Joseph F. Baumeister

Glenn Research Center, Cleveland, Ohio

RL10 Engine Ability to Transition From

Atlas to Shuttle/Centaur Program

NASA/TM—2015-218736

April 2015
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Joseph F. Baumeister

Glenn Research Center, Cleveland, Ohio

RL10 Engine Ability to Transition From

Atlas to Shuttle/Centaur Program

NASA/TM—2015-218736

April 2015

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Glenn Research Center

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NASA/TM—2015-218736 1 

RL10 Engine Ability to Transition From 

Atlas to Shuttle/Centaur Program 

Joseph F. Baumeister 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

A key launch vehicle design feature is the ability to take advantage of new technologies while 

minimizing expensive and time consuming development and test programs. With successful space launch 

experiences and the unique features of both the National Aeronautics and Space Administration (NASA) 

Space Transportation System (Space Shuttle) and Atlas/Centaur programs, it became attractive to 

leverage these capabilities. The Shuttle/Centaur Program was created to transition the existing Centaur 

vehicle to be launched from the Space Shuttle cargo bay. This provided the ability to launch heaver and 

larger payloads, and take advantage of new unique launch operational capabilities. A successful 

Shuttle/Centaur Program required the Centaur main propulsion system to quickly accommodate the new 

operating conditions for two new Shuttle/Centaur configurations and evolve to function in the human 

Space Shuttle environment. This paper describes the transition of the Atlas/Centaur RL10 engine to the 

Shuttle/Centaur configurations; shows the unique versatility and capability of the engine; and highlights 

the importance of ground testing. Propulsion testing outcomes emphasize the value added benefits of 

testing heritage hardware and the significant impact to existing and future programs. 

Nomenclature 

LH2 Liquid hydrogen 

LOX Liquid oxygen 

O/F Oxidizer fuel ratio 

ONPSP Oxidizer net positive suction pressure 

OPIP Oxidizer pump inlet pressure 

TDRSS Tracking and Data Relay Satellite System 

Introduction 

NASA’s Centaur vehicle was the first of a new generation of space vehicle upper-stage rockets that 

pioneered the use of liquid hydrogen fuel for space flight. The Centaur upper-stage was developed by 

NASA Lewis Research Center, now Glenn Research Center in Cleveland, Ohio. The Centaur became the 

world’s first high-energy upper-stage, burning liquid hydrogen (LH2) and liquid oxygen (LOX) to place 

payloads in geosynchronous orbits or to provide escape velocity for interplanetary space probes. In 

combination with the RL10 engine (currently manufactured by Aerojet Rocketdyne), the Centaur became 

the highest performing upper-stage and one of the most prominent launch vehicles in America’s 

exploration of space. The vehicle has more than 200 launches that included sending the Voyager 

spacecraft to the outer planets, Viking landers to the surface of Mars, Cassini to Saturn, and the New 

Horizons probe to Pluto. The Centaur Vehicle was developed and manufactured by General Dynamics/ 

Astronautics under the direction of NASA’s Lewis Research Center (Ref. 1). 

 
NASA/TM—2015-218736 2 

The Centaur Program started in 1958 with its first successful flight in November 27, 1963 at 

Cape Canaveral Air Force Station in Florida. Since then the vehicle has undergone numerous 

evolutionary upgrades and modification to improve its performance, operability, and reliability. From 

1966 to 1989, the Centaur-D configuration was used as the upper-stage for 63 Atlas rocket launches 

(55 successful launches). Today the Atlas/Centaur has evolved into a suite of Atlas supported vehicles 

launching from both Cape Canaveral Air Force Station and Vandenberg Air Force Base. 

Hydrogen in its natural state is a gas and is the lightest and simplest of all elements. Liquid hydrogen 

is colorless, odorless, and very lightweight. To remain in a liquid state, hydrogen must be kept 

supercooled to –423 °F (–253 °C). For the Centaur to accommodate liquid hydrogen, new technologies 

were required. This included developing a highly integrated propellant storage and utilization system that 

included thermal insulation, propellant feed system and propulsion capable engines. It also required 

appropriate hydrogen ground handling and controls. This lightweight Centaur vehicle with engines that 

utilized liquid hydrogen and oxygen mixture could then deliver about 35 percent more thrust per pound of 

propellant than other conventional kerosene-type fuels. 

The RL10 propulsion engine was a major sub-system in the successful Centaur Program and continues 

to provide sustained reliability and performance. Its efficiency and capabilities have kept the Centaur in 

the forefront of space flight. Integrating the existing Centaur RL10 engine within the Space Shuttle was a 

cornerstone of the Shuttle/Centaur Program. The RL10 operational flexibility enabled easy transition with 

minimal modifications and the ability to transition to a human space flight environment. We will explore 

the differences between the Atlas/Centaur and Shuttle/Centaur vehicles, impacts on the propulsion 

system, and the role ground testing provided. 

Atlas/Centaur Vehicle Configuration 

The Atlas/Centaur is an expendable launch vehicle derived from the SM-65D Atlas intercontinental 

ballistic missile in the late 1950s. The Atlas boosters were developed by the United States Air Force and 

were repurposed for NASA’s Atlas/Centaur Program. The Atlas first stage booster rocket was mated to 

the new Centaur liquid-hydrogen and liquid-oxygen cryogenic fueled upper stage. This high-energy 

Centaur upper-stage was capable of placing much heavier payloads into higher earth orbits, unmanned 

lunar trajectories, or planetary trajectories by incorporating the new RL10 propulsion system. This 

Atlas/Centaur design is shown in Figures 1 and 2. The Atlas/Centaur-2 launch demonstrated a successful 

Centaur vehicle separation and flight with the first RL10 engine inflight burn of liquid hydrogen and 

liquid oxygen. The Centaur vehicle had an extraordinary operational success record and has been called 

“America’s Workhorse in Space”. 

