L- 75^1 iNASA Technical Memorandum 105897

Performance and Heat Transfer Characteristicsof a Carbon Monoxide/OxygenRocket Engine

Diane L. LinneLewis Research CenterCleveland, Ohio

February 1993

NASA

Performance and Heat Transfer Characteristics of aCarbon Monoxide/Oxygen Rocket Engine

Diane L. LinneNational Aeronautics and Space Administration

Lewis Research CenterCleveland, Ohio 44135

ABSTRACT

The combustion and heat transfercharacteristics of a carbon monoxide andoxygen rocket engine were evaluated. The testhardware consisted of a calorimeter combustionchamber with a heat sink nozzle and aneighteen element concentric tube injector.Experimental results are given at chamberpressures of 1070 and 2070 kPa, and over amixture ratio range of 0.3 to 1.0. ExperimentalC* efficiency was between 95 and 96.5percent. Heat transfer results are discussedboth as a function of mixture ratio and axialdistance in the chamber. They are alsocompared to a Nusselt number correlation forfully developed turbulent flow.

INTRODUCTION

The future exploration of the solarsystem will require the launch of large massesfrom the surface of the Earth. If chemicalpropulsion is used for these travels, then asignificant percentage of this launch mass willconsist of propellants for the outbound trip andthe journey home. One proposal to reducelaunch mass requirements is to eliminate theneed to launch the return propellants (and someof the outbound propellant used to carry thereturn propellants) by producing thesepropellants at the site of exploration. Thisutilization of indigenous materials forpropulsion has recently garnered muchattention among mission planners, who showthat in situ propellants can reduce the Earthlaunch mass for a lunar or Mars mission by 30to 66 percent (refs. 1-7). The propulsiontechnology base for some of the proposedpropellants, however, needs to be enlargedbefore an actual engine can be developed.

The atmosphere of Mars consists ofover 95 percent carbon dioxide. One proposedmethod for utilizing this resource is bydissociating the CO2 into oxygen and carbonmonoxide, and then recombining thesepropellants in a rocket engine. Althoughcarbon monoxide has been burned in severalapplications, such as the catalytic converter inan automobile engine, specific experimentationto obtain the information necessary to design aflight engine has only been performed to alimited extent (refs. 8,9). The ignitioncharacteristics of a dry carbonmonoxide/oxygen mixture in a spark torchigniter were studied, and an ignition rangeidentified (ref. 8). Additionally, steady-statecombustion has been demonstrated in heat sinkhardware, and some preliminary combustionefficiencies obtained (ref. 9).

Because of the limited database, testswere conducted with a calorimeter combustionchamber to study the heat transfercharacteristics of the oxygen and carbonmonoxide propellant combination. Theobjectives of the experiment were to measurethe combustion efficiency with a newlydesigned 18 element concentric tube injector,and to obtain hot-gas-side heat flux data. Thecombustion efficiency is compared with resultsfrom a previous experiment (ref. 9), and withtheoretical predictions of real-engine losses.The heat transfer results are shown togetherwith a Nusselt number correlation for fullydeveloped turbulent flow.

TEST APPARATUS AND PROCEDURE

Test Facility

The experimental tests for this studywere performed in Cell 21 of the Rocket Lab atthe NASA Lewis Research Center. This

facility contains a low thrust rocket engine teststand with supporting fluid systems that allowprecise flow control. Four separate gaseouspropellant lines were used for this researchprogram: one oxygen supply line to theengine, one oxygen supply line to the sparktorch igniter, one carbon monoxide fuel supplyline to the engine, and one hydrogen fuelsupply line to the igniter.

The flow rate of each of the gases in thesystem described above was controlled with asonic orifice. Inserted as a component of thepropellant line, each orifice insured a constantflow rate of gas, independent of downstreampressure perturbations. By measuring the linepressure and temperature at a point justupstream of each sonic orifice, gas flow rateswere calculated. Different diameter orificescould be easily interchanged in the system sothat the gas flow rate range could be variedthroughout the test program.

