Space Shuttle Main Engine test firing. (The bright area at the bottom of the picture is a Mach disk.) |
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| Country of origin | United States |
|---|---|
| First flight | 12 April 1981 (STS-1) |
| Manufacturer | Pratt & Whitney Rocketdyne |
| Associated L/V | Space Shuttle Space Launch System |
| Status | Inactive since STS-135 |
| Liquid-fuel engine | |
| Propellant | Liquid oxygen / Liquid hydrogen |
| Cycle | Staged combustion |
| Configuration | |
| Nozzle ratio | 77 |
| Performance | |
| Thrust (Vac.) | 512,300 lbf (2.279 MN) |
| Thrust (SL) | 418,000 lbf (1.86 MN) |
| Chamber pressure | 2,994 psi (20.64 MPa) |
| Isp (Vac.) | 452.3 seconds |
| Isp (SL) | 363 seconds |
| Dimensions | |
| Length | 168 inches (4.3 m) |
| Diameter | 96 inches (2.4 m) |
| Dry weight | 7,774 pounds (3,526 kg) |
| References | |
| References | [1][2] |
| Notes | Data is for RS-25D Block II at 109% throttle. |
The RS-25, otherwise known as the Space Shuttle Main Engine (SSME), is a liquid-fuel cryogenic rocket engine built in the United States of America by Pratt & Whitney Rocketdyne and used on the Space Shuttle and Space Launch System (SLS). The RS-25 burns cryogenic liquid hydrogen and liquid oxygen from the Space Shuttle external tank or SLS core stage and has two main variants, the reusable RS-25D used in the Space Shuttle (in clusters of 3 engines) and the expendable RS-25E which will be used in the SLS (in clusters of up to 5 engines). The engines are used for propulsion during the spacecraft's ascent in addition to two solid rocket boosters and, on the Space Shuttle, the orbiter's Orbital Maneuvering System.
Each engine generates almost 1.8 meganewtons (400,000 lbf) of thrust at liftoff, produces a specific impulse (Isp) of 453 seconds in a vacuum, or 363 seconds at sea level (effective exhaust velocities of 4,440 m/s and 3,560 m/s respectively) and has a mass of approximately 3.5 tonnes (7,700 pounds). The engines are capable of throttling between 65% and 109% of their Rated Power Level in one percent increments,[3] and operate at extreme temperatures, with the liquid hydrogen fuel being stored at −253 °C (−423 °F) while the temperature in the combustion chamber reaches 3,300 °C (5,970 °F), higher than the boiling point of iron. Each engine consumes 1,340 L (350 US gal) of propellant per second.[4]
Contents |
The design of the RS-25 engine consists of various pumps, valves and other components which work in concert with one another in order to produce the thrust required to launch the spacecraft. Fuel and oxidizer from the external tank entered the orbiter at the an umbilical disconnect, and from there flow through the orbiter's main propulsion system (MPS) feed lines. In the Space Launch System, LH2 and LOX from the rocket's core stage will flow directly into the MPS lines. Once in the MPS lines, the fuel and oxidizer each branch out into several parallel paths (three on the Space Shuttle, up to five on the SLS), one to each engine. In each branch, prevalves then allow the propellants to enter the engine..[5]
Once in the engine, fuel and oxidizer flow through low-pressure turbopumps (LPFTP and LPOTP), which then feed the propellants into high-pressure turbopumps (HPFTP and HPOTP). From these HPTPs the propellants take different routes through the engine. Oxidizer is split into four separate paths; the oxidizer heat exchanger (which then split into the fuel tank pressurization and pogo suppression systems), the low pressure oxidiser turbopump (LPOTP), the high pressure oxidizer preburner (from which it is split into the HPFTP turbine and HPOTP before being reunited in the hot gas manifold and sent on to the main combustion chamber), or directly into the main combustion chamber (MCC) injectors. Fuel, meanwhile, flows through the main fuel valve into regenerative cooling systems for the nozzle and MCC or through the chamber coolant valve. Fuel passing through the MCC cooling system then passes through the LPOTP turbine before being split into the tank pressurization system or hot gas manifold cooling system (from where it passes into the MCC). Fuel in the nozzle cooling and chamber coolant valve systems is then sent via preburners into the HPFTP turbine and HPOTP before being reunited again in the hot gas manifold and from there passing into the MCC injectors. Once in the injectors, the propellants are mixed and injected into the main combustion chamber where they are ignited. The burning propellant mixture is then ejected through the throat and bell of the engine's nozzle, the pressure of which creates the thrust required to launch the spacecraft.[5]
The Low Pressure Oxidizer Turbopump (LPOTP) is an axial-flow pump driven by a six-stage turbine powered by liquid oxygen which operates at approximately 5,150 rpm. It boosts the liquid oxygen's pressure from 0.7 to 2.9 MPa (100 to 420 psi), with the flow from the LPOTP then being 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, which measures approximately 450 by 450 mm (18 by 18 in), is connected to the vehicle propellant ducting and supported in a fixed position by the launch vehicle's structure.[5]
