NASA Technical Memorandum 105967

AIAA 92-3390_ _

//v---_V_.O

/+,o77b

Performance Comparison of Axisymmetric

and Three-Dimensional Hydrogen Film

Coolant Injection in a 1 !0N Hydrogen/

Oxygen Rocket _ _

__3 ....

Lynn A. Arrington ":

Sverdrup Technology, Inc.

Lewis Research Center_ Group

Brook Park, Ohio

and

Brian D. Reed

Lewis Research Center

Cleveland, Ohio

Prepared for the

28th Joint Propulsion Conference and Exh_ibit

sponsored by the AIAA/SAE/ASME and ASEE

Nashville, Tennessee, July 6-8, !992

N/ A

(NASA-TM-105967) PERFORMANCE

COMPARISON OF AXISYMMETRIC AND

THREE-DIMENSIONAL HYDROGEN FILM

COOLANT INJECTION IN A 110N

HYDROGEN/OXYGEN ROCKET (NASA)

22 p

G3/ZO

N93-16714

Unclas

0140770
r

_ • _ 7_
Performance Comparison of Axisymmetric

and

Three-Dimensional Hydrogen Film Coolant Injection

in a 110 N Hydrogen/Oxygen Rocket

Lynn A. Arrington

Sverdrup Technology Inc.

NASA Lewis Research Group

2001 Aerospace Parkway

Brook Park, OH 44142

ki

BrianD. Reed

NASA

NASA Lewis Research Center

21000 Brookpark Road

Cleveland, OH 44135

Abstract

An experimental performance comparison of two geometrically different fuel film coolant

injection sleeves was conducted on a 110 N gaseous hydrogen/oxygen rocket. One sleeve had

slots milled axially down the walls and the other had a smooth surface to give axisymmetric

flow. The comparison was made to investigate a conclusion in an earlier study that attributed

a performance underprediction to a simplifying modeling assumption of axisymmetric fuel film

flow. The smooth sleeve had higher overall performance at one film coolant percentage and

approximately the same or slightly better at another. The study showed that the lack of modeling

of three-dimensional effects was not the cause of the performance underprediction as speculated

in earlier analytical studies.

Introduction

Low thrust propulsion is required on every launch vehicle, satellite, and spacecraft. Low

thrust propulsion functions include apogee insertion, attitude control, stationkeeping, rendezvous,

docking separation, midcourse correction and planetary retro. Currently, the bulk of low thrust

propulsion functions are provided by small chemical rockets with thrust levels, depending on the

function, ranging from 450 mN to 4500 N.

=

Small rocket flowfields differ from those of larger thrust class rockets in several ways. The

flows are more strongly influenced by viscous effects because of the relatively small size and

corresponding large surface-to-volume ratio in small rockets. Also, a substantial percentage of

the fuel is usually required for film cooling of the walls. The large amounts of fuel film cooling

and low Reynolds number flow can lead to significant mixing and boundary layer losses in small

rockets. Furthermore, in the mixing layer, a secondary combustion zone can exist between the

fuel film cooling flow and the core flow. This chemically reacting viscous dominated flowfield

This paper is declared a work of the U.S. Government and

is not subject to copyright protection in the United States.
has resulted in considerable difficulty in the prediction of performance and thermal behavior in

small rockets.

To accurately model the mixing and heat transfer in small rocket flowfields, an effort was

undertaken to apply numerical methods developed in the aeronautical community to rocket

flowfield modeling. The RPLUS code,'- originally developed to study supersonic combustion of

hydrogen in air for ramjets and scramjets, was adapted to model small rocket flowfields. The

RPLUS code numerically solves the coupled set of Navier-Stokes and species transport equations

in axisymmetric coordinates, over the entire flowfield.

In the preliminary assessment of RPLUS, two versions 3"4were developed that compared the

measured performance from a 110 N, gaseous hydrogen/gaseous oxygen rocket to the performance calculated by RPLUS. In the code, the rocket geometry was simplified to fit the

axisymmetric assumption. The actual rocket utilized a chamber sleeve insert that fit against the

injector to divide the hydrogen flow. Part of the hydrogen was diverted into the injector for

ignition with the oxygen, while the rest was ducted down milled slots in the chamber sleeve

insert for wall film cooling. The exit of the sleeve, then, was composed of a core flow of

hydrogen and oxygen combustion products and a coannular flow of hydrogen. By starting at the

sleeve exit, the modeling was simplitied to that of a single element coaxial injector, with a

precombusted oxidizer-rich core, surrounded by an outer annular flow of gaseous hydrogen

blanketing the wall.