Test Hardware

The test hardware for this experimentconsisted of an igniter, injector, calorimeterchamber spool piece, and converging-divergingnozzle. Figure 1 shows a schematic of the testapparatus and figure 2 shows the hardware onthe test stand.

A hydrogen-oxygen spark torch igniterwas used to initiate combustion. Gaseousoxygen and gaseous hydrogen were injectedinto the igniter chamber at an oxygen-to-fuelmixture ratio (O/F) of approximately 40, wherea standard spark plug initiated combustion.The hot gases then travelled down a tubethrough the center of the injector body and intothe combustion chamber. At the exit of theigniter tube, additional gaseous hydrogen,which had been used to cool the outside of theigniter tube, was added to the hot gases toincrease the flame temperature. This additionalhydrogen lowered the total igniter mixture ratioat the exit of the igniter tube to approximately6.0.

An 18 element concentric tube injectorwas used for the combustion chamber. Each

element injected oxygen through the centerorifice and carbon monoxide through the outerannulus. The elements were arranged in twocircles centered around the igniter, with sixelements on the inner circle and twelveelements on the outer circle (figure lb).Chamber pressure was measured by means of apressure tap on the face of the injector.

The calorimeter test chamber was 14.9cm (5.875 inches) long and had a 6.6 cm (2.6inches) inside diameter. The inner copper linerwas cooled by 46 circumferential coolingchannels. Coolant water was supplied to thesechannels by 22 inlet tubes. Heat flux wascalculated at each of the 22 stations bymeasuring the coolant inlet and outlettemperatures, the coolant flow rate, and thechamber wall area cooled by those channels.In order to obtain an axial temperature profile,two thermocouples were inserted into thecopper liner to within 0.0762 cm (0.030inches) of the hot-gas-side wall at 11 axiallocations. Table I lists the axial location of thecoolant circuits and the thermocouples.

A copper heat sink converging-diverging nozzle was used. The nozzle had athroat diameter of 1.143 cm (0.45 inches) andan exit area ratio of 2.997. The divergingnozzle contour was a cone, with an exit half-angle of 15 degrees.

Test Procedure

To insure a uniform run profilethroughout the duration of the test program,each firing of the engine was sequenced by aprogrammable line controller. Each test runstarted with the initiation of the oxygen andhydrogen flows to the igniter, followed by theoxygen and carbon monoxide flows to the maincombustion chamber after the igniter spark wasstarted. After initiation of main combustion,the igniter was stopped, and the test continuedfor approximately 6 seconds with no hydrogenflowing. This sequencing allowed forhydrogen to be present during start-up of theengine to aid in the ignition of the dry carbonmonoxide and oxygen mixture.

2

The oxygen flow rate was varied from21 to 94 g/sec (.046 to .21 lbm/sec). Thecarbon monoxide flow rate was varied from 47to 144 g/sec (.104 to .317 lbm/sec). The totalflow rate was held relatively constant at 95 and186 g/sec (.21 and .41 lbm/sec) whichprovided actual chamber pressures ofapproximately 1070 and 2070 kPa (155 and300 psia).

Experimental data was gathered duringthe test runs by a high-speed data acquisitionsystem. In addition to the instrumentation onthe hardware, pressure transducers andthermocouples were applied to the facility feedsystems to properly measure the propellantflow rates and temperatures. A total of 100instrumentation channels were each scanned atthe rate of 100 times per second. Each valuequoted in this analysis is an average of 10readings of the instrument by the data system.The data reduction was performed by aFORTRAN 77 computer program hosted on aVAX cluster.

DESCRIPTION OF COMPUTER CODES

Two computer codes were used duringevaluation of the experimental results. TheLiquid Propellant Program (LPP) computercode (ref. 10) was used for characteristicvelocity (C*) efficiency and heat fluxcomparisons. The Rocket Engine HeatTransfer Evaluation Program (REHTEP) (anunpublished, NASA Lewis Research Centercomputer code) was also used for heat transfercharacteristics comparisons.