Figure 1.—Atlas/Centaur Configuration and 

Atlas/Centaur-2 Launch. 

Figure 2.—Early Centaur Configuration. 
NASA/TM—2015-218736 3 

Figure 3.—Atlas/Centaur Being Lowered into 

NASA Lewis Research Center Test Facility. 

The Centaur D-1 vehicle was 29 ft in length and had a diameter of 10 ft with all avionics and payload 

mounted on the forward adapter. It had no internal support structure and was pressurized to maintain it 

shape and strength. The thrust was obtained from two RL10 engines that generate 15,000 nominal pounds 

of thrust each. These engines were capable of being shut-down and restarted during flight to increase 

mission capability and were gimbaled to provide spacecraft directional control. Figure 3 shows the 

Centaur engine configuration as the vehicle is being lowered into a NASA Lewis Research Center Test 

facility. In addition, smaller hydrogen peroxide engines provided additional thrust for propellant settling 

and attitude control during transition and low thrust coast periods. This vehicle had the capability of 

lifting over 8000 lb payloads into near Earth orbits, 2300 lb to the Moon, and boosting 1300 lb on 

interplanetary trajectories. Unique Centaur design features and upgrades included: 

 Thin wall stainless steel tanks (lightweight). Structural rigidity was provided by propellant 

pressure (prior to propellant loading, nitrogen gas pressurization provided structural support) 

 A common double-bulkhead to separate the LOX and LH2 tanks 

 Multiple engine restart vehicle 

 Hydrazine monopropellant reaction control system (previously a hydrogen peroxide system) 

 Tank pressure from RL10 boost pumps (previous hydrogen peroxide system) 

 Centaur powered by one or two RL10 rocket engines 

Titan-Centaur Vehicle Configuration 

The Centaur was also successfully integrated with the United States Air Force Titan vehicle known as 

Titan IIIE. The Centaur was modified to be used as the third stage on the Titan IIIE vehicle. From 1974 to 

1977, the Centaur D-1T configuration launches included the Viking 1 and 2, Voyager 1 and 2, and 

Helios 1 and 2. Figure 4 shows the Titan/Centaur configuration. 

Centaur RL10 Engines 

The RL10 engine was designed as a regeneratively cooled system (hydrogen cooled thrust chamber) 

which included a turbo-pump that utilized an expander cycle (thrust chamber expanded hydrogen to 

power turbo-pump) (Ref. 2). The original Centaur engines flown on the first successful Atlas/Centaur 

flight were designated as the RL10A-3 engines. Over the years, prior to the Shuttle/Centaur Program, the 

engines were upgraded to enhance vehicle performance, shown in Table 1. The engine is shown in Figure 5. 
NASA/TM—2015-218736 4 

Figure 4.—Titan/Centaur Vehicle (Left), Centaur and Payload (Middle), 

August 20, 1975 Titan Centaur Launch of Viking 1 Spacecraft (Right). 

TABLE 1.—EARLY RL10 ENGINE PERFORMANCE UPGRADES 

RL10 model number 

RL10A-1 RL10A-3 RL10A-3-1 RL10A-3-3 RL10A-3-3A 

Thrust, lb (vacuum) 15,000 15,000 15,000 15,000 16,500 

ISP, sec (5.0 mixture) 424 429 433 444 446.4 

Chamber pressure, psia 300 300 300 395 465 

Throat diameter, in. 6.0 6.0 6.0 5.1 4.9 

Expansion ratio 40:1 40:1 40:1 57:1 61:1 

Engine envelope, in. 70 long 

40 wide Unchanged Unchanged Unchanged Unchanged 

Figure 5.—RL10A-3-3A Engine Side Views. 

 
NASA/TM—2015-218736 5 

Each Centaur RL10A-3-3A engine at space vacuum conditions produced a rated thrust of 16,500 lb 

and a 444.42.5 sec nominal specific impulse at a nominal propellant oxidizer to fuel mixture ratio of 

5.0:1. The engines also gimbal which provides flight control, are capable of inflight restarts, and provide a 

small amount of hydrogen gas for vehicle fuel tank pressurization. Solenoid valves which control the flow 

of vehicle helium to pressure actuated valves provide engine control for component chill down, engine 

start, and shutdown. Chill down of the engine pumps is required prior to engine start (Ref. 3). Additional 

features include: 

 Turbo-pump supplies both hydrogen and oxygen to the combustion chamber 

 Thrust chamber has 180 short and 180 tapered tubes for nozzle cooling and expands hydrogen to 

power the turbo-pump 

 Engine thrust is regulated by the thrust control valve 

 Propellant mixture ratio regulated by the oxidizer flow control valve 

 Engines may gimbal 4

Shuttle/Centaur Program 

Integrating the Space Shuttle with the high-energy Centaur upper-stage provided a broad range of 

new space flight possibilities. This included launching heavier and larger spacecraft, the potential to 

checkout and service the spacecraft prior to separation, and even return the spacecraft to earth if 

necessary. Figure 6 represents a Centaur deployment and Figure 7 highlights the major Shuttle/Centaur 

systems. The Shuttle/Centaur Program began in 1982 as a requirement for the upcoming Galileo and 

European Space Agency’s Ulysses spacecraft missions in 1986 and resulted in a partnership between 

NASA and the United State Air Force. The first planned missions included Galileo, Ulysses, Magellan, 

International Solar Polar Mission and Department of Defense missions. With the Shuttle/Centaur 

increased payload capabilities and unique mission flexibility, it positioned the United States to continue 

its leadership in launch vehicles. If not for the tragic Space Shuttle Challenger, the Shuttle/Centaur may 

have changed spaceflight as we know it today. 

