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28-607: The original LE-5 was built as a second stage engine for the H-I launch vehicle. It used a fairly conventional gas generator cycle . The LE-5A was a heavily redesigned version of the LE-5 intended for use on the new H-II launch vehicle's second stage. The major difference is that the operation of the engine was switched from the gas generator to expander bleed cycle . The LE-5A was the world's first expander bleed cycle engine to be put into operational service. Cryogenic liquid hydrogen fuel for

56-414: 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 . Typically, propellant flows through two kinds of combustion chambers;

84-423: 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. Its counterpart in

112-499: 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

140-505: 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, up to 1000 flights are expected for Raptor from SpaceX. Further, the full-flow cycle eliminates the need for an interpropellant turbine seal normally required to separate oxidizer-rich gas from

168-466: 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;

196-415: Is currently ongoing. Gas-generator cycle (rocket) The gas-generator cycle , also called open cycle , is one of the most commonly used power cycles in bipropellant liquid rocket engines. Propellant is burned in a gas generator (or "preburner") and the resulting hot gas is used to power the propellant pumps before being exhausted overboard and lost. Because of this loss, this type of engine

224-583: 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. Staged combustion ( Замкнутая схема )

252-411: Is termed open cycle . The gas generator cycle exhaust products pass over the turbine first. Then they are expelled overboard. They can be expelled directly from the turbine, or are sometimes expelled into the nozzle (downstream from the throat) for a small gain in efficiency. The main combustion chamber does not use these products. This explains the name of the open cycle. The major disadvantage

280-399: Is that this propellant contributes little to no thrust because they are not injected into the combustion chamber. The major advantage of the cycle is reduced engineering complexity compared to the staged combustion (closed) cycle . Staged combustion cycle The staged combustion cycle (sometimes known as topping cycle , preburner cycle , or closed cycle ) is a power cycle of

308-734: 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

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336-631: 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, SpaceX's flight capable methalox Raptor engine first test-fired in February 2019, and the methalox engine developed for the first stage of the Stoke Space Nova vehicle in 2024. The first flight test of

364-476: 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. 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

392-796: 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 . Hydrogen peroxide / kerosene powered engines may use

420-403: The combustion chamber as opposed to both the chamber and the nozzle in the 5A. Alterations to the combustion chamber cooling passages and constituent materials were made with special emphasis on effective heat transfer to allow this method to be successful. After flight F5 of H-IIA on March 28, 2003 resulted in severe (although not damaging) vibration of the upper stage during LE-5B firing, work

448-473: The cycle is drawn through tubes and passages in both the engine's nozzle and combustion chamber where the hydrogen heats up incredibly while simultaneously cooling those components. The heating of the initially cold fuel causes it to expand, and it is utilized to drive the turbine for the propellant pumps. The LE-5B  [ ja ] was a further modified version of the LE-5A. The changes focused on lowering

476-435: The engine controller were to be replaced with modern components that could be reliably sourced for years to come, and the manufacturing method for the combustion chamber was to be likewise updated for similar reasons. The liquid hydrogen turbopump and turbine nozzle were to be updated for H3's longer mission duration times, and the performance of the liquid oxygen turbopump and fuel mixer was to be improved. The first example of

504-452: 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 . The main advantage

532-423: The fuel turbopump or fuel-rich gas from the oxidizer turbopump, 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 . Full gasification of components leads to faster chemical reactions in the combustion chamber, allowing

560-413: 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. The Russian RD-180 engine also employs a staged-combustion rocket engine cycle. Lockheed Martin began purchasing

588-450: The per-unit cost of the engine while continuing to increase reliability. The modifications veered towards simplification and cheaper production where possible at the cost of actually lowering the specific impulse to 447 seconds, the lowest of all three models. However, it produced the highest thrust of the three and was significantly cheaper. The primary change from the 5A model was that the 5B's expander bleed system circulated fuel around only

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616-508: The preburner. 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. The fuel turbopump is driven by the fuel-rich preburner, and the oxidizer turbopump is driven by the oxidizer-rich preburner. Benefits of the full-flow staged combustion cycle include turbines that run cooler and at lower pressure, due to increased mass flow, leading to

644-474: 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

672-651: The unsuccessful Lunar N1 rocket . The non-cryogenic N 2 O 4 / 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

700-566: The updated design was test fired in March, 2017. On the H3 launch vehicles first flight on March 7, 2023, the first stage, consisting of two SRB-3 , and two LE-9 engines, performed nominally up until stage separation. Following separation, ignition of the LE-5B-3 could not be confirmed, and velocity started dropping significantly. At L+ 00:14:50, a self-destruct command was sent to H3. An investigation

728-456: The upper stage by half. For the new H3 launch vehicle, the veteran design of the LE-5B was once again revisited. To meet the requirements of the H3 and to ensure a stable supply of parts over H3's lifetime, performance was to be improved and costs were to be lowered, all while keeping development risk as low as possible. Obsolete parts that were becoming hard to acquire such as the electronics in

756-566: 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. 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

784-468: Was initiated on an upgraded version of the LE-5B. The upgraded engine, named LE-5B-2, was first flown on a H-IIB on September 10, 2009. The main fixes were adding flow-laminarizing plates in the expander manifold, a new mixer of gaseous and liquid hydrogen in the hydrogen feed line, and a new injector plate with 306 smaller coaxial injectors (versus 180 in LE-5B). The upgrade reduced the vibrations produced by

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