Staged combustion cycle

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Fuel-rich staged combustion cycle. Here, all of the fuel and a portion of the oxidizer are fed through the preburner, generating fuel-rich gas. After being run through a turbine to power the pumps, the gas is injected into the combustion chamber and burned with the remaining oxidizer. Staged combustion rocket cycle.svg
Fuel-rich staged combustion cycle. Here, all of the fuel and a portion of the oxidizer are fed through the preburner, generating fuel-rich gas. After being run through a turbine to power the pumps, the gas is injected into the combustion chamber and burned with the remaining oxidizer.

The staged combustion cycle (sometimes known as topping cycle, preburner cycle, or closed cycle) is a power cycle of a bipropellant rocket engine. In the staged combustion cycle, propellant flows through multiple combustion chambers, and is thus combusted in stages. The main advantage relative to other rocket engine power cycles is high fuel efficiency, measured through specific impulse, while its main disadvantage is engineering complexity.

Contents

Typically, propellant flows through two kinds of combustion chambers; the first called preburner and the second called main combustion chamber. In the preburner, a small portion of propellant, usually fuel-rich, is partly combusted under non-stoichiometric conditions, increasing the volume of flow driving the turbopumps that feed the engine with propellant. The gas is then injected into the main combustion chamber and combusted completely with the other propellant to produce thrust.

Tradeoffs

The main advantage is fuel efficiency due to all of the propellant flowing to the main combustion chamber, which also allows for higher thrust. The staged combustion cycle is sometimes referred to as closed cycle, as opposed to the gas generator, or open cycle where a portion of propellant never reaches the main combustion chamber. The disadvantage is engineering complexity, partly a result of the preburner exhaust of hot and highly pressurized gas which, particularly when oxidizer-rich, produces extremely harsh conditions for turbines and plumbing.

History

Staged combustion (Замкнутая схема) was first proposed by Alexey Isaev in 1949. The first staged combustion engine was the S1.5400 (11D33) used in the Soviet Molniya rocket, designed by Melnikov, a former assistant to Isaev. [1] About the same time (1959), Nikolai Kuznetsov began work on the closed cycle engine NK-9 for Korolev's orbital ICBM, GR-1. Kuznetsov later evolved that design into the NK-15 and NK-33 engines for the unsuccessful Lunar N1 rocket. The non-cryogenic N2O4/UDMH engine RD-253 using staged combustion was developed by Valentin Glushko circa 1963 for the Proton rocket.

After the abandonment of the N1, Kuznetsov was ordered to destroy the NK-33 technology, but instead he warehoused dozens of the engines. In the 1990s, Aerojet was contacted and eventually visited Kuznetsov's plant. Upon meeting initial skepticism about the high specific impulse and other specifications, Kuznetsov shipped an engine to the US for testing. Oxidizer-rich staged combustion had been considered by American engineers, but was not considered a feasible direction because of resources they assumed the design would require to make work. [2] The Russian RD-180 engine also employs a staged-combustion rocket engine cycle. Lockheed Martin began purchasing the RD-180 in circa 2000 for the Atlas III and later, the V, rockets. The purchase contract was subsequently taken over by United Launch Alliance (ULA--the Boeing/Lockheed-Martin joint venture) after 2006, and ULA continues to fly the Atlas V with RD-180 engines as of 2022.

The first laboratory staged-combustion test engine in the West was built in Germany in 1963, by Ludwig Boelkow. [3]

Hydrogen peroxide/kerosene powered engines may use a closed-cycle process by catalytically decomposing the peroxide to drive turbines before combustion with the kerosene in the combustion chamber proper. This gives the efficiency advantages of staged combustion, while avoiding major engineering problems.

The RS-25 Space Shuttle main engine is another example of a staged combustion engine, and the first to use liquid oxygen and liquid hydrogen. [4] Its counterpart in the Soviet shuttle was the RD-0120, which had similar specific impulse, thrust, and chamber pressure, but with some differences that reduced complexity and cost at the expense of increased engine weight.

Variants

Oxidizer-rich turbine exhaust from a SpaceX Raptor preburner shown during a 2015 sub-system test on a test stand at Stennis Space Center. In the full-flow rocket engine, the preburner exhaust is fed into a turbine and then into the main combustion chamber. SpaceX's Raptor oxygen preburner testing at Stennis (2015).jpg
Oxidizer-rich turbine exhaust from a SpaceX Raptor preburner shown during a 2015 sub-system test on a test stand at Stennis Space Center. In the full-flow rocket engine, the preburner exhaust is fed into a turbine and then into the main combustion chamber.

