Timothy D. Smith, Mark D. Klem, and Kenneth L. Fisher

Glenn Research Center, Cleveland, Ohio

Propulsion Risk Reduction Activities for

Nontoxic Cryogenic Propulsion

NASA/TM—2010-216820

October 2010

AIAA–2010–8680
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Timothy D. Smith, Mark D. Klem, and Kenneth L. Fisher

Glenn Research Center, Cleveland, Ohio

Propulsion Risk Reduction Activities for

Nontoxic Cryogenic Propulsion

NASA/TM—2010-216820

October 2010

AIAA–2010–8680

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

Space 2010 Conference and Exposition

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Anaheim, California, August 30 to September 2, 2010
Acknowledgments

The authors gratefully acknowledge the leadership afforded this project from the Exploration Technology Development 

Program (ETDP); Frank Peri, Program Manager; Richie Law, Integration Manager; and Diane Hope, Program Element 

Manager. The authors also want to acknowledge the center teams: from NASA Glenn Research Center (GRC),

Todd Peterson, Lori Aponte-Richter, and Marilyn Groff; from NASA Marshall Space Flight Center (MSFC),

David Stephenson, Tony Kim, Joel Robinson, and David Chaffee; and from NASA

Johnson Space Center (JSC), Kris Romig and Howard Flynn.

Available from

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Trade names and trademarks are used in this report for identifi cation 

only. Their usage does not constitute an offi cial endorsement, 

either expressed or implied, by the National Aeronautics and 

Space Administration.

Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2010-216820 1 

Propulsion Risk Reduction Activities for 

Nontoxic Cryogenic Propulsion 

Timothy D. Smith, Mark D. Klem, and Kenneth L. Fisher 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

The Propulsion and Cryogenics Advanced Development (PCAD) Project’s primary objective is to develop 

propulsion system technologies for nontoxic or “green” propellants. The PCAD project focuses on the development 

of nontoxic propulsion technologies needed to provide necessary data and relevant experience to support informed 

decisions on implementation of nontoxic propellants for space missions. Implementation of nontoxic propellants in 

high performance propulsion systems offers NASA an opportunity to consider other options than current hypergolic 

propellants. The PCAD Project is emphasizing technology efforts in reaction control system (RCS) thruster designs, 

ascent main engines (AME), and descent main engines (DME). 

PCAD has a series of tasks and contracts to conduct risk reduction and/or retirement activities to demonstrate 

that nontoxic cryogenic propellants can be a feasible option for space missions. Work has focused on 1) reducing the 

risk of liquid oxygen/liquid methane ignition, demonstrating the key enabling technologies, and validating 

performance levels for reaction control engines for use on descent and ascent stages; 2) demonstrating the key 

enabling technologies and validating performance levels for liquid oxygen/liquid methane ascent engines; and 3) 

demonstrating the key enabling technologies and validating performance levels for deep throttling liquid 

oxygen/liquid hydrogen descent engines. The progress of these risk reduction and/or retirement activities will be 

presented. 

Nomenclature 

C* characteristic velocity 

EPW electronic pulse width 

Hz hertz 

in. inch(es) 

Isp specific impulse 

lbf pounds force 

lbm pounds mass 

L* characteristic length 

min minutes 

R degrees Rankine 

sec seconds 

I. Introduction 

The PCAD Project’s primary objective is to develop propulsion system technologies for exploration missions. 

The PCAD project is funded by the Exploration Technology Development Program in NASA’s Exploration Systems 

Mission Directorate. PCAD has concentrated its activities on nontoxic or green propellants to meet near term 

Constellation Program decision gates. Implementation of green propellants in high performance propulsion systems 

offers NASA an opportunity to consider other options than current hypergolic propellants. The PCAD Project is 

emphasizing efforts in reaction control system (RCS) thruster designs, ascent main engines (AME) for lunar 

missions, and descent main engine (DME) for lunar missions. PCAD has developed the following specific 

objectives: 

 Perform cryogenic and noncryogenic RCS design, ignition testing, and performance testing 

 Perform cryogenic ascent main engine design, ignition testing, and performance testing 

 Perform cryogenic descent main engine design and performance testing 
NASA/TM—2010-216820 2 

II. Liquid Oxygen (LOx) – Liquid Methane (LCH4) Propulsion 

In support of the U.S. Space Exploration Policy for returning to the Moon and beyond, NASA and its partners 

are developing and testing cryogenic propulsion system technologies that will meet the need for high-performance 

propulsion systems on long-duration missions. In particular, the lunar ascent module propulsion systems are critical 

performance drivers, due to the high “gear ratio” (ratio of mass launched to delivered mass to the Moon) associated 

with elements that are utilized through the late phases of the mission. However, due to the relatively small size of 

the Ascent Module, multiple propulsion system options exist. System trades for both lunar and Mars missions have 

indicated that LOx/LCH4 is a promising option, due to the approximate 600- to 800-lbm savings in overall systems 

mass over more conventional hypergolic systems. Because the Ascent Module is taken to the lunar surface, the 

indicated mass savings would be converted directly to lunar surface payload. LOx/LCH4 propulsion for Ascent Main 

and Ascent/Descent Reaction Control Propulsion is currently conceded as a critical enhancing technology, due to the 

potential increase of lunar surface payload. The primary technology risks, as determined by the PCAD project team, 

associated with LOx/LCH4 propulsion are the following: 

1. Reliable/ignition pressure fed LOx/LCH4 Reaction Control Engines (RCE) 

2. Meeting minimum performance and life requirements of LOx/LCH4 RCE and Main Engines with integrated 

testing 

3. Reliable/ignition pressure fed LOx/LCH4 Main Engine 

The PCAD project focus’ on the development of cryogenic propulsion technologies needed to provide necessary 

data and relevant experience to support informed decisions on potential implementation of cryogenic propellants in 

the Altair architecture. 

A. LOx/LCH4 Reaction Control Engine Development 

Since 2005, the PCAD project has invested in technologies leading to pre-prototype development of LOx/LCH4

reaction control engines (RCE) with the release of contract request for proposals (RFPs). The focus of the activities 

were originally to support the Service Module, however in 2006 the activity was steered to support a lunar lander. The 

top three risks identified for RCE technology are: 1) reliable ignition; 2) Performance (vacuum specific impulse – Isp); 

and 3) Repeatable pulse width. To address the risks, PCAD undertook a combination of in-house and contract 

activities. 