The trends in performance were correctly predicted by both versions of RPLUS. However,

the measured performance values were underpredicted by both versions by three to four percent

(figure l a,b3). This result was unexpected, since the analyses assumed equilibrium composition

for the core flow and adiabatic walls, both of which have the effect of increasing the calculated

performance values. The low predictions of performance seemed to indicate inaccurate modeling

of the core flow/fihn flow mixing. Specifically, it was suspected that there were three possible

causes for the underprediction. First, it was thought that the turbulence modeling in the

combustion shear layer did not accurately represent the mixing between the core and film flows.

However, when RPLUS was run with and without the turbulence modeling, little difference was

found in the calculated flowfields. Second, the results of the turbulence modeling, coupled with

an estimated shear layer Reynolds number between 400 to 600, 4 indicated an unsteady laminar

flow. The Reynolds number was based on the width of the film cooling annulus. A laminar

shear layer would provide less mixing than the measured performance numbers suggest. If

laminar flow then prevails in the small rocket, the unaccounted for mixing could be due to

unsteady effects not well predicted by turbulence models. Finally, it was proposed that the

three-dimensional mixing effects at the film coolant injection point not modeled by RPLUS were

a possible cause. This report describes an experimental evaluation of the last proposed cause,

since the model results are insensitive to large variations in assumed turbulent mixing and

unsteady laminar flow phenomena are extremely difficult to evaluated in real rockets.

This study compares the experimental performance of a three-dimensional fuel fihn injection

geometry with an axisymmetric fuel film injection geometry. Pertbrmance tests were conducted
with the same 110 N, gaseous hydrogen/oxygen rocket used in the previous studies. The slotted

sleeve used previously and a second smooth sleeve of equivalent flow area were tested at the

same operating conditions. The slotted and smooth sleeves correspond to the three-dimensional

and axisymmetric injection of the fuel film coolant, respectively. An attempt was made to

operate over the same range of fuel film coolant(FFC) percentages as used in previous studies.

However, this could not be done due to problems with the thruster temperatures. This will be

discussed below. While no direct performance comparisons could be made to previous studies,

the results provide insight into the effect that fuel film injection geometry can have upon

performance, and the requirements for accurately modeling low thrust rockets. The purpose of

this study was to measure the difference in performance between an axisymmetric and a

three-dimensional film coolant injection geometry to determine if the lack of modeling of

three-dimensional effects was the cause of the performance underprediction. In the following

sections the RPLUS code is discussed first, followed by a description of the thruster and test

facility used to test the axisymmetric assumption of the RPLUS model. The test results were

then discussed, followed by the concluding remarks.

RPLUS

The RPLUS code was originally developed at NASA Lewis Research Center to analyze

supersonic combustion of scramjets and ramjets. 5 It was adapted to analyze subsonic-supersonic

flows in rockets, 2'6and applied to a 1000:1 area ratio nozzle 7 and to the 33:1 area ratio rocket that

is the subject of this study. 3 Another version of the code was developed at Pennsylvania State

University 4 for the same 33:1 area ratio rocket. That version included a modified turbulence

model and a solution algorithm to speed up convergence in the subsonic region.

RPLUS solves the full Navier-Stokes equations and the species equations in a coupled

manner, using the lower-upper symmetric successive over-relaxation (LU-SSOR) numerical

scheme. The hydrogen-oxygen combustion process is modeled by an 8-species, 18-step

finite-rate reaction mechanism. For the rocket used in this study, a grid composed of 202 axial

and 60 radial lines was used which was clustered in the region of high gradients in the flow.

Convergence was determined as the point when the mass flow rate was conserved to within two

percent. A more detailed description of the code can be found in references 2-4.

The plane of the chamber sleeve exit serves as the inflow surface or starting point for the

RPLUS calculation. The inflow surface is assumed to be composed of a pre-combusted,

oxygen-rich core flow, with a coannular flow of hydrogen. The RPLUS input includes the Mach

number, pressure, temperature and species mass fractions of the film and core flow streams.

There are assumptions which are known to both over and under predict performance. The

accuracies of the predictions were not quantified in the previous study.

The measured chamber pressure is assumed to be the pressure of the core and film flows.

The core flow temperature and species mass fractions are found from the Chemical Equilibrium

Composition (CEC) computer program 8, using the pressure and mixture ratio of the core as

inputs. The use of CEC, which computes the equilibrium composition of combustion products,
impliesan 100percentcorecombustionefficiency. This leadsto anoverestimation of the inlet

enthalpy in the core which should result in an overprediction of performance.

The current version of RPLUS also assumes adiabatic walls. The measured film temperature

accounts for the enthalpy that is added to the hydrogen from regenerative heating in the nozzle

and combustion chamber, but does not account for heat losses from the thruster. The adiabatic

wall assumption should result in higher computed performance values.