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 4,400 psi) while operating at approximately 28,120 rpm with a power of 23,260 hp. 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 pressurize the liquid oxygen tank. Another path enters the HPOTP 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 35 in). It is attached by flanges to the hot-gas manifold.[5]
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 can create a hazard and, to prevent this, the two sections are separated by a cavity that is continuously purged by the engine's 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.[5]
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 in). It is connected to the vehicle propellant ducting and is supported in a fixed position by the launch vehicle's structure 180 degrees from the LPOTP.[5]
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 with a power of 71,140 hp. The discharge flow from the turbopump is routed to and through the main valve and then split 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 liquid hydrogen tank to maintain 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 1,100 mm (22 by 43 in). It is attached by flanges to the hot-gas manifold.[5]
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 was 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.[5]
The speed of the HPOTP and HPFTP turbines depends on the position of the corresponding oxidizer and fuel preburner oxidizer valves. These valves were 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 6.03:1 propellant mixture ratio.[1]
The main oxidizer valve and the main fuel valve control the flow of liquid oxygen and liquid hydrogen into the engine and were controlled by each engine controller. When an engine was operating, the main valves were fully open.[5]
RS-25 thrust (or power level) can be throttled between 67 to 109% of rated thrust. Space Shuttle launches used 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.[6]
| Sea level | Vacuum | |
|---|---|---|
| 100% thrust | 1,670 kN (380,000 lbf) | 2,090 kN (470,000 lbf) |
| 104.5% thrust | 1,750 kN (390,000 lbf) | 2,170 kN (490,000 lbf) |
| 109% thrust | 1,860 kN (420,000 lbf) | 2,280 kN (510,000 lbf) |
Specifying power levels over 100% may seem nonsensical, but there was a logic behind it. The 100% level does not mean the maximum physical power level attainable. Rather it was a specification, decided on early during engine development, for the expected rated power level. Later studies indicated the engine could operate safely at levels above 100%, which was now the norm. Maintaining the original relationship of power level to physical thrust helps reduce confusion. It created 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.[7]
Engine power level affects engine reliability. Studies indicated the probability of an engine failure increased rapidly with power levels over 104.5%, which was why those were retained for contingency use only.[8]
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.[5]
The engine's nozzle is 121 in (3.1 m) long with a diameter of 10.3 in (0.26 m) at its throat and 90.7 in (2.30 m) at its exit.[9] The nozzle takes the shape of a bell-shaped extension bolted to the main combustion chamber, referred to as a de Laval nozzle. To aid the engine's operation at sea level, the RS-25 nozzle has an unusually large expansion ratio (about 77.5:1) for the chamber pressure.[7] 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. Rocketdyne engineers, however, 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 (32 and 39 kPa), and prevents flow separation. The inner part of the flow is at much lower pressure, around 2 psi (14 kPa) or less.[10]
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. On the Space Shuttle, a support ring welded to the forward end of the nozzle was the engine attach point to the orbiter-supplied heat shield. Thermal protection was necessary because of the exposure portions of the nozzles experience during the launch, ascent, on-orbit and entry phases of a mission. The insulation consisted of four layers of metallic batting covered with a metallic foil and screening.[5]
Each engine is equipped with a Main Engine Controller (MEC), an integrated computer which controls all of the engine's functions (through the use of valves) and monitors its performance. Built by Honeywell Aerospace, each MEC originally comprised two redundant Honeywell HDC-601 computers,[11] later being upgraded to a system composed of two double-redundant Motorola 68000 (M68000) processors (for a total of 4 M68000s per controller).[12] Having the controller installed on the engine itself greatly simplifies the wiring between the engine and the launch vehicle, because all the sensors and actuators are connected directly to only the controller, each MEC then being connected to the orbiter's General Purpose Computers (GPCs) or the SLS's avionics suite via its own Engine Interface Unit (EIU).[13] Using a dedicated system also simplifies the software and thus improves its reliability.
