Rocket engine

Rocket propellant is mass that is stored, usually in some form of propellant tank, or within the combustion chamber itself, prior to being ejected from a rocket engine in the form of a fluid jet to produce thrust.

Chemical rocket propellants are the most commonly used.These undergo exothermic chemical reactions producing hot gas which is used by the rocket for propulsive purposes. Alternatively, a chemically inert reaction mass can be heated using a high-energy power source via a heat exchanger, and then no combustion chamber is used.

Solid rocket propellants are prepared as a mixture of fuel and oxidising components called 'grain' and the propellant storage casing effectively becomes the combustion chamber.

Liquid-fuelled rockets force separate fuel and oxidiser components into the combustion chamber, where they mix and burn. Hybrid rocket engines use a combination of solid and liquid or gaseous propellants. Both liquid and hybrid rockets use injectors to introduce the propellant into the chamber. These are often an array of simple jets – holes through which the propellant escapes under pressure; but sometimes may be more complex spray nozzles. When two or more propellants are injected, the jets usually deliberately cause the propellants to collide as this breaks up the flow into smaller droplets that burn more easily.

For chemical rockets the combustion chamber is typically cylindrical, and flame holders, used to hold a part of the combustion in a slower-flowing portion of the combustion chamber, are not needed. The dimensions of the cylinder are such that the propellant is able to combust thoroughly; different rocket propellants require different combustion chamber sizes for this to occur.

This leads to a number called ${\displaystyle L^{*}}$:

${\displaystyle L^{*}={\frac {V_{c}}{A_{t}}}}$

where:

• ${\displaystyle V_{c}}$ is the volume of the chamber
• ${\displaystyle A_{t}}$ is the area of the throat of the nozzle.

L* is typically in the range of 25–60 inches (0.64–1.52 m).

The combination of temperatures and pressures typically reached in a combustion chamber is usually extreme by any standard. Unlike in airbreathing jet engines, no atmospheric nitrogen is present to dilute and cool the combustion, and the propellant mixture can reach true stoichiometric ratios. This, in combination with the high pressures, means that the rate of heat conduction through the walls is very high.

In order for fuel and oxidizer to flow into the chamber, the pressure of the propellant fluids entering the combustion chamber must exceed the pressure inside the combustion chamber itself. This may be accomplished by a variety of design approaches including turbopumps or, in simpler engines, via sufficient tank pressure to advance fluid flow. Tank pressure may be maintained by several means, including a high-pressure helium pressurization system common to many large rocket engines or, in some newer rocket systems, by a bleed-off of high-pressure gas from the engine cycle to autogenously pressurize the propellant tanks[2][3] For example, the self-pressurization gas system of the SpaceX Starship is a critical part of SpaceX strategy to reduce launch vehicle fluids from five in their legacy Falcon 9 vehicle family to just two in Starship, eliminating not only the helium tank pressurant but all hypergolic propellants as well as nitrogen for cold-gas reaction-control thrusters.[4]

Rocket thrust is caused by pressures acting in the combustion chamber and nozzle. From Newton's third law, equal and opposite pressures act on the exhaust, and this accelerates it to high speeds.

The hot gas produced in the combustion chamber is permitted to escape through an opening (the "throat"), and then through a diverging expansion section. When sufficient pressure is provided to the nozzle (about 2.5–3 times ambient pressure), the nozzle chokes and a supersonic jet is formed, dramatically accelerating the gas, converting most of the thermal energy into kinetic energy. Exhaust speeds vary, depending on the expansion ratio the nozzle is designed for, but exhaust speeds as high as ten times the speed of sound in air at sea level are not uncommon. About half of the rocket engine's thrust comes from the unbalanced pressures inside the combustion chamber, and the rest comes from the pressures acting against the inside of the nozzle (see diagram). As the gas expands (adiabatically) the pressure against the nozzle's walls forces the rocket engine in one direction while accelerating the gas in the other.

The four expansion regimes of a de Laval nozzle: • under-expanded • perfectly expanded • over-expanded • grossly over-expanded

The most commonly used nozzle is the de Laval nozzle, a fixed geometry nozzle with a high expansion-ratio. The large bell- or cone-shaped nozzle extension beyond the throat gives the rocket engine its characteristic shape.

The exit static pressure of the exhaust jet depends on the chamber pressure and the ratio of exit to throat area of the nozzle. As exit pressure varies from the ambient (atmospheric) pressure, a choked nozzle is said to be

• under-expanded (exit pressure greater than ambient),
• perfectly expanded (exit pressure equals ambient),
• over-expanded (exit pressure less than ambient; shock diamonds form outside the nozzle), or
• grossly over-expanded (a shock wave forms inside the nozzle extension).

