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RS-25: Difference between revisions

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This article is about the space shuttle engines. For the services field, see Service Science, Management and Engineering.

The Space Shuttle Main Engines (SSMEs) are the three main engines on the Space Shuttle orbiter. They are constructed by Pratt & Whitney's Rocketdyne Division. The SSME is also designated as the RS-25 for engineering purposes.

The Space Shuttle main engines are very sophisticated power plants that burn liquid hydrogen and liquid oxygen from the Space Shuttle external tank. They are used for propulsion during its ascent, in addition to the two more powerful Solid Rocket Boosters. Each engine can generate almost 1.8 MN (400,000 pounds) of thrust at liftoff. The engines are capable of generating a specific impulse (Isp) of 453 seconds in a vacuum, or 363 seconds at sea level (exhaust velocities of 4440 m/s and 3560 m/s respectively). Overall, a space shuttle main engine weighs approximately 3.2 t (7,000 pounds). The engines are removed after every flight and taken to the Space Shuttle Main Engine Processing Facility (SSMEPF) for inspection and replacement of any necessary components.

The Space Shuttle's rocket engines are capable of operating at extreme temperatures. The liquid hydrogen fuel is stored at −253 degrees Celsius (−423 degrees Fahrenheit). However, when burned with liquid oxygen, the temperature in the combustion chamber reaches 3,300 °C (6,000 °F), higher than the boiling point (thus beyond the melting point) of iron. If the main engines pumped water instead of liquid oxygen and liquid hydrogen, an average family-sized swimming pool could be drained in 25 seconds.

Apart from the three main engines, the orbiter has 44 smaller rockets around the orbiter's surface, which are part of the Orbital Maneuvering System and Reaction Control System, used to provide steering, pointing, and altitude adjustment capability while in orbit.

The engines perform as follows: Fuel and oxidizer from the external tank enters the orbiter at the orbiter/external tank umbilical disconnect and then the orbiter's main propulsion system feed lines. There the fuel and oxidizer each branch out into three parallel paths, to each engine. In each branch, prevalves must be opened to permit flow to the low-pressure fuel or oxidizer turbopump.

The Low Pressure Oxidizer Turbopump (LPOTP) is an axial-flow pump driven by a six-stage turbine powered by liquid oxygen. It boosts the liquid oxygen's pressure from 0.7 to 2.9 MPa (100 to 422 psia). The flow from the LPOTP is supplied to the High-Pressure Oxidizer Turbopump (HPOTP). During engine operation, the pressure boost permits the High Pressure Oxidizer Turbine to operate at high speeds without cavitating. The LPOTP operates at approximately 5,150 rpm. The LPOTP, which is approximately 450 by 450 mm (18 by 18 inches) , is connected to the vehicle propellant ducting and supported in a fixed position by the orbiter structure.

The HPOTP consists of two single-stage centrifugal pumps (a main pump and a preburner pump) mounted on a common shaft and driven by a two-stage, hot-gas turbine. The main pump boosts the liquid oxygen's pressure from 2.9 to 30 MPa (422 to 4,300 psi) while operating at approximately 28,120 rpm. The HPOTP discharge flow splits into several paths, one of which is routed to drive the LPOTP turbine. Another path is routed to and through the main oxidizer valve and enters into the main combustion chamber. Another small flow path is tapped off and sent to the oxidizer heat exchanger. The liquid oxygen flows through an anti-flood valve that prevents it from entering the heat exchanger until sufficient heat is present to convert the liquid oxygen to gas. The heat exchanger utilizes the heat contained in the discharge gases from the HPOTP turbine to convert the liquid oxygen to gas. The gas is sent to a manifold and is then routed to the external tank to pressurize the liquid oxygen tank. Another path enters the HPOT second-stage preburner pump to boost the liquid oxygen's pressure from 30 to 51 MPa (4,300 psia to 7,420 psia). It passes through the oxidizer preburner oxidizer valve into the oxidizer preburner and through the fuel preburner oxidizer valve into the fuel preburner. The HPOTP is approximately 600 by 900 mm (24 by 36 inches). It is attached by flanges to the hot-gas manifold.

The HPOTP turbine and HPOTP pumps are mounted on a common shaft. Mixing of the fuel-rich hot gas in the turbine section and the liquid oxygen in the main pump could create a hazard. To prevent this, the two sections are separated by a cavity that is continuously purged by the MPS engine helium supply during engine operation. Two seals minimize leakage into the cavity. One seal is located between the turbine section and the cavity, and the other is between the pump section and cavity. Loss of helium pressure in this cavity results in an automatic engine shutdown.