In 2006 PCAD awarded two RCE contracts to Northrop Grumman and Aerojet respectively. Each contract was 

focused on the development and delivery of a 100-lbf thrust pre-prototype engine subsystem. The key performance 

requirements in the contracts were: 1) 317-sec vacuum Isp; 2) 4 lbf-sec minimum impulse bit (Ibit); 3) 80-msec 

electronic pulse width (EPW); 4) 25,000 valve cycles and 5) operation over a range of inlet conditions from gas to 

liquid for start. The engine concepts put forward by each company were different in approach to meeting the 

contract requirements. 

Aerojet put forward a concept with foundations in previous work on LOx/ethanol and internally funded 

activities. The first engines tested were originally LOx/ethanol 870-lbf thrusters that were modified to accommodate 

LOx/LCH4 (Ref. 1). The modified units were successfully tested on the Auxiliary Propulsion System Test Bed 

(APSTB) in the NASA White Sands Test Facility (WSTF) Test Stand (TS) 401. The proposed 100-lbf engine 

concept consisted of a compact integral exciter/spark plug system, a dual coil direct-acting solenoid valve for 

oxidizer and fuel, an integral igniter and injector, and a columbium chamber/nozzle with an expansion area ratio of 

80:1. 

Over the course of several contract option periods, multiple injector patterns were developed and manufactured 

using Aerojet’s platelet technology, Figure 1(a). Flow control for both the main chamber and igniter were controlled 

by a single set of dual coil valves. The valves were demonstrated to over 55,000 cryogenic cycles in liquid nitrogen, 

exceeding the 25,000 cycle life. Ignition was accomplished with the use of a spark torch igniter. Over the duration of 

the contract, a series of igniter and injector concepts were tested at sea level to examine engine performance. The 

result of the testing was an impinging injector design that successfully met all key performance criteria either by 

demonstration or calculations based on test data. Aerojet conducted over 1300 engine pulse tests at a variety of duty 

cycles for over 1900-sec total of sea level testing during the engine development (Refs. 2 and ,

3). Specifically 

Aerojet was able to meet 317-sec Isp calculated based on estimated nozzle losses and exceeded the 80-msec EPW 

requirement by demonstrating 40-msec EPW. As a result, Aerojet was able to provide five engine units to NASA for 

multiple engines testing on the APSTB at WSTF and two units for testing at the NASA Glenn Research Center 

(GRC) in the Altitude Combustion Stand (ACS). 

 
NASA/TM—2010-216820 3 

Sea level (Ref. 4) and altitude performance testing (Ref. 5) has been conducted at GRC with the Aerojet engines. 

Figure 1(b) shows the Aerojet engine during test at GRC. A total of 60 altitude hot-fire tests were completed with 

the Aerojet 100-lbf LOx/LCH4 engine and propellant conditioning feed systems (PCFS) (Refs. 6 and 7). The PCFS, 

as shown in Figure 2, was used to obtain conditions over the range of nominal (204 °R LOx/204 °R LCH4), 

cold/cold (160 °R LOx/170 °R LCH4), to warm/warm (224 °R LOx/224 °R LCH4). The PCFS uses a combination of 

cooling loops and heaters to vary the propellant conditions. Test results demonstrated that propellant conditions 

could be controlled to within 5 °R for a given set point. Altitude performance testing was conducted using a 45:1 

area ratio columbium radiation cooled nozzle. The main goal of the testing was to develop specific impulse 

performance curves as a function of mixture ratio. Testing was also conducted over a wide range of propellant inlet 

conditions (pressure and temperature), to simulate operation in a variety of space environments. The engine 

demonstrated that meeting the required 317-sec performance is feasible for the 80:1 nozzle based on the results with 

a 45:1 nozzle. 

 (a) (b) 

Figure 1.—(a) Aerojet 100-lbf LOx/LCH4 reaction control engine, (b) Aerojet 100-lbf LOx/LCH4 reaction control 

engine in test at GRC. 

Figure 2.—Propellant Conditioning Feed System skid (PCFS) at GRC. 
NASA/TM—2010-216820 4 

Northrop Grumman put forward a concept with foundations in 

previous work on hypergol engines. The concept was regeneratively 

cooled with both oxygen and methane through the combustion 

chamber and part of the nozzle (Fig. 3) (Ref. 8). The full engine area 

ratio (120:1) was completed with a columbium nozzle extension. 

Flow control for both the main chamber and igniter was controlled by 

a single set of single coil valves. Ignition was accomplished with the 

use of a spark torch igniter. A series of hardware configurations were 

tested, starting with workhorse hardware, to develop the engine 

cooling circuit. During the course of the contract Northrop Grumman 

ran into a number of design and manufacturing issues which slowed 

progress. As a result, budget limitations required changes to the scope 

of the contract which eliminated the planned four pre-prototype 

deliverables. However, Northrop Grumman was able to develop a 

single pre-prototype unit that was tested in vacuum conditions at their 

Capistrano test facility. Test results indicate that the engine concept 

was able to meet the performance specifications in the contract, 

including exceeding the specific impulse requirement. The measured 

Isp was approximately 331-sec, which exceeded the specification 

requirement of 317-sec. NASA currently has one pre-prototype unit 

available for further in-house testing. 

B. LOx/LCH4 Reaction Control Engine Integrated Testing 

Once developed, the plan was to integrate the 

RCE thrusters into a four engine cluster which 

would simulate a vehicle engine configuration. 

The Auxiliary Propulsion System Test Bed 

(APSTB), Fig. 4, at WSTF was modified with a 

high vacuum bell jar which serves as the engine 

cluster simulator. In the bell jar all propellant feed 

lines and valves were mounted in a way similar to 

a space craft system. The feed system was also 

fitted with a thermodynamic vent system (TVS) 

to condition the propellant delivery to the 

engines. A total of five engines were delivered 

from Aerojet for the APSTB testing. Engines 

were installed and tested at each position. 

Approximately 2500 pulses were conducted over 

a sequence of 145 tests. In one test, a total of 380 

consecutive pulses were completed. Also, an 

additional 90 pulses were conducted with two 

engines firing simultaneously. The engines 

performed as expected, however the testing did 

uncover issues with the feed system design. A number of tests suffered from high flow spikes or water hammer 

which resulted in a number of pressure transducer failures. The data is now being used to develop improvements to 

feed system models. A complicating factor to the feed system was the APSTB design. Because the rig was originally 

designed for the Space Shuttle systems development, the rig was significantly oversized. As a result, PCAD has 

undertaken the development of the Integrated Propulsions System Test Bed (IPSTB). The IPSTB, like the APSTB, 

will be a propulsion system simulator with propellant tanks, feed lines and an engine cluster. However, the IPSTB 

will be designed with smaller propellant tanks and with the flexibility to change component locations or vary feed 

line lengths. The goal of the testing will be to examine system interactions with a number of feed system designs and 

to obtain the data for comparison with state of the art fluid models. Currently the IPSTB will utilize the current 

inventory of Aerojet and Northrop Grumman engines. 