Figure 6.—Shuttle/Centaur Deployment. 
NASA/TM—2015-218736 6 

Figure 7.—Shuttle/Centaur Hardware Configuration. 

TABLE 2.—SHUTTLE/CENTAUR AND CISS CONFIGURATIONS 

Shuttle/Centaur 

G Prime, 

ft 

Shuttle/Centaur 

G, 

ft 

Vehicle length 30 20 

Hydrogen tank diameter (increase) 14 14 

Oxygen tank diameter (unchanged) 10 10 

Centaur and CISS diameter 15 15 

Available payload length 30 40 

Major Shuttle/Centaur Program requirements were: 

 General Requirements 

o Design and develop a high-energy upper-stage for use with the Space Shuttle 

o Develop two versions to maximize launch capabilities 

 NASA Specific Requirements 

o Meet new interplanetary velocity requirements 

o Accommodate a 30 ft payload 

o Support the Galileo and International Solar Polar missions in 1986 

 United States Air Force Specific Requirements 

o 10,000 lb to geosynchronous orbit 

o 11,500 lb to 12-hr orbit (Molniya orbits) 

o Accommodate a 40 ft payload in orbiter bay 

In order to take full advantage of the Space Shuttle platform, the Centaur (based on the Centaur D-1A 

vehicle) needed to be resized, develop a Space Shuttle deployment system, make appropriate Centaur 

modifications, and ensure the Centaur addresses all human space flight requirements. This resulted in 

developing two new vehicles, the Centaur G and Centaur G-prime vehicles. Table 2 shows both 

integrated Centaur and Centaur Integrated Support System (CISS) configurations and dimensions. Each 

vehicle provided payload flexibility depending on the mission needs. The G-Prime vehicle was designed 

to hold 54 percent more propellant than Centaur D-1A. 
NASA/TM—2015-218736 7 

Centaur modifications were implemented to accommodate Space Shuttle integration and to take 

advantage of new technologies, which included: 

 New 170 in. diameter LH2 tank cylindrical section and insulation 

 New conical transition to LH2 tank 

 LH2 and LOX tank lengths are dependent on vehicle and engine burn mixture ratio 

 New TDRSS-compatible S-band transmitter and RF system 

 Dual failure-tolerant arm/safe unit 

 Added star scanner (optional) 

 Forward adapter (composite stub adapter) 

 New propellant dump system for aborts 

 New aft adapter (composite skin) and separation ring 

 Integrate the Centaur and Space Shuttle with the Centaur Integrated Support System (CISS) 

The CISS was a major systems development activity for the program. The CISS was integrated with 

the Space Shuttle to provide the means to support the Centaur during Space Shuttle flights, the ability to 

control and release the Centaur, and supported all functions in the event of a return to earth scenario. This 

included supporting all the pre-launch operations, controlling Centaur pressurization, and providing a 

means for propellant dumps in the event of a Centaur Launch Abort. Figure 8 provides the 

Shuttle/Centaur Program integrated schedule. 

Figure 8.—Shuttle/Centaur Integrated Schedule. 

 
NASA/TM—2015-218736 8 

Key to the Shuttle/Centaur Program was the propulsion system’s ability to transition with minimal 

changes. This included accommodating the new Space Shuttle environments, new operating constraints, 

supporting two new Centaur configurations, new Space Shuttle launch loads, two different engine 

propellant mixture ratios, modifying prelaunch cooling, and accommodating new vehicle interfaces 

features. Table 3 lists significant nominal vehicle and engine characteristics (Ref. 4). 

To accommodate the Space Shuttle safety requirement of no propellants at the engine interface, the 

Shuttle/Centaur Program modified the ducts and added valves for propellant isolation. This required 

introducing two new valves with demanding temperature and leakage requirements, Figure 9. 

TABLE 3.—NOMINAL VEHICLE AND ENGINE CHARACTERISTICS 

 

Atlas/Centaur 

(D-1) 

Shuttle/Centaur 

G Prime 

Shuttle/Centaur 

G 

 RL10A-3-3A RL10A-3-3A RL10A-3-3B 

Propellant Tanks:

Liquid Oxygen, lb 24,700 38,000 25,000 

Liquid Hydrogen, lb 5,100 7,500 4,000 

Engine Characteristics:

Engine Thrust, lb (vacuum) 16,500300 16,500300 15,000300 

Specific Impulse sec (vacuum) 444.42.5 444.42.5 436.02.7 

Nominal Mixture Ratio (O/F) 5.0:1 5.0:1 6.0:1 

Mixture Ratio Range 4.4:1 to 5.6:1 4.4:1 to 5.6:1 5.4:1 to 6.7:1 

Liquid Hydrogen Flow 6.2 lb/sec 6.2 lb/sec 4.9 lb/sec 

Liquid Oxygen Flow 31.0 lb/sec 31.0 lb/sec 29.3 lb/sec 

Nominal Chamber Pressure, psia 474 474 474 

Engine Start Conditions:

Prelaunch Cooldown—Turbo Pump Below 75 R None None 

Prelaunch Cooldown—Oxidizer Pump Below 270 R None None 

Figure 9.—Centaur Ducts and Pre-Valve Modifications. 

 
NASA/TM—2015-218736 9 

Integrating the prevalves into the Shuttle/Centaur configuration required a new method of efficiently 

providing necessary temperature conditioning prior to engine start (to prevent pump cavitation during 

engine start, referred to as cooldown). This resulted in a major change in the engine start cooldown 

process and inlet conditions. 