The film temperature is assumed to be the average of two thermocouples, located 180 degrees

apart, at the sleeve exit (position TCI on figure 2). The thermocouple tips are located flush with

chamber wall. The film is assumed to be 100 percent hydrogen. The input value of the Mach

number of the core flow is found from one-dimensional, isentropic relations. The current version

of RPLUS assumes the same Mach number for the core and film flows to facilitate the

calculation. This assumption could lead to an underestimation of the enthalpy in the sleeve and

lower computed performance values.

Apparatus and Procedure

Thruster

The RPLUS code models the regeneratively cooled gaseous hydrogen/gaseous oxygen thruster

built by the Gencorp Aerojet Propulsion Division under contract for NASA Lewis Research

Center. The thruster is designed to run at a nominal chamber pressure of 517 kPa giving a

nominal chamber thrust of 110 N with an overall mixture ratio of 8:1. The thruster has an

overall length of 24.8 cm, a combustion chamber diameter of 2.54 cm and a throat diameter of

1.27 cm. The Rao optimized 9 nozzle is bell shaped with a 33:1 area ratio. A schematic of the

thruster is shown on figure 3 and a photograph of the injector components is shown in figure 4.

Hydrogen enters the regeneratively cooled nozzle at the exit plane were it flows through the

milled passages within the nozzle and chamber walls towards the injector. Within the injector,

hydrogen enters a manifold where a flow splitting washer divides it into a core flow and film

cooling flow. The slot dimensions in the washer determines the percentage of hydrogen used as

fuel film coolant. The core hydrogen is injected radially just downstream of the spark plug tip.

All of the oxygen is injected radially through a platelet stack upstream of the spark plug tip. The

spark plug excites the oxygen as it flows past the tip where it mixes and ignites with the radially

injected hydrogen to form an oxygen rich core flow. The film cooling hydrogen first travels

down the sleeve insert between the combustion chamber and the regeneratively cooled wall. The

hydrogen is then dumped into the combustion chamber, where it forms a film coolant which

mixes with the core flow by shear layer interaction to increase the thruster performance.

The slotted sleeve insert originally designed for the Aerojet thruster, provided one of the test

cases for the present study. The second sleeve, referred to as the smooth sleeve, was designed

to have an equivalent coolant flow area without the slots, while having the same inside radius

so as not to change the combustion chamber size. The motive for having the same flow area,
wasin theory to have the same film velocities in both cases. Geometries of the two sleeves are

shown in figure 5 and both sleeves are pictured in figure 6. The thruster could not be tested at

the lower FFC levels used in the previous studies because of higher temperature profiles that

resulted in nozzle overheating. An injector modification was completed between the previous

studies and this one which was suspected as the cause of the higher temperatures. The injector

was modified by milling it out at the base of the spark plug to decrease the chance of the spark

arcing at the base instead of the tip of the spark plug; thereby improving the number of

successful ignitions.

The injector, nozzle, and combustion chamber are instrumented with static pressure taps and

thermocouples. Pressures measured include chamber, hydrogen injection, and oxygen injection

pressures. Thermocouples measure both internal and external wall temperatures. Internal wall

thermocouple positions are shown in figure 3. The exter,ml wall temperatures are not reported

in this work. The thruster is described in detail in reference 10.

Test Facility

The tests were conducted at the NASA Lewis Research Center in the low thrust propulsion

test facility. 1_ The facility was designed as a testbed for low thrust rockets that use gaseous

hydrogen and gaseous oxygen for propellants. It became operational in March of 1989 and can

be used for a range of testing programs from long duration steady state to short duration cyclic.

The facility was designed to test the 22-220 N thrust class rockets. All data are recorded on the

stand alone data acquisition system. During testing, real-time calculated performance parameters

are available.

The rocket hardware was installed in the cylindrical test tank that maintains a pressure

equivalent to an altitude of 36.6 km or a back pressure on the nozzle of 1.4 to 2.1 kPa. The

pressure in the tank, is maintained by a two-stage air ejector system. The nozzle axis was

oriented horizontally in the thrust stand. The thruster was mounted on flexible plates to insure

freedom of movement along the thrust axis. Also, all the pressure and propellant lines were

constructed from rigid tubing. The lines were mounted perpendicular to the thrust axis to

minimize the effect on the thrust measurement. The plume from the nozzle was fired into a

water cooled diffuser. After the diffuser, the exhaust enters a spray cart, where it is cooled by

water spray prior to entering the air ejectors. The flow rate of water spray is controlled by the

exiting exhaust temperature. The exhaust is finally vented through a pair of mufflers. A more

detailed facility description is available in reference 10.