In practice, perfect expansion is only achievable with a variable-exit area nozzle (since ambient pressure decreases as altitude increases), and is not possible above a certain altitude as ambient pressure approaches zero. If the nozzle is not perfectly expanded, then loss of efficiency occurs. Grossly over-expanded nozzles lose less efficiency, but can cause mechanical problems with the nozzle. Fixed-area nozzles become progressively more under-expanded as they gain altitude. Almost all de Laval nozzles will be momentarily grossly over-expanded during startup in an atmosphere.[5]

Nozzle efficiency is affected by operation in the atmosphere because atmospheric pressure changes with altitude; but due to the supersonic speeds of the gas exiting from a rocket engine, the pressure of the jet may be either below or above ambient, and equilibrium between the two is not reached at all altitudes (see diagram).

For optimal performance, the pressure of the gas at the end of the nozzle should just equal the ambient pressure: if the exhaust's pressure is lower than the ambient pressure, then the vehicle will be slowed by the difference in pressure between the top of the engine and the exit; on the other hand, if the exhaust's pressure is higher, then exhaust pressure that could have been converted into thrust is not converted, and energy is wasted.

To maintain this ideal of equality between the exhaust's exit pressure and the ambient pressure, the diameter of the nozzle would need to increase with altitude, giving the pressure a longer nozzle to act on (and reducing the exit pressure and temperature). This increase is difficult to arrange in a lightweight fashion, although is routinely done with other forms of jet engines. In rocketry a lightweight compromise nozzle is generally used and some reduction in atmospheric performance occurs when used at other than the 'design altitude' or when throttled. To improve on this, various exotic nozzle designs such as the plug nozzle, stepped nozzles, the expanding nozzle and the aerospike have been proposed, each providing some way to adapt to changing ambient air pressure and each allowing the gas to expand further against the nozzle, giving extra thrust at higher altitudes.

When exhausting into a sufficiently low ambient pressure (vacuum) several issues arise. One is the sheer weight of the nozzle—beyond a certain point, for a particular vehicle, the extra weight of the nozzle outweighs any performance gained. Secondly, as the exhaust gases adiabatically expand within the nozzle they cool, and eventually some of the chemicals can freeze, producing 'snow' within the jet. This causes instabilities in the jet and must be avoided.

On a de Laval nozzle, exhaust gas flow detachment will occur in a grossly over-expanded nozzle. As the detachment point will not be uniform around the axis of the engine, a side force may be imparted to the engine. This side force may change over time and result in control problems with the launch vehicle.

Advanced altitude-compensating designs, such as the aerospike or plug nozzle, attempt to minimize performance losses by adjusting to varying expansion ratio caused by changing altitude.

Typical temperature (T), pressure (p), and velocity (v) profiles in a de Laval Nozzle

For a rocket engine to be propellant efficient, it is important that the maximum pressures possible be created on the walls of the chamber and nozzle by a specific amount of propellant; as this is the source of the thrust. This can be achieved by all of:

• heating the propellant to as high a temperature as possible (using a high energy fuel, containing hydrogen and carbon and sometimes metals such as aluminium, or even using nuclear energy)
• using a low specific density gas (as hydrogen rich as possible)
• using propellants which are, or decompose to, simple molecules with few degrees of freedom to maximise translational velocity

Since all of these things minimise the mass of the propellant used, and since pressure is proportional to the mass of propellant present to be accelerated as it pushes on the engine, and since from Newton's third law the pressure that acts on the engine also reciprocally acts on the propellant, it turns out that for any given engine, the speed that the propellant leaves the chamber is unaffected by the chamber pressure (although the thrust is proportional). However, speed is significantly affected by all three of the above factors and the exhaust speed is an excellent measure of the engine propellant efficiency. This is termed exhaust velocity, and after allowance is made for factors that can reduce it, the effective exhaust velocity is one of the most important parameters of a rocket engine (although weight, cost, ease of manufacture etc. are usually also very important).

For aerodynamic reasons the flow goes sonic ("chokes") at the narrowest part of the nozzle, the 'throat'. Since the speed of sound in gases increases with the square root of temperature, the use of hot exhaust gas greatly improves performance. By comparison, at room temperature the speed of sound in air is about 340 m/s while the speed of sound in the hot gas of a rocket engine can be over 1700 m/s; much of this performance is due to the higher temperature, but additionally rocket propellants are chosen to be of low molecular mass, and this also gives a higher velocity compared to air.