Atlas/Centaur Shuttle/Centaur 

 Ducts full (cold) 

 Engine inlet valves mounted immediately 

upstream of turbo-machine 

 Prelaunch cooldown (ground supplied liquid He) 

 Engine start:~4.5 min into launch 

 Short prestart cooldown 

 Ducts empty (warm) 

 Pre-valves added upstream of engine 

 No prelaunch cooldown 

 Engine start: range from hours to days 

 Utilize propellant to cool ducts and 

turbo-machinery 

 Long first burn prestart cooldown 

In addition, the RL10 engine had to accommodate the Space Shuttle Program human space flight 

requirements that included: 

 Space shuttle program requirements 

 Payloads safety policy and requirements 

 Safety, reliability, maintainability and quality provisions 

 Requirements for soldered electrical connections 

 Launch site accommodations. 

 Contamination control requirements 

Major Launch Event Differences Between the 

Atlas and Shuttle/Centaur Propulsion Systems 

The following highlights major Atlas and Shuttle/Centaur engine flight event and procedure 

differences. 

Prelaunch Cooldown 

The Atlas/Centaur Program utilized unique engine hardware and ground support equipment to 

introduce ground supplied cold helium gas to cool the turbo-pump assembly. This was designed to 

minimize inflight cooldown time and propellants. Prior to launch, the fuel and oxidizer pumps were 

maintained below 75 and 270 R, respectively and the fuel and oxidizer ducts were cold from the 

propellants in the ducts. The Shuttle/Centaur Program did not accommodate these features. 

Boost Flight Phase 

Atlas/Centaur had propellant and pneumatic power at its engine interfaces. Due to Space Shuttle 

safety requirements, the Shuttle/Centaur engines were inactive and all propellants and pneumatic power 

were not permitted at engine interfaces. This required the Shuttle/Centaur Program to place additional 

valves in the propellant feed system to isolate the propellants and engines. 
NASA/TM—2015-218736 10 

Separation 

Atlas/Centaur immediately began prestart procedures after stage separation. Following 

Shuttle/Centaur separation, the Centaur coasts for 45 min before pneumatic power and propellants were 

provided to the engine interface. 

Prestart 

Atlas/Centaur engine inlet valves were opened immediately after staging with the oxidizer flowing for 

nine seconds and the fuel for five seconds to provide cooldown prior to engine start. This rapid 

conditioning was needed while the vehicle was under low thrust and influence of the earth’s gravity prior 

to reaching orbit. The Space Shuttle delivers the Shuttle/Centaur vehicle into a stable obit so rapid engine 

conditioning was not a requirement. It did however; require additional propellant to properly condition 

the duct and turbo-pumps due to higher temperatures. 

Start 

Both the Atlas and Shuttle/Centaur had common engine start procedures. The Shuttle/Centaur added a 

dual pressure switch ignition system in the ignition exciter box to provide redundancy for indicating if 

acceptable internal pressure was maintained. Table 4 identifies the RL10 nominal engine requirements for 

both the G Prime and G vehicle. 

Restart 

Both the Atlas and Shuttle/Centaur had similar restart sequences with the ability to accommodate 

different cooldown times. The G-Prime vehicle required only one engine start (one engine burn to boost 

payloads into high energy interplanetary trajectories). The G vehicle required two engine starts (place 

payloads into geosynchronous orbit). Additional engine restart capabilities were available if needed. 

TABLE 4.—NOMINAL SHUTTLE/CENTAUR ENGINE START REQUIREMENTS 

Shuttle/Centaur 

G Prime 

Shuttle/Centaur 

G 

Engine: RL10A-3-3A RL10A-3-3B 

Number of Engine Starts 1 2 

Burn Time 650 sec 450 sec 

Engine start, Nominal: 

Fuel Pump Inlet Temperature 38.5 R 38.5 R 

Fuel Pump Inlet Pressure 29.8 psia 36.0 psia 

Oxidizer Pump Inlet Temperature 175.7 R 174.5 R 

Oxidizer Pump Inlet Pressure 47.7 psia 39.0 psia 

Steady state, Nominal: 

Liquid Fuel Flow 6.2 lb/sec 4.9 lb/sec 

Liquid Oxidizer Flow 31.0 lb/sec 29.3 lb/sec 
NASA/TM—2015-218736 11 

Mission Aborts 

The Space Shuttle provided the ability to abort the Centaur separation, dump propellants, and return 

the payload and vehicle back to earth. The RL10 engines were required to be reusable following a return 

to earth abort. 

With the program dependent on meeting the upcoming mission launch windows, program success 

was directly tied to timely and successful testing to verify and validate engine performance. NASA 

management initiated a variety of testing activities to accommodate Shuttle/Centaur propulsion changes. 

The testing focused on the Shuttle/Centaur Program, but it also set the stage for future advanced engine 

developments, applications and unique operations. 

Shuttle/Centaur Engine Testing 

The Space Shuttle payload volume provided the opportunity to maximize the Centaur’s mission 

capabilities by accommodating two Centaur configurations with different RL10 engine configurations and 

operating conditions. The vast majority of the Atlas/Centaur to Shuttle/Centaur transition required 

minimal hardware modifications. But management’s decision to perform a series of program related 

engine testing proved to be extremely valuable. The following summarizes major ground testing activities 

performed for the program. 

RL10 Pre Start Engine Conditioning Testing 

Whereas the Atlas/Centaur vehicle benefited from both cold propellants in the ducts and prelaunch 

cooldown procedures, the Shuttle/Centaur now had empty ducts while maintaining the ducts and engines 

at higher temperatures. With the need to condition the propellant lines and engine pump temperatures 

prior to engine start, a test program was conducted to demonstrate the ability to operate under these new 

conditions (Ref. 5). Testing focused on the new expanded inlet conditions and new cooldown 

requirements while trying to minimize propellant consumption. The expanded inlet test program was 

conducted to define the RL10A-3-3A engine start characteristics for the Shuttle/Centaur G-Prime vehicle 

and the RL10A-3-3B engine start characteristics for the Shuttle/Centaur G vehicle. It also presented a new 

set of testing issues to accurately represent flight conditions during ground testing of flow through empty 

propellant ducts with gravity effects. Therefore, verification would require a combination of analysis with 

component and engine testing. 

