Timothy D. Smith, Mark D. Klem, and Kevin J. Breisacher

Glenn Research Center, Cleveland, Ohio

Shahram Farhangi and Robert Sutton

Boeing, Rocketdyne Division, Canoga Park, California

Experimental Evaluation of a Subscale Gaseous

Hydrogen/Gaseous Oxygen Coaxial

Rocket Injector

NASA/TM—2002-211982

November 2002
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Timothy D. Smith, Mark D. Klem, and Kevin J. Breisacher

Glenn Research Center, Cleveland, Ohio

Shahram Farhangi and Robert Sutton

Boeing, Rocketdyne Division, Canoga Park, California

Experimental Evaluation of a Subscale Gaseous

Hydrogen/Gaseous Oxygen Coaxial

Rocket Injector

NASA/TM—2002-211982

November 2002

National Aeronautics and

Space Administration

Glenn Research Center
Available from

NASA Center for Aerospace Information

7121 Standard Drive

Hanover, MD 21076

National Technical Information Service

5285 Port Royal Road

Springfield, VA 22100

Available electronically at http://gltrs.grc.nasa.gov
NASA/TM—2002-211982 1

Experimental Evaluation of a Subscale Gaseous Hydrogen / Gaseous

Oxygen Coaxial Rocket Injector

Timothy D. Smith

Mark Klem

Kevin J. Breisacher

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Shahram Farhangi

Robert Sutton

Boeing Rocketdyne Division

Canoga Park, California

Abstract

The next generation reusable launch vehicle may utilize a Full-Flow Stage Combustion (FFSC) rocket

engine cycle. One of the key technologies required is the development of an injector that uses gaseous

oxygen and gaseous hydrogen as propellants. Gas-gas propellant injection provides an engine with

increased stability margin over a range of throttle set points. This paper summarizes an injector design and

testing effort that evaluated a coaxial rocket injector for use with gaseous oxygen and gaseous hydrogen

propellants. A total of 19 hot-fire tests were conducted up to a chamber pressure of 1030 psia, over a range

of 3.3 to 6.7 for injector element mixture ratio. Post-test condition of the hardware was also used to assess

injector face cooling. Results show that high combustion performance levels could be achieved with gasgas propellants and there were no problems with excessive face heating for the conditions tested.

Symbols

A Area, in2

Cp specific heat based upon constant pressure, BTU/lbm-°R

C* characteristic exhaust velocity, ft/s

C*EXP experimental characteristic exhaust velocity, ft/s

C*MA calculated mass average characteristic exhaust velocity, ft/s

C*Th(x) theoretical, one-dimensional characteristic exhaust velocity obtained from the CEA program, ft/s

g proportionality constant (32.2 lbm-ft/lbf-s)

h enthalpy, BTU/lbm

Kc empirical injector design function, dimensionless

m propellant mass flow, lbm/s

mT total propellant mass flow, lbm/s

MRT injector inlet momentum ratio (fuel divided by oxidizer), dimensionless

O/F propellant mixture ratio (oxidizer flow divided by fuel flow), dimensionless

Pc combustion chamber pressure, psia

R gas constant, lbf-ft/lbm-°R

T propellant temperature, °R
NASA/TM—2002-211982 2

V Velocity, ft/s

VR injector inlet velocity ratio (fuel divided by oxidizer), dimensionless

ηC* characteristic exhaust velocity efficiency, percent

γ ratio of specific heats

ρ density, lbm/in3

Subscripts

BLC boundary layer hydrogen

C injector core flow

H2 hydrogen

O2 oxygen

t rocket throat

T results based on total massflow

Background

The Full-Flow Stage Combustion (FFSC) rocket engine cycle continues to be a promising system for

the next generation of space launch vehicles. The FFSC cycle provides the potential for increased specific

impulse efficiency over a wide range of power settings due to improved throttling stability margin. The

improved stability margins result from the use of gaseous propellants, which provide consistent pressure

drop ratios above stability criteria levels. More importantly, the FFSC cycle can improve engine reliability

and reduce complexity. The key to the cycle is the use of both fuel-rich and oxidizer-rich gases to power

the fuel and oxidizer turbopumps respectively.1 Turbine inlet temperatures are lowered because all of the

propellants are routed through preburners to power the turbomachinery. This is unlike other systems, such

as fuel rich staged combustion, where the turbomachinery is primarily driven by the amount of fuel flow.