Expansion in the rocket nozzle then further multiplies the speed, typically between 1.5 and 2 times, giving a highly collimated hypersonic exhaust jet. The speed increase of a rocket nozzle is mostly determined by its area expansion ratio—the ratio of the area of the exit to the area of the throat, but detailed properties of the gas are also important. Larger ratio nozzles are more massive but are able to extract more heat from the combustion gases, increasing the exhaust velocity.

Vehicles typically require the overall thrust to change direction over the length of the burn. A number of different ways to achieve this have been flown:

• The entire engine is mounted on a hinge or gimbal and any propellant feeds reach the engine via low pressure flexible pipes or rotary couplings.
• Just the combustion chamber and nozzle is gimballed, the pumps are fixed, and high pressure feeds attach to the engine.
• Multiple engines (often canted at slight angles) are deployed but throttled to give the overall vector that is required, giving only a very small penalty.
• High-temperature vanes protrude into the exhaust and can be tilted to deflect the jet.

The most important metric for the efficiency of a rocket engine is impulse per unit of propellant, this is called specific impulse (usually written ${\displaystyle I_{sp}}$). This is either measured as a speed (the effective exhaust velocity ${\displaystyle v_{e}}$ in metres/second or ft/s) or as a time (seconds). For example, if an engine producing 100 pounds of thrust runs for 320 seconds and burns 100 pounds of propellant, then the specific impulse is 320 seconds. The higher the specific impulse, the less propellant is required to provide the desired impulse.

The specific impulse that can be achieved is primarily a function of the propellant mix (and ultimately would limit the specific impulse), but practical limits on chamber pressures and the nozzle expansion ratios reduce the performance that can be achieved.

Below is an approximate equation for calculating the net thrust of a rocket engine:[7]

${\displaystyle F_{n}={\dot {m}}\;v_{e}={\dot {m}}\;v_{e-opt}+A_{e}(p_{e}-p_{amb})}$

where:
${\displaystyle {\dot {m}}}$	=  exhaust gas mass flow
${\displaystyle v_{e}}$	=  effective exhaust velocity (sometimes otherwise denoted as c in publications)
${\displaystyle v_{e-opt}}$	=  effective jet velocity when Pamb = Pe
${\displaystyle A_{e}}$	=  flow area at nozzle exit plane (or the plane where the jet leaves the nozzle if separated flow)
${\displaystyle p_{e}}$	=  static pressure at nozzle exit plane
${\displaystyle p_{amb}}$	=  ambient (or atmospheric) pressure

Since, unlike a jet engine, a conventional rocket motor lacks an air intake, there is no 'ram drag' to deduct from the gross thrust. Consequently, the net thrust of a rocket motor is equal to the gross thrust (apart from static back pressure).

The ${\displaystyle {\dot {m}}\;v_{e-opt}\,}$ term represents the momentum thrust, which remains constant at a given throttle setting, whereas the ${\displaystyle A_{e}(p_{e}-p_{amb})\,}$ term represents the pressure thrust term. At full throttle, the net thrust of a rocket motor improves slightly with increasing altitude, because as atmospheric pressure decreases with altitude, the pressure thrust term increases. At the surface of the Earth the pressure thrust may be reduced by up to 30%, depending on the engine design. This reduction drops roughly exponentially to zero with increasing altitude.

Maximum efficiency for a rocket engine is achieved by maximising the momentum contribution of the equation without incurring penalties from over expanding the exhaust. This occurs when ${\displaystyle p_{e}=p_{amb}}$. Since ambient pressure changes with altitude, most rocket engines spend very little time operating at peak efficiency.

Since specific impulse is force divided by the rate of mass flow, this equation means that the specific impulse varies with altitude.

Due to the specific impulse varying with pressure, a quantity that is easy to compare and calculate with is useful. Because rockets choke at the throat, and because the supersonic exhaust prevents external pressure influences travelling upstream, it turns out that the pressure at the exit is ideally exactly proportional to the propellant flow ${\displaystyle {\dot {m}}}$, provided the mixture ratios and combustion efficiencies are maintained. It is thus quite usual to rearrange the above equation slightly:[8]

${\displaystyle F_{vac}=C_{f}\,{\dot {m}}\,c^{*}}$

and so define the vacuum Isp to be:

${\displaystyle v_{evac}=C_{f}\,c^{*}\,}$

where:

${\displaystyle c^{*}}$  =  the speed of sound constant at the throat

${\displaystyle C_{f}}$  =  the thrust coefficient constant of the nozzle (typically about 2)

And hence:

${\displaystyle F_{n}={\dot {m}}\,v_{evac}-A_{e}\,p_{amb}}$

Rockets can be throttled by controlling the propellant combustion rate ${\displaystyle {\dot {m}}}$ (usually measured in kg/s or lb/s). In liquid and hybrid rockets, the propellant flow entering the chamber is controlled using valves, in solid rockets it is controlled by changing the area of propellant that is burning and this can be designed into the propellant grain (and hence cannot be controlled in real-time).