LiQuid Propellant Program

The LPP code uses a chamber andnozzle geometry together with thermodynamicsand kinetics to calculate various performancelosses that an actual engine may experience innormal operation. The code consists of severalmodules, each of which models a different typeof performance loss. All modules assumecomplete combustion in the chamber, that is,no loss in energy release caused by slowvaporization or nonuniform mixing.

The Mass Addition Boundary Layer(MABL) module was used to calculate C*efficiencies and chamber heat flux. This is aboundary layer module that models the growthof the viscous boundary layer in the chamberand nozzle. For this analysis, the start of theboundary layer was assumed to be at theinjector. The MABL module uses output fromthe previous modules, especially the TwoDimensional Kinetics (TDK) module, whichpredicts the inviscid, two-dimensionalexpansion of the gaseous combustion productsassuming finite-rate kinetics. To simulate theexpected wall conditions, the actual walltemperature profile measured in theexperimental tests was used as input. For thetheoretical analyses presented in this paper,MABL calculates the displacement thicknessfor the actual chamber and nozzle geometry anduses this to obtain a displaced, or inviscid, wallcontour. The TDK module is then rerun withthe new contour. A new mass flow rate isobtained and this mass flow is then used alongwith the actual or geometric throat area toobtain a predicted value of C*. This value ofC* is divided by the theoretical ideal value ofC* to obtain a theoretical C* efficiency.MABL also calculates the heat flux at thechamber wall, and these values were used forcomparison to the experimental results.

Rocket En ig ne Heat Transfer EvaluationPro gram

The REHTEP code uses a chamber andnozzle geometry together with specificationsfor axial coolant passages to evaluate heattransfer characteristics and cooling capabilitiesof various propellant combinations andcoolants. The program calculates theconditions (combustion products, temperature,and other thermodynamic properties) in theengine using a one dimensional equilibriumsubroutine. With these values, a heat transfercoefficient is calculated using a Nusselt numbercorrelation for fully developed turbulent flow

Nu= k*D= C8 *Re g *Pr-3(1)

8

3

where h g is the heat transfer convectioncoefficient, D is the diameter of the chamber,kg is the conductivity of the combustion gases,C g is the correlation coefficient (0.026 wasused in this study), Re is the Reynoldsnumber, and Pr is the Prandtl number. In thisprogram, the transport properties are evaluatedat Eckert's reference enthalpy (ref. 11). Theheat flux at the chamber wall is then calculatedusing an assumed wall temperature and the heattransfer equation

Q = hg *(T' , —Tg.) (2)

where Q is the heat flux per unit area, TQ,,, is theadiabatic wall temperature, and TgW is thetemperature of the chamber wall. Thetheoretical adiabatic wall temperature iscorrected for combustion efficiency, which isan input to the program. The wall temperatureis then iterated upon until an energy-balancebetween the coolant side and the combustionside of the chamber is achieved.

RESULTS AND DISCUSSION

Two types of experimental data,performance and heat transfer, were of interestto meet the objectives of the test program.Experimental C* efficiency results from thistest series are compared to previously obtainedexperimental data, and to efficiencies obtainedfrom the LPP program. The heat transferresults included chamber wall temperatures andchamber heat flux. Experimental results arediscussed as functions of mixture ratio andaxial distance in the combustion chamber andare also compared to results obtained from thetwo computer codes.

C* Efficiency

In a previous experimental test program(ref. 9), C* efficiencies of 89 to 92 percentwere obtained. In those tests an eight-elementtriplet injector design was used. For this testprogram, an eighteen-element, concentric tubeinjector was designed in an effort to increasethe combustion C* efficiency. The injectorface is shown schematically in figure 1 b.

