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Space Shuttle main engine

By Wikipedia,
the free encyclopedia,

http://en.wikipedia.org/wiki/Space_shuttle_main_engine

SSME redirect here. For the services field, see Service Science, Management and Engineering
Space Shuttle Main Engine

Space Shuttle Main Engine test firing
Country of Origin United States
First flight February 18, 1977
Manufacturer Pratt & Whitney Rocketdyne
Status flying
Liquid-fuelled engine
Propellant LOX / Liquid hydrogen
Cycle Staged combustion
Nozzle Area ratio 77
Performance
Thrust(Vac) 490,850 pounds at 104.5% of Design Thrust
Thrust(SL) 400,000 pounds
Chamber pressure 2,747 psi (18.9 MPa) at 100% power
Isp(Vac) 452.5 seconds
Isp(SL) 363 seconds

Space Shuttle main engines (SSMEs) are reusable liquid-fuel rocket engines built by Rocketdyne. Each Space Shuttle ascent to orbit is propelled by three of the fourteen SSME engines currently used by the NASA Space Shuttle program. After each flight, the three SSMEs are removed from the Space Shuttle orbiter, inspected and refurbished in preparation for reuse on a subsequent flight. The SSME is also designated as the RS-24 for engineering purposes.

Introduction

The Space Shuttle main engines 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 and sometimes the Orbital Maneuvering System. Each engine can generate almost 1.8 meganewtons (MN) or 400,000 lbf 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 lb). 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 of iron. The main engines collectively consume 3,917 liters (1,035 gallons) of propellant per second. If the main engines pumped water instead of liquid oxygen and liquid hydrogen, an average-sized swimming pool could be drained in 25 seconds.

Apart from the three main engines, the orbiter has 44 smaller rockets around its 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.

Oxidizer system


Major components of the Space Shuttle main engine
Major components of the Space Shuttle main engine

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 420 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 measures 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 (420 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 7,400 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 measures 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.

Hydrogen fuel system


Orbiter main propulsion system
Orbiter main propulsion system

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.

Pre-burners and thrust control system


Main Engine #1 being installed into an orbiter in one of the Orbiter Processing Facilities (OPF)
Main Engine #1 being installed into an orbiter in one of the Orbiter Processing Facilities (OPF)

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 spark igniter 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.

Cooling control system

A coolant 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.


Space Shuttle Atlantis' main engines visibly in operation during the launch of STS-117.
Space Shuttle Atlantis' main engines visibly in operation during the launch of STS-117.

Combustion chamber and nozzle

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 brazed stainless steel tube-wall 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.

For a nozzle able to run at sea level, the SSME nozzle has an unusually large expansion ratio (about 77) for the chamber pressure. A nozzle that large would normally undergo flow separation of the jet from the nozzle which would cause control difficulties and could even mechanically damage the vehicle. Instead the Rocketdyne engineers varied the angle of the nozzle, reducing it near the exit. This raises the pressure just around the rim to between 4.6 and 5.7 psi, and prevents flow separation. The inner part of the flow is at much lower pressure, around 2 psi or less.

Main valves

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.

Gimbal

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 specifications

SSME thrust (or power level) can be throttled between 67 to 109% of rated thrust. Current launches use 104.5%, 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 the absence of atmospheric effects.

  • 100% thrust (sea level / vacuum): 1670 kN / 2090 kN (375,000 lbf / 470,000 lbf)
  • 104.5% 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)

Specifying power levels over 100% may seem confusing, but there is a logic behind it. The 100% level does not mean the maximum physical power level attainable. Rather it is a specification, decided on early during SSME development, for the "normal" rated power level. Later studies indicated the engine could operate safely at levels above 100%, which is now the norm. Maintaining the original relationship of power level to physical thrust helps reduce confusion. It creates an unvarying fixed relationship, so that test data, or operational data from past or future missions can be easily compared. If each time the power level was increased, that value was made 100%, then all previous data and documentation would either require changing, or cross-checking against what physical thrust corresponded to 100% power level on that date.

SSME power level affects engine reliability. Studies indicate the probability of an engine failure increases rapidly with power levels over 104.5%, which is why those are retained for contingency use only.

The SSME after the Shuttle era

<-img src="http://upload.wikimedia.org/wikipedia/commons/thumb/2/22/Space-shuttle-engine-NASA.jpg/200px-Space-shuttle-engine-NASA.jpg"/>
Main engine of a US Space Shuttle

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, time-consuming, and weight-intensive 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 six 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. If, however, the DIRECT Jupiter family replaces the Ares rockets due to delays and cost overruns, the core stage will use 3 or 4 SSMEs.

Specifications

Specifications as listed in the Encyclopedia Astronautix

  • Engine Model: SSME
  • Manufacturer Name: RS-25
  • Other Designations: RS-24
  • Designer: Rocketdyne
  • Developed in: 1972
  • Propellants: Lox/LH2
  • Thrust(vac): 2,278.000 N (512,114 lbf)
  • Thrust(sl): 1,817.400 N (408,568 lbf)
  • Isp: 453 sec
  • Isp (sea level): 363 sec
  • Burn time: 480 sec
  • Mass Engine: 3,177 kg (7,004 lb)
  • Diameter: 1.63 m (5.36 ft)
  • Length: 4.24 m (13.92 ft)
  • Chambers: 1
  • Chamber Pressure: 204.08 bar
  • Area Ratio: 77.50
  • Oxidizer to Fuel Ratio: 6.00
  • Thrust to Weight Ratio: 73.12
  • Country: USA. Status: In production
  • First Flight: 1981

Other Specifications as previously listed on Wikipedia:

  • Design altitude = 60,000 feet (18,300 m)
  • Nozzle Mach number = 5.05 (calculated)
  • Throat area = 93 square inches (600 cm²)
  • Nozzle area = 50.265 square feet (4.6698 m)
  • Chamber pressure = 2,747 pounds per square inch (18,940 kPa) at 100% power
  • Exit pressure = 1.049 pounds per square inch (7.23 kPa) (calculated)
  • Burn time = 520 seconds
  • Vacuum Isp = 452.5 seconds
  • Vacuum thrust per engine = 490,850 pounds-force (222,650 kgf) at 104.5% of design thrust

See also

  • MPTA-098 - the SSME test article used in Shuttle development



Text from Wikipedia is available under the Creative Commons Attribution/Share-Alike License; additional terms may apply.


Published - July 2009














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