Fuel enters the orbiter at the liquid hydrogen feed line disconnect valve, then flows into the orbiter liquid hydrogen feed line manifold and branches out into three parallel paths to each engine. In each liquid hydrogen branch, a prevalve permits liquid hydrogen to flow to the low-pressure fuel turbopump when the prevalve is open.

The Low Pressure Fuel Turbopump (LPFTP) is an axial-flow pump driven by a two-stage turbine powered by gaseous hydrogen. It boosts the pressure of the liquid hydrogen from 30 to 276 psia (0.2 to 1.9 MPa) and supplies it to the High-Pressure Fuel Turbopump (HPFTP). During engine operation, the pressure boost provided by the LPFTP permits the HPFTP to operate at high speeds without cavitating. The LPFTP operates at approximately 16,185 rpm. The LPFTP is approximately 450 by 600 mm (18 by 24 inches). It is connected to the vehicle propellant ducting and is supported in a fixed position by the orbiter structure 180 degrees from the LPOTP.

The HPFTP is a three-stage centrifugal pump driven by a two-stage, hot-gas turbine. It boosts the pressure of the liquid hydrogen from 1.9 to 45 MPa (276 to 6,515 psia). The HPFTP operates at approximately 35,360 rpm. The discharge flow from the turbopump is routed to and through the main valve and then splits into three flow paths. One path is through the jacket of the main combustion chamber, where the hydrogen is used to cool the chamber walls. It is then routed from the main combustion chamber to the LPFTP, where it is used to drive the LPFTP turbine. A small portion of the flow from the LPFTP is then directed to a common manifold from all three engines to form a single path to the external tank to maintain liquid hydrogen tank pressurization. The remaining hydrogen passes between the inner and outer walls to cool the hot-gas manifold and is discharged into the main combustion chamber. The second hydrogen flow path from the main fuel valve is through the engine nozzle (to cool the nozzle). It then joins the third flow path from the chamber coolant valve. The combined flow is then directed to the fuel and oxidizer preburners. The HPFTP is approximately 550 by 1100 mm (22 by 44 inches). It is attached by flanges to the hot-gas manifold.

The oxidizer and fuel preburners are welded to the hot-gas manifold. The fuel and oxidizer enter the preburners and are mixed so that efficient combustion can occur. The augmented sparkigniter is a small combination chamber located in the center of the injector of each preburner. The two dual-redundant spark igniters, which are activated by the engine controller, are used during the engine start sequence to initiate combustion in each preburner. They are turned off after approximately three seconds because the combustion process is then self-sustaining. The preburners produce the fuel-rich hot gas that passes through the turbines to generate the power to operate the high-pressure turbopumps. The oxidizer preburner's outflow drives a turbine that is connected to the HPOTP and the oxidizer preburner pump. The fuel preburner's outflow drives a turbine that is connected to the HPFTP.

The speed of the HPOTP and HPFTP turbines depends on the position of the corresponding oxidizer and fuel preburner oxidizer valves. These valves are positioned by the engine controller, which uses them to throttle the flow of liquid oxygen to the preburners and, thus, control engine thrust. The oxidizer and fuel preburner oxidizer valves increase or decrease the liquid oxygen flow, thus increasing or decreasing preburner chamber pressure, HPOTP and HPFTP turbine speed, and liquid oxygen and gaseous hydrogen flow into the main combustion chamber, which increases or decreases engine thrust, thus throttling the engine. The oxidizer and fuel preburner valves operate together to throttle the engine and maintain a constant 6-1 propellant mixture ratio.

The main oxidizer valve and the main fuel valve control the flow of liquid oxygen and liquid hydrogen into the engine and are controlled by each engine controller. When an engine is operating, the main valves are fully open.

Acoolant control valve is mounted on the combustion chamber coolant bypass duct of each engine. The engine controller regulates the amount of gaseous hydrogen allowed to bypass the nozzle coolant loop, thus controlling its temperature. The chamber coolant valve is 100 % open before engine start. During engine operation, it will be 100 % open for throttle settings of 100 to 109 % for maximum cooling. For throttle settings between 65 to 100 %, its position will range from 66.4 to 100 % open for reduced cooling.

Each engine main combustion chamber receives fuel-rich hot gas from a hot-gas manifold cooling circuit. The gaseous hydrogen and liquid oxygen enter the chamber at the injector, which mixes the propellants. A small augmented spark igniter chamber is located in the center of the injector. The dual-redundant igniter is used during the engine start sequence to initiate combustion. The igniters are turned off after approximately three seconds because the combustion process is self-sustaining. The main injector and dome assembly is welded to the hot-gas manifold. The main combustion chamber also is bolted to the hot-gas manifold.