 

Figure 3.—Northrop Grumman 100-lbf 

LOx/LCH4 reaction control engine. 

Figure 4.—APSTB at WSTF showing RCE and AME test 

positions. 
NASA/TM—2010-216820 5 

(a) (b) (c) 

Figure 5.—LOx/LCH4 Altitude Ignition Testing at GRC (a) Test Cell 21 configuration; (b) WASK spark torch igniter 

during test; (c) ascent main engine class igniter during test. 

C. LOx/LCH4 Ignition Risk Reduction 

To address the highest risk for LOx/LCH4 propulsion 

systems, reliable ignition, NASA has conducted numerous inhouse experimental efforts to examine the issue. The work 

has been completed at both RCE and ascent main engine 

(AME) scales. Figure 5(a) shows the basic altitude test 

configuration in Cell 21 at GRC. The majority of the work 

has been conducted with spark torch igniters (Refs. 9, 10, 11, 

12, and 13). however there has been work done with 

microwave, (Refs. 14 and 15) piezoelectric, spark torch/glow 

plug combination (Ref. 16), and catalytic ignitions systems. 

Figure 5(b) shows a spark torch configuration from WASK 

Inc. and Figure 5(c) shows a NASA breadboard 

configuration. Overall there have been no significant issues 

identified that would prohibit the reliable ignition over a range 

of conditions with LOx/LCH4. One of the last ignition 

specific activities completed was the demonstration of 30,000 ignition cycles on a spark torch ignition system at 

vacuum conditions (Ref. 17). Completion of this activity did not identify any issues with the hardware or designs for 

long duration applications. The work to date has identified issues with spark plug durability and the reliability of power 

exciter units. In both cases, PCAD has worked additional technology tasks to address the issues. There appear to be 

viable solutions in work to reduce the risk. 

In particular, advancements have been made on the exciter where Aerojet and Unison have developed a single 

compact exciter (Ref. 18) unit, as shown in Figure 6, to replace the current state of the art exciter box and high 

voltage power lines. NASA has also successfully completed altitude testing with a compact exciter developed under 

a Small Business Innovative Research (SBIR) Phase II task with Alphaport, Inc. Many of the issues remaining with 

LOx/LCH4 ignition are related to the specific requirements and duty cycles that will be imposed on the systems or 

with the final spaceflight qualification of the units. One general area that would still require investigation is ignition 

in the cold thermal environment of space where both the hardware and propellants have been exposed to those 

conditions for a significant period of time before being required to operate. 

D. LOx/LCH4 Ascent Main Engine Development 

As with RCE, the PCAD project has invested in technologies leading to pre-prototype development of 

LOx/LCH4 main engine since 2005. The focus of the activities were originally to support the Service Module, 

however in 2006 the activity was steered to support the lunar lander. The top three risks identified for RCE 

technology are: 1) reliable ignition; 2) performance (vacuum specific impulse – Isp); and 3) fast start (90% thrust in 

0.5-sec). To address the risks, PCAD undertook a combination of in-house and contract activities. 

In 2006 PCAD awarded two main contracts to ATK and KT Engineering (KTE) respectively. Each contract was 

focused on the development and delivery of a 7,500-lbf thrust pre-prototype engine. The key performance targets for 

the activity were: 1) 7,500-lbf thrust, 355-sec vacuum Isp; 2) 90% rated thrust within 0.5 sec; 3) total of 24 restarts; 

and 5) operation over a range of inlet conditions from gas to liquid for start. The engine concepts put forward by 

each company were different in approach to meeting the contract requirements. 

Figure 6.—Prototype Unison compact exciters 

configured for use with Aerojet 870-lbf RCE. 
NASA/TM—2010-216820 6 

ATK teamed with XCOR to develop a pressure-fed engine concept that was actively cooled with methane 

(Refs. 19, 20, and 21). To enhance the engine life, liquid methane passed through coolant channels machined into 

the combustion chamber. The warm methane is then injected into the engine where it mixes with liquid oxygen, 

creating the combustion mixture which provides the engine thrust. As part of the project execution, the ATK/XCOR 

team developed a “trombone” combustion chamber and injector to conduct early ground testing to examine 

combustion performance (C* efficiency). The trombone chamber was a water cooled thrust chamber designed to 

accommodate multiple length configurations to determine an optimum. The data was then used to fabricate a 

methane cooled workhorse combustion chamber. Sea level testing was conducted with both the trombone and 

workhorse combustion chambers at XCOR facilities in Mojave, California. 

The second contractor, KTE, chose an ablative combustion chamber in attempts to meet the contract 

requirements. An ablative material is simply a thick chamber lining that slowly chars away as the engine operates. In 

this configuration, oxygen and methane are injected into the combustion chamber as liquids. KTE also chose to 

conduct smaller ignition risk reduction activities at Purdue University on both spark initiated torch igniters (SITI) 

and catalytic initiated torch igniters (CITI) systems. Both systems were tested successfully at sea level conditions 

and expected to be used in the larger engines during ground test. As part of the engine development, KTE planned to 

use a water-cooled combustion chamber for initial injector performance tests. A handful of hot fire sea level tests 

were conducted with the hardware. 

To meet the Altair engine requirements, NASA issued a new RFP for a workhorse engine. Work under this 

contract would primarily be focused with demonstrating the main requirements of 1) 5,500-lbf thrust, 355-sec 

vacuum Isp; 2) 90% rated thrust within 0.5 sec; 3) total of 24 restarts; and 5) operation over a range of inlet 

conditions from gas to liquid for start. However, since the hardware was designated workhorse; weight and certain 

component developments such as valves, were omitted. From the competitive process, Aerojet was selected as the 

contractor. Aerojet put forward an ablative engine concept with liquid oxygen/liquid methane injection (Ref. 22). 