Rockets can usually be throttled down to an exit pressure of about one-third of ambient pressure[9] (often limited by flow separation in nozzles) and up to a maximum limit determined only by the mechanical strength of the engine.

In practice, the degree to which rockets can be throttled varies greatly, but most rockets can be throttled by a factor of 2 without great difficulty;[9] the typical limitation is combustion stability, as for example, injectors need a minimum pressure to avoid triggering damaging oscillations (chugging or combustion instabilities); but injectors can be optimised and tested for wider ranges. For example, some more recent liquid-propellant engine designs that have been optimised for greater throttling capability (BE-3, Raptor) can be throttled to as low as 18–20 percent of rated thrust.[10][3] Solid rockets can be throttled by using shaped grains that will vary their surface area over the course of the burn.[9]

Rocket vehicle mechanical efficiency as a function of vehicle instantaneous speed divided by effective exhaust speed. These percentages need to be multiplied by internal engine efficiency to get overall efficiency.

Rocket engine nozzles are surprisingly efficient heat engines for generating a high speed jet, as a consequence of the high combustion temperature and high compression ratio. Rocket nozzles give an excellent approximation to adiabatic expansion which is a reversible process, and hence they give efficiencies which are very close to that of the Carnot cycle. Given the temperatures reached, over 60% efficiency can be achieved with chemical rockets.

For a vehicle employing a rocket engine the energetic efficiency is very good if the vehicle speed approaches or somewhat exceeds the exhaust velocity (relative to launch); but at low speeds the energy efficiency goes to 0% at zero speed (as with all jet propulsion). See Rocket energy efficiency for more details.

Rockets, of all the jet engines, indeed of essentially all engines, have the highest thrust to weight ratio. This is especially true for liquid rocket engines.

This high performance is due to the small volume of pressure vessels that make up the engine—the pumps, pipes and combustion chambers involved. The lack of inlet duct and the use of dense liquid propellant allows the pressurisation system to be small and lightweight, whereas duct engines have to deal with air which has around three orders of magnitude lower density.

Of the liquid propellants used, density is lowest for liquid hydrogen. Although this propellant has the highest specific impulse, its very low density (about one fourteenth that of water) requires larger and heavier turbopumps and pipework, which decreases the engine's thrust-to-weight ratio (for example the RS-25) compared to those that do not (NK-33).

The combustion may display undesired instabilities, of sudden or periodic nature. The pressure in the injection chamber may increase until the propellant flow through the injector plate decreases; a moment later the pressure drops and the flow increases, injecting more propellant in the combustion chamber which burns a moment later, and again increases the chamber pressure, repeating the cycle. This may lead to high-amplitude pressure oscillations, often in ultrasonic range, which may damage the motor. Oscillations of ±200 psi at 25 kHz were the cause of failures of early versions of the Titan II missile second stage engines. The other failure mode is a deflagration to detonation transition; the supersonic pressure wave formed in the combustion chamber may destroy the engine.[25]

Combustion instability was also a problem during Atlas development. The Rocketdyne engines used in the Atlas family were found to suffer from this effect in several static firing tests, and three missile launches exploded on the pad due to rough combustion in the booster engines. In most cases, it occurred while attempting to start the engines with a "dry start" method whereby the igniter mechanism would be activated prior to propellant injection. During the process of man-rating Atlas for Project Mercury, solving combustion instability was a high priority, and the final two Mercury flights sported an upgraded propulsion system with baffled injectors and a hypergolic igniter.

The problem affecting Atlas vehicles was mainly the so-called "racetrack" phenomenon, where burning propellant would swirl around in a circle at faster and faster speeds, eventually producing vibration strong enough to rupture the engine, leading to complete destruction of the rocket. It was eventually solved by adding several baffles around the injector face to break up swirling propellant.

More significantly, combustion instability was a problem with the Saturn F-1 engines. Some of the early units tested exploded during static firing, which led to the addition of injector baffles.

In the Soviet space program, combustion instability also proved a problem on some rocket engines, including the RD-107 engine used in the R-7 family and the RD-216 used in the R-14 family, and several failures of these vehicles occurred before the problem was solved. Soviet engineering and manufacturing processes never satisfactorily resolved combustion instability in larger RP-1/LOX engines, so the RD-171 engine used to power the Zenit family still used four smaller thrust chambers fed by a common engine mechanism.

