX/liquid methane system. This effortconcentrated on lightweight, high efficiency, reliability,and thermal compatibility. Such an injector technologywill have a potential use in the future Mars Ascentengines (MAE).

A considerable number of investigators have examinedthe LOX/CHt system for rocket engine applications.This propellant combination has been successfullydemonstrated3'4. The material compatibility, fuel cooling,and combustion instability issues5'6'7 of the CHVLOXsystem have also been investigated. Overall, the resultsshowed a positive impact of using such propellants forrocket engines. It should be noted, however, that most ofthese studies were aimed at applying the propellantsystem to earth-to-orbit boost engines. The advantagesof CHt/LOX are cleaner exhaust products, lowerpropellant costs, and higher performance compared withcurrent propellant systems. Hence, the operatingconditions for these investigations are at high chamberpressures, where CH4 and LOX are in a supercriticalpressure state. Therefore, the data may not be directlyrelevant to the current work since the intended operatingpressure is below the critical pressure of both LOX andca,.This paper will address the results of the liquidmethane/LOX injector study conducted at MSFC. A totalof four impinging injector configurations were testedunder combustion conditions in a modular combustor testarticle (MCTA), equipped with optically accessiblewindows. A series of forty hot-fire tests, which covered awide range of engine operating conditions with thechamber pressure ranging from 343 to 513 psia and themixture ratio from 1.2 to 3.5, were performed. Theinjector performance has been assessed from the testdata, which were collected from a number ofmeasurement techniques: optical diagnostic, flowvisualization, and high-frequency measurements.


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INJECTOR DESIGN

Engine BaselineTechnology developed from this project will be utilizedto develop a flight demonstrated cryogenic propulsionsystem that will meet the requirements of the plannedMars sample return (MSR) mission8. The originalintention was to offer a mission option that incorporatesin-situ propellant production and utilization for theascent stage. A propulsion system study9 conducted atMSFC has shown that a pressure-fed system is suitablefor a cryogenic ascent stage for the MSR mission. Thissystem was selected based on several factors includingweight minimization, packaging efficiency, andoperational simplicity. While a regulated pressurizationsystem is planned for the first-stage of the Mars AscentVehicle, a blow-down pressurization system is baselinedfor the second-stage.

Table 1 summarizes key MAE operating conditionsderived from the MSFC system study. The injectortechnology program uses the data shown in the thirdcolumn as baseline design conditions. The last columnof table 1 shows a range of operating conditions for theblow-down situation. Both the LOX (Pcriticai=731 psia)and liquid methane (Pcrincai=668 psia) are in thesubcritical pressure regime throughout the expectedrange of chamber pressures.

and oxidizer orifice sizes. This arrangement is suitablefor the MAE since the baseline O/F mixture ratio is 3.

An unlike-doublet (F-O) was selected as an alternativeconfiguration. At first glance, this configurationresembles the split-triplet, if two unlike-doublet elementsare arranged on the injector face with a back-to-backposition, (F-O) and (O-F). However, this unlike-doubletgrouping may result in an oxygen-rich region betweenthe two injection elements. To avoid this situation, theelements, as shown in Figure 2, are oriented on theinjector face as (F-O) and (F-O).

LOX LOX

Figure 1: Split Triplet Injector (F-OO-F)(a) Injector Face, (b) Injection Element

Figure 2: Unlike doublet Injector (F-O-F-O)(a) Injector Face, (b) Injection Element

In this study, full-size injectors (2.4-inches in diameter)have been designed and fabricated. Injectors weredesigned to achieve the objectives of lightweight andhigh performance while ensuring thermal compatibilityand combustion stability. In order to optimize the

Table 1: MAE Operating ConditionsParameter

Chamber PressureMixture RatioThrustVacuum Specific ImpulseExit/Throat Area RatioMass Flow Rate

UnitPsia

LbfSecond

Lbm/sec

Baseline2503

6003461001.86

Range

100 - 550(N/A)

300-1000(N/A)(N/A)