The FFSC cycle also reduces system complexity by eliminating the need for inner propellant seals on the

oxygen turbopump.

One of the key components to the performance of a FFSC rocket engine is the main injector. A first

look would indicate that the development of gaseous hydrogen-gaseous oxygen injectors would be a

straightforward process when compared to liquid systems. The elimination of atomization and vaporization

from the combustion process should dramatically simplify the design process. However, as with any rocket

component development, this type of injector has its own unique set of design challenges. As discussed

in reference 2, propellant mixing is a critical driver to development of gas-gas injectors. Not only does

propellant mixing drive the performance level, it also has a significant impact on face cooling and injector

durability.

Over the years work has been carried out to investigate the use of gaseous propellants for main

injectors2345. The bulk of work has been focused on the investigation of gaseous hydrogen and liquid

oxygen. However, that work scale has been on small thrusters for on orbit applications, not large-scale

launch engines. Aerojet performed a comprehensive study6 of gas-gas rocket injectors under NASA

contract in the 1970’s. The program evaluated a wide-range of both single- and multi-element injector

configurations in both cold-flow and hot-fire tests. The results also include a design guide based upon the

empirical results. However, much of the data was generated for smaller scale elements, which operated at

lower flowrates and pressures than those required by current FFSC systems.

The search for the next generation of reusable launch vehicles has refocused efforts to examining the

use of the FFSC. To address the need for further development of gas-gas injectors, NASA initiated a

cooperative effort between government, industry, and academia to investigate injector element concepts7

.

As part of this program, both single-element and multi-element testing was conducted on a variety of
NASA/TM—2002-211982 3

injector concepts. Single element testing was used to obtain species measurements using Raman laser

techniques8

. Results from that project further enhanced the database by providing multi-element data at

chamber pressures up to 1200 psia.

The objective of the current gas-gas injector test program is to build upon the previous results and

examine the performance of a coaxial injector during hot-fire. Testing was conducted with a multi-element

injector up to a chamber pressure of 1050 psia using ambient temperature hydrogen and oxygen. Pressure

and temperature measurements were used to determine combustion performance. Post-test condition of the

hardware was also used to address injector face cooling.

Test Facility

Testing was conducted in Cell 32 of the Research Combustion Laboratory9. The cell is part of a highly

flexible facility that encompasses several test cells for both combustion research and material evaluation.

Cell 32 is capable of testing with a variety of propellants including gaseous oxygen and gaseous hydrogen,

liquid oxygen and liquid hydrogen, and hydrocarbon propellants. A 1500 psia water system is also

available for combustor cooling applications. The facility can accommodate combustors that generate thrust

levels up to 2000 lbf with chamber pressures to 1200 psia. The propellant systems consisted of a gaseous

oxygen circuit and a gaseous hydrogen circuit fed from 2400 psia roadable tube trailers. The gaseous

hydrogen circuit was split into two legs inside of the test cell to supply both the main propellant and

boundary layer cooling propellant. Each propellant circuit used calibrated subsonic venturis for flow

measurement.

Test Hardware

Rocketdyne, as part of a cooperative test program, provided the combustion chamber, including

injector, window, and nozzle housings. Figure 1 shows a test firing with the windows installed during a

previous program.
NASA/TM—2002-211982 4

The combustion chamber is a heat sink configuration made of stainless steel with zirconia coated

nozzle inserts. It can be observed in the photograph that there is significant thermal mass available for heat

dissipation. Table 1 provides the relevant dimensions for the test configuration.

TABLE 1.—COMBUSTION CHAMBER

DIMENSIONS

Dimension

Chamber diameter, in. 4.00

Throat diameter, in. 1.18

Contraction ratio 11.49

Exit diameter, in. 1.84

Exit area ratio 2.42

Chamber length, in. 10.60

The hardware consists of a number of interchangeable parts that can be modified depending upon the

test requirements. As shown in figure 2, the hardware can accommodate two large side windows for

optical access along with four smaller windows on the top and bottom for optical access.

Figure 2.—Schematic of heat sink combustion chamber assembly.

Injector Nozzle

TCW1

TCW5

TCW2 TCW3 TCW4

Top

Bottom

Ignition was provided by a spark plug located in the top, downstream window location. The

combustion chamber nozzle consisted of two parts, the main nozzle housing which attached to the chamber

body, and a smaller throat insert. The throat insert could be varied to provide a wide variety of throat

diameters. The injector consisted of two parts: a main housing and an injector insert. The injector housing

was recessed into the combustion chamber so the injector face plane was just upstream of the large side

windows. Oxygen was fed to the injector through a single line located straight back from the center of
NASA/TM—2002-211982 5

the injector face. Both main and boundary layer hydrogen were fed through multiple feed lines to provide

even distribution around the face. The modular design of the combustion chamber provides a high level of

flexibility for injector evaluation.