Figure 3 shows the experimental C*efficiencies as a function of mixture ratioobtained from both injectors, along with a C*efficiency obtained from the boundary layermodule of the Liquid Propellant Program(LPP). The computer program calculated realengine losses for the test hardware, andpredicted C* efficiencies between 96 and 97percent over the mixture ratio range of 0.3 to1.3 (stoichiometric mixture ratio= 0.571). Theexperimental efficiencies obtained with theconcentric tube injector were between 95 and96.5 percent. The figure shows that theexperimental efficiencies were higher at thelower mixture ratios where the fuel to oxygeninjection velocity ratio was higher, providingfor better mixing between the carbon monoxideand oxygen. The figure also shows that theefficiencies obtained with the concentric tubeinjector are a significant improvement overthose obtained in the previous test program.

Heat Transfer Characteristics

The heat transfer characteristics of thetest hardware were evaluated in two manners.First, experimental chamber heat flux wasexamined as a function of mixture ratio and as afunction of axial distance from the injector.Second, the experimental heat flux results werecompared to heat flux calculated from boundarylayer theory and from a Nusselt numbercorrelation for fully developed turbulent flow ina constant area duct.

Experimental Results. Theexperimental heat flux, q, in the chamber wascalculated based on the temperature rise of thecoolant water and the coolant water mass flowrate

__ m,*CP*(T. —T^,)(3)

qA

where m c is the coolant water mass flow ratethrough that circuit, Cp is the specific heat ofwater, Teo is the coolant outlet temperature, Ti

is the coolant inlet temperature, and A is thechamber wall area cooled by that circuit. This

4

method of calculation assumes only radialconduction in the copper liner. To reduce theamount of axial conduction, the coolant waterflow rates for each coolant circuit were adjusteduntil the chamber wall temperatures wereapproximately equal between adjoiningstations. Figure 4 shows the walltemperatures, as measured by the ribthermocouples, for a typical test run for eachchamber pressure. The figure shows that forthe tests run at a chamber pressure of 1070kPa, the wall temperatures were within fivedegrees of 355 K, with the exception of thefirst and last thermocouples (station 2 andstation 21). Because the heat flux at thesestations was significantly lower than the rest ofthe stations, it was difficult to adjust the coolantwater flow rate low enough to obtain the samewall temperature. Similar results were obtainedfor the tests run at a chamber pressure of 2070kPa, where the wall temperatures were withinfive degrees of 360 K.

The experimental heat flux as a functionof mixture ratio is shown in figure 5 at an axiallocation of 5.268 cm (2.074 inches, coolingstation 8) and of 14.16 cm (5.574 inches,cooling station 22). Cooling station 8 was theapproximate location of highest heat flux, andcooling station 22 was the last station in thechamber. The figure shows a similar pattern atboth stations and for both chamber pressures.At both stations the heat flux is significantlylower at the lower mixture ratios, but relativelyeven between mixture ratio of 0.50 and 0.70.It is clear from the figures, however, that theheat flux at station 8 was much higher than atstation 22. For a chamber pressure of 2111kPa, the maximum heat flux at station 8 was5000 kW/m2 compared to only 2650 kW/m 2 atstation 22. Similarly, for a chamber pressureof 1063 kPa, the maximum heat flux at station8 was 3200 kW/m2 compared to only 1450kW/m2 at station 22.

To evaluate the axial variation in heatflux, figure 6 shows the experimental heat fluxas a function of axial distance from the injectorface for both chamber pressures. The mixtureratio for these curves was approximately 0.55.This mixture ratio resulted in the highest heat

fluxes, and is also a likely operating point foran actual engine. The figure shows that theheat flux increases steadily from the beginningof the chamber until about 5 to 6 cm into thechamber, where the heat flux then decreasesthrough the rest of the chamber. The graphs ofheat flux as a function of axial location wereused to determine the location of the end of thecombustion zone for each mixture ratio. Thiswas defined as that point showing maximumheat flux. By using this point, it was assumedthat the further growth of the boundary layerhas a small effect on the location of maximumheat flux.