The existing RL10 engine test stand required integrating the new Shuttle/Centaur prevalves and 

building the propellant ducts to best simulate flight hardware. The oxidizer duct was identical to the flight 

unit and the fuel Y-duct configuration was only a single path duct. Each maintained flight design wall 

thickness, gimbal joints and bellows, and appropriate foam and Mylar insulation, as shown in Figures 10 

and 11. Testing of cooldown time and propellant consumption was performed for various inlet conditions 

and compared to calculated values. 

Testing revealed that the effect of gravity decreased the duct cooldown efficiency, particularly on the 

oxidizer duct, due to propellant traveling along the bottom surface of the horizontal duct and vapor 

cooling the upper surface. This resulted in a temperature variation of over 150 R and gas being trapped. 

This required retesting in a rig in a vertical orientation. 

The final test outcomes for both engine configurations were: 

 Successful engine start over the expanded inlet conditions for both missions 

 Reduced propellant consumption if engine cooldown flow areas were significantly reduced 

(reducing cooldown valves flow areas) and flow time was increased 

 New engine start characteristics where successfully tested with cooldown changes, Figure 12 

 Cooldown changes on engine acceleration were acceptable, Figure 13 

 Water-hammer tests for worst case conditions were acceptable 

 Qualification testing was defined 
NASA/TM—2015-218736 12 

Figure 10.—Fuel Duct Instrumented with Skin Thermocouples (Left), 

Insulated (Right). 

Figure 11.—Oxidizer Duct Instrumented with Skin Thermocouples (Left), 

Insulated (Right). 

Figure 12.—Fuel and Oxidizer Pump Start Envelopes. 

Figures 12 and 13 reveal that Centaur vehicle and mission modifications along with small engine 

changes resulted in a variety of significant engine outcomes. Figure 13 shows the RL10A-3-3A and 

RL10-3-3B engine time to accelerate impacts on all Centaur vehicles for both the original and reduced 

cooldown areas. The reduced cooldown area changes improved engine acceleration times. Note in 

Figure 13 that the time to accelerate for engines with the reduced cooldown area correlated with Oxidizer 

Pump Inlet Pressure (OPIP) rather than Oxidizer Net Positive Suction Pressure (ONPSP) as it did for the 

original cooldown areas. 
NASA/TM—2015-218736 13 

Figure 13.—Predicted Shuttle/Centaur Engine Acceleration Envelope, Original Cooldown Areas (Left), Reduced 

Cooldown Areas (Right). 

With the Centaur G vehicle potentially having very long coast times prior to geosynchronous orbit 

boost, the heating rates for the oxidizer tank (no orbit vent capability) caused a large range of possible 

oxidizer inlet temperatures (may cool or heat during the orbit duration). This produced much larger inlet 

start conditions and required longer cooldown times with impractical propellant consumption levels. For 

the G vehicle, instead of having a single cooldown time to produce the temperature profile for the 

oxidizer, it was proposed that the cooldown time would be dependent on the tank saturation pressure at 

prestart. The cooldown time would then decrease as the inlet saturation pressure increased which 

minimized the propellant flow times. 

Shuttle/Centaur engine restart cooldown times could be shorted due to propellants being in the ducts 

and the metal parts being near operating temperatures. These conditions were similar to the Atlas/Centaur 

restart operations. 

This prestart cooldown test program resulted in redefining the engine start conditions and small 

hardware changes to allow for consistent engine start while consuming the smallest amount of propellant. 

This resulted in the recommendation for a minimal first flight cooldown time with a 10 percent margin as 

listed in Table 5. It is worth noting that the flow of cooldown propellant provides a small impulse to the 

vehicles. These cooldown times were considerably longer than the Atlas/Centaur with oxidizer and fuel 

flowing for 9 and 5 sec, respectively. 
NASA/TM—2015-218736 14 

TABLE 5.—NOMINAL SHUTTLE/CENTAUR ENGINE START REQUIREMENTS 

Shuttle/Centaur 

G Prime 

Shuttle/Centaur 

G 

Fuel 45 sec 

29.9 lb 

65 sec 

44 to 66 lb 

Oxidizer 245 sec 

91.6 lb 

330 sec 

68 to 139 lb 

RL10A-3-3B Qualification Testing 

The RL10A-3-3B Qualification Test was conducted during the period of March 1985 through January 

1986. Testing was performed on two engines using similar procedures as the RL10A-3-3A qualification 

test program (Ref. 6). One engine performed hot fire testing and one structural limit testing (high 

structural loads if the thrust control failed at maximum mixture ratio). The qualification test plan was 

created to qualify the RL10A-3-3B engine, this included: 

 Reduce thrust level, 15,000 lb instead of 16,500 lb 

 Operate at a higher nominal mixture ratio (O/F) of 6.0:1 instead of 5.0:1 

 Operate over a wider range of propellant inlet conditions 

 Increase cooldown times with smaller flow areas in both fuel and oxidizer systems to provide 

more efficient cooldown 

The qualification program included three series of tests using a single development engine. 

Series I: Hardware durability and limits testing (23 firings) 

Series II: Preliminary qualification testing (26 firings) 

Series III: Qualification testing (26 firings) 

Testing included pre and post calibration tests. At the end of testing, the engine was disassembled, 

inspected, and calibrated. Quality engineering reviewed the test data. Figure 14 displays the engine 

mounted in the RL10 hot fire test stand. 