Measurement Uncertainties

Hydrogen and oxygen mass flow rates were calculated using the measured inlet pressures,

the measured inlet temperatures and the discharge coefficient of the critical flow venturis with

corrections for real gas effects)'- Thrust was measured with a 440 N thrust measurement system.

Thrust calibrations were performed in-situ, at altitude conditions and with pressurized propellant

lines.
Themeasurementuncertaintiesin the performance values are determined using the JANNAF

recommended procedure, t3 With each measured quantity (pressure, temperature, etc.), there is

a random precision error and a bias error associated with the calibrations of the measuring

instrument and with the data acquisition of the measurement. The uncertainty of a performance

parameters, vacuum specific impulse, characteristic velocity and thrust coefficient, are a

combination of the precision and bias errors of the measured quantities, propagated to the

performance values. The largest contributors to the precision errors in this test series were the

venturi inlet pressure and venturi discharge coefficient calibration errors used in the mass

flowrate calculations. A zero drift in the thrust measurement load cell (probably due to bending

of the thruster flange under thermal loading) caused some error is thrust measurements. This was

reflected as a bias in the thrust measurement, calculated as the difference between the pretest and

posttest zero readings of the thrust.

The uncertainty of the vacuum specific impulse values ranges from 1.7 percent to 2.5 percent

in the positive direction (the direction of thrust bias) and typically 1.7 percent in the negative

direction. The uncertainty of the characteristic velocity values are typically +/- 1.7 percent. The

uncertainty for the thrust coefficient is 1.5 percent or less in positive direction, and 0.8 percent

or less in the negative direction.

Test Program

During testing the chamber pressures were held at 535 +/- 7 kPa, and the hydrogen mass

flow rates were approximately the same, for both the slotted and smooth sleeves. Two fuel film

cooling washers, 74.7 and 79.1%, were used over a mixture ratio range of 3 to 5 for both sleeves.

The test runs were conducted on consecutive test days. Changing the sleeve and %FFC washer

required that only the injector be removed from the nozzle/chamber portion, which remained in

the thrust stand. Tank pressures were generally between 1.4 and 1.8 kPa at the start of a test,

then dropped to 1.3 kPa once the ignition and flow through the diffuser were established.

Results and Discussion

The performance test data are presented sequentially in Table 1. The characteristic

velocity(often referred to as C*), specific impulse, and thrust coefficient for the 74,7 and 79.1%

FFC for both smooth and slotted sleeves are plotted in figures 7-10. Multiple tests were run at

each mixture ratio and the data were found to be repeatable.

Figures 7 and 8 show that the thrust coefficients are the same for the slotted and smooth

sleeves for the 74.7% and 79.1% fuel fihn cooling cases. Since the nozzle performed the same

over the different test cases, specific impulse and characteristic velocity can then be discussed

interchangeably. These performance parameters are plotted in figures 9a and 9b for the 74.7%

FFC case, and 10a and 10b for the 79.1% FFC case. Both performance parameter curves show

the same trends at a given FFC percentage. The specific impulse and characteristic velocity

graphs (figures 9a and 9b) show that the smooth sleeve had better performance than the slotted

sleeve for the 74.7% FFC. At a mixture ratio of four the performance is greater by only two
percent, but at a mixture ratio of five the performance was seven percent greater. The seven

percent is well above the uncertainties for this experiment. For the 79.1% FFC case (figures 10a

and 10b), the performance difference between smooth and slotted sleeves is within the range of

the experimental uncertainty. At a mixture ratio of three, the slotted sleeve performs better by

approximately two percent. At a mixture ratio of four and five, the smooth sleeve performs

better by four and two percent, respectively. The results of this performance study show that the

thruster with the three-dimen-sional sleeve did not have the greater measured performance.

Therefore the underprediction of performance by the axisymmetric RPLUS code is not due to the

axisymmetric "smooth" assumption for the FFC percentages tested.

Thermocouple measurements obtained inside the combustion chamber at the locations shown

in figure 3 are shown in figures I l and 12. For the 74.7% FFC case, the slotted sleeve is hotter

at mixture ratios of four and five than the smooth sleeve. At a mixture ratio of three the

temperature profiles are almost the same. For the FFC case of 79.1%, the slotted sleeve has

higher temperatures only for a mixture ratio of five. At the mixture ratio of four the smooth

sleeve experiences higher temperatures and, at a mixture ratio of three the profiles are almost the

same. Temperature results, at the TCI position, are used to calculate the hydrogen film coolant

velocity to core velocity ratios, shown in Figures 13 and 14. Note that the velocity ratios are

near unity and that they are generally higher for the higher film coolant percentages. Also,

except for the mixture ratio of four, the smooth wall generally ran cooler than the slotted wall

sleeve, making it difficult to correlate the temperature profiles and the performance.