0.86 - 2.87

Injection ConceptsJet impinging injectors have commonly been used inLOX/liquid hydrocarbon rocket engines. For smallrocket engines, triplet and unlike-impingingconfigurations are widely employed. Typically theyprovide higher performance than other configurationssuch as like-on-like impingement and shower head. Inthis study, a split-triplet (F-O-O-F) arrangement, whichwas introduced by Pavli10, has been selected as thebaseline. This configuration, as shown in Figure 1, issimilar to a conventional triplet (F-O-F) impinger; theonly difference is an additional oxidizer orifice on thesplit-triplet for reducing the disparity between the fuel

manifold system and assure a uniform mass distribution,the injection elements are equally distributed in thecircumferential direction. For the split-triplet, only asingle ring of injection elements fit into the injector. Onthe other hand, two rings of unlike impinging elements fitin the injector face. In all cases the outermost ring oforifices injects fuel in order to prevent an oxygen-richenvironment near the chamber walls.

A total of ten versions of the two basic injectorconfigurations were designed and fabricated. Variationsin the impinging angle, orifice size, and injection elementarrangement are utilized in these injector face designs.For the unlike doublet injectors (F-O-F-O), an oxygen-rich condition may exist at the centerline region of the


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injector face. To minimize this condition, the fuelorifices on the inner element ring are twisted at an anglewith respect to the radial direction (figure 2). Thisorientation permits fuel to penetrate into the core regionto provide additional propellant mixing. In the test seriespresented in this paper, only four injectors consisting oftwo unlike and two split-triplet configurations were

of the combustion chamber can be altered by insertingblank modules of various lengths.

The injector assembly consists of an annular injectionring, an acoustic cavity tuning block, and a main injector.Nitrogen (GN2) will be introduced as a film coolantalong the chamber wall through a series of small orifices

Table 2: Injector ConfigurationsInj.No.

1234

Injector Type

UNLIKE DOUBLETUNLIKE DOUBLETSPLIT TRIPLETSPLIT TRIPLET

Config.ID. No.

FOFO#2FOFO#6FOOF#1FOOF#10

#Elem.

30422416

Orifice Dia.LOX

(in)0.050.0420.0420.046

CH4

(in)0.0350.030.030.034

TwistedAngle

(deg)0

43(N/A)(N/A)

Outer RingImping.Angle(deg)28.428.430

28.5

Inner RingImping.Angle(deg)21.826.430

28.5

selected for testing. Table 2 shows the configurations ofthese four injectors.

Combustion Instability ConsiderationInjector orifices were sized primarily based oncombustion stability considerations. To size fuel orifices,Hewitt's correlation11 has been employed to scaleinjector configurations from Pavli's data10. Oxidizerorifice sizes were selected to be similar to the fuelorifices, while keeping their momentum ratio very closeto the values at optimum mixing. In addition, severalacoustic cavity tuning blocks have been designed toaccommodate two quarter-wave acoustic cavity sizes,0.5" and 1" in depth. These tuning blocks were used tocharacterize the combustion stability behavior of thecandidate injectors.

TEST HARDWARE AND FACILITY

Test ArticleThe test hardware includes the MSFC ModularCombustion Test Article (MCTA), and the injectorhardware built specifically for this program. TheMCTA, as shown in figure 3, is a "workhorse"combustion chamber. It is composed of several modules,which are held together by four high-strength tie rods.The copper throat section contains drilled passages forcounter-flowing cooling water. The cylindrical portionof the 4-inch diameter chamber consists of severalmodules, including a window module and anigniter/instrumentation module. The window moduleprovides optical access to the chamber for photographingthe flow field or for performing non-intrusive opticaldiagnostics. There are several ports in theigniter/instrumentation module, one of which is used fora conventional gaseous hydrogen/gaseous oxygen torchigniter system. Pressure transducers and thermocouplesare installed into other ports in this module. The length

in the annular injection ring. A space between theannular injection ring and the main injector will serve asan acoustic cavity to enhance combustion stability.Different cavity configurations can be tested by changingtuning blocks. The main injector issues liquid methaneand liquid oxygen into the combustion chamber. Theinjector face, which is made of copper, is brazed to amanifold that has several concentric channels todistribute the propellants.