The RL10A-3-3B successfully completed qualification testing, but testing did produce one engine 

start anomaly. The objective of that particular hot fire test was to “relight 2 min after shutdown of 

previous run start and steady state inlet conditions”. The engine did not light within the allowable time 

limits. The ignition problem was found to be related to the small cooldown flow area in the oxidizer flow 

control and inlet conditions which delayed ignition beyond the test-stand no-ignition abort time of 

0.49 sec. Shuttle/Centaur configuration reduced the cooldown flow to minimize propellant usage for the 

smaller tank configuration. This resulted in additional margin being incorporated for the RL10A-3-3B 

engine for the Shuttle/Centaur Program. It was recommended that a new ignition system be design, built 

and qualified to eliminate the no-light condition. Subsequent engine testing demonstrated that the no-light 

encountered during the Qualification Test would have lit on actual flight. 

With the exception of the single ignition problem, the engine operated successfully at the higher 6.0:1 

mixture within the conditions identified in the engine specification document. 

 
NASA/TM—2015-218736 15 

Figure 14.—RL10A-3-3B Engine Mounted in the Test Stand. 

Figure 15.—RL10A-3-3B Engine Mounted in the Test Stand. 

RL10 Vibration Testing 

The environmental loads in the Space Shuttle cargo bay during a launch were significantly different 

from the well-known Atlas or Titan vehicle loads. A vibration test program was developed to gather 

vibration data on the engine, the hydraulic actuators, and the engine nozzle plug (Shuttle/Centaur unique) 

(Ref. 7). Testing was conducted during the period of May 1985 through October 1985 and was performed 

on a single engine using a configuration similar to the Space Shuttle interface, Figure 15. The test plan 

was created to: 

 Perform sinusoidal and random motion vibration tests on flight like hardware 

 Obtain data for input into NASTRAN model analysis 

 Verify engine durability that included pre and post hot fire testing 

For the sinusoidal vibration testing, a variety of shaker tests were performed with different 

configurations of the engine actuators that consisted of sine sweeps from 500 to 5 Hz (Space Shuttle pitch 

direction). With the data results provided for the NASTRAN model analysis, engine stresses and actuator 

loads were deemed well within the limits of the RL10 specification. Therefore, no additional qualification 

testing was required for the sine vibration. Testing did reveal that self-locking nuts on the nozzle support 

plug were coming loose and were replaced with safety wire nuts. 
NASA/TM—2015-218736 16 

For the random motion vibration test, the fixture was modified and consisted of random excitations 

over the frequency range of 50 to 2000 Hz for 3 min per axis. For this testing NASA provided detailed 

directions. Based on the test data results, the engine would survive the proposed random vibration 

environment. Visual inspection and pre and post hot fire tests showed the engine operated within its 

expected limits. 

The RL10A-3-3A engine successfully completed all vibration testing and verified the engine’s ability 

to operate in the demanding Space Shuttle cargo bay environment for launch and ascent. 

RL10 Gimbal Friction Testing 

The RL10 engine gimbal provided the engine thrust load path and bearing system for thrust vectoring 

to maintain vehicle flight control. This flight control is dependent on the engine actuators, gimbal 

hardware, and flight control software. With the Shuttle/Centaur introducing a new class of payloads and 

the 15,000 lb thrust engine, a test program was developed to measure the friction characteristics of the 

engine gimbal friction forces to address flight control concerns (Ref. 8). A test fixture was designed to 

simulate engine flight thrust load conditions in a thermal vacuum environment. Testing was performed at 

NASA Lewis Research Center during the period of September 1985 through November 1985, shown in 

Figure 16. The test plan was created to: 

 Determine the coulomb friction characteristics (breakaway torque prior to gimbal sliding) of the 

gimbal system for pitch and yaw axes 

 Simulate 15,000 lb thrust load 

 Simulate liquid oxygen tank temperature (175 R) 

 Vacuum up to 10–8 torr 

The gimbal assembly consisted of a conical engine mount, a pedestal, and a spider block 

incorporating a dry-lubricated journal bearing. After a total of 32,000 cycles were run, the test fixture 

performed as expected. Final results showed that the gimbal had elastic characteristics and the coulomb 

friction levels were significantly lower (40 to 60 ft-lb for production gimbals) from the RL10A-3-3A 220 

ft-lb design specification. The decrease in environmental conditions also had a small effect on decreasing 

torque required to initiate gimbal sliding. Upon completion, low amplitude and low frequency gimbal 

friction data was incorporated into autopilot performance studies. 

Figure 16.—Engine Mount Gimbal Assembly (Left) and Dummy Gimbal Install in Bell Jar 

Text Fixture (Right). 
NASA/TM—2015-218736 17 

RL10 Ignition Limits Testing 

With the Shuttle/Centaur Program introducing new propellant inlet conditions, a proposed change to 

provide consistent propellant flow to the engine igniter over the broader range of inlet conditions was 

introduced. This was the result of the no-light conditions during RL10A-3-3B testing (Ref. 9). To address 

this concern a test plan was created to: 

 Perform ignition limits testing to characteristics of the engine igniter mixture ratio over a wide 

range of propellant inlet conditions 

 Investigate methods to correct potential problems 

Final results showed that there was a need to increase the engine igniter mixture ratio (hardware 

modifications to increase igniter oxygen flow) to ensure that the igniter performed successfully. This 

resulted in igniter pluming and igniter leakage control modifications and successful engine ignition 

throughout the broader propellant inlet range. 

Testing also provided valuable insight to warrant further investigation on cooldown valves discharge 

conditions that provided a more accurate engine start condition. During typical hot fire tests, the 

cooldown valves were not operated at vacuum conditions due to the possibility of water contamination in 

the engine from the steam ejector that provided the vacuum. It became evident that hot fire testing, when 

the cooldown values were vented to a vacuum (Shuttle/Centaur flight environment), was more demanding 

and could induce no light conditions. Testing revealed that a more favorable start condition was present 

without the vacuum. Additional tests were to be performed, but were not fully completed due to the 

cancellation of the Shuttle/Centaur Program. 