The axisymmetric sleeve was built to match the film flow area of the one modeled in

previous analytical studies. The intent was to compare two injection geometries of the same flow

area. However, the results indicate that the edge thickness of the sleeves may play a more

significant role in the mixing than the injector geometry. As shown in figure 6, in order to

maintain the same flow area for both sleeves the edge or splitter plate thicknesses had to be

different. At the sleeve edge or flow splitter where the film coolant flow and core flow meet,

the slotted sleeve was relatively thin, i.e. 0.076 cm, in comparison to the smooth wall sleeve

with an edge thickness of 0.132 cm. The thicker edge of the smooth sleeve may be creating

larger recirculations zones between the film coolant and core flows, which enhances mixing and

thus performance. The significance of this geometry change is shown in the ratio of the edge

thickness to film coolant gap. This ratio is much greater for the smooth sleeve compared to the

slotted sleeve, 1.21 and 0.46, respectively, allowing the unsteady flow or turbulence to interact

with more of the film coolant. The effect of the taper on the end of the splines of the slotted

sleeve is uncertain, because it should result in a thinning and slowing down of the film at the

edge of the sleeve. This effect could be studied by tapering a smooth sleeve to the same

dimension as a slotted sleeve, and conduct a performance study on both sleeves. In this way

perhaps some of the difference in performance between the smooth and slotted sleeves could be

explained.

Concluding Remarks

An experimental performance comparison was conducted on a 110 N(25 Ibf) gaseous
hydrogen oxygen rocket to establish the effect of hydrogen film coolant sleeve design on rocket

performance. One sleeve had milled slots running axially down the surface, the other was

designed with a smooth surface. Both were designed to give the same flow area and therefore,

the same film velocities.

Previous work at Pennsylvania State University 4 compared the experimental performance

characteristics of the same slotted sleeve and thruster combination with the analytical results

obtained from the RPLUS code. In the code the film coolant sleeve was modeled as a smooth

sleeve to simplify calculations. The RPLUS results underpredicted the experimental performance

results by three to four percent. In the study, it was suggested that the three-dimensional effects

of the slotted sleeve increased the mixing, which was not accounted for in the axisymmetric

model.

Test were conducted over a mixture ratio range of 3 to 5 at FFC values of 74.7% and

79.1% to determine if the slotted sleeve was actually responsible for the increased performance

observed in the model. Results showed that the smooth sleeve configuration had greater

characteristic velocity and specific impulse, for the 74.7% FFC case and approximately the same

or slightly better performance for the 79.1% FFC case. These results showed the three-dimensional slots were not responsible for the underprediction by the RPLUS code. Because of the

difference in geometries between sleeves, the smooth sleeve having the thicker edge at the core

and film coolant flow interface may enhance the mixing between the two flows, thus, improving

the performance.

.

.

.

,

5_

.

References

Richter, P.G., and Price, H.G., "Proven Long-Life Hydrogen/Oxygen Thrust Chambers

for Space Station Propulsion", NASA TM-88822, 1986 JANNAF Propulsion Meeting,

August 1986.

Kim, S.C., and VanOverbeke, T.J., "Calculations of Gaseous Hydrogen/Oxygen

Thruster", AIAA Paper 90-2490, July 1990.

Reed, B.D., Penko P.F., Schneider S.J., and Kim, S.C., "Experimental and Analytical

Comparison of Flowfields in a l10N(251bf) H202 Rocket.", AIAA-91-2283, June 1991.

Weiss, J.M., and Merkle, C.L., "Numerical Investigation of Reacting Flowfields in

Low-Thrust Rocket Engine Combustors", AIAA-91-2080, June 1991.

Shuen, J.S., and Yoon, S., "Numerical Study of Chemically Reacting Flows Using a

Lower-Upper Symmetric Overrelaxation Scheme", AIAA Paper 88-0436, January 1988.

VanOverbeke, T.J., and Sheun, J.S., "A Numerical Study of Chemically Reacting Flow in

Nozzles", AIAA Paper 89-2793, July 1989.
7. Kim, S., "Numerical Study of High-Area Ratio H2/02 Rocket Nozzles", AIAA-91-2434, June

1991.

o Gordon, S., and McBride B., "Computer Program for Calculation of Complex Chemical

Equilibrium Composition, Rocket Performance, Incident and Reflect Shocks, and ChapmanJoguet Detonations", NASA SP 273, March 1976.

9. Nickerson, Gary R., Dang Anthony L., and Dunn Stuart S., "The Rao Method Optimum

Nozzle Contour Program", NAS8-26863, SN91, February 15, 1988.