The overall activity was broken into two phases. The first phase involved Aerojet fabrication and sea level testing of 

multiple injector designs. The second phase was NASA taking delivery of the engines and conducting altitude 

performance testing at WSTF. Under the contract, three injectors were fabricated and tested at Aerojet (Ref. 23). A 

total of 48 tests were completed with both 8-. and 10-in. length ablative combustion chambers. Most of the tests 

were conducted at between 10 to 20 sec; however, one was conducted at 110-sec duration. Performance levels were 

lower than expected due to excessive film cooling along the combustion chamber wall. To improve performance, 

two additional injectors were fabricated. The second injector incorporated an alternate injector pattern than the first 

injector. A total of seven tests were completed before testing was stopped due to high heat release near the injector 

face resulting in excessive ablative erosion. Due to heating issues and low overall performance, this injector was not 

a viable candidate for altitude testing. The third injector was an iteration of the first injector, only with a lower 

percentage of film coolant. Testing was cut short due to excessive heating at the injector face. 

Testing at WSTF proceeded with the first injector from the Aerojet AME contract. While the sea level testing 

performance levels were lower than desired, it was felt the altitude testing could still provide useful information. In 

particular, the team was interested in developing a correlation between the sea level results and altitude tests. The 

tests results would also provide key data to use in validating nozzle performance analysis, including quantifying 

potential loss parameters. Testing (Ref. 24) was conducted with an 8-in. long ablative combustion chamber and a 

radiation cooled columbium Space Shuttle OMS-E nozzle extension, which provides an area ratio of 129:1. Design 

area ratio for the vision prototype engine design is 150:1. A total of 187 sec of run time was achieved on the engine 

including seven 20-sec tests and one 40-sec test. Figure 7 shows clearly the nozzle heating of the AME during 

testing, from left to right, at 5-, 10-, 15-, and 20-sec. The injector, chamber and nozzle were all in good physical 

condition after the testing. Calculated vacuum specific impulse numbers for the test program averaged 

approximately 344 lbf-sec/lbm and peaked at 345.3 lbf-sec/lbm with the 129:1 area ratio OME nozzle. Extrapolating 

to 150:1 conceptual flight design point a Isp~ 348-sec could be achieved. This is within 2% of the target. This result 

higher than expected based on pretest predictions from the sea level test results. Predictions were done with the well 

characterized Two Dimensional Kinetics (TDK) (Ref. 25) computer code. Characteristic exhaust velocity 

efficiencies were estimated to be between 94 and 95%. 
NASA/TM—2010-216820 7 

 5-sec 10-sec 15-sec 20-sec 

Figure 7.—Aerojet LOx/LCH4 ascent main engine during altitude testing at WSTF. 

Figure 8.—LOx/LCH4 injector sea level test at MSFC. 

E. LOx/LCH4 Ascent Main Engine Component Development 

In parallel to the contract efforts, NASA conducted in-house injector development on oxygen/methane injectors. 

Tests were conducted on both 2- and 6-in. diameter chambers at the NASA Marshall Space Flight Center (MSFC) 

(Refs. 26, 27, 28, 29, and 30). Figure 8 shows a MSFC in-house injector during sea level testing. Testing has been 

focused on the performance and stability characteristics of a swirl coaxial injector with multiple combustion 

chamber lengths. The in-house tests have been able to demonstrate 98%+ C* efficiencies with a 20-in. long 

combustion chamber. The testing has also collected heat transfer data with use of a water cooled combustion 

chamber; combustion stability data for model comparison; and chamber length correlations to obtain performance 

levels. In addition, work has been successful in demonstrating microwave and spark torch ignition systems in sea 

level and altitude tests. 
NASA/TM—2010-216820 8 

A pressure fed methane regeneratively cooled engine 

could be used to meet a lunar lander mission. One area 

identified from the ATK testing is flow instabilities in the 

coolant channels with methane at subcritical conditions. 

NASA is conducting in-house experiments (Ref. 31) with a 

heated tube facility to simulate a methane coolant channel 

to examine flow stability and characterize heat transfer 

properties. 

To address the key risk of a main engine ignition at 

vacuum and to provide a pathfinder engine for WSTF 

altitude testing, NASA and Pratt & Whitney Rocketdyne 

(PWR) tested an unmodified RS-18 engine with 

LOx/LCH4 and a spark torch igniter, in altitude conditions 

at WSTF TS401(Fig. 9(a)) (Ref. 32). Because the injector 

was not modified from the original configuration used for 

the hypergolic propellant combination of NTO/Aerozine 

50, it was not expected to provide a high C* efficiency. 

However, three successful main engine vacuum ignitions 

were conducted which met the main objective of the test. 

In conjunction with the Innovative Partnership Program 

(IPP) and PCAD, work began at the NASA Johnson Space 

Center (JSC) with Armadillo Aerospace on the testing of a 

1,500-lbf thrust-class LOx/LCH4 rocket engine (Refs. 33 

and 34). Sea level testing was conducted at the Armadillo 

facilities in Caddo Mills, Texas, and simulated altitude 

tests were conducted at WSTF (Fig. 9(b)). Testing 

examined engine performance and ignition, both gas torch 

and pyrotechnic, at altitude conditions. The rocket engine 

was designed to be configured with three different nozzle 

configurations, including a dual-bell nozzle geometry. A 

total of 10 hot-fire ignition and dual-bell nozzle tests were 

conducted at WSTF. 

III. Liquid Oxygen (LOx) – Liquid Hydrogen (LH2) Propulsion 

One of the mission enabling technologies to support future lunar missions is the development of a LOx – LH2

deep throttling descent engine. The descent main engines must be able to throttle and remain controlled by the crew 

to provide a soft landing or to maneuver to a different landing site. Rocket engines typically have a fixed point 

design that does not allow power levels to throttle over a wide range of operating conditions. If not designed 

properly, throttling a rocket engine can create low frequency instability in engine pressure, which can cause a 

reduction in performance or even damage to the engine or vehicle. As currently defined, deep throttling for the lunar 

missions is a 10:1 ratio, or an engine that can stably throttle from 100 to 10% power. The PCAD project is exploring 

three options through contracted efforts to develop deep throttling technologies. The first is with the Common 

Extensible Cryogenic Engine (CECE) (Ref. 35), a modified RL10 from PWR. A second effort is technology 

development for an expander cycle engine with Northrop Grumman Aerospace Systems (NG) based on the Pintle 

injector. The third option is a throttling injector concept being developed by Aerojet. Along with the contracted 

efforts, NASA is exploring in-house technology efforts with the development of an expander cycle test bed at 

MSFC. 