Figure 7 shows the location of the endof the combustion zone in the chamber for themixture ratios tested. At lower oxygen to fuelmixture ratios, the combustion zone is shorter,and then becomes longer at the higher mixtureratios. The figure indicates that the combustionzone ended between 2.7 and 6.5 cm (coolingstations 4 and 10). The duration of thecombustion could be affected by severalparameters, including kinetic reactions andinjection velocity ratio. The reaction rate forthe CO and 02 reaction is known to be slow. Itis possible that the slow kinetics of the systemrequire more time for enough collisions tooccur with sufficient energy to form completecombustion. Specific injector parameters mayalso affect combustion. With a concentric tubeinjector, optimum mixing occurs at high fuel-to-oxygen injection velocity ratios. At the lowermixture ratios, more fuel is injected through theouter annulus, and it is therefore injected at ahigher velocity. Similarly, at the low mixtureratios, less oxygen is injected through thecenter orifice, and it is therefore injected at alower velocity. This combination produces ahigher fuel-to-oxygen injection velocity ratio atthe lower mixture ratios, promoting mixing,and allowing combustion to be completed morequickly.

Theoretical Analysis. In figure 8 theexperimental heat flux as a function of axialdistance for a mixture ratio of 0.55 and achamber pressure of 1070 kPa is shown again.Also included in the figure is the heat fluxcalculated from two different computer codes.

It can be seen that the experimental heat flux issignificantly higher than that calculated by bothcodes. LPP predicts a high heat flux at theinjector face which then decreases throughoutthe chamber as the boundary layer grows. TheREHTEP code uses a Nusselt numbercorrelation for fully developed turbulent flow(equations (1) and (2)) to calculate heat flux.Both codes assume that all energy release in thecombustion chamber occurs at the injector.

To further compare the experimentalheat transfer results with fully developedturbulent flow, the experimental heat transfercorrelation coefficient is shown in figure 9.The dashed line in the figure represents theempirically derived correlation coefficient of0.026 for cooling in a duct with fullydeveloped turbulent flow. The figure showsthat the experimental correlation coefficient isalways higher than the 0.026 value. Thelargest difference occurs at the same locationwhere the highest heat flux was observed(where the combustion zone ends), with anexperimental value five times higher than0.026. These high correlation coefficientswould seem to indicate that the combustionprocess has caused the experimental heattransfer characteristics to behave in a mannerdifferent from fully developed turbulent flow.By the end of the chamber, however, theexperimental correlation coefficient is onlyslightly higher than the 0.026 value, and it ispossible that at this point the flow in thechamber is more like fully developed turbulentflow.

CONCLUSIONS

The combustion and heat transfercharacteristics of carbon monoxide and oxygencombustion were evaluated in a calorimetercombustion chamber with a heat sink nozzleand an eighteen element concentric tubeinjector. The experimental C* efficiency wasbetween 95 and 96.5 percent over a mixtureratio range of 0.3 to 1.0. This was asignificant improvement over a triplet injectordesign evaluated in a previous test program.Maximum heat flux was approximately 3200kW/m2 at a chamber pressure of 1063 kPa and

4900 kW/m 2 at a chamber pressure of 2111kPa. Using the location of maximum heat fluxas an indicator, the end of the combustion zoneoccurred between 2.7 and 6.5 cm downstreamof the injector.

The experimental heat flux andcorrelation coefficients were much higher thanthose calculated using a Nusselt typecorrelation for cooling in a duct with fullydeveloped turbulent flow. Moreexperimentation is needed to further evaluatethe results obtained in the calorimeter chambertests. Experiments with various injectordesigns and varying chamber lengths areneeded to determine the effects of the injectionvelocity ratio and the slow kinetic reactions ofthe carbon monoxide and oxygen combination.

REFERENCES

Teeter, R.R.; and Crabb, T.M.: LunarSurface Base Propulsion System Study.NASA CR-171982—VOL-1, 1987.