10. Robinson, Phillip J., "Space Station Auxiliary Thrust Chamber Technology", Final Report

2210-90-FR, NASA Contractor Report 185296, July 1990.

11. Arrington, L., and Schneider, S., "Low Thrust Rocket Test Facility", AIAA-90-2503, July

1990.

12. "Fluid Meters: Their Theory and Application", American Society of Mechanical Engineers,

1971.

13. "Handbook for Estimating the Uncertainty in Measurements Made with Liquid Propellant

Rocket Engine Systems", Chemical Propulsion Information Agency Document Number 180,

April 1969.
230O

Figure la:

22OO

v

2100

<

2000

w

o

180o

OPEN- EXPERIMENTALDATA

SOLID- RPLUSRESULT

FUEL FILM COOLING %

O 55.4

n 5t.g

A 4g.7

• • 47,0

J -' ". i , i i ! l i

4 5 6 7 8 9 10

MIXTURE RATIO

Experimental and RPLUS Comparison of Characteristic Exhaust Velocity vs. Mixture Ratio

420 -

400 '

380  ̧

',2_ 3so,

340,

320

Figure lb:

O

OPEN- EXPERIMENTALDATA

SOL:D- RPLUSRESULT

FUEL FILM COOLING %

O 55 4

• [] 51 ,g

A 49,7

• 470

• i I I I I , l - i

4 5 6 7 8 9 10

MIXTURE RATIO

Experimental and RPLUS Comparison of Specific Impulse vs. Mixlure Ratio

10
7.62 cm _ T_

8.89 ¢m TC3 "

I

_-_ 9.66 on -_--_ TO4

10.29 cm------_ TC5

Figure 2. Axial Positions of Internal Thermocouples.

Fuel Inlet

Spark Plug _

g=

- " Sleeve

nl

Figure 3. Aerojet 110 N(25 lbf) Gaseous H2/O2 Thruster.

11
c_o00P000o

,4

/

m

12
A

EUd

EU

O0

03o

c)

EE

UU

'3>

_D L_>W

LJ

_J

(/J

t,i

FI-- 0

_]

\l

>

(/)

bJ

bJ/

b') T

E3

E3

_E

b')

(_

E Ca U D_. --6

E

EEU C)

UU _-__ -H

"° c5_o

c_m c_

II II _p _- _ m).,_] ,v > > c c C" 0 c cqD Z.

(_ _C _C_

(,,') (10 (_) (_D (_ r_O

CO "_" LIn _D E E U U

OJ

',De')

°cd cd

II

>>

.--_ C'O

@DF

_"h C©

r'--

;_._J

L_

]3
14

0
Test Run

#

4578

12

15

16

17

18

21

22

24

25

26

27

30

32

33

34

35

36

37

41

42

44

47

48

4g

50

51

55

56

57

56

59

60

61

63

64

66

67

68

69

70

71

72

75

76

77

78

79

80

81

82

84

87

88

8g

90

91

Mixture

Ratio

3.33

4.13

5.26

5.15

4.15

3.24

4.20

5.28

5.30

4.14

4.17

3.31

3.28

4.15

5.33

3.30

3.23

4.16

4.19

4.24

4.15

4.18

3.18

3.22

3.25

3.25

5.19

5.21

5.25

5.23

4.17

4.12

4.13

4.14

4.13

3.08

3.05

3.09

3.09

3.07

3.07

5.16

5.18

5.21

5.21

5.22

4.06

4.01

4.06

4.04

5.06

5.11

5.12

5.12

5.15

3.22

3.27

3.27

3.28

3.26

Table 1. Experimental Test Data

Specific Characteristic Thrust

Impulse Exhaust Velocity Coefficient

(sec) (n_'sec)