A. Descent Engine Technology Contracts 

The CECE contract with PWR was initiated in July 2005 with the development of the Demo 1.0 activity. The 

primary focus of Demo 1.0 was to assemble a deep throttling technology demonstrator from existing expander cycle 

RL10 parts. A number of key components were changed to develop the demonstrator including the fabrication of a 

fixed-geometry, high pressure drop injector, change out of turbine bypass (TCV) and oxidizer control valves (OCV), 

adding a larger turbine bypass valve (TBV), and a variable area cavitating venturi (VACV). 

(a) 

(b) 

Figure 9.—LOx/LCH4 engine testing at WSTF. (a) 

PWR RS18; (b) Armadillo Aerospace dual bell 

nozzle engine. 
NASA/TM—2010-216820 9 

The first test series, Demo 1.0, completed four test 

runs between April and May 2006 at the PWR E6 

facility in West Palm Beach, Florida. Figure 10 shows 

the CECE during altitude testing in the E6 facility. The 

testing was able to obtain baseline performance and 

stability data from 20 to 90%. To meet the requirements, 

testing was completed down to 10% power. However, at 

16% power, lower power chugging oscillations were 

detected. Despite the chugging the tests were successful 

because it quantified the baseline operating boundaries 

and provided valuable data to update performance and 

operations models. It was determined from the data 

analysis (Ref. 36) that the chugging was the result of 

vapor formation in the injector oxygen manifold. A 

second series of tests, Demo 1.5, were conducted in 

March and April 2007 with the same engine 

configuration as Demo 1.0. A total of four tests 

accumulated a total of 1162-sec of run time. The testing 

explored the boundaries of the chug instability over a 

range of mixture ratios and chamber pressures. During 

the testing additional technology challenges were 

identified, in particular, 1 Hz instability in the fuel 

system due to film boiling at low power. There was also 

a 4000 Hz, 1T combustion oscillation observed between 

30 to 40% power. Testing was also conducted at throttle 

rates from 100 percent/sec down to 2.5 percent/sec. 

Overall the Demo 1.0 and Demo 1.5 (Ref. 37) testing 

developed a wide ranging set of baseline performance 

data down to 10% power and identified key technology 

needs for future efforts. 

The Demo 1.6 (Ref. 38) test campaign was designed to evaluate mitigations for the low frequency combustion 

instability (“chug”) observed at low power conditions during the Demo 1.0 and Demo 1.5 test programs. To 

eliminate the oxygen manifold film boiling, a new injector was designed which incorporated a thermal barrier 

coating on oxygen side of the inner propellant plate. The goal was to reduce the heat transfer from the warm 

hydrogen into liquid oxygen and prevent the film boiling. To mitigate the chug, the Demo 1.6 injector was modified 

from the previous configurations to include a spray-on insulation to reduce heat transfer to the LOx manifold, which 

was believed to be a significant contributor to the low power instability. In addition, gaseous helium injection into 

the LOx manifold was used as a means to stabilize the system. Also explored in this test series was mitigation for a 

low power 1 Hz fuel system oscillation caused by sub-critical hydrogen boiling in the chamber cooling jacket. 

Reduced area gas venturis were utilized to avoid the 1 Hz fuel-size oscillation by keeping the cooling jacket 

supercritical down to lower engine power levels. 

The final test of the CECE engine, Demo 1.7 (Ref. 39) was designed to test the ability of starting the engine at 

low power and to demonstrate closed loop control of a throttling engine. Demo 1.7 testing (Ref. 40) successfully 

demonstrated a number of engine modes of operation including chamber pressure and mixture ratio closed-loop 

control over a wide range of throttled power levels, fast throttle ramp rates, minimum power down to a smooth start 

to 10% power, eleven rapid relights demonstrated (many achieved as 2 relights within the same test matrix run), and 

high power, high mixture ratio operation. Finally the testing demonstrated low power stability, including chug-free 

operation down to 5.9% power. This represents a 17.6:1 overall cryogenic deep throttling ratio in a complete 

expander cycle engine system with all system-level interactions which greatly enhanced the value of the technology 

database acquired. Total Demo 1.7 engine testing has concluded with a total run time of 2,403.0-sec (40.0-min). 

Total CECE demonstrator engine run time has concluded with 7,435.8-sec (123.9-min). 

The second contracted effort developing deep throttling LOx/LH2 engine technologies is with Northrop 

Grumman Aerospace Systems (NGAS) on the TR202 contract (Refs. 41 and 42). The work with NGAS was started 

in June 2005 and is also focused on an expander cycle engine. The focal point of the NGAS engine concept is the 

variable area pintle injector, which is similar to the injector used on the Apollo Lunar Module Descent Engine. The 

first phase of the contract was focused on the design and development of a test-bed pintle injector. The injector 

Figure 10.—PWR CECE during altitude testing in PWR 

E6 test stand. 
NASA/TM—2010-216820 10 

design has a oxidizer centered pintle where the oxygen 

flows through a central passage and is injected radially 

through individual orifices into the combustion chamber. 

The fuel is injected through an annular sleeve around the 

center pintle post. The fuel creates a sheet that impinges 

with the radial oxygen flow. The injector throttling is 

controlled by articulating the fuel sleeve along the length 

of the pintle to either increase or decrease the oxygen 

flow area. Figure 11 shows the throttling pintle injector: 

high thrust setting (a) and low thrust setting (b) during 

water flow testing. For a flight engine the sleeve would 

be controlled by an actuator based on throttle inputs 

from the flight profile. For ground testing the fuel 

sleeve/throttle position did not have a position actuator. 

Testing was conducted with both ablative and water 

cooled combustion chambers (Refs. 43 and 44). The 

ablative chamber test series encompassed 22 tests at a 

nominal mixture ratio of 6, and thoroughly explored 

injector momentum rate ratio design space; confirmed 

expectations for excellent high performance potential 

over the high-end of the throttle power range; 

demonstrated stable deep-throttle combustion 

performance at 25 and 10% power conditions; and, 

validated the thermal integrity of the hardware design. A 

total of six Pintle configurations were tested using two 

fuel injection ring sizes. The ablative chamber test series 

yielded sufficient understanding and confidence in the 

injector design to justify change over to calorimeter 

chamber hardware, which enables accurate 

determination of performance and heat transfer 

characteristics in a follow-on test series. Testing with the 

calorimeter was successful in meeting all primary and 

secondary technical objectives including high 

performance (>98% C* (combustion) efficiency); stable 

10:1 deep throttling; measurement of heat transfer 

characteristics; evaluation of off-nominal oxidizer to fuel 

mixture ratio (MR) sensitivities; and evaluation of L* 

sensitivity. The majority of the test program was devoted 

to an extensive Design of Experiments (DOE) for 

optimized injector performance in which the major 

influencing parameters were characterized. After 

extensive testing, the team arrived at an optimized high-performance injector design. Testing of the optimized 

injector demonstrated stable combustion over the full 10:1 throttle range, and heat transfer characteristics were 

within anticipated ranges. 