Test Facility

GN2IWECTOR RING FASTENING RODS

INSTRUMENTATIONMODULE

COOLANT MANIFOLD(NOZZLE PLATE!

THROAT MODULE

INJECTOR MANIFOLD TUNING BLOCK

Figure 3: MCTA with an Injector AssemblyPropellants can be provided at a variety of conditions atTest Stand 115. For the MAE program, LOX wassupplied from a 500 gallon run tank at pressures up to3000 psi. Liquid methane was stored in a 2200 gallonvacuum-jacketed vessel. A foam-insulated fuel run tankhaving a capacity of 20 gallons was used for this testprogram. Should the long test duration be required, anexisting 500 gallon, 3000 psi tank can be utilized.High pressure gases are available on the test stand forpurging and pressurization (helium and nitrogen) and forigniter propellants (oxygen and hydrogen). De-ionizedwater for cooling the MCTA throat section is suppliedfrom a 500 gallon, 3000 psi tank.


AIAA2000-3119

RESULTS

Assessment of Injector Configurations and HardwareA total of forty tests with a cumulative run time of 260seconds at main-stage were performed in this program. Aphotograph of a typical hot-fire test is shown in figure 4.As mentioned previously, the full-scale MAE injectorshaving a dimension of 2.4 inches in diameter were testedin the existing 4-inch diameter MCTA chamber. Theacoustic cavity ring and the annular injection ring ofGN2 film cooling were inserted in a gap between themain injector and the chamber wall. Strong re-circulatinghot-gas flows existed in the gap region. That causeddamage to the GN2 and acoustic cavity ring surfaces.Cumulative overheated damage for the first thirty-onetests is depicted in figure 5. A majority of theoverheating on the ring surfaces occurred in the first fewtests. A considerable amount of the GN2 film coolingwas required in the later tests to eliminate additionaldamage.

Post-test photographs of the four injector faces areshown in figure 6. Some normal discolorization wasspotted on the injector surfaces due to the minor sootdeposition and Lox-rich conditions; however, nooverheating indication was observed. The unlikeimpinging injection scheme was used in injector # 1,figure 6.a, with the elements in the inner ring impingingalong the radial direction. Since the Lox orifices are inthe innermost row, Lox-rich conditions were expected inthe core at the near injector surface region. The brightcolor in the center area of the injector, as shown in figure6.a, was an indication of these Lox-rich conditions.Injector # 2 was designed with the inner ring elementsimpinging at a twisted angle of 43 degrees with respectto the radial direction. Hence, fuel was able to entrain tothe core region with this arrangement. It should be notedthat the bright color no longer appeared on injector face# 2, as depicted figure 6.b.

Injectors # 3 and 4, as shown in figures 6.c and 6.d, have24 and 16 split-triplet impinging elements, respectively.Oxidizer was injected at a straight outward directionfrom the two center orifice rings; while the fuel jets fromthe outermost and innermost orifice rings impinged onthe oxidizer jets at the angles described in table 2. Thetwo injectors were designed to deliver the same massflow rate. Hence, the propellant was injected at a highermomentum per injection element in injector # 4 ascompared to injector #3. Consequently, the hot-gas re-circulation, as described previously, was quite strong. Itshould be noted that the hot-fire tests with injector # 4had more damage on the GN2 ring and acoustic cavityring surfaces than with the other three injectors. The last

test of the series was run with injector # 4 for 10 secondsat main stage. Some overheating on the MCTA windowregion was observed on this test, although there was nodamages to the MCTA chamber as on the previous tests.

Main Chamber IgnitionFor a given chamber pressure condition, theLox/methane ignition may take place at a certain mixtureratio level. In his study of the fuel-rich preburner for theLox/natural gas system, Bailey3 spent considerable effortin lighting the preburner chamber. He concluded that theignition of the Lox/methane system would require a highpropellant mixture ratio for low chamber pressureconditions. The conditions are described in figure 73.One-dimensional equilibrium combustion temperatureand experimental oxygen-methane minimum pressureignition limits as functions of mixture ratio are shown inthis plot. The ignition pressure curve is drawn from onlythree data points of ref. 12. Although the conditions fromwhich the data was obtained were not representative ofthe rocket chamber environments, nevertheless, thetrends should be applicable to the start-up ignitionprocess.