Two independent, dual-CPU computers, A and B, formed the controller, giving redundancy to the system. The failure of controller system A would automatically switch to controller system B without impeding operational capabilities; the subsequent failure of controller system B would have provided a graceful shutdown of the engine. Within each system (A and B), the two M68000s operate in "lock-step", thereby enabling each system to detect failures by comparing the TTL signal levels of the buses of the two M68000 processors within that system. If differences are encountered between the two buses, then an interrupt was generated and control turned over to the other system. Because of subtle differences between M68000s from Motorola and the second source manufacturer TRW, each system uses M68000s from the same manufacturer (for instance system A would have two Motorola CPUs while system B would have two CPUs manufactured by TRW). Memory for both Block I and Block II Main Engine Controllers were of the "plated-wire" type. Plated-wire memory functions in a manner similar to magnetic core memory and retains data even after power was turned off.[14]
The controllers were designed from the beginning to be tough enough to survive the forces of launch and proved to be extremely resilient to damage when, during the investigation of the Challenger accident, the two MECs (from engines 2020 and 2021) that were recovered from the seafloor were delivered to Honeywell Aerospace for examination and analysis. One controller was broken open on one side, and both were severely corroded and damaged by marine life. Both units were disassembled and the memory units flushed with deionized water. After they were dried and vacuum baked, data from these units was retrieved for forensic examination.[15]
To control the engine's function, the MEC operates five hydraulically actuated propellant valves on each engine; the oxidizer preburner oxidizer, fuel preburner oxidizer, main oxidizer, main fuel, and chamber coolant valves. In an emergency, the valves can be fully closed by using the engine's helium supply system as a backup actuation system.[5]
In the Space Shuttle, the main oxidizer and fuel bleed valves were used after shutdown to dump any residual propellant, with residual liquid oxygen venting overboard through the engine and residual liquid hydrogen venting through the liquid hydrogen fill and drain valves. After the dump was completed, the valves closed and remain closed for the remainder of the mission.[5]
A coolant control valve was mounted on the combustion chamber coolant bypass duct of each engine. The engine controller regulated the amount of gaseous hydrogen allowed to bypass the nozzle coolant loop, thus controlling its temperature. The chamber coolant valve was 100% open before engine start. During engine operation, it was 100% open for throttle settings of 100 to 109% for maximum cooling. For throttle settings between 65 to 100%, its position ranged from 66.4 to 100% open for reduced cooling.[5]
| External videos | |
|---|---|
| SSME gimbal test | |
Each engine is installed with a gimbal bearing, a universal ball and socket joint which is bolted to the launch vehicle by its upper flange and to the engine by its lower flange. It represents the thrust interface between the engine and the launch vehicle, supporting 7,480 lb (3,390 kg) of engine weight and withstanding over 500,000 lb (230,000 kg) of thrust. As well as providing a means to attach the engine to the launch vehicle, the gimbal bearing allows the engine to be pivoted (or 'gimballed') around two axes of freedom with a range of ±12.5°. This motion allows the engine's thrust vector to be altered, thus steering the vehicle into the correct orientation. The bearing assembly is approximately 290 by 360 mm (11 by 14 in), has a mass of 105 lb (48 kg), and is made of titanium alloy.[6]
The low-pressure oxygen and low-pressure fuel turbopumps were 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, and also to prevent damage to the pumps when loads were applied to them. The liquid hydrogen line from the LPFTP to the HPFTP was insulated to prevent the formation of liquid air.[5]
In addition to fuel and oxidizer systems, the launch vehicle's Main Propulsion System is also equipped with a helium system consisting of ten storage tanks in addition to various regulators, check valves, distribution lines, and control valves. The system is used in-flight to purge the engine, and it provides pressure for actuating engine valves within the propellant management system and during emergency shutdowns. During entry, on the Space Shuttle, any remaining helium was used to purge the engines during reentry and for repressurization.[5]
The history of the RS-25 traces back to the 1960s, when NASA's Marshall Space Flight Center and Rocketdyne were conducting a series of studies on high-pressure engines developed from the successful J-2 engine used on the Saturn V upper stages during the Apollo program. The studies were conducted under a program to upgrade the Saturn V engines which produced a design for a 350,000 lbf upper-stage engine known as the HG-3.[16] As funding levels for Apollo wound down, however, the HG-3 was cancelled alongside the replacement for the F-1 engine, the M-1 (the development for which ended in 1968).[17] It was the design for the HG-3 that would form the basis for the RS-25.[18]