Several variants of the staged combustion cycle exist. Preburners that burn a small portion of oxidizer with a full flow of fuel are called fuel-rich, while preburners that burn a small portion of fuel with a full flow of oxidizer are called oxidizer-rich. The RD-180 has an oxidizer-rich preburner, while the RS-25 has two fuel-rich preburners. The SpaceX Raptor has both oxidizer-rich and fuel-rich preburners, a design called full-flow staged combustion.

Staged combustion designs can be either single-shaft or twin-shaft. In the single-shaft design, one set of preburner and turbine drives both propellant turbopumps. Examples include the Energomash RD-180 and the Blue Origin BE-4. In the twin-shaft design, the two propellant turbopumps are driven by separate turbines, which are in turn driven by the outflow of either one or separate preburners. Examples of twin-shaft designs include the Rocketdyne RS-25, the JAXA LE-7, and Raptor. Relative to a single-shaft design, the twin-shaft design requires an additional turbine (and possibly another preburner), but allows for individual control of the two turbopumps. Hydrolox engines are typically twin-shaft designs due to greatly differing propellant densities.

In addition to the propellant turbopumps, staged combustion engines often require smaller boost pumps to prevent both preburner backflow and turbopump cavitation. For example, the RD-180 and RS-25 use boost pumps driven by tap-off and expander cycles, as well as pressurized tanks, to incrementally increase propellant pressure prior to entering the preburner.

Full-flow staged combustion cycle

Full-flow staged combustion rocket cycle Full flow staged rocket cycle.png
Full-flow staged combustion rocket cycle

Full-flow staged combustion (FFSC) is a twin-shaft staged combustion fuel cycle design that uses both oxidizer-rich and fuel-rich preburners where the entire supply of both propellants passes through the turbines. [5] The fuel turbopump is driven by the fuel-rich preburner, and the oxidizer turbopump is driven by the oxidizer-rich preburner. [6] [5]

Benefits of the full-flow staged combustion cycle include turbines that run cooler and at lower pressure, due to increased mass flow, leading to a longer engine life and higher reliability. As an example, up to 25 flights were anticipated for an engine design studied by the DLR (German Aerospace Center) in the frame of the SpaceLiner project, [5] up to 1000 flights are expected for Raptor from SpaceX. [7] Further, the full-flow cycle eliminates the need for an interpropellant turbine seal normally required to separate oxidizer-rich gas from the fuel turbopump or fuel-rich gas from the oxidizer turbopump, [8] thus improving reliability.

Since the use of both fuel and oxidizer preburners results in full gasification of each propellant before entering the combustion chamber, FFSC engines belong to a broader class of rocket engines called gas-gas engines. [8] Full gasification of components leads to faster chemical reactions in the combustion chamber, allowing a smaller combustion chamber. This in turn makes it feasible to increase the chamber pressure, which increases efficiency.

Potential disadvantages of the full-flow staged combustion cycle include more stringent materials requirements, and the increased engineering complexity and parts count of the two preburners, relative to a single-shaft staged combustion cycle.

As of 2024, four full-flow staged combustion rocket engines have been tested on test stands; the Soviet storable propellant RD-270 project at Energomash in the 1960s, the US government-funded hydrolox Integrated Powerhead Demonstrator project at Aerojet Rocketdyne in the mid-2000s, [8] SpaceX's flight capable methalox Raptor engine first test-fired in February 2019, [9] and the methalox engine developed for the first stage of the Stoke Space Nova vehicle in 2024. [10]

The first flight test of a full-flow staged-combustion engine occurred on 25 July 2019 when SpaceX flew their Raptor methalox FFSC engine on the Starhopper test rocket, at their South Texas Launch Site. [11] As of 2024, the Raptor is the only FFSC engine that has flown on a launch vehicle.

Applications

Oxidizer-rich staged combustion

Fuel-rich staged combustion

Full-flow staged combustion

SpaceX Raptor FFSC rocket engine, sample propellant flow schematic, 2019 Raptor Engine Unofficial Combustion Scheme.svg
SpaceX Raptor FFSC rocket engine, sample propellant flow schematic, 2019

Past and present applications of staged-combustion engines

Future applications of staged-combustion engines

See also

Related Research Articles

A tripropellant rocket is a rocket that uses three propellants, as opposed to the more common bipropellant rocket or monopropellant rocket designs, which use two or one propellants, respectively. Tripropellant systems can be designed to have high specific impulse and have been investigated for single-stage-to-orbit designs. While tripropellant engines have been tested by Rocketdyne and NPO Energomash, no tripropellant rocket has been flown.