As stated previously, there is no large database to use when designing or analyzing gas-gas rocket

engine injectors. To initiate the design, ground rules were established based on facility pressure, thrust,

and flow rate limits. Nominal design was based upon a chamber pressure of 1000 psia at a mixture ratio of

six. The designs were compared to the gas-gas injector design procedure found in reference 7. However,

since that work was at low chamber pressures this analysis can be considered an extrapolation. The two

parameters of interest, based on reference 7, are the mixing efficiency and the energy release efficiency.

Following the design procedure, the mixing efficiency was estimated to be 90.5 percent and the energy

release efficiency to be above 99 percent for the current design.

The injector tested during this effort was a seven-element shear coaxial design with oxygen fed though

the main post and hydrogen through the surrounding annulus. Figure 3 is a pretest photograph of the multielement injector.

The main injector design had a single element located in the center and 6 elements equally spaced on

a diameter of 1.1 in. The oxygen posts were 0.197-in. diameter, with a 0.025-in. wall thickness. Hydrogen

was injected through a straight, 0.0175-in. annular gap around the oxygen post. A porous sintered wire

mesh faceplate was used to provide face cooling. Boundary layer cooling (BLC) was provided through

the 120, 0.050-in. diameter cooling jets located along the outer face diameter. The boundary layer coolant

and main hydrogen propellant circuits were fed and measured separately. A plate is located behind the face,

at the same diameter as the face screws, to keep the gas supplies from mixing. As with the combustion

chamber, the part was manufactured from stainless steel.

Figure 4 is a schematic showing propellant feed system to the hardware along with instrumentation

locations.
NASA/TM—2002-211982 6

Figure 4.—Schematic showing propellant circuits and instrumentation.

Main gaseous hydrogen

Main gaseous oxygen

BLC gaseous hydrogen

T P T P

P

T P T P

P P

P

T P T P

P

Combustor

Venturi Valve

Valve

Valve

Venturi

Venturi

Redundant pressure and temperature measurements were made at the injector inlet connections to

the feed lines on each circuit. Chamber pressure was measured at two axial locations, 1.9-in. and 7.4-in.

downstream of the injector face. The measurements were separated radially by approximately 90 degrees.

The combustion chamber also could accommodate a variety of thermocouple measurement locations. For

this test series, five locations (figure 2) were used to measure gas temperature along the internal wall of the

combustion chamber. Each thermocouple was mounted such that the tip was slightly into the combustion

zone. The location of the thermocouples is described in table 2.

TABLE 2.—COMBUSTION CHAMBER

THERMOCOUPLE LOCATIONS

Thermocouple Axial

location,

in.

Radial

location,

deg

TCW1 1.9 152.5

TCW2 1.9 27.5

TCW3 3.4 27.5

TCW4 5.5 45

TCW5 5.5 315

Experimental Results

Performance

A total of 19 hot-fire tests were completed with the coaxial injector design over a range of chamber

pressure and mixture ratio conditions. All propellants were injected at ambient temperature (529 to 555 °R).

The injector survived the testing with no significant discoloration or erosion to either the faceplate or

oxygen posts.

The key values measured for each run were chamber pressure, propellant flowrates, and combustion

chamber wall temperatures. The target mixture ratio levels were based upon injector element (core) mixture

ratio as opposed to total mixture ratio, which includes the boundary layer coolant. These parameters were

defined by the following:

H BLC

O

T m m

m

F

O

& &

&

+

 = 

 





2

2 (1)
NASA/TM—2002-211982 7

2

2

H

O

C m

m

F

O

&

&  = 

 



 (2)

For the current series of tests, the main figure of merit for injector performance is characteristic exhaust

velocity (C*EXP). By definition,

m

PcA g C t

EXP & = * (3)

To obtain C* efficiency, the calculated C* EXP values were divided by the theoretical, one-dimensionalequilibrium values obtained from the Chemical Equilibrium Composition and Applications10 (CEA)

program. Inputs to the CEA code included chamber pressure, mixture ratio, chamber contraction ratio,

and propellant inlet temperature. Propellant inlet temperatures were input so the code could calculate the

corresponding inlet enthalpy. The characteristic exhaust velocity efficiency is defined as:

*

( )

*

*

Th ODE

C C

C EXP η = (4)

Due to the abundance of boundary layer cooling present in the system, it is difficult to determine the true

performance of the injection elements. The flow within the combustion chamber could vary between fully

mixed and completely stratified, where the core and boundary layer stay in streamtubes. To examine the

differences in performance, three sets of analysis were conducted to bound the performance range. The

analysis was conducted based on the assumption of fully mixed flow, core flow with ambient temperature

boundary layer cooling flow, and core flow with 2000 °R boundary layer cooling flow.

For the analysis of the fully mixed conditions, combustion performance was based upon measured

chamber pressure and total mixture ratio. This analysis assumes complete mixing between the core flow

from the injector elements and the boundary layer coolant. The input values to CEA for mixture ratio

would be from equation 1 and combustion performance calculated from equations 3 and 4.

The stratified flow analysis breaks the flow up into two streamtubes, one based on the core flow

from the injector, and the other a boundary layer flow of pure hydrogen at specified temperatures. The

theoretical C* values must be calculated for each streamtube. The hydrogen boundary layer C* can be

calculated from the following equation based upon the assumed temperature (TBLC), gas constant (°R),

and ratio of specific heats (γ).

1

1

*

1

2

2

−

+





 





+

=

γ

γ

γ

γ

γRH TBLC g

C BLC (5)

For hydrogen the gas constant is 766.4 ft-lbf/lbm-°R and the ratio of specific heats is 1.4. From

equation 5, for a 2000 °R hydrogen boundary layer the C* is 10 253 ft/s and over the range of ambient

inlet temperatures is 5270 to 5405 ft/s. The core C* is calculated from the CEA program based upon the

measured chamber pressure and mixture ratio through the injector elements (equation 2).

A new theoretical C* is then calculated on a mass average basis from the core C* and the boundary

layer C*. For the case where the boundary layer streamtube was at ambient temperature, there is no heat

transfer assumed from the core flow. However, for the case where streamtube flow is heated to 2000 °R,
NASA/TM—2002-211982 8

heat must be transferred from the core flow to the boundary layer flow. The amount of heat transferred

from the core flow to the boundary layer can be determined by calculating the change in enthalpy required

to raise the incoming hydrogen flow to 2000 °R.

( ) ∆hBLC = CpBLC −TH2Inlet = −∆hC 2000 (6)

The new core C* was calculated from CEA by adjusting the inlet enthalpy of the hydrogen based upon

the reduced core enthalpy calculated from equation 6.

* *

( )

*

, BLC

T

BLC

Th C

T

C

Th MA C

m

m

C

m

m

C &

&

&

& = + (7)

Combustion efficiency for the stratified flow cases can then be calculated by modifying equation 6 as

follows:

*

,

*

*

Th MA

Exp

C

C

C

MA η = (8)

Table 3 presents a summary of the conditions, flowrates, and results for the test series.