Development of the RS-25 itself began in 1970 when NASA released a request for proposals to design a main engine for the Space Shuttle. The request was based on the then-current design of the Space Shuttle which featured two reusable stages, the orbiter and a manned fly-back booster, and required one engine which would be able to power both vehicles via two different nozzles (12 booster engines with 550,000 lbf sea level thrust each and 3 orbiter engines with 632,000 lbf vacuum thrust each). By the time the contract was awarded, however, budgetary pressures meant that the shuttle's design had changed to its final orbiter, external tank and two boosters configuration and so the engine was only required to power the orbiter during ascent. The request also specified that the engine had to have a conventional bell-type nozzle (rather than an aerospike), a performance level which required a "high chamber pressure staged combustion cycle" to "force an advancement of rocket engine technology". Three companies (Rocketdyne, Pratt & Whitney and Aerojet General) submitted bids in April 1971 to develop the engine (following a year of preliminary studies), with Rocketdyne being awarded the contract on 13 July 1971, although work did not begin on engine development until 31 March 1972 due to a legal challenge from Pratt & Whitney.[7][19]
Following the awarding of the contract, a Preliminary Design Review was carried out in September 1972, followed by a Critical Design Review in September 1976 after which point the engine's design was baselined and construction of the first set of flightworthy engines began. Final review of all the Space Shuttle's components, including the engines, was conducted in 1979. The design reviews operated in parallel with several test milestones, initial tests consisting of individual engine components which identified shortcomings with various areas of the design, including the HPFTP, HPOTP, valves, nozzle and fuel preburners. The individual component tests were followed by the first test of a complete engine (0002) on 16 March 1977. NASA specified that, prior to the Shuttle's first flight, the engines must have undergone at least 65,000 seconds of testing, a milestone that was reached on 23 March 1980, with the engine having undergone 110,253 seconds of testing by the time of STS-1 both on test stands at Stennis Space Center and installed on the Main Propulsion Test Article (MPTA). The first set of engines (2005, 2006 and 2007) were delivered to Kennedy Space Center in 1979 and installed on Columbia, before being removed in 1980 for further testing and reinstalled on the orbiter. The engines, which were of the FMOF configuration and certified for operation at 100% Rated Power Level (RPL), were operated in a 20-second Flight Readiness Firing on 20 February 1981 and, after inspection, declared ready for flight.[7]
Each Space Shuttle was propelled by three SSMEs and, after each flight, the engines would be removed from the Space Shuttle orbiter, inspected and refurbished in the Space Shuttle Main Engine Processing Facility (SSMEPF) in preparation for reuse on a subsequent flight.[20] A total of 46 reusable RS-25D engines were flown as part of the Space Shuttle programme.[21]
During the course of the Space Shuttle program, a total of 46 RS-25 engines were used (with one extra RS-25D being built but never used), flying aboard all 135 missions for a total of 405 individual 'engine missions'. Pratt & Whitney Rocketdyne reports a 99.95% reliability rate, with the only in-flight SSME failure occurring during Space Shuttle Challenger's STS-51-F mission, when the centre engine shut down at T+5:43, leading to an Abort To Orbit.[1]
The engine went through a series of upgrades during the 30 year long Shuttle program, including combustion chamber changes,[22] improved welds and turbopump upgrades.[23] As a result, several versions of the RS-25 were used during the program:[8][24][25]
During the period preceding final Space Shuttle retirement various plans for the several dozen engines were proposed, ranging from them all kept by NASA, to them all being given away (or sold for $400,000—800,000 each) to various institutions such as museums and universities.[26] This policy followed changes to the planned configurations of the Constellation program's Ares V cargo-launch vehicle and Ares I crew-launch vehicle rockets, which had been planned to use the RS-25 in their first and second stages respectively.[27] Whilst these configurations had initially seemed worthwhile as they would use then-current technology following the shuttle's retirement in 2010, the plan had several drawbacks:[27]
Following several design changes to the Ares I and Ares V rockets, the RS-25 was to be replaced with a single J-2X engine for the Ares I second stage and six modified RS-68 engines (which was based on both the SSME and Apollo-era J-2 engine) on the Ares V core stage, meaning the RS-25 would be retired along with the space shuttle fleet.[27] In 2010, however, NASA was directed to halt the Constellation program and, with it, development of both the Ares I and Ares V, instead focusing on building a new heavy lift launcher.[28]
Following the retirement of the Space Shuttle, NASA announced on 14 September 2011 that it would be developing a new launch vehicle, known as the Space Launch System (SLS), to replace the shuttle fleet.[29] The design for the SLS features the RS-25 on its core stage, with four different versions of the rocket - Block 0, I, IA & II - being installed with varying numbers of engines (3, 4, 4-5 & 5 respectively).[30] The initial flights of the new launch vehicle will make use of flown Block II RS-25D engines, with NASA keeping the remaining such engines in a "purged safe" environment, "along with all of the ground systems required to maintain them."[31] The first two launches (SLS-1 and SLS-2) will also make use of the Main Propulsion Systems from Space Shuttles Atlantis and Endeavour, which will be removed from the orbiters as part of their retirement processing.[32] Once the remaining RS-25Ds are used up, they are to be replaced with a cheaper, single-use variant designated the RS-25E ('E' for expendable).[33] Propellants will be supplied from the rocket's core stage, which will consist of a modified Space Shuttle External Tank with the MPS plumbing and engines at its aft and an interstage structure at the top.[33]
| External videos | |
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| STS-49 Flight Readiness Firing | |
| Time-lapse video of STS-135 SSME installation | |
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