380.9 2086 1.7g

377.6 2117 1.75

359.5 2008 1.76

364.7 2017 1.77

380.3 2126 1.76

383.1 21 08 1.78

378.5 21 08 1.76

360.5 2000 1.77

362.1 2001 1.78

380.6 21 04 1.78

379.7 2120 1.76

383.5 2123 1.77

381.3 2123 1.76

379.5 2101 1.77

361.3 2019 1.76

382.8 2114 1.78

403.1 2231 1.77

386.1 2138 1.77

387.1 21 53 1.76

387.8 2162 1.76

386.5 2169 1.75

388.1 2159 1.76

404.9 2232 1.78

405.5 2239 1.78

402.7 2222 1.78

401.6 2235 1.76

391.8 2162 1.78

391.1 21 56 1.78

393.1 2162 1.78

388.5 21 47 1.78

376.7 21 05 1.76

376.3 2093 1.76

376.5 2090 1.77

376.5 2087 1.77

378.0 2090 1.77

385.9 2150 1.76

383.4 2137 1.76

382.4 2131 1.76

382.8 2119 1.77

384.2 2126 1.77

381.9 21 28 1.76

361.3 1998 1.77

361.9 2016 1.76

358.5 2005 1.75

359.6 2005 1.76

359.6 2003 1.76

388.5 2169 1.76

392.3 2169 1.78

388.8 21 78 1.75

389.1 21 72 1.76

367.0 2040 1.77

365.1 2014 1.78

373.8 2035 1.80

373.5 2035 1.80

370.1 2033 1.79

378.9 2117 1.76

379.4 2078 1.79

378.8 2078 1.79

376.7 2073 1.78

373.7 2070 1.77

15

FFC

(%)

74.7

74.7

74.7

74.7

74.7

74.7

74.7

74.7

74.7

74.7

74.7

74.7

74.7

74.7

74.7

74.7

74.7

74.7

74.7

74.7

74.7

74.7

74.7

74.7

74.7

74.7

74.7

74.7

74.7

74.7

79.1

79.1

79.1

7g.1

79.1

7g.1

79.1

79.1

79.1

79.1

79.1

79.1

79.1

79.1

79.1

79.i

79.1

79.1

79.1

79.1

79.1

79.1

79.1

79.1

79.1

79.1

79.1

79.1

79.1

79.1

Sleeve

Configuration

slotted

slotted

slotted

slotted

slotted

slotted

slotted

slotted

slotted

slotted

slotted

slotted

slotted

slotted

slotted

slotted

smooth

smooth

smooth

smooth

smooth

smooth

smooth

smooth

smooth

smooth

smooth

smooth

smooth

smooth

slotted

slotted

slotted

slotted

slotted

slotted

slotted

slotted

slotted

slotted

slotted

slotted

slotted

slotted

slotted

slotted

smooth

smooth

smooth

smooth

smooth

smooth

smooth

smooth

smooth

smooth

smooth

smooth

smooth

smooth
E

_J

=

J=

[-

2.00

1.95

1.90

1.85

1.80

1.75

1.70

1.65

1.60

1.55

1.50

0 Slotted

[] Smooth

3 4 5 6

Mixture Ratio

Figure 7. Comparison of Thrust Coefficient Versus Mixture Ratio for

74.7% FFC at a Chamber Pressure of 535 +1- 7 kpa (77 +/- 1 psia).

e_

E

_J

2.00

1.95

.90

1.85

1.80

[]

,IB ,-,

1.75

1.70

1.65

; .60

1.55

1.50

3

A

0 Slotted

[] Smooth

r

I !

4 5 6

Mixture Ratio

Figure 8. Comparison of Thrust Coefficient Versus Mixture Ratio for

79.1% FFC at a Chamber Pressure of 535 +/- 7 kpa (77 +/- 1 psia).

16
g

2600

2500

2400

23OO

._ 2200

2100

M

2OOO

1900

9_p

¢o o

o SlotP..d

[] Smooth

I I

3 4 5 6

Mixture Ratio

Figure 9a. Comparison of Characteristic Velocity Versus Mixture Ratio for

74.7% FFC at a Chamber Pressure of 535 +/- 7 kpa (77 +/- 1 psia).

l=

¢d

470

460 _

450 _

440 _

43O _

420 _

410 _

400 _

390 _

380 _

370 _

360

35O

3

o

o Slotted

[] Smooth

4 5 6

Mixture Ratio

Figure 9b. Comparison of Specific Impulse Versus Mixture Ratio for

74.7% FFC at a Chamber Pressure of 535 +/- 7 kpa (77 +/- 1 psia).

17
==

s=

26O0

2500

2400

!

2300

2200

._ 2100

2OO0

Elm

0 Slotted

[] Smooth

19oo I !

3 4 5 6

Mixture Ratio

Figure 10a. Comparison of Characteristic Velocity Versus Mixture Ratio for

79.1% FFC at a Chamber Pressure of 535 +/- 7 kpa (77 +/- 1 psia).

H

J.q

470

460

450

440

430 -

42O -

410

4OO

39O

380

37O

36O

350

g

n

%

", ! |

3 4 5

Mixture Ratio

o Slotted

D Smooth

Figure 10b. Comparison of Specific Impulse Versus Mixture Ratio for

79.1% FFC at a Chamber Pressure of 535 +/- 7 kpa (77 +/- 1 psia).

18
E

900

800

700

600

500

400

300

6

! ! .... i I " "

!