In this program, the tests were designed to run at fuel-rich conditions during the startup transient. First, themain fuel valve was opened; then a portion of Lox masswas introduced into the chamber approximately 0.2seconds later. The full amount of Lox was injected afterthe chamber pressure rose to a prescribed value. Since arange of mixture ratio was covered in the test matrix, thepropellant feed-line system and the test sequence weresetup to provide similar chamber flow conditions for alltests at the start-up transient. The first nine tests of thetest matrix were dedicated to characterize thehydrogen/oxygen torch igniter and the main chamberignition process. Main chamber ignition did not occur intest # 8, in which the initial mixture ratio and thechamber pressure were approximately 0.88 and 29 psia,respectively. The chamber was lit in test # 9 with a smallincrease in the mixture ratio and the chamber pressure bythe means of adding more propellant mass into thechamber. The mixture ratio and the chamber pressure atthe ignition conditions for tests # 8, 9, and otherrepresentative tests were also plotted in figure 7.

Optical MeasurementsBy taking advantage of the MCTA chamber equippedwith the optically accessible windows, Raman scatteringsignals of the combusting species and images of thechamber flow field were recorded through the use of alaser diagnostic system, infrared camera, and a high-speed film camera. These efforts were to demonstrate theapplicability of the measurement techniques for the


AIAA2000-3119

Lox/methane combustion system as well as to provide aqualitative assessment of the combustion flow field.

The Raman measurement techniques have been used forother propellant systems, such as hydrogen/oxygen, butthey have not ever been applied to high-pressurehydrocarbon combustion due to the signal interference ofthe species present in this combustion system. Recentdevelopment efforts13'14'15 of a Raman measurementmethod for hydrocarbon combustion has shown that thereis a strong feasibility of applying such a technique foractual rocket engine environments. Detailed discussionand results of the measurement demonstration from thisproject was reported in ref. 16. A short description ofthis method will be presented here for completeness. Forthis project, a tunable KrF excimer laser produced pulsedultraviolet light that was sent through the chamberwindows to the combustion products. Two simultaneousUV Raman images, vertically polarized and horizontallypolarized signals, were recorded by an imagingspectrograph through another window. The spectrographis equipped with a calcite crystal that separates theRaman and fluorescence signals into two polarization-resolved images. The difference between these imagesprovided the net Raman signal, which is then free ofunwanted broadband fluorescence interference caused byhydrocarbon species excited by the laser. The signalintensity images of the GN2 purge just before the firecommand and of the combustion flow field at the mainstage of Test # 25 are presented in figure 8. The spatiallocation is shown as the vertical component of the plot;while the wavelengths of the signals, to identify thespecies, are the horizontal components. The data weretaken along ±4 mm from the chamber centerline. Theupper and lower parts of the plot are the horizontally andvertically polarized signals, respectively. When thesignals are calibrated, the color intensity after subtractingthe upper part from the lower part will represent thespecies concentration. Nevertheless, figure 8.b shows aspectrum of the species presented in the flow field.Furthermore, the images also indicate that the specieswere well mixed in the measured region.

For the infrared imaging measurement, a Pulnix CCDcamera was set up to detect the infrared signals throughthe chamber window. The images were then recorded ona JVC 1-inch video recorder at a rate of 30frames/second. The intention of this effort was tomeasure the injector face temperature during the hot-firetest. The system had been successfully demonstrated withthe hydrogen/oxygen combustion conditions17, and it wascalibrated to yield temperature in the range from 930 to1560°F. However, the camera did not successfully viewthe injector face because of the carbon present in theoxygen/methane reactive flow. Instead, the images have

revealed considerable details of the flow field, as well ascapturing the hot spots on the GN2 film-cooling andacoustic cavity ring surfaces, as shown in figure 9. Theinfrared image in this figure was taken during the mainstage of test #21 . False color has been added to theimage to enhance contrast. Black being the coolest, thenblue, purple, red and white is the hottest. Note the hotspot (white), on the lower right side of figure 9. This isthe GN2 filmcooling and acoustic cavity region that hadbeen overheated. Finally, a high-speed film at a rate of500 frames/second was used to record the flow field atthe near-injector face. The objective for this effort wasflow visualization.