2 Wickman, J.H.; Oberth, A.E.;Mockenhaupt, J.D.: Lunar Base Space-craft Propulsion with Lunar Propel-lants. AIAA Paper 86-1763, 1986.

3. Repic, E., et al.: If We're going toMars, Why Stop at the Moon? Pre-sented at the Conference, "The Case forMars IV, The International Explorationof Mars," Boulder, CO, June 1990.

4. Stancati, M.L., et al: In Situ PropellantProduction for Improved Sample Re-turn Mission Performance. AAS Paper79-177, Astrodynamics 1979; Proceed-ings of the Conference, P. Penzo, et al.,eds., Pt. 2, Univelt, Inc., San Diego,CA, 1980, pp. 909-921.

5. French, J.R.: Rocket Propellants fromMartian Resources. J. Br. Interplan.Soc., vol. 42, Apr. 1989, pp. 167-170.

6. Zubrin, R.M.; Baker, D.A.; andGwynne, O.: Mars Direct: A Coherent

6

Architecture for the Space ExplorationInitiative. AIAA Paper 91-2333,1991.

7. Giudici, B.: A Get Started Approachfor Resource Processing. AAS Paper87-262, The Case for Mars III: Strate-gies for Exploration—Technical, C.R.Stoker, ed., Univelt, Inc., San Diego,CA 1989, pp. 469-478.

Linne, D.L.; Roncace, J.; and Groth,M.F.: Mars In Situ Propellants: Car-bon Monoxide and Oxygen IgnitionExperiments. NASA TM-103202,1990 (Also, AIAA Paper 90-1894,1990).

9. Linne, D.L.: Carbon Monoxide andOxygen Combustion Experiments: ADemonstration of Mars In Situ Propel-lants. NASA TM-104473,1991 (Also,AIAA Paper 91-2443, 1991).

10. Nickerson, G.R.; Dang, A.L.; andCoates, D.E.: Engineering and Pro-gramming Manual: Two-DimensionalKinetic Reference Computer Program(TDX). NASA CR-178628, 1985.

11. Eckert, E.R.G.; and Drake, R.M., Jr.:Heat and Mass Transfer. Seconded.,McGraw-Hill Book Co., Inc., 1959.

Table 1. - Location of Coolant Stations and Rib Thermocouples

Coolant Station Number Axial Location(Distance from Injector Face)

(cm)

Thermocouples Inserted inCopper Liner

1 0.546 No2 1.458 Yes3 2.093 No4 2.728 No5 3.363 Yes6 3.998 No7 4.633 No8 5.268 Yes9 5.903 No10 6.538 Yes11 7.173 No12 7.811 Yes13 8.443 No14 9.078 Yes15 9.713 Yes16 10.35 Yes17 10.98 Yes18 11.62 Yes19 12.25 No20 12.89 No21 13.52 Yes22 14.16 No

TH2/02 18 ElementSpark-Torch Concentric Tube

Igniter Injector

Calorimeter Chamber Converging-DivergingHeat-Sink Nozzle

PrimaryOxygen

14.9 cm

Carbon Monoxide)T`Water In

Rib Thermocouple(typical)

11 Locations

Water Out

(a) Calorimeter engine assembly.

Igniter Outlet

O. .^. ^..... . .. ...•^^::a:o::...

I 1Z,

Carbon MonoxideAnnulus

(.0165 cm gap)

Oxygen Orifice(.109 cm Dia.)

(b) 18 element concentric tube injector face.

Figure 1.—CO/0 2 calorimeter experimental test hardware schematic.

8

370 r n Pc = 2070 kPaO Pc = 1070 kPa

Y365

El

a 360Ea)m 3553N

350

a 345x

n

n

n O O O 0O

O

O

i

Figure 2.—Calorimeter chamber on test stand.

100

q Liquid propellant programn Concentric tube injector• Triplet injector

98

q qno qn ^ n n nn n96 n n

W nav 94c

92

•

90•^7^•

88' ' ' ' ' 1

0.3 0.5 0.7 0.9 1.1 1.3

Mixture ratio (O/F)Figure 3.—Comparison of combustion chamber C' efficiencies

for triplet and concentric tube injectors.