In 2009, Aerojet was also awarded a contract to develop deep throttling injector (Ref. 45) technologies. The 

contract builds upon an internal research project the company conducted to demonstrate 10:1 throttling with a 1,500-

lbf injector (Ref. 46). The current effort will focus on 10:1 throttling with a 9,000-lbf thrust injector. The engine 

system envisioned is an expander cycle LOx/LH2 engine. The injector is anticipated to be sea level tested in 2011 

with a hydrogen regenerative cooled combustion chamber supplied under a Space Act. 

B. NASA In-House Component Development 

NASA is conducting several complementary component development activities in-house. Development of the inhouse technologies will be conducted on the Lunar Lander Descent Engine Testbed (LLDETB) on Test Stand 500 at 

MSFC. This sea-level rig is a flexible system to accommodate change out of injectors, combustion chambers, and 

turbomachinery. As part of the test rig build-up a number of individual components have been fabricated and tested 

independently. One of the first components tested was a dual oxygen–inlet swirl coaxial element deep throttling 

(a) 

(b) 

Figure 11.—Northrop Grumman throttling pintle injector: 

high thrust setting (a) and low thrust setting (b) during 

water flow testing. 
NASA/TM—2010-216820 11 

injector (Ref. 47). The dual-inlet injector has two fixed area 

oxygen manifolds to maintain sufficient pressure drop 

across a wide range of throttle conditions. Each manifold 

has fixed inlet areas to the oxygen posts of the injector and 

flow can be independently controlled with shutoff valves. 

For high power cases oxygen would flow through both 

manifolds, however at low power, flow to the secondary 

manifold would be cut off. The sea level tests like that 

shown in Figure 12, provided all data needed to calculate 

C* efficiency, heat flux, and other information such as high 

speed pressure data (Ref. 48). The injector achieved very 

high C* efficiency numbers and stable operation at the high 

power levels. There were some low frequency (chug) 

instabilities at the lower power levels. These chug modes 

are currently being attributed to the LOx supply 

temperatures which were warmer than ideal. Results from 

the testing will contribute to future development of a twostage injector concept or any deep throttling technology. 

An important technology in the control of deep 

throttling engines is the ability to control the cooling flow 

from the combustion chamber to the fuel turbo pump. To 

examine improved control, work under an Innovative 

Partnership Program (IPP) with Vacco Industries developed 

an advanced turbine bypass valve (ATBV). The goal of 

testing was to determine the effective flow area versus 

valve position at nine equally spaced points in the valve 

travel and exercise the valve under engine conditions to 

examine seal performance. Figure 13 shows the ATBV in a 

test position at MSFC. The test program consisted of two 

tests series to determine the flow coefficient versus position 

and evaluate the ATBV design while operating in simulated 

engine temperature, flow rate, and pressure conditions. The 

team was also able to operate the valve in various simulated 

engine environments to fully characterize the performance 

of the ATBV design. 

IV. Conclusion 

The Propulsion and Cryogenic Advanced Development (PCAD) Project Team led by the NASA Glenn Research 

Center (GRC) in partnership with the NASA Marshall Space Flight Center (MSFC), the NASA Johnson Space 

Center (JSC), the NASA White Sands Test Facility (WSTF), and industrial partners, is conducting a focused 

technology development effort to advance high performance cryogenic propulsion systems. Over the last 5 years this 

team has been a model for cross center collaboration. To date the team has made great strides in reducing the 

primary risk of LOx/LCH4 ignition. At the beginning of PCAD, concerns were expressed that the ignition of 

LOx/LCH4 was not feasible. However, with a combination of in-house and contractor activities, the PCAD team has 

shown that LOx/LCH4 can be reliably ignited over a wide range of conditions. Also, under contract, PCAD has 

demonstrated that reaction control engines can be developed to the pre-prototype level which meets mission 

requirements. Due to the nature of pulsed operation, it can be argued that the LOx/LCH4 reaction control engine was 

the most challenging problem facing the team. However, despite the team’s successes, new challenges have arisen 

during the course of the project. For the reaction control engines, system interactions and operations in a cluster 

proved to be difficult. The engine and flow system operation were sensitive to system design and operation, hence 

the requirement to move forward with the Integrated Propulsion System Test Bed (IPSTB). There is also still 

individual work to be done with the reaction control engines with additional vacuum testing. Much of the 

performance work was done at sea level at single set point flow inlet conditions. PCAD is planning to do extensive 

testing to evaluate engine performance across a wide range of propellant inlet pressure and temperatures. Testing 

will also be conducted to simulate the hot and cold variations the engine will see during space operations. 

Figure 12.—NASA two-stage throttling LOx/LH2

injector during sea level testing with water cooled 

calorimeter at MSFC. 

Figure 13.—ATBV during performance testing at 

MSFC.
NASA/TM—2010-216820 12 

The ascent main engine has not had as much success as the reaction control. While the RCE work has done much 

to reduce the risks associated with the propellant combination, ultimately it is the performance of the ascent main 

engine which will determine if LOx/LCH4 is a viable candidate for the lunar ascent vehicle. Based on the system 

studies, the success is tied to the ability to demonstrate the highest level of vacuum specific impulse, with 355-sec 

being the current target. The amount of weight savings to the vehicle is directly tied to the Isp level achieved by the 

main engine. A lower specific impulse will result in a lower mass savings for the LOx/LCH4 option versus the 

current hypergolic baseline. The current effort with the Aerojet design is to see just how close the team can get a 

main engine to that goal of 355-sec. Once successful, the next step will be to develop the main engine technologies 

with a pre-prototype engine. This engine could be either ablative or regeneratively cooled. 