Combustion Efficiency:As previously mentioned, a considerable amount of GN2film cooling, from 40% to 60% of the total mass flowrate, was used in the hot-fire tests. This resulted in alarge variation in estimated values of combustionefficiency. Theoretical C* used in the combustionefficiency prediction was calculated from two separateconditions as follows: In the first case, the GN2 filmcooling was assumedly mixed completely with the hotgas core. In the second case, the GN2 was unmixed withthe hot gas core in. The corresponding efficiency valuesare reported in figure 10. The actual C* efficiency couldbe anywhere between the two cases.

As shown in the plot, the C* efficiency at the mixtureratio below 2.0 has a value greater than 100% even forthe GN2 well-mixed assumption. The values are inquestion due to several factors. The efficiencycalculation had not considered the Rayleigh effects andthe nozzle discharge loss; the chamber pressure wasmeasured 6.0 inches downstream of the injector face,while the chamber length is 13 inches. Furthermore, thetheoretical C* was calculated based on the tank enthalpyconditions rather than the injector manifold conditions.To resolve these issues, assumptions in calculating C*efficiency were examined. For the Rayleigh and nozzledischarge losses, data of test # 25 and 30 were randomlyselected for the study. It should be noted that thecontraction area ratio of the MCTA and the chamberMach number are approximately 8.9 and 0.07,respectively. Consequently, the results showed that theseeffects contribute only up to 0.33% of the difference inthe C* efficiency value. The same test data were alsoused to evaluate the inlet propellant enthalpy variations.By using the manifold conditions, only a slight a changein the C* value is observed. In conclusion, the high C*efficiency value may be attributed to the mass flow ratemeasurements. Nevertheless, the results showed that thecombustion efficiency value is reduced at an increase ofthe propellant mixture ratio. The combustion efficiencyvalue also has the same trend when it is plotted against


AIAA2000-3119

the LOX/CH4 momentum ratio as shown in figure 11.Moreover, the efficiency values did not significantly varyamong the injector configurations.

Several tests were also dedicated to investigate theeffects of the chamber length. Two chamber lengths, 9.0and 13.0 inches, were selected for this study. The resultsshowed that no significant variation in the combustionefficiency was observed.

Combustion Instability Measurement:The configuration of a 2.4-inch diameter injector insideof a 4-inch diameter chamber prompted some uniquecombustion instability concerns. Hence, several acousticcavity tuning-rings were designed to accommodate ITand IR-mode combustion instabilities. Out of which, anacoustic cavity ring consisting of three 1-inch depthcavities and three 0.5-inch depth cavities correspondingto 1-T and 1-R modes, respectively, was selected for. thetests. In this project, two high-frequency (25Khz)pressure transducers were installed on the MCTAchamber wall at a location of 6 inches away from theinjector face. No combustion instability was encounteredin any of the tests. Figure 12 shows the results of test #38 at main stage. Similar peak-to-peak chamber pressurefluctuations were also observed in the other tests. Themaximum peak-to-peak pressure variations in tests #21,26, 34, and 38, corresponding with injectors #3, 1,2,and 4, respectively, were less than 3% and are reportedin figure 13. These tests were run approximately at thebase-lined mixture ratio of 3. It should be noted that theacoustic cavity ring was replaced later with a solid blankblock in some of the final tests. No difference in thepressure fluctuation between with and without acousticcavity cases was observed.

CONCLUSION

A total of forty hot-firings on four impingement injectorsfor the CHt/LOX system were conducted in thisprogram. The optical diagnostic measurements showedthat the propellants were well mixed. No thermalcompatibility issues on the injector faceplates wereobserved.