TABLE 3.—HOT-FIRE TEST RESULTS FOR GASEOUS HYDROGEN / GASEOUS OXYGEN

COAXIAL INJECTOR

RDG Chamber

pressure,

psia

Total

mixture

ratio

Injector

mixture

ratio

Oxygen

flow,

lbm/s

Injector

H2

flow,

lbm/s

BLC

H2

flow,

lbm/s

ηC*

mixed,

percent

ηC*

TBLC=Amb.,

percent

ηC*

TBLC=2000 °R,

percent

1 513.0 1.39 6.15 1.31 0.21 0.73 98.4 115.8 103.9

2 513.6 1.40 6.21 1.33 0.21 0.73 97.4 114.9 104.1

3 471.1 0.96 3.67 1.02 0.28 0.79 101.2 110.8 106.4

4 587.6 1.39 5.35 1.48 0.28 0.79 99.1 114.1 105.4

5 471.4 0.97 3.77 1.03 0.27 0.78 101.0 111.1 106.4

6 464.2 1.12 5.81 1.08 0.19 0.78 100.1 117.1 106.7

7 712.1 1.28 3.43 1.71 0.50 0.84 101.3 109.5 105.2

8 707.8 1.26 3.29 1.70 0.52 0.83 100.9 108.6 104.6

9 712.3 1.48 4.68 1.83 0.39 0.84 99.7 111.8 105.0

10 752.6 1.80 6.03 2.08 0.34 0.81 98.5 115.1 106.1

11 749.2 1.80 5.99 2.07 0.34 0.80 98.5 115.0 106.0

12 866.8 1.66 3.61 2.28 0.63 0.74 100.9 108.4 104.5

13 899.7 1.37 3.63 2.23 0.61 1.01 100.6 109.2 104.5

14 898.8 1.39 3.68 2.24 0.61 1.00 100.6 109.3 104.6

15 952.4 1.91 6.74 2.71 0.40 1.01 97.6 116.4 106.4

16 917.7 1.80 6.34 2.54 0.40 1.01 98.3 115.9 106.2

17 1030.3 1.86 4.89 2.84 0.58 0.94 99.7 111.3 105.2

18 1022.1 1.97 5.34 2.88 0.54 0.93 99.2 112.3 105.4

19 1000.1 2.03 6.39 2.88 0.45 0.97 97.8 114.9 106.0
NASA/TM—2002-211982 9

From table 3 it is apparent that combustion performance varied over the range of conditions tested.

The C* efficiency for the fully mixed values greater than 100 percent can be attributed to uncertainty in

the measured data. An uncertainty analysis was conducted for the tests completed based upon the method

described in reference 12. Results from the analysis show an uncertainty range of 4 to 6 percent for C*

efficiency. Therefore, it is best to view the performance results in terms of general trends as opposed to

a quantitative assessment. The streamtube calculation methods are significantly higher than 100 percent

efficiency, which indicates that the propellants were mixing well before the chamber throat and thus

the fully mixed analysis is the best measure to evaluate trends in the data. Clearly the assumption of an

ambient temperature boundary layer flow is not physically relevant for this type of system. However, it

was felt the analysis would provide an interesting point of comparison with the other methods. Figure 5

presents C* efficiency for each calculation method as a function of total mixture ratio.

95

0.9 1.0 1.1 1.2 1.3 1.4 1.5 1.6 1.7 1.8 1.9 2.0 2.1

100

105

110

115

120

Total mixture ratio

C* efficiency, percent

Figure 5.—C* efficiency as a function of total mixture ratio.

Fully mixed

TBLC = 2000R

TBLC = AMB

The results show a divergence in trends between the three methods of analysis. The fully mixed case

demonstrated a decreasing trend in performance based upon total mixture ratio, while the ambient boundary

layer assumption provided an increasing trend. The ambient boundary layer results have significantly

more scatter than the other methods presented, which may be a bi-product of the non-physical nature of

the assumption, but more likely uncertainty in the flowrate measurements. The use of a 2000 °R hydrogen

boundary layer resulted in a flat trend across the mixture ratio range. Average calculated performance level

decreases with increasing boundary layer temperature assumptions.

Due to the amount of scatter and lack of information regarding mixing between the boundary layer

and core flow, no definite conclusion can be reached on combustion efficiency based upon the results.

However, it does appear, despite the large amount of boundary layer cooling, that combustion efficiencies

of greater than 97 percent are achievable.

To further examine the data, C* efficiency was plotted as a function of core injector mixture ratio,

and the results are plotted in figure 6. As with total mixture ratio, the results show multiple trends in

performance with an increasing injector mixture ratio. The disparity in results is more pronounced based

upon this method as evident from the slopes for both the ambient boundary layer assumption and the

fully mixed assumption. This indicates that scatter in the boundary layer mass flow measurement has a

significant effect on the results. As with the previous method, the results from the 2000 °R boundary layer

analysis are flat across the core mixture ratio range.
NASA/TM—2002-211982 10

95

3.0 3.5 4.0 4.5 5.0 5.5 6.0 6.5 7.0

100

105

110

115

120

Core mixture ratio

C* efficiency, percent

Figure 6.—C* efficiency as a function of injector core mixture ratio.

Fully mixed

TBLC = 2000R

TBLC = AMB

The effect on C* efficiency from the amount of boundary layer flow as a percent of the total mass flow

was also examined. As shown in figure 7, a significant amount of boundary layer flow (20 to 38 percent)

was used to cool the chamber walls. The results show that effects due to the percentage of boundary layer

flow is relatively flat with increasing percentage. One interesting feature on the fully mixed results is that

the linear curve fit of the results falls completely below the 100 percent level, whereas for the mixture ratio

analysis the results showed a trend of slightly increasing performance (above 100 percent) with reduced

mixture ratio.

