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, and the increasing volume flow is used to drive 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.[ citation needed ]

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.[ citation needed ] 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 the greatly different densities of the propellants.

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 cycle that uses both oxidizer-rich and fuel-rich preburners. The cycle allows full flow of both propellants through the turbines; hence the name. [3] The fuel turbopump is driven by the fuel-rich preburner, and the oxidizer turbopump is driven by the oxidizer-rich preburner. [4] [3]

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, [3] up to 1000 flights are expected for Raptor from SpaceX. [5] 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, [6] 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. [6] 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 increased engineering complexity of two preburners, relative to a single-shaft staged combustion cycle, as well as an increased parts count.

As of 2019, only three full-flow staged combustion rocket engines had ever progressed sufficiently to be tested on test stands; the Soviet Energomash RD-270 project in the 1960s, the US government-funded Aerojet Rocketdyne Integrated powerhead demonstration project in the mid-2000s, [6] and SpaceX's flight capable Raptor engine first test-fired in February 2019. [7]

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. [8] As of August 22, 2023, the Raptor is the only FFSC engine to have 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

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<span class="mw-page-title-main">RP-1</span> Highly refined form of kerosene used as rocket fuel

RP-1 (alternatively, Rocket Propellant-1 or Refined Petroleum-1) is a highly refined form of kerosene outwardly similar to jet fuel, used as rocket fuel. RP-1 provides a lower specific impulse than liquid hydrogen (H2), but is cheaper, is stable at room temperature, and presents a lower explosion hazard. RP-1 is far denser than H2, giving it a higher energy density (though its specific energy is lower). RP-1 also has a fraction of the toxicity and carcinogenic hazards of hydrazine, another room-temperature liquid fuel.

<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 utilizes 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">NK-33</span> Soviet rocket engine

The NK-33 and NK-43 are rocket engines designed and built in the late 1960s and early 1970s by the Kuznetsov Design Bureau. The NK designation is derived from the initials of chief designer Nikolay Kuznetsov. The NK-33 was among the most powerful LOX/RP-1 rocket engines when it was built, with a high specific impulse and low structural mass. They were intended for the ill-fated Soviet N1F Moon rocket, which was an upgraded version of the N1. The NK-33A rocket engine is now used on the first stage of the Soyuz-2-1v launch vehicle. When the supply of the NK-33 engines are exhausted, Russia will supply the new RD-193 rocket engine. It used to be the first stage engines of the Antares 100 rocket series, although those engines are rebranded the AJ-26 and the newer Antares 200 and Antares 200+ rocket series uses the RD-181 for the first stage engines, which is a modified RD-191, but shares some properties like a single combustion chamber unlike the two combustion chambers used in the RD-180 of the Atlas V and the four combustion chambers used in the RD-170 of the Energia and Zenit rocket families, and the RD-107, RD-108, RD-117, and RD-118 rocket engines used on all of the variants of the Soyuz rocket.

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

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RD-270 (Russian: Раке́тный дви́гатель 270, Rocket Engine 270, 8D420) was a single-chamber liquid-bipropellant 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 unsymmetrical dimethylhydrazine (UDMH) and nitrogen tetroxide (N2O4). 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.

The YF-75 is a liquid cryogenic rocket engine burning liquid hydrogen and liquid oxygen in a gas generator cycle. It is China's second generation of cryogenic propellant engine, after the YF-73, which it replaced. It is used in a dual engine mount in the H-18 third stage of the Long March 3A, Long March 3B and Long March 3C launch vehicles. Within the mount, each engine can gimbal individually to enable thrust vectoring control. The engine also heats hydrogen and helium to pressurize the stage tanks and can control the mixture ratio to optimize propellant consumption.

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

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

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

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

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

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

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References

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