<span class="mw-page-title-main">Expander cycle</span> Rocket engine operation method

The expander cycle is a power cycle of a bipropellant rocket engine. In this cycle, the fuel is used to cool the engine's combustion chamber, picking up heat and changing phase. The now heated and gaseous fuel then powers the turbine that drives the engine's fuel and oxidizer pumps before being injected into the combustion chamber and burned.

<span class="mw-page-title-main">Liquid-propellant rocket</span> Rocket engine that uses liquid fuels and oxidizers

A liquid-propellant rocket or liquid rocket uses a rocket engine burning liquid propellants. (Alternate approaches use gaseous or solid propellants.) Liquids are desirable propellants because they have reasonably high density and their combustion products have high specific impulse (Isp). This allows the volume of the propellant tanks to be relatively low.

<span class="mw-page-title-main">Rocketdyne J-2</span> Rocket engine

The J-2, commonly known as Rocketdyne J-2, was a liquid-fuel cryogenic rocket engine used on NASA's Saturn IB and Saturn V launch vehicles. Built in the United States by Rocketdyne, the J-2 burned cryogenic liquid hydrogen (LH2) and liquid oxygen (LOX) propellants, with each engine producing 1,033.1 kN (232,250 lbf) of thrust in vacuum. The engine's preliminary design dates back to recommendations of the 1959 Silverstein Committee. Rocketdyne won approval to develop the J-2 in June 1960 and the first flight, AS-201, occurred on 26 February 1966. The J-2 underwent several minor upgrades over its operational history to improve the engine's performance, with two major upgrade programs, the de Laval nozzle-type J-2S and aerospike-type J-2T, which were cancelled after the conclusion of the Apollo program.

<span class="mw-page-title-main">Gas-generator cycle</span> Rocket engine operation method

The gas-generator cycle, also called open cycle, is one of the most commonly used power cycles in bipropellant liquid rocket engines.

<span class="mw-page-title-main">NK-33</span> Soviet rocket engine

The NK-33 and its vacuum-optimized variant, the NK-43, are rocket engines developed in the late 1960s and early 1970s by the Kuznetsov Design Bureau for the Soviet space program's ill-fated N1 Moon rocket. The NK-33 is among the most powerful LOX/RP-1 powered rocket engines ever built, noted for its high specific impulse and low structural mass.

<span class="mw-page-title-main">RD-8</span> Soviet rocket engine

The RD-8 is a Soviet / Ukrainian liquid propellant rocket engine burning LOX and RG-1 in an oxidizer rich staged combustion cycle. It has a four combustion chambers that provide thrust vector control by gimbaling each of the nozzles in a single axis ±33°. It was designed in Dnipropetrovsk by the Yuzhnoye Design Bureau as the vernier thruster of the Zenit second stage. As such, it has always been paired with the RD-120 engine for main propulsion.

<span class="mw-page-title-main">Aerojet M-1</span> One of the largest rocket engines to be designed

The Aerojet M-1 was one of the largest and most powerful liquid-hydrogen-fueled liquid-fuel rocket engines to be designed and component-tested. It was originally developed during the 1950s by the US Air Force. The M-1 offered a baseline thrust of 6.67 MN and an immediate growth target of 8 MN. If built, the M-1 would have been larger and more efficient than the famed F-1 that powered the first stage of the Saturn V rocket to the Moon.

RD-270 was a single-chamber liquid-fuel rocket engine designed by Energomash (USSR) in 1960–1970. It was to be used on the first stages of proposed heavy-lift UR-700 and UR-900 rocket families, as well as on the N1. It has the highest thrust among single-chamber engines of the USSR, 640 metric tons at the surface of Earth. The propellants used are a hypergolic mixture of unsymmetrical dimethylhydrazine (UDMH) fuel with dinitrogen tetroxide oxidizer. The chamber pressure was among the highest considered, being about 26 MPa. This was achieved by applying full-flow staged combustion cycle for all the incoming mass of fuel, which is turned into a gas and passes through multiple turbines before being burned in the combustion chamber. This allowed the engine to achieve a specific impulse of 301 s (2.95 km/s) at the Earth's surface.

<span class="mw-page-title-main">YF-100</span> Chinese rocket engine

The YF-100 is a Chinese liquid rocket engine burning LOX and kerosene in an oxidizer-rich staged combustion cycle.

<span class="mw-page-title-main">Chemical Automatics Design Bureau</span> Russian rocket engine manufacturer

Chemical Automatics Design Bureau (CADB), also KB Khimavtomatika, is a Russian design bureau founded by the NKAP in 1941 and led by Semyon Kosberg until his death in 1965. Its origin dates back to a 1940 Moscow carburetor factory, evacuated to Berdsk in 1941, and then relocated to Voronezh city in 1945, where it now operates. Originally designated OKB-296 and tasked to develop fuel equipment for aviation engines, it was redesignated OKB-154 in 1946.