The combustion instabilities can be provoked by remains of cleaning solvents in the engine (e.g. the first attempted launch of a Titan II in 1962), reflected shock wave, initial instability after ignition, explosion near the nozzle that reflects into the combustion chamber, and many more factors. In stable engine designs the oscillations are quickly suppressed; in unstable designs they persist for prolonged periods. Oscillation suppressors are commonly used.

Periodic variations of thrust, caused by combustion instability or longitudinal vibrations of structures between the tanks and the engines which modulate the propellant flow, are known as "pogo oscillations" or "pogo", named after the pogo stick.

Three different types of combustion instabilities occur:

This is a low frequency oscillation at a few Hertz in chamber pressure usually caused by pressure variations in feed lines due to variations in acceleration of the vehicle.[26]^:261 This can cause cyclic variation in thrust, and the effects can vary from merely annoying to actually damaging the payload or vehicle. Chugging can be minimised by using gas-filled damping tubes on feed lines of high density propellants.

This can be caused due to insufficient pressure drop across the injectors.[26]^:261 It generally is mostly annoying, rather than being damaging. However, in extreme cases combustion can end up being forced backwards through the injectors – this can cause explosions with monopropellants.

This is the most immediately damaging, and the hardest to control. It is due to acoustics within the combustion chamber that often couples to the chemical combustion processes that are the primary drivers of the energy release, and can lead to unstable resonant "screeching" that commonly leads to catastrophic failure due to thinning of the insulating thermal boundary layer. Acoustic oscillations can be excited by thermal processes, such as the flow of hot air through a pipe or combustion in a chamber. Specifically, standing acoustic waves inside a chamber can be intensified if combustion occurs more intensely in regions where the pressure of the acoustic wave is maximal.[27][28][29][26] Such effects are very difficult to predict analytically during the design process, and have usually been addressed by expensive, time-consuming and extensive testing, combined with trial and error remedial correction measures.

Screeching is often dealt with by detailed changes to injectors, or changes in the propellant chemistry, or vaporising the propellant before injection, or use of Helmholtz dampers within the combustion chambers to change the resonant modes of the chamber.

Testing for the possibility of screeching is sometimes done by exploding small explosive charges outside the combustion chamber with a tube set tangentially to the combustion chamber near the injectors to determine the engine's impulse response and then evaluating the time response of the chamber pressure- a fast recovery indicates a stable system.

For all but the very smallest sizes, rocket exhaust compared to other engines is generally very noisy. As the hypersonic exhaust mixes with the ambient air, shock waves are formed. The Space Shuttle generated over 200 dB(A) of noise around its base. To reduce this, and the risk of payload damage or injury to the crew atop the stack, the mobile launcher platform was fitted with a Sound Suppression System that sprayed 1.1 million litres (290,000 US gal) of water around the base of the rocket in 41 seconds at launch time. Using this system kept sound levels within the payload bay to 142 dB.[30]

The sound intensity from the shock waves generated depends on the size of the rocket and on the exhaust velocity. Such shock waves seem to account for the characteristic crackling and popping sounds produced by large rocket engines when heard live. These noise peaks typically overload microphones and audio electronics, and so are generally weakened or entirely absent in recorded or broadcast audio reproductions. For large rockets at close range, the acoustic effects could actually kill.[31]

More worryingly for space agencies, such sound levels can also damage the launch structure, or worse, be reflected back at the comparatively delicate rocket above. This is why so much water is typically used at launches. The water spray changes the acoustic qualities of the air and reduces or deflects the sound energy away from the rocket.

Generally speaking, noise is most intense when a rocket is close to the ground, since the noise from the engines radiates up away from the jet, as well as reflecting off the ground. Also, when the vehicle is moving slowly, little of the chemical energy input to the engine can go into increasing the kinetic energy of the rocket (since useful power P transmitted to the vehicle is ${\displaystyle P=F*V}$ for thrust F and speed V). Then the largest portion of the energy is dissipated in the exhaust's interaction with the ambient air, producing noise. This noise can be reduced somewhat by flame trenches with roofs, by water injection around the jet and by deflecting the jet at an angle.

The Rocketdyne H-1 engine, used in a cluster of eight in the first stage of the Saturn I and Saturn IB launch vehicles, had no catastrophic failures in 152 engine-flights. The Pratt and Whitney RL10 engine, used in a cluster of six in the Saturn I second stage, had no catastrophic failures in 36 engine-flights.[notes 1] The Rocketdyne F-1 engine, used in a cluster of five in the first stage of the Saturn V, had no failures in 65