95

20 22 24 26 28 30 32 34 36 38 40

100

105

110

115

120

Boundary layer cooling, percent

C* efficiency, percent

Figure 7.—C* efficiency as a function of percent boundary layer flow.

Fully mixed

TBLC = 2000R

TBLC = AMB

Figure 8 shows the results of C* efficiency for the fully mixed analysis again plotted against core

mixture ratio. However, in this plot the data has been divided into groups based upon chamber pressure

ranges of 100 psia increments from 450 to 1050 psia. The results show no effect due to chamber pressure

with in the ranges of mixture ratios.
NASA/TM—2002-211982 11

95

3.0 3.5 4.0 4.5 5.0 5.5 6.0 6.5 7.0

96

97

98

99

101

100

Core mixture ratio

C* efficiency, percent

450-550

650-750

850-950

950-1050

Pc

Figure 8.—C* efficiency as a function of core mixture ratio for the fully mixed analysis with data grouped

 by chamber pressure range.

Velocity ratio through the injector elements was also examined to determine if there was any

correlation with C* efficiency. The calculation discounts any potential effects from the boundary layer

flow upon performance. Velocity was calculated at the exit of the coaxial elements, before any combustion

occurs. For the purposes of this study, velocity ratio was defined as:

2

2

O

H

V

V

VR = (9)

where the velocities were calculated by:

A

m V

ρ

& = (10)

Velocity for the hydrogen was based on the simplifying assumption that all core hydrogen flow was

injected through the elements. Density was calculated based upon measured chamber pressure and

propellant inlet temperature. Results of C* efficiency as a function of velocity ratio for is plotted in

figure 9 for the three methods of C* efficiency calculations. Results show the diverging trends between

the various methods. Results for the fully mixed analysis are consistent with the core mixture ratio

comparison of figure 6 because velocity ratio increases with a decrease in mixture ratio for the same

injector element.
NASA/TM—2002-211982 12

95

4 5 6 7 8 9 10 11 12

100

105

110

115

120

Velocity ratio

C* efficiency, percent

Figure 9.—C* efficiency as a function of velocity ratio.

Fully mixed

TBLC = 2000R

TBLC = AMB

A calculation was also performed to examine a possible correlation of performance as a function of

momentum ratio, which is defined as the following.

2 2

2 2

*

*

O O

H H

m V

m V

MRT &

& = (12)

As shown in figure 10, the results follow the same general trend as velocity ratio. Combustion

efficiency for the fully mixed analysis increases with increasing momentum ratio.

95

0 1 2 3 4 5 6 7 8

100

105

110

115

120

Momentum ratio per unit area

C* efficiency, percent

Figure 10.—C* efficiency as a function of momentum ratio.

Fully mixed

TBLC = 2000R

TBLC = AMB

Thermal environment

One of the key areas of interest in the development of gas-gas injectors is cooling of the injector face.

To alleviate some of the cooling issues a porous metal faceplate was used to provide transpiration cooling

of hydrogen through the face. In the current configuration this is a reasonable approach since the hydrogen

is still relatively cool at ambient room temperature. However, in an operational system the hydrogen

through the face will actually be hydrogen rich gas generator exhaust at an elevated temperature, which has

significantly less cooling capacity. The current testing did not experience any significant problems with

face cooling. Figures 11(a) and 11(b) show the posttest condition of the injector face. There is minimal

distress on the face and the main injector elements are all in good shape. However, there was some melting

on two of the face screws; while not indicative of a fundamental cooling problem, it does show the

sensitivity of the issue.
NASA/TM—2002-211982 13

Combustion chamber wall temperature measurements were obtained at five locations along the length

of the chamber. Location of the thermocouples was shown previously in figure 2 and table 2. The results

shown in figures 12 to 14 give some indication of the location of combustion in the chamber. From figure

12, the thermocouples within the first 3.4 inches downstream from the injector face (TCW 1–3) are

consistent, with temperatures in the range of 750 to 1100 °R. However, the two thermocouples located

5.5 inches downstream from the injector face (TCW 4 and 5), shows a temperature increase, which

indicates additional heat release at that location. However, this could be due to either additional heat

release from the core flow or to more complete mixing between the core and boundary layer flows. The

thermocouple at TCW5 reads significantly higher in some cases, which may be an indication of combustion

non-uniformity or simply that the thermocouple tip was extended into the flow farther than the others.