RL10 Ignition System Environmental Testing 

The RL10 engine ignition system contained an exciter assembly (box) that is hermetically sealed with 

an internal pressure between 20 and 35 psia to prevent electrical breakdown when exposed to vacuum 

conditions. This pressure is monitored by micro-switches within the exciter box. The Shuttle/Centaur dual 

pressure switch system is electronically similar to the Atlas/Centaur single pressure system configuration. 

With the Shuttle/Centaur Program functioning under new environmental operating conditions, a test 

program was developed to verify the ignition pressure switches and system performance (Ref. 10). The 

test plan was created to: 

 Determine how the system would operate at temperatures down to –300 F (design requirement is 

–180 F) 

 Identify temperature conditions the pressure switches will open at (normally closed when internal 

pressures are greater than 20 psia) 

 Identify internal box pressure when exciter fails 

 Correlate exciter box external temperature to internal pressure 

 Use a surplus development test unit that was manufactured in March 1966 (20 years old) 

The goal of the testing was to provide a much broader understanding of the pressure switch 

operations (open and close conditions) and exciter assembly. As testing was underway, several factors 

became evident that influenced the results and complicated data interpretation. This included: 

 System testing of hardware outside its design specifications made it hard to draw conclusions 

from its unpredictable operations 

 The 20 year old unit appeared to operate unpredictably (exciter open and close lagged with both 

temperature and pressure) 
NASA/TM—2015-218736 18 

 Pressure switch performance below the specified –160 F requirement was less than operational 

minimums 

 Exciter spark rate operated at –225 F but was below specification at –275 F and barely fired at 

–300 F. This was presumed to be from the worn condition of the gas tube rectifier 

 The test unit ultimately failed when internal pressured dropped to 3.2 psia (well below 

specification) 

 Data collected to correlate exciter box external temperature to internal pressure was erratic and 

inconclusive 

The testing had no impact on the program but did provide a better understanding of the hardware and 

the testing procedures. Additional testing was recommended if the engine ignition system may operate 

below –225 F. 

Shuttle/Centaur Program Impacts on the RL10 Engine 

Developing the RL10 engine for the Shuttle/Centaur Program presented the manufacturer with the 

challenges of supporting multiple vehicle platforms, different engine configurations, and new human 

space flight requirements. The Shuttle/Centaur Program resulted in two new RL10 Engine Specification 

Documentations, a new RL10A-3-3A that accommodated both Atlas/Centaur and Shuttle/Centaur 

applications and the RL10A-3-3B for the unique Shuttle/Centaur G application. Even though both the 

Atlas/Centaur and Shuttle/Centaur refer to a common RL10A-3-3A engine, different hardware and 

operating conditions were required. 

Designing and ground testing the RL10 for the Shuttle/Centaur Program resulted in several small 

hardware modifications and operational changes for the engine. These changes included adding a 

redundant pressure switch in the ignition system, reducing the oxidizer flow control cooldown area, and 

changing the helium plumbing to improve fuel pump cooldown efficiency. Even with small hardware and 

operational modifications, considerable changes were introduced from the heritage Atlas/Centaur 

proficiency, Table 6. 

TABLE 6.—SHUTTLE/CENTAUR PROPULSION CHANGES 

Engine related elements Effect on 

Shuttle/Centaur 

Performance, steady-state thrust and specific impulse (RL10A-3-3B) Change 

Propellant inlet conditions Change 

Prelaunch propellant engine interface Change 

Prelaunch cooldown Change 

Boost flight phase Change 

Shuttle/Centaur separation/deployment New 

Prestart Change 

Start No-Change 

Steady state and shutdown No Change 

Restart No Change 

Mission aborts New 

Electrical power, for propellant valve control No Change 

Ignition Change 

Pneumatic power, for engine operations No Change 

Propellants, propellant feed systems operation limits No Change 

Thermal control (precooling prior to engine start) Change 

Instrumentation No Change 

Hydrogen gas bleed for fuel tank pressurization No Change 

Ventage and leakage for propellants and pneumatics interfaces Change 

Support equipment, nozzle support plugs and removable desiccant containers New 

External cleanliness New 

Propellant contaminant limits (adjusted for larger tank and operating time) Change 
NASA/TM—2015-218736 19 

Shuttle/Centaur Termination 

On January 28, 1986, shortly before the first Shuttle/Centaur launch, Space Shuttle Challenger 

mission STS-51L broke apart 73 sec into its flight, leading to the tragic loss of its seven crew members. 

The disaster resulted in a lengthy accident investigation and a 32-month hiatus in the Space Shuttle 

Program. As a result of this accident, the Shuttle/Centaur Program was terminated. With the program 

termination, both NASA and the United States Air Force lost the ability to launch heavier and larger 

spacecraft and take advantage of a unique deployment system. With the United States Air Force’s need 

for increased launch capabilities, they took over the Shuttle/Centaur G-Prime development and applied it 

to the Titan booster. This resulted in the new Titan IV (401A/B) vehicle with the 14 ft diameter hydrogen 

tank Centaur. 

Summary 

The key Shuttle/Centaur Program design feature was to transition the existing Atlas/Centaur RL10 

engines to meet this bold and radical concept of launching a Centaur rocket from the Space Shuttle cargo 

bay. To optimize the Space Shuttle launch capabilities, two Shuttle/Centaur configurations were 

developed (Centaur G and G-Prime). Utilizing the Atlas/Centaur heritage hardware and expertize 

provided the program a path with minimal risk. With the repurposing of existing flight hardware and 

demanding launch window requirements, justifications that minimized hardware testing could have 

resulted. Management’s decision to verify and validate key components and systems is a testament to 

truly understanding that heritage hardware for new application has built in uncertainties and risks. The 

versatility of the RL10 engine resulted in the qualification of two new RL10 engine configurations in 

record time and with human rated flight capabilities. 