The four subject injectors have similar combustionefficiency values for the chamber pressure and mixtureratio ranging from 343 to 513 and 1.2 to 3.5,respectively. The efficiency value increases as themixture ratio is reduced. The results also showed that theC* efficiency value did not vary when the chamberlength changed from 9.0 to 13.0 in. Finally, nocombustion instability was observed for all the testingconditions.

ACKNOWLEDGMENTS

The author would like to acknowledge the contributionsof a number of MSFC personnel and contractors on thisproject. The hot-fire tests could not have beensuccessfully executed without the technical consultationof John Cramer. Andy Hissam, Kevin Baker, and MikeShadoan have dedicated their efforts to the design andfabrication of the MCTA and the injector hardware.Cynthia Lee and her test team have labored many hoursto prepare the test facility and to perform the hot-firetests. Brad Bullard has provided significant technicalsupports of the post test data analysis. Marvin Rockerand Tom Nesman have performed stability analysis forthe MAE injectors designs and evaluated the test results.David McDaniels, Herbert Zollar, Hai Nguyen and JeffLin have conducted cold-flow tests, thermal analyses,and combustion analyses, respectively. Finally, Dr.Joseph Wehrmeyer of Vanderbilt University and PerryGray of Micro Craft Incorporate have spent much time toperform combustion measurements with the laserdiagnostic and the infrared camera, respectively.

REFERENCES

1. Kos, L., "The Human Mars Mission: TransportationAssessment," Space Technology and ApplicationsInternational Forum (STAJP) conference,Conference Proceedings Part III, p. 1206,Albuquerque, NM, January 1998.

2. Bailey, C.R., High Pressure Lox/Natural Gas StagedCombustion Technology, 1984 JANNAF PropulsionMeeting, New Orleans, Louisiana, Feb. 8, 1984.

3. Kirby, F. M., "LOX/CH4 For Reusable HighPerformance Booster Engines," AJAA Paper

4. Claflin, S. E., Volkmann, J. C., "MaterialCompatibility and Fuel Cooling Limit Investigationfor Advanced LOX/Hydrocarbon Thrust Chamber,"AIAA 90-2185, 26th AIAA/ASME/SAE/ASEE JointPropulsion Conference, July 13-15, 1990.

5. Breisacher, K. J., Priem, R. J., "Analysis of 5 KHzCombustion Instabilities in 40K Methane/LoxCombustion Chambers," 25th JANNAF CombustionMeeting, Huntsville, Alabama, 10/1988.

6. Rosenberg, S. D., Gage, M. L., Homer, G. D.,Franklin, J. E., "Hydrocarbon-Fuel/CopperCombustion Chamber Liner Compatibility,Corrosion Prevention, Refurbishment," J. ofPropulsion and Power, Vol. 8, No. 6, Pages 1200-1207, Nov.-Dec. 1992.


AIAA2000-3119

7. Sanders, J., "In-situ Propellant Production (ISPP) &Propulsion Coordination," presentation charts of2/12/1998, NASA/Johnson Space Center.

8. Mueller, P. J., Plachta, D.W., Peters, T., Whitehead,J.C., "Subscale Precursor to a Human Mars MissionUsing in Situ Propellant Production," A1AA-98-3301, 34th AIAA/ASME/SAE/ASEE JointPropulsion Conference, July 13-15, 1998.

9. Chapman, J. M., "Mars Sample Return AscentStage," presentation charts of 2/25/98,NASA/Marshall Space Flight Center.

10. Pavli, A. J., "Design and Evaluation of HighPerformance Rocket Engine Injectors for Use withHydrocarbon Fuels," NASA Technical Memo79319, Prepared for the 16th JANNAF CombustionConference, Monterey, CA, 9/10-14, 1979.

11. Yang, V., Anderson W., Liquid Rocket EngineCombustion Instability. 4, Vol. 169, Progressin Astronautics and Aeronautic, 1995.