7 8 9 10

Distance along Thruster axis from Injector End , cm

1

o MR=3.24,slolled

g ' MR==4.15,slotted

& MR=5.15,slotted

• MR=3.22,smooth

= MR=4.16,smooth

• MR--5.19,srnooth

Figure 11. Internal Temperatures versus Axial Position

from Injector End for 74.7% FFC.

E

¢d

[-

900

800

700

60O

500

400

300

' J ' 1_0 ' 7 8 9

Distance along Thruster AxEs from Injector End , cm

Figure 12. Internal Temperatures versus Axial Distance

from Injector End for 79.1% FFC.

o MR=3.09,slotted

[] MR=4.14,slotted

A MR=5.21,slotted

• MR=3.27,smooth

• MR=4.06,smooth

• MR=5.15,smooth

i9
1.7

,-;,

m

O

w

e

1.6'

1.5'

1.4

1,3

1.2

1.1

1,0

0.9

3

i-I

o

oo _P

0 Slotted Sleeve

o t3 Smooth Sleeve

! !

4 5 6

Mixture Ratio

Figure 13. The Ratio of Film Velocity to Core Velocity versus

Mixture Ratio for 74.4% Film Flow Cooling at a Chamber

Pressure of 535 +/- 7 kpa (77 +/. 1 psia).

,2'

O

>.

e

1.7

1.6

1.5

1.4

1.3

1.2 q_ _

1,1

1.0

0.9

3

aa

eo

O SlottedSleeve

13 SmoothSleeve

I !

4 5 6

Mixture Ratio

Figure 14. The Ratio of Film Velocity to Core Velocity versus

the Mixture Ratio for 79.1% Film Cooling at a Chamber

Pressure of 535 +/- 7 kpa (77 +/- 1 psia).

2O
f _ "1]
Form Approved

REPORT DOCUMENTATION PAGE OM8 No.0Z04-0188

Public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing Instructions, searching existing data sources.

gathedng and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this

collection of information, including suggestions for reducing this burden, to Washington Headquarters Services, Dk'ectorate for information Operations and Reports. 1215 Jefferson

Davis Highway. Suite 1204. Arlington, VA 22202-4302, and to the Office of Management and Budget, Papecwork Reduction Project (0704-0188), Washington, DC 20503.

1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

December 1992 Technical Memorandum

4. TITLE AND SUBTITLE

Performance Comparison of Axisymmetric and Three-Dimensional Hydrogen

Film Coolant Injection in a 110N Hydrogen/Oxygen Rocket

6. AUTHOR(S)

Lynn A. Arrington and Brian D. Reed

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Lewis Research Center

Cleveland, Ohio 44135-3191

9. SPONSORING/MONITORING AGENCY NAMES(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, D.C. 20546-0001

5. FUNDING NUMBERS

WU-506--42-31

8. PERFORMING ORGANIZATION

REPORT NUMBER

E-7490

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA TM- 105967

11. SUPPLEMENTARY NOTES

Prepared for the 28th Joint Propulsion Conference and Exhibit sponsored by the AIAA/SAE/ASME and ASEE, Nashville, Tennes_e,

July 6--8, 1992. Lynn A. Arrington, Sverdrup Technology, Inc., Lewis Research Center Grou _, Brook Park, Ohio 44142 and Brian D.

Reed, NASA Lewis Research Center. Responsible person, Brian D. Reed, (216) 433-7489).

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified - Unlimited

Subject Category 20

12b. DISTRIBUTION CODE

la. ABSTRACT (Maximum 200 words)

An experimental performance comparison of two geometrically different fuel film coolant injection sleeves was conducted on a lION gaseous hydrogen/oxygen rocket. One sleeve had slots milled axially down the walls and the other had

a smooth surface to give axisymmetric flow. The comparison was made to investigate a conclusion in an earlier study

that attributed a performance underprediction to a symplifying modeling assumption of axisymmetric fuel film flow. The

smooth sleeve had higher overall performance at one film coolant percentage and approximately the same or slightly

better at another. The study showed that the lack of modeling of three-dimensional effects was not the cause of the

performance underprediction as speculated in earlier analytical studies.

14. SUBJECT TERMS

Thruster; Film flow cooling; Hydrogen/oxygen; Low thrust; Hydrogen injection cooling;

RPLUS

17. SECURITY CLASSIFICATION

OF REPORT

Unclassified

18. SECURITY CLASSIFICATION

OF THIS PAGE

Unclassified

NSN 7540-01-280-5500

19. SECURITYCLASSIFICATION

OF ABSTRACT

Unclassified

15. NUMBER OF PAGES

22

16. PRICE CODE

A03

20.' LIMITATION OF ABS_CT

Standard Form 298 (Rev. 2-89)

Prescribedby ANSI Std.Z39-1B

298-102