3400 2 4 6 8 10 12 14 16

Axial distance from injector face (cm)

Figure 4.—Comparison of hot-gas side wall temperatures asmeasured by thermocouples at a depth of 0.076 cm.

9

5500

5000 n n

n4500 n

n n Pc = 2111 kPa4000 n O Pc = 1063 kPa

3500 n

3000 0 0 §00 00 O O

ON 00E 2500

(a)Coolant station 8 (x = 5.268 cm, 2.074 in.)K

7

2900mx

2500 n _n nn

n n2100 n

n1700 n

n O (900 00 1300 O

go O 0 0

9000.3 0.5 0.7 0.9 1.1 1.3

Mixture ratio (O/F)(b)Coolant station 22 (x = 14.16 cm, 5.574 in.)

Figure 5.—Experimental heat flux as a function of mixture ratiofor both chamber pressures.

5000 • • Pc = 2070 kPa• • q Pc = 1070 kPa

•

4000 • • ••

E • •

Y 3000 El E] El El e •qq

El q• • •2000

CU

q • •El q • •

_ q qq • q

1000 q q q q

0 2 4 6 8 10 12 14 16Axial distance from injector face (cm)

Figure 6.—Experimental heat flux for both chamber pressures(mixture ratio = 0.55).

1.2O

n1.0 0

n

o ON0 0.8 OE2

o.s

0.4 IN O n Pc = 2111 kPa000 O Pc = 1063 kPa

n0.2

0 2 4 6 8 10 12 14 1 16

Injectorface Axial location of maximum Q (cm) End of

chamberFigure 7.—Location of maximum heat flux for varying mixture

ratio.

10

n Pc = 1070 kPaq Liquid propellant programO Rocket engine heat transfer

3500evaluation program

3000 n nnn

2500 n n nnn

2000 nnX

1500 n n n ni6 n

= 1000 n n ME500 g qq®

qq L1q Oq O 00000 00

Ll q qq qqq qq

0 2 4 6 8 10 12 14 16Axial distance from injector face (cm)

Figure 8.—Comparison of experimental and theoretical heat fluxin the chamber (Pc = 1070 kPa, mixture ratio = 0.558).

n Pc = 2070 kPaO Pc = 1070 kPa

0.14 ---- Fully developedturbulent flow

CU

0.12 n nmoU n U U0 0.10 n0

n

m 02 0.08 O o

On

U

0.06 nCro

n oe

ro 0.04m

O O

0.020 2 4 6 8 10 12 14 16

Axial distance from injector face (cm)Figure 9.—Experimental heat transfer correlation coefficient in

the combustion chamber.

11

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February 1993 Technical Memorandum4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

Performance and Heat Transfer Characteristics of a Carbon Monoxide/OxygenRocket Engine

W U —506-42-726. AUTHOR(S)

Diane L. Linne

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National Aeronautics and Space AdministrationLewis Research Center E-7521Cleveland, Ohio 44135-3191

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Responsible person, Diane L. Linne, (216) 977-7512.

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13- ABSTRACT (Maximum 200 words)

The combustion and heat transfer characteristics of a carbon monoxide and oxygen rocket engine were evaluated. Thetest hardware consisted of a calorimeter combustion chamber with a heat sink nozzle and an eighteen element concentrictube injector. Experimental results are given at chamber pressures of 1070 and 3070 kPa, and over a mixture ratio rangeof 0.3 to 1.0. Experimental C* efficiency was between 95 and 96.5 percent. Heat transfer results are discussed both as afunction of mixture ratio and axial distance in the chamber. They are also compared to a Nusselt number correlation forfully developed turbulent flow.

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Propellent combustion, Extraterrestrial resources, Carbon monoxide, Heat Transfer 1216. PRICE CODE

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