<span class="mw-page-title-main">RD-0120</span> Soviet rocket engine

The Soviet RD-0120 (also designated 11D122) was the Energia core rocket engine, fueled by LH2/LOX, roughly equivalent to the Space Shuttle Main Engine (SSME). These were attached to the Energia core rather than the orbiter, so were not recoverable after a flight, but created a more modular design (the Energia core could be used for a variety of missions besides launching the shuttle). The RD-0120 and the SSME have both similarities and differences. The RD-0120 achieved a slightly higher specific impulse and combustion chamber pressure with reduced complexity and cost (but it was single-use), as compared to the SSME. It uses a fuel-rich staged combustion cycle and a single shaft to drive both the fuel and oxidizer turbopumps. Some of the Russian design features, such as the simpler and cheaper channel wall nozzles, were evaluated by Rocketdyne for possible upgrades to the SSME. It achieved combustion stability without the acoustic resonance chambers that the SSME required.

The RL60 was a planned liquid-fuel cryogenic rocket engine designed in the United States by Pratt & Whitney, burning cryogenic liquid hydrogen and liquid oxygen propellants. The engine runs on an expander cycle, running the turbopumps with waste heat absorbed from the main combustion process. This high-efficiency, waste heat based combustion cycle combined with the high-performance liquid hydrogen fuel enables the engine to reach a very high specific impulse of up to 465 seconds in a vacuum. The engine was planned to be a more capable successor to the Aerojet Rocketdyne RL10, providing improved performance and efficiency while maintaining the installation envelope of the RL10.

<span class="mw-page-title-main">Cryogenic rocket engine</span> Type of rocket engine which uses liquid fuel stored at very low temperatures

A cryogenic rocket engine is a rocket engine that uses a cryogenic fuel and oxidizer; that is, both its fuel and oxidizer are gases which have been liquefied and are stored at very low temperatures. These highly efficient engines were first flown on the US Atlas-Centaur and were one of the main factors of NASA's success in reaching the Moon by the Saturn V rocket.

The RD-701 is a liquid-fuel rocket engine developed by Energomash, Russia. It was briefly proposed to propel the reusable MAKS space plane, but the project was cancelled shortly before the end of USSR. The RD-701 is a tripropellant engine that uses a staged combustion cycle with afterburning of oxidizer-rich hot turbine gas. The RD-701 has two modes. Mode 1 uses three components: LOX as an oxidizer and a fuel mixture of RP-1 / LH2 which is used in the lower atmosphere. Mode 2 also uses LOX, with LH2 as fuel in vacuum where atmospheric influence is negligible.

<span class="mw-page-title-main">Aerojet LR87</span> American rocket engine family used on Titan missile first stages

The LR87 was an American liquid-propellant rocket engine used on the first stages of Titan intercontinental ballistic missiles and launch vehicles. Composed of twin motors with separate combustion chambers and turbopump machinery, it is considered a single unit and was never flown as a single combustion chamber engine or designed for this. The LR87 first flew in 1959.

<span class="mw-page-title-main">Rocket propellant</span> Chemical or mixture used in a rocket engine

Rocket propellant is used as reaction mass ejected from a rocket engine to produce thrust. The energy required can either come from the propellants themselves, as with a chemical rocket, or from an external source, as with ion engines.

<span class="mw-page-title-main">SpaceX rocket engines</span> Rocket engines developed by SpaceX

Since the founding of SpaceX in 2002, the company has developed four families of rocket engines — Merlin, Kestrel, Draco and SuperDraco — and since 2016 developed the Raptor methane rocket engine and after 2020, a line of methalox thrusters.

<span class="mw-page-title-main">RD-0110</span> Soviet (later Russian) rocket engine

The RD-0110 is a rocket engine burning liquid oxygen and kerosene in a gas generator combustion cycle. It has four fixed nozzles and the output of the gas generator is directed to four secondary vernier nozzles to provide attitude control for the stage. It has an extensive flight history with its initial versions having flown more than 64 years ago.

<span class="mw-page-title-main">SpaceX Raptor</span> SpaceX family of liquid-fuel rocket engines

Raptor is a family of rocket engines developed and manufactured by SpaceX. It is the third rocket engine in history designed with a full-flow staged combustion (FFSC) fuel cycle, and the first such engine to power a vehicle in flight. The engine is powered by cryogenic liquid methane and liquid oxygen, a mixture known as methalox.

References

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