The descent main engine activities have successfully demonstrated stable throttling to 10% thrust or less with 

multiple injector concepts using liquid oxygen and liquid hydrogen propellants. The Pratt & Whitney Rocketdyne 

CECE demonstrator engine test series concluded with 7,435.8-sec (123.9-min) of total run time. The testing 

demonstrated chamber pressure and mixture ratio closed-loop control over a wide range of throttled power levels, 

fast throttle ramp rates, minimum power down to a smooth start to 10% power, eleven rapid relights demonstrated, 

and high power, high mixture ratio operation. The testing also demonstrated low power stability, including chug-free 

operation down to 5.9% power. This represents a 17.6:1 overall cryogenic deep throttling ratio in a complete 

expander cycle engine system with all system-level interactions which greatly enhanced the value of the technology 

database acquired. Testing with a pintle injector from Northrop Grumman was successful in meeting all primary and 

secondary technical objectives including high performance (>98% C* (combustion) efficiency); stable 10:1 deep 

throttling; measurement of heat transfer characteristics; evaluation of off-nominal oxidizer to fuel mixture ratio 

(MR) sensitivities; and evaluation of L* sensitivity. Finally, a NASA in-house developed dual oxygen manifold 

injector was also able to demonstrate stable throttling to a 10% power level. 

The PCAD team continues to build upon the success to date and strives to provide timely and relevant data to 

NASA mission study teams so an informed decision can be made on the direction of the next propulsion system for 

exploration missions. 

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Subcommittee Meeting, Dec. 2008. 

7. Grasl, S., Nguyen, and Skaff, A., “Liquid Methane/Liquid Oxygen Propellant Conditioning Feed System 

(PCFS) Test Rigs - Preliminary Test Results,” 5th JANNAF Liquid Propulsion Meeting, Colorado Springs, CO, 

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LOX/LCH4 Reaction Control Engine,” 4th JANNAF Liquid Propulsion Meeting, December 2008. 

9. Studak, J.W., Schneider, S.J., LOX/LCH4 Ignition Results with Aerojet’s RCS LOX/ethanol Igniter. 54th

JANNAF Propulsion Meeting, Denver, CO, May 2007. 

10. Schneider, S.J., John, J.W., and Zoeckler, J.G., “Design, Fabrication and Test of a LOX/LCH4 RCS Igniter,” at 

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215038. 

11. Reynolds, D.C., “Oxygen/Methane Torch Igniter Design and Testing,” NASA/TM—2008-215247, Feb. 2008. 

12. Breisacher, K. and Ajmani, K., “LOX/Methane Main Engine Igniter Tests and Modeling,” AIAA Joint 

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NASA/TM—2010-216820 13 

13. Wright and Wendell, “Design, Fabrication and Test of LO2/LCH4 Augmented Spark Igniter for 100 lbf 

Reaction Control Engine,” 5th JANNAF Liquid Propulsion Subcommittee Meeting, May 2010. 

14. Reynolds, D.C., “Microwave Igniter–Gaseous Oxygen/Gaseous Methane Vacuum Pressure Tests,” 

NASA/TM—2008-215195, Feb. 2008. 

15. Osborne, R., Peschel, W., Elam, S., Sprader, Bell, and Nichols, “Liquid Oxygen/Liquid Methane Impinging and 

Microwave Igniter Vacuum Testing at TS115,” NASA/TM—2010-216374, Mar. 2010. 

16. Breisacher, K., and Ajmani, K., “LOX/Methane Main Engine Glow Plug Igniter Tests and Modeling,” 

4th JANNAF Liquid Propulsion Subcommittee Meeting, Dec. 2008. 

17. Schneider, S., “LOX/LCH4 Igniter Pulse Durability—Test Results Combined Phase I and Phase II,” 

5th JANNAF Liquid Propulsion Subcommittee Meeting, May 2010. 

18. Robinson, P.J., Veith, E.M., Cochran, M.J., et al., “Development of a Flight-Type Exciter for a Spark-Initiated 

Torch Igniter,” 4th JANNAF Liquid Propulsion Subcommittee Meeting, Dec. 2008. 

19. DeLong, D., Greason, J., Rodway-McKee, K., “Liquid Oxygen/Liquid Methane Rocket Engine Development,” 

SAE 2007 AeroTech Congress and Exhibition, Los Angeles, CA, Sep. 2007. 

20. Engers, R., Balepin, V., Greason, J., and Pierce, C., “Liquid Oxygen/Methane Rocket Engine Development,” 

54th JANNAF Propulsion Meeting, Denver, Colorado, May 2007. 

21. Rutley., Balepin, V., Tilakos, N., Greason, J., and DeLong, “NASA Crew Exploration Vehicle Updated 

Version,” NASA/CR—2009-216267, Dec. 2009. 

22. Robinson, P.J., Veith, E.M., Linne, D.L., and Robinson, J.W., “Conceptual Design of a 5,500-lbf LOX/LCH4 

Lunar Ascent Main Engine,” 4th JANNAF Liquid Propulsion Subcommittee Meeting, Dec. 2008. 

23. Robinson, P., Veith, E.M., and Linne, D.L., “Design and Development of a 5,500-lbf LOX/LCH4 Ascent Main 

Engine,” 5th JANNAF Liquid Propulsion Subcommittee Meeting, May 2010. 

24. Stiegemeier, B., Williams, G., Melcher, J.C., and Robinson, J., “Altitude Testing of an Ascent Stage LOX/ 

Methane Main Engine,” 5th JANNAF Liquid Propulsion Subcommittee Meeting, May 2010. 

25. TDK, Two Dimensional Kinetics, version TDK’08TM, Software Engineering Associates, Carson City, NV. 

26. Elam, S.K., and Sprader, C.L., “Test Summary Report for Test Program P2514-Six-Inch Liquid Oxygen/Liquid 

Methane Injector Demonstration,” NASA/TM—2008-215470, Aug. 2008. 

27. Elam, S.K., Protz, C., and Reynolds, D., Design and Demonstration of a Liquid Oxygen/Liquid Methane 

(LOX/LCH4) Impinging Injector, 54th JANNAF Propulsion Meeting, Denver, Colorado, May 2007. 

28. Elam, S.K., and Sprader, C.L., “LOX/LCH4 Technology Demonstration - P8006A,” NASA/TM—2009-

216268, Dec. 2009. 

29. Elam, S.K, and Sprader, C.L., “LOX/LCH4 Technology Demonstration - P8006B,” NASA/TM—2009-216269, 

Dec. 2009. 

30. Elam, S.K., and Sprader, C.L., “Test Report for Test Program P7061: Liquid Oxygen/Liquid Methane 

Demonstration with Coaxial Injectors,” NASA/TM—2009-215743, Apr. 2009. 