750

0.8 0.9 1.0 1.1 1.2 1.3 1.4 1.5 1.6 1.7 1.8 1.9 2.0 2.1

1000

1250

1500

1750

2250

2000

Total mixture ratio

Temperature, °R

Figure 12.—Wall temperatures as a function of total mixture ratio.

TCW1

TCW2

TCW3

TCW4

TCW5

The results from the total mixture ratio analysis when compared to injector mixture ratio show a

consistent pattern to the temperature distribution. In most cases the thermocouple located 5.5 inches

downstream of the injector face read higher than the other measurements.
NASA/TM—2002-211982 14

750

2.5 3.0 3.5 4.0 4.5 5.0 5.5 6.0 6.5 7.0

1000

1250

1500

1750

2250

2000

Total mixture ratio

Temperature, °R

Figure 13.—Wall temperatures as a function of injector mixture ratio.

TCW1

TCW2

TCW3

TCW4

TCW5

Finally, the results were plotted as a function of chamber pressure. There is no relevant trend based

upon the plot, as the excess boundary layer flow was successful in reducing wall temperature through the

entire test matrix.

750

400 500 600 700 800 1000 1100

1000

1250

1500

1750

2250

2000

Chamber pressure, psia

Temperature, °R

Figure 14.—Wall temperatures as a function of chamber pressure.

TCW1

TCW2

TCW3

TCW4

TCW5

900

Summary

An experimental test program was conducted to evaluate the performance of a rocket engine injector

designed for use with gaseous hydrogen and gaseous oxygen propellants. A total of 19 hot-fire tests were

completed with a 7-element coaxial injector design in a heat sink combustion chamber with a 1.18-in.

diameter throat. A comparison was made of the current design with the best available design correlation.

All propellants were injected at ambient temperature (529 to 555 °R). Combustion chamber walls were

cooled with a layer of hydrogen injected through 120 orifices located around the injector periphery.

Test firings were conducted at combustion chamber pressures of 470 to 1030 psia, total mixture ratios of

0.96 to 2.03, and injector mixture ratios of 3.3 to 6.7. The injector survived the testing with no significant

discoloration or erosion to either the faceplate or oxygen posts at the conditions tested.

Injector performance was based upon calculated values for characteristic exhaust velocity (C*)

efficiency. Measurements were obtained for combustion chamber pressure and mass flow. Temperature

measurements were obtained at 5 locations along the length of the combustion chamber. Due to the amount

of boundary layer coolant present, three methods were used to calculate C* efficiency based on fully mixed

and streamtube flow. The fully mixed analysis assumed that all propellants were completely mixed to the
NASA/TM—2002-211982 15

total mixture ratio whereas the streamtube analysis assumed the boundary layer coolant did not mix with

the core flow and either stayed at injection ambient temperature or was heated to 2000 °R. Unfortunately,

the presence of large amounts of boundary layer flow prevented any quantitative conclusions to be obtained

for injector performance. The C* efficiency range for the three methods of analysis was 97 to 117 percent,

with the ambient streamtube analysis providing the high range of results and the fully mixed analysis

providing the low range of results. Values greater than 100 percent for the fully mixed results can be

attributed to experimental uncertainty, which was in the range of 4 to 6 percent. Values up to 117 percent

for the streamtube analysis indicate that there was significant mixing between the streams. Therefore, the

trends from the streamtube analysis, particularly the ambient temperature boundary layer method, are not

indicative of the phenomena observed in the experiment. However, a qualitative analysis of the results can

lead to the conclusion that a coaxial gas-gas injector design can provide a high level of performance based

upon the fact that all performance levels were greater than 97 percent despite the significant presence of

excess hydrogen. A comparison of general trends shows decreasing performance with increasing mixture

ratio based upon the fully mixed analysis. However, uncertainty with the experimental results indicate that

to obtain detailed performance data on specific main injector element designs, the presence of any

boundary layer or secondary cooling inside the chamber should be avoided. Future testing to evaluate

gas-gas injector performance should be conducted with either heat sink or actively cooled combustion

chambers. Additional fidelity can be obtained by obtaining thrust measurements to corroborate results

calculated from chamber pressure measurements.

