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Główne silniki promu kosmicznego (ang. Space Shuttle Main Engines - SSME) to trzy główne silniki orbitera promu kosmicznego. Są one konstruowane przez Rocketdyne, oddział firmy Pratt & Whitney. Silnik oznaczany jest także jako RS-24.
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Wstęp
Główne silniki promu kosmicznego to bardzo wyrafinowane elementy przetwarzania energii, które spalają ciekły wodór i ciekły tlen z zbiornika zewnętrznego promu (ET). Używane są jako napęd w czasie wznoszenia, uzupełniając dwie rakiety dodatkowe na paliwo stałe (SRB) o większej mocy. Czasem SSME używane są także przy manewrach orbitalnych. Każdy z głównych silników generuje prawie 1.8 MN ciągu w momencie wyniesienia (oderwania się promu od ziemi w momencie startu). Silniki są zdolne do wygenerowania impulsu właściwego (Iwł) długości 453 sekund w próźni lub 363 sekund na poziomie morza (prędkości wylotowe odpowiednio 4440 m/s i 3560 m/s). Ogólnie, główny silnik promu kosmicznego waży około 3,2 t. Silniki są wymontowywane po każdym locie i transportowane do ośrodka przetwarzania silników (ang. Space Shuttle Main Engine Processing Facility - SSMEPF) w celu dokonania sprawdzenia i wymian komponentów, które tego wymagają.
Silniki rakietowe promu kosmicznego są zdolne do pracy w ekstremalnych temperaturach. Ciekły wodór jest przechowywany w temperaturze -253 °C. Jednakże kiedy jest on spalany wraz z ciekłym tlenem, temperatura w komorze spalania osiąga do 3 300 °C, więcej niż temperatura wrzenia żelaza.
Poza trzema głównymi silnikami, orbiter posiada 44 mniejsze rakiety na swoim pokładzie, będące częścią systemu manewrowania orbitalnego (OMS) i systemu kontroli reakcji (RCS). Systemy te pozwalają na sterowanie, zmianę pozycji w przestrzeni (obroty) i zmianę wysokości promu w czasie pozostawania na orbicie.
W czasie pracy silnika, paliwo i utleniacz ze zbiornika zewnętrznego pompowane jest na pokład orbitera poprzez złącze, a dalej do głównych linii przesyłowych systemu napędowego orbitera. Tam paliwo i utleniacz rozgałęziają się na trzy równoległe linie, do każdego z silników. Każde rozgałęzienie posiada osobny zawór, zabezpieczający przed swobodnym wpływaniem paliwa lub utleniacza do odpowidniej turbopompy.
System utleniacza
Niskociśnieniowa turbopompa utleniacza (ang. The Low Pressure Oxidizer Turbopump - LPOTP) to pompa śmigłowa napędzana przez sześciostopniową turbinę zasilaną przez ciekły tlen. Zwiększa ona ciśnienie ciekłego tlenu z 0,7 do 2.9 MPa. Przepływ z LPOTP jest kierowany do wysokociśnieniowej turbopompy utleniacza (ang. High-Pressure Oxidizer Turbopump - HPOTP). W czasie działania silnika, zwiększenie ciśnienia pozwala pompie wysokociśnieniowej na pracę z dużą prędkością bez kawitacji. Pompa niskociśnieniowa działa z prędkością około 5 150 obrotów na minutę. Pompa ta, o wymiarach ok. 450 na 450 mm, jest połączona z instalacją przesyłu paliwa i miejsce jej zamocowania jest dodatkowo chronione przez strukturę orbitera.
Pompa wysokociśnieniowa składa się z dwóch jednostopniowych pomp odśrodkowych (pompa główna i pompa przedspalania) zamocowanych na wspólnym wale i zasilanych przez dwustopniową turbinę gazową.
Główna pompa podwyższa ciśnienie ciekłego tlenu z 2,9 do 30 MPa pracując przy średnio 28 120 rpm (obrotach na minutę). Strumień wychodzący z HPOTP jest dzielony na kilka ścieżek, z których jedna jest kierowana do zasilania turbiny LPOTP. Inna ścieżka kierowana jest poprzez główny zawór utleniacza do komory spalania. Kolejny mały strumień jest redukowany i wprowadzany do wymiennika ciepła utleniacza. Ciekły tlen wpływa przez zawór przeciwzalewowy, który blokuje dostęp do wymiennika ciepła, jeśli nie jest on wystarczająco ciepły aby zamienić tlen w gaz. Wymiennik ciepła wykorzystuje ciepło zawarte w gazach wylotowych turbiny HPOTP. Gaz jest kierowany do kolektora a później do zbiornika zewnętrznego, aby utrzymać ciśnienie w zbiorniku ciekłego tlenu. Kolejny strumień wpływa do pompy przedspalającej drugiego stopnia HPOT, przez co ciśnienie ciekłego tlenu zwiększane jest z 30 do 51 MPa. Przepływa przez komorę wstępnego spalania utleniacza i dalej do komory wstępnego spalania paliwa.
HPOTP ma wymiary około 60×90 cm. Jest zamocowana przez kołnierze do kolektora gorących gazów. Turbina HPOTP i pompy HPOTP są zamontowane na wspólnym wale.
Mieszanie bogatego w paliwo gorącego gazu w sekcji turbinowej i ciekłego tlenu w głównej pompie mogło stwarzać zagrożenie eksplozją. Aby temu zapobiec, te dwie sekcje są odseparowane wnęką, która w czasie pracy silnika jest stale przeczyszczana przez wodór z zasobnika głównego systemu napędowego. Dwie plomby minimalizują wycieki do wnęki. Jedna plomba mieści się między sekcją turbiny a wnęką, druga między sekcją pompy a wnęką. Utrata ciśnienia wodoru we wnęce powoduje automatyczne wyłączenie silnika.
Hydrogen fuel 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
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.
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.[1]
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 no 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.[1]
The SSME after the Shuttle era
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.
Specifications
- 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 at 100% power
- 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
See also
- MPTA-098 - the SSME test article used in Shuttle development
- Zależności energii i mocy w SSME
Źródła
- NASA Shuttle Press Kit SSME Reference.
- Space Shuttle Main Engine.
- Space Shuttle Main Engine Enhancements.
- The Roar of Innovation.
- Space Shuttle Main Engine - incredible facts.
- Space Shuttle Main Engine The First Ten Years.
- NSTS 1988 News Reference Manual.
- "Boeing Liquid Propellant Rocket Systems", Rocketdyne Propulsion & Power, Pub. 573-A-100 9/99, page 26. (SSME was designated the RS-24, not RS-25 as commonly believed. The RS-25 was a misnomer for the expendable version of the SSME that adopted early in the Ares program, later replaced by the RS-68 and J-2X.)
- Encyclopedia Astronautix, reference SSME / RS-24.
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