The Shuttle/Centaur RL10 engine ground test program provided valuable information that included: 

 A new understanding of engine and duct prestart cooldown conditions and constraints. Hardware 

modifications were implemented 

 The engine mixture ratio change did result in an engine start anomaly. This was related to 

cooldown flow areas; hardware modifications were implemented 

 Vibration testing revealed self-locking nuts on the new nozzle support plug were coming loose. 

Hardware modifications were implemented 

 Gimbal assembly coulomb friction levels were significantly lower for the new RL10A-3-3B 

engine. Software changes were implemented 

 Engine igniter mixture ratio needed to be increased to ensure igniter performance for all RL10 

engines 

 Identified a small igniter leakage that resulted in a design change 

 Uncovered engine test stand testing conditions that misrepresented flight conditions 

 Provided a broader understanding of the exciter box pressure switch operations and testing 

procedures 

The majority of these findings required implementing engine modifications. But testing also provided 

the confidence that the RL10 engine was well suited to meet the Shuttle/Centaur missions. Ground testing 

provided value added insight into the new engine operating conditions, whereas a flight demonstration 

test would not have. 

Conclusion 

To meet the bold and radical concept of launching a Centaur rocket from the Space Shuttle cargo bay, 

the Shuttle/Centaur Program reconfigured the Centaur vehicle to be stowed in the Space Shuttle cargo bay 

and deployed in low earth orbit. Incorporating the Centaur vehicle into the Space Shuttle provided an 

attractive and relatively simple approach to enhanced boost capacity and provided both NASA and 
NASA/TM—2015-218736 20 

United States Air Force with unique deployment capabilities. To meet the Shuttle/Centaur Program goals 

that included cost, safety, reliability, schedule and technical requirements, the integration of the Centaur 

with the Space Shuttle needed to retain the vast majority of the existing Centaur hardware. A key program 

design feature was the ability to transition the existing Atlas/Centaur RL10 engine. 

Upgrading the Centaur to accommodate the Space Shuttle human space flight requirements was the 

primary concern. Leveraging from the successful Atlas/Centaur vehicle heritage was a factor in reducing 

program development time, testing, and risk. The ability of the RL10 engine to transition to the 

Shuttle/Centaur configurations with limited modifications showed the engine architecture’s unique 

versatility and capability. 

The Shuttle/Centaur Program required the engine to quickly accommodate new operating and 

environmental conditions with minimal hardware changes. The RL10 engines long proven flight record 

and what may have seemed like trivial engine modifications was extremely valuable for program 

conceptualization and definition. But the added value and risk reduction of actual component and system 

ground testing provided critical information that a flight demo would have missed. This was evident from 

the Shuttle/Centaur RL10 testing outcomes. It also provided the opportunity to truly learn how the 

hardware performed outside its original design specifications and identified new capabilities and 

operating limits. It also gave both new and experienced engineers testing and operating experiences aswell-as exposure to good engineering practices. It also provided valuable hands-on experience and 

confidence to apply knowledge and lessons learned for time critical launch decisions. 

In hindsight, the RL10 engine was well suited to accommodate the Shuttle/Centaur missions with its 

capabilities, flexibility and durability. But testing of heritage hardware applied to new missions provided 

a valuable learning opportunity and ideal approach to accurately comprehend both verification and 

validation requirements. Without the ground test activities, the highest levels of safety and reliable would 

have never been achieved and future applications would have lost the benefits of knowledge gained. 

References 

1. Stofan, A.J., “A High Energy Stage for the National Space Transportation System,” 

NASA TM-83795, 1984. 

2. “RL10 Liquid Rocket Engine, Service Manual, Model RL10A-3-3A,” Pratt and Whitney Aircraft, 

Feb. 15, 1982. 

3. “Model Specification Rocket Engine, Liquid Propellant Pratt & Whitney RL10A-3-3A 

Atlas/Centaur and Shuttle/Centaur Application,” Pratt and Whitney Aircraft, April 11, 1985. 

4. “Model Specification Rocket Engine, Liquid Propellant Pratt & Whitney RL10A-3-3B 

Shuttle/Centaur Installation,” Pratt and Whitney Aircraft, April 11, 1985. 

5. “NASA Final Report on all Shuttle/Centaur Engine and Nozzle Support Plug Vibration Testing,” 

United Technologies Corporation, Pratt & Whitney, NASA Contract NAS3-23791, May 1987. 

6. Vogel T., Varella D., Smith C., “NASA RL10A-3-3B High Mixture Ratio Qualification Program, 

Final Report,” United Technologies Corporation, Pratt & Whitney, NASA Contract NAS3-22902, 

May 1987. 

7. “NASA Final Report on all Shuttle/Centaur Engine and Nozzle Support Plug Vibration Testing,” 

United Technologies Corporation, Pratt & Whitney, NASA Contract NAS3-23791, May 1987. 

8. “NASA RL10 Gimbal Friction Test, Final Report,” United Technologies Corporation, Pratt & 

Whitney, NASA Contract NAS3-23313, May 1987. 

9. “NASA RL10 Ignition Limits Test for Shuttle/Centaur, Final Report,” United Technologies 

Corporation, Pratt & Whitney, NASA Contract NAS3-22902, May 1987. 

10. “NASA RL10 Ignition System Environmental Testing, Final Report,” United Technologies 

Corporation, Pratt & Whitney, NASA Contract NAS3-23791, May 1987.