The inner surface of each combustion chamber, as well as the inner surface of each nozzle, is cooled by liquid hydrogen flowing through coolant passages. The nozzle assembly is a bell-shaped extension bolted to the main combustion chamber. The nozzle is 2.9 m (113 inches) long, and the outside diameter of the exit is 2.4 m (94 inches). A support ring welded to the forward end of the nozzle is the engine attach point to the orbiter-supplied heat shield. Thermal protection is necessary because of the exposure portions of the nozzles experience during the launch, ascent, on-orbit and entry phases of a mission. The insulation consists of four layers of metallic batting covered with a metallic foil and screening.

The five propellant valves on each engine (oxidizer preburner oxidizer, fuel preburner oxidizer, main oxidizer, main fuel, and chamber coolant) are hydraulically actuated and controlled by electrical signals from the engine controller. They can be fully closed by using the MPS engine helium supply system as a backup actuation system.

The main oxidizer valve and fuel bleed valve are used after shutdown. The main oxidizer valve is opened during a propellant dump to allow residual liquid oxygen to be dumped overboard through the engine, and the fuel bleed valve is opened to allow residual liquid hydrogen to be dumped through the liquid hydrogen fill and drain valves overboard. After the dump is completed, the valves close and remain closed for the remainder of the mission.

The gimbal bearing is bolted to the main injector and dome assembly and is the thrust interface between the engine and orbiter. The bearing assembly is approximately 290 by 360 mm (11.3 by 14 inches).

The low-pressure oxygen and low-pressure fuel turbopumps are mounted 180 degrees apart on the orbiter's aft fuselage thrust structure. The lines from the low-pressure turbopumps to the high-pressure turbopumps contain flexible bellows that enable the low-pressure turbopumps to remain stationary while the rest of the engine is gimbaled for thrust vector control. The liquid hydrogen line from the LPFTP to the HPFTP is insulated to prevent the formation of liquid air.

SSME thrust can be throttled between 67 to 109% of rated thrust. Current launches use 104%, with 106 or 109% available for abort contingencies. Thrust can be specified as sea level or vacuum thrust. Vacuum thrust will be higher due to no atmospheric effects.

• 100% thrust (sea level / vacuum): 1670 kN / 2090 kN (375,000 lbf / 470,000 lbf)
• 104% thrust (sea level / vacuum): 1750 kN / 2170 kN (393,800 lbf / 488,800 lbf)
• 109% thrust (sea level / vacuum): 1860 kN / 2280 kN (417,300 lbf / 513,250 lbf)

Originally, the SSME was to see service in the post-Shuttle era as the main engines for the unmanned Ares V cargo-launch vehicle and as a second-stage engine for the manned-rated Ares I crew-launch vehicle. Although the use of the SSME seemed good on paper, as it would use current Shuttle technology after the Shuttle's retirement in 2010, it had several drawbacks:

• It would not be reusable, as they would be permanently attached to the discarded stage(s).
• It would have to undergo a flight-readiness firing (FRF) before installation – the so-called "Main Engine Test" that NASA conducted with each new Orbiter and prior to the STS-26 flight.
• It would be expensive and time-consuming to convert the ground-started SSME to an air-started version for the Ares I second stage.

With several design changes to the Ares I and Ares V rockets, the SSME will be replaced with a single J-2X engine for the Ares I second stage. The Ares V will use five modified RS-68 engines (which is based on both the SSME and Apollo-era J-2 engine) for its core stage. Hence the SSMEs will be retired along with the Shuttle fleet.

• MPTA-098 - the SSME test article used in Shuttle development
• SSME energy and power relationships

Design Altitude = 60,000 feet
Nozzle Mach Number = 5.05 (calculated)
Throat Area = 93 square inches
Nozzle Area = 50.265 square feet
Chamber Pressure = 2747 psi
Exit Pressure = 1.049 psi (calculated)
Burn Time = 520 seconds
Vacuum Isp = 452.5 seconds
Vacuum Thrust per Engine = 490,850 pounds at 104.5% of Design Thrust

• "NASA Shuttle Press Kit SSME Reference" (PDF).
• "Space Shuttle Main Engine". Boeing.
• "Space Shuttle Main Engine Enhancements". NASA.
• "The Roar of Innovation". NASA.
• "Space Shuttle Main Engine - incredible facts".
• "Space Shuttle Main Engine The First Ten Years" (PDF).
• "NSTS 1988 News Reference Manual".