References

1 Farhangi, S., Yu, T.I., Rojas, L., Sprouse, K., and McKinnon, J.; “Gas-Gas Injector Technology for Full

Flow Stage Combustion Cycle Application,” AIAA 99-2757, 35th AIAA Joint Propulsion Conference, June

20–23, 1999, Los Angeles, CA

2 Farhangi, S., Yu, T.I., Rojas, L., Sprouse, K., and Fisher, S.: “Gas-Gas Main Injector Development for

Full-Flow Staged Combustion Cycle,” 12th Annual PERC Symposium on Propulsion, Cleveland, OH,

October 26–27, 2000.

3 Paster, R.D.; “Hydrogen-Oxygen APS Engines,” Final Report, Volume 1 and 2, Rockwell International,

Rocketdyne Division, NASA CR–120805 and CR–120806, February 173.

4 Heckert, B.J.; Yu, T.I.; Allums, S.L.; and Carrasquillo, E.A.: “25-lbf Gaseous Oxygen/Gaseous Hydrogen

Thruster for Space Station Application,” JANNAF Propulsion Meeting, New Orleans, LA, August 1986.

5 Robinson, P.J.: “Space Station Auxiliary Thrust Chamber Technology,” Aerojet Final Report, NASA

CR–185296, July 1990.

6 Jones, R.E.; Morren, W.E.; Sovey, J.S.; and Tacina, R.R.: “Space Station Propulsion,” NASA TM–

100216, 1987.

7 Calhoon, D., Ito, J., and Kors, D., “Investigation of Gaseous Propellant Combustion and Associated

Injector-Chamber Design Guidelines,” Aerojet Liquid Rocket Company, NASA CR–121234, Contract

NAS3–13379, July 1973.

8 Tucker, P.K., Klem, M.D., Smith, T.D., Farhangi, S., Fisher, S.C., and Santoro, R.J.; “Design of Efficient

GO2/GH2 Injectors: A NASA, Industry and University Cooperative Effort,” AIAA 97–3350, 33rd AIAA

Joint Propulsion Conference, July 6–9, 1997, Seattle, WA.

9 Faust, M.J.; Deshpande, M.; Pal, S.; Ni, T.; Merkle, C.L.; and Santoro, R.J.: “Experimental and Analytical

Characterization of a Shear Coaxial combusting GO2/GH2 Flowfield.” AIAA 96–0646, 32ndAIAA

Aerospace Sciences Meeting & Exhibit, January 15–18, 1996.

10 Zoeckler, J.G., Green, J.I., Raitano, P. “A New Facility for Advanced Rocket Propulsion Research,”

NASA TM–106193, June 1993.

11 Gordon, S.; and McBride, B.J.: Computer Program for Calculation of Complex Chemical Equilibrium

Compositions and Applications. NASA RP–1311, 1994.

12 Davidian, K.J.; Dieck, R.H.; and Chuang, I.: A Detailed Description of the Uncertainty Analysis for High

Area Ratio Rocket Nozzle Tests at the NASA Lewis Research Center. 24th JANNAF Combustion Meeting,

NASA TM–100203, October 1987.
This publication is available from the NASA Center for AeroSpace Information, 301–621–0390.

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21

Experimental Evaluation of a Subscale Gaseous Hydrogen/Gaseous Oxygen

Coaxial Rocket Injector

Timothy D. Smith, Mark D. Klem, Kevin J. Breisacher, Shahram Farhangi, and

Robert Sutton

Rocket engines; Injectors; Gaseous rocket propellants; Rocket engine design

Unclassified -Unlimited

Subject Category: 20 Distribution: Nonstandard

Timothy D. Smith, Mark D. Klem, and Kevin J. Breisacher, NASA Glenn Research Center; Shahram Farhangi and

Robert Sutton, Boeing, Rocketdyne Division, Canoga Park, California. Responsible person, Timothy D. Smith,

organization code 5830, 216–977–7546.

The next generation reusable launch vehicle may utilize a Full-Flow Stage Combustion (FFSC) rocket engine cycle. One

of the key technologies required is the development of an injector that uses gaseous oxygen and gaseous hydrogen as

propellants. Gas-gas propellant injection provides an engine with increased stability margin over a range of throttle set

points. This paper summarizes an injector design and testing effort that evaluated a coaxial rocket injector for use with

gaseous oxygen and gaseous hydrogen propellants. A total of 19 hot-fire tests were conducted up to a chamber pressure of

1030 psia, over a range of 3.3 to 6.7 for injector element mixture ratio. Post-test condition of the hardware was also used

to assess injector face cooling. Results show that high combustion performance levels could be achieved with gas-gas

propellants and there were no problems with excessive face heating for the conditions tested.