Encyclopedia Astronautica
SSME

SSME
Credit: Boeing / Rocketdyne

The Space Shuttle Main Engines was the only high-pressure closed-cycle reusable cryogenic rocket engine ever flown. Three of these engines were mounted in the base of the American space shuttle, and were fed liquid hydrogen and liquid oxygen propellants from the Space Shuttle external tank during ascent to orbit. The external tank was released when the desired orbit was attained, and the shuttle returned the engines to earth for reuse. The original design points for the engines were ten flights between overhaul and a vacuum specific impulse of 455 seconds. Neither goal was achieved - vacuum impulse was 453 seconds, and the engines had to be pulled, inspected, and refurbished after each flight. In the end, the shuttle proved to a very expensive method of recovering reusable engines that perhaps cost more than expendable ones�.

Fuel and oxidizer were pumped from the external tank by high pressure oxidizer and fuel turbopumps (HPOT and HPFT) mounted to each engine assembly. To prevent cavitation, the flow was boosted before entry into these pumps by low pressure oxidizer and fuel turbopumps (LPOT and LPFT). The propellant flow was tapped off during operation for a variety of purposes, including pressurization of the external fuel and oxidizer tanks and operating the LPOT and LPFT. Liquid hydrogen was pumped through the combustion chamber and engine nozzle to cool them before being discharged into the combustion chamber.

The now-gaseous hydrogen and liquid oxygen entered the chamber at the injector, which mixed the propellants. A dual-redundant igniter was used during the engine start sequence to initiate combustion.

The SSME could be throttled between 67% and what NASA called "109%" of its rated thrust. Launches normally used 104% (2170 kN) as a maximum, with 109% (2280 kN) reserved for abort emergency situations.

After NASA's decision to retire the Shuttle fleet in 2010, there was an attempt to keep the SSME in production as the second stage engine of the Ares I launch vehicle and the booster stage engine of the Ares V. However the uneconomical nature of continued SSME use became apparent, and NASA finally settled on an expendable J-2X engine for the Ares I and the RS-68 for the Ares V.

The SSME was naturally considered for the many shuttle derivatives and upgrades proposed during its forty-year development and service life. In the end, it appeared to be a technological bridge too far - the specified weight, reliability, durability, and reusability simply could not be met in a single engine using existing or foreseen technology and materials.

Thrust (sl): 1,817.400 kN (408,568 lbf). Thrust (sl): 185,330 kgf. Engine: 3,177 kg (7,004 lb). Chamber Pressure: 204.08 bar. Area Ratio: 77.5. Thrust to Weight Ratio: 73.1197829645898. Oxidizer to Fuel Ratio: 6. Coefficient of Thrust vacuum: 1.90710733043191. Coefficient of Thrust sea level: 1.52735733043191.

AKA: RS-24; RS-25.
Status: In production.
Unfuelled mass: 3,177 kg (7,004 lb).
Height: 4.24 m (13.92 ft).
Diameter: 1.63 m (5.36 ft).
Thrust: 2,278.00 kN (512,114 lbf).
Specific impulse: 453 s.
Specific impulse sea level: 363 s.
Burn time: 480 s.
First Launch: 1972.
Number: 351 .

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• Shuttle American winged orbital launch vehicle. The manned reusable space system which was designed to slash the cost of space transport and replace all expendable launch vehicles. It did neither, but did keep NASA in the manned space flight business for 30 years. Redesign of the shuttle with reliability in mind after the Challenger disaster reduced maximum payload to low earth orbit from 27,850 kg to 24,400 kg. More...

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• Lox/LH2 Liquid oxygen was the earliest, cheapest, safest, and eventually the preferred oxidiser for large space launchers. Its main drawback is that it is moderately cryogenic, and therefore not suitable for military uses where storage of the fuelled missile and quick launch are required. Liquid hydrogen was identified by all the leading rocket visionaries as the theoretically ideal rocket fuel. It had big drawbacks, however - it was highly cryogenic, and it had a very low density, making for large tanks. The United States mastered hydrogen technology for the highly classified Lockheed CL-400 Suntan reconnaissance aircraft in the mid-1950's. The technology was transferred to the Centaur rocket stage program, and by the mid-1960's the United States was flying the Centaur and Saturn upper stages using the fuel. It was adopted for the core of the space shuttle, and Centaur stages still fly today. More...

• Kudryavtseva, V M, ed., Zhidkostnikh Raketnikh Dvigatley, Visshaya Shkola, Moscow, 1993.

• Ares Stage 1 Lox/LH2 propellant rocket stage. Loaded/empty mass 787,700/64,200 kg. Thrust 8,705.99 kN. Vacuum specific impulse 453 seconds. Core vehicle proposed by NASA for Project Constellation exploration of moon and Mars. It would use shuttle external tank tooling. All masses estimated. More...

• Cargo LV Stage 1 Lox/LH2 propellant rocket stage. Loaded/empty mass 1,093,330/88,449 kg. Thrust 10,441.06 kN. Vacuum specific impulse 452 seconds. Core vehicle proposed by NASA for Project Constellation exploration of moon and Mars. Originally it would use shuttle external tank tooling. This version was proposed by Thiokol prior to Constellation decision. Modification of shuttle external tank. Includes 28.6 tonne SSME engine pod. More...

• CLV Stage 2 Lox/LH2 propellant rocket stage. Loaded/empty mass 160,000/16,000 kg. Thrust 2,333.00 kN. Vacuum specific impulse 465 seconds. Second stage originally proposed by Thiokol for launch of the CEV into low earth orbit. Also could be used as trans-Mars injection stage on the Cargo LV. Nominal single engine; alternatively 7 RL10-derived engines. All masses estimated. More...

• Magnum Core Lox/LH2 propellant rocket stage. Loaded/empty mass 830,000/70,000 kg. Thrust 9,112.38 kN. Vacuum specific impulse 453 seconds. Alternative configurations used 2 to 3 RS-68 engines More...

• Shuttle Orbiter Lox/LH2 propellant rocket stage. Loaded/empty mass 99,318/99,117 kg. Thrust 6,834.30 kN. Vacuum specific impulse 455 seconds. More...

• Starlifter Lox/LH2 propellant rocket stage. Loaded/empty mass 42,630/19,955 kg. Thrust 5,217.00 kN. Vacuum specific impulse 455 seconds. More...