31. Van Noord, J., A Heat Transfer Investigation of Liquid and Two-Phase. 5th JANNAF Liquid Propulsion 

Subcommittee Meeting, May 2010. 

32. Melcher, J.C., Allred, J., and Cabiran, M., Liquid Oxygen/Liquid Methane Testing of the RS-18 Lunar Ascent 

Engine at Simulated Altitude Conditions at NASA White Sands Test Facility. JANNAF Liquid Propulsion 

Meeting, Dec. 2008. 

33. Collins, J., Hurlbert, E., Romig, K.R., Melcher, J.C., Hobson, A., and Eaton, “Sea-Level Flight Demonstration 

& Altitude Characterization of a LO2/LCH4 Based Ascent Propulsion Lander,” AIAA–2009–4948, Aug. 2009. 

34. Collins, J., Melcher, J.C., Hurlbert, E., and Eaton, “Sea-Level Flight Demonstration and Altitude 

Characterization of a LO2/LCH4 Based Ascent Propulsion Lander,” 5th JANNAF Liquid Propulsion 

Subcommittee Meeting, May 2010. 

35. Giuliano, V.J., Leonard, T.G., Lyda, R.T., and Kim, T.S., “CECE Deep Throttling Technology Demonstrator 

Engine Development Status,” 4th JANNAF Liquid Propulsion Subcommittee Meeting, Dec. 2008. 

36. Leonard, T., and Giuliano, V., “Deep Throttling Common Extensible Cryogenic Engine (CECE), CECE 

Demonstrator 1 Data Mining Task,” NASA/CR—2008-215409, May 2008. 

37. Leonard, T., and Giuliano, V., “Deep Throttling Common Extensible Cryogenic Engine (CECE), CECE 

Demonstrator 1.5 Data Mining Task,” NASA/CR—2008-215251, Mar. 2008. 

38. Leonard, T., and Giuliano, V., “Deep Throttling Common Extensible Cryogenic Engine (CECE)-Demonstrator 

1.6 Data Mining Task,” NASA/CR—2009-215958. 

39. Giuliano, V., Lyda, R., and Kim, T.S., “CECE Deep Throttling Technology Demonstrator Engine Development 

Status,” 5th JANNAF Liquid Propulsion Subcommittee Meeting, May 2010. 
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40. Devine, Casiano, M., Hulka, J., Adamski, W., Brown, J., and Fang, J.J., “Throttling Characteristics of the RL 10 

Derivative Common Extensible Cryogenic Engine—Demo 1.6 and 1.7 Test Results,” 5th JANNAF Liquid 

Propulsion Subcommittee Meeting, May 2010. 

41. Majamaki, A.N., Chianese, S.G., and Kim, T.S., “TR202 Deep Throttling Lunar Descent Engine Pintle Injector 

Technology Development Status,” 4th JANNAF Liquid Propulsion Subcommittee Meeting, Dec. 2008. 

42. Gromski, Majamaki, A.N., Chianese, S., Weinstock, V., and Kim T.S., Northrop Grumman TR202 LOX-LH2 

Deep Throttling Engine Project Status, 5th JANNAF Liquid Propulsion Subcommittee Meeting, May 2010. 

43. Weinstock, V., Chianese, S., Majamaki, A.N., and Lichford, R., “Northrop Grumman TR202 LOX/GH2 Deep 

Throttling Pintle Injector Fabrication and Demonstration Testing,” 5th JANNAF Liquid Propulsion 

Subcommittee Meeting, May 2010. 

44. Chianese, S., Gromski, Weinstock, V., Majamaki, A.N., and Lichford, R., “Northrop Grumman TR202 

LOX/GH2 Deep Throttling Pintle Injector Performance, Stability, and Heat Transfer Measurements,” 5th

JANNAF Liquid Propulsion Subcommittee Meeting, May 2010. 

45. Hayes, W., Hewitt, R., Veith, E.M., Robinson, J., and Jones, G., “Development of a Vacuum Compression 

Brazed Combustion Chamber and a LO2/GH2 Deep Throttling Injector for the Marshall Space Flight Center 

Liquid Engine Test Bed,” 5th JANNAF Liquid Propulsion Subcommittee Meeting, May 2010. 

46. Hernandez, et al., “Liquid Oxygen/Hydrogen Subscale 10:1 Throttling Engine,” 4th JANNAF Liquid Propulsion 

Subcommittee Meeting, Dec. 2008. 

47. Jones, G.W., Baker, K.J., Hensley, N.J., Litchford, R.L., and Hulka, J.R., “Design of a Dual-Inlet Swirl Coaxial 

Injector for Deep-Throttling Applications,” 4th JANNAF Liquid Propulsion Subcommittee Meeting, Dec. 2008. 

48. Jones, G., and Hulka, J., “Testing of a 9k Deep Throttling LOX/GH2 Dual-Inlet Swirl Coaxial Injector,” 5th

JANNAF Liquid Propulsion Subcommittee Meeting, May 2010.
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Propulsion Risk Reduction Activities for Nontoxic Cryogenic Propulsion 

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14. ABSTRACT

The Propulsion and Cryogenics Advanced Development (PCAD) Project’s primary objective is to develop propulsion system technologies 

for nontoxic or “green” propellants. The PCAD project focuses on the development of nontoxic propulsion technologies needed to provide 

necessary data and relevant experience to support informed decisions on implementation of nontoxic propellants for space missions. 

Implementation of nontoxic propellants in high performance propulsion systems offers NASA an opportunity to consider other options than 

current hypergolic propellants. The PCAD Project is emphasizing technology efforts in reaction control system (RCS) thruster designs, 

ascent main engines (AME), and descent main engines (DME). PCAD has a series of tasks and contracts to conduct risk reduction and/or 

retirement activities to demonstrate that nontoxic cryogenic propellants can be a feasible option for space missions. Work has focused on 1) 

reducing the risk of liquid oxygen/liquid methane ignition, demonstrating the key enabling technologies, and validating performance levels 

for reaction control engines for use on descent and ascent stages; 2) demonstrating the key enabling technologies and validating performance 

levels for liquid oxygen/liquid methane ascent engines; and 3) demonstrating the key enabling technologies and validating performance 

levels for deep throttling liquid oxygen/liquid hydrogen descent engines. The progress of these risk reduction and/or retirement activities 

will be presented. 

15. SUBJECT TERMS

Liquid propellant rocket engines; Engine tests; Hydrogen oxygen engines; Methane; Throttling 

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