12. Coward, H.F., Jones, G.W., Limits of Flammabilityof Gases and Vapors, Bureau of Mines Bulletin 503,1952.

13. Osborne, R. J., Wehrmeyer, J. A., Pitz R. W., "AComparison of UV Raman and Visible RamanTechniques for Measuring Non-Sooting PartiallyPremixed Hydrocarbon Flames," AIAA-2000-0776,

38th Aerospace Sciences Meeting & Exhibit, January2000.

14. J. Wehrmeyer, H. Trinh, R. Hartfield, C. Dobson,and R. Eskridge, "Raman Gas SpeciesMeasurements in Hydrocarbon-Fueled RocketEngine Injector Flows," AIAA2000-3391, the 36thAIAA/ASME/SAE/ASEE Joint Propulsion, July2000

15. Wehrmeyer, J.A., Cramer, J.M, Eskridge, R.H.,Dobson, C.C., UV Raman Diagnostics for RocketEngine Injector Development, AIAA Paper 97-2843,1997.

16. Hartfield, R., Dobson, C., Eskridge, R., Wehrmeyer,J., Development of a Technique for SeparatingRaman Scattering Signals from BackgroundEmission with Signle-Shot Measurement Potential,AIAA paper 97-3357, 1997.

17. Moser, M.D., Measurement of Injector FaceTemperature Using Optical Diagnostic Techniques,Final report, Contract # NAS8-97095, Task # H-28520D, Oct. 25, 1999

GN2 Film-CoolingInjection Ring

Acoustic CavityRing

Figure 4: A Typical MAE Injector Hot-FireTest in the MCTA Chamber

Figure 5: Cumulative Overheated Damage onGN2 and Cavity Rings After 31 tests


AIAA2000-3119

Injector # 1After Test #

Bright Color Due ToCombustion

No Appearance of Bright Colqcase or the Twisted Injection

the Twisted Injection


•y'/A



Figure 6: Post-Test Photographs of The

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0.0 0.5 1.0 1.5 2.0

— Temp. (Ref. 8) Mixture Ratio (0/F) ^ No-light, Test # 8— Pressure (Ref. 8) * Light, Test #8,16,36,40

Figure 7: Fuel Rich Ignition Limits andTemperature of


AIAA2000-3119

(a) (b)

Vertical ILocation!

(mm)

VERTICALPOLARIZATION

HORIZONTALPOLARIZATION

250 02 N2CH4WAVELENGTH, nm

O2 N2CH4 H2QWAVELENGTH, nm

250

Figure 8: Raman Scattering Images of Combustion Flow Field of Test # 25(a): GN2 Purge at 0 second; (b): Main Stage at 2 Seconds

Relatively ColdPropellant Jets

Overheated Region onthe GN2 Injection Ring

FlowDirection

Figure 9: Infrared Image of Injector Face and CombustionFlow Field of Test # 21 at Main Stage

Injector Face

130 -120-

? 110

o 100

•§ 90| 80

" 70

601

IflS

*

*

! I——— «« ———— I ———— I ————H-J ———————— ̂ ———

1 ». . 1! r * *i i

!

0 1.5 2.0 2.5 3.0 3.5 4.0

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Figure 10: Combustion Efficiency Values for SeveralRepresentative Tests of All Four Injectors

130 -,~ 120t 110.<T 100* nn0 90

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Figure 11: C* Efficiency Vs. LOX/CH4Momentum Ratio

(c)2000 American Institute of AeronauticsJcs & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization.

AIAA2000-3119

Raw Man TQ-CLQO n-11.00 — m«38..P31.02 Ba*MeanTO-O.OQTI.t 1.00— Chamber press.— Lox press.— Fuel press.

Figure 12 Chamber Pressure Fluctuation of Test #38 Measured from TwoHigh-Frequency (25Khz) Pressure Transducers

#2* #3 #4*

Injector Number (*) NO Acoustic Cavity

Figure 13 Maximum Peak-to-Peak Chamber Pressure Fluctuation ofFour Injectors at MR-3.0

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[american institute of aeronautics and astronautics 36th aiaa/asme/sae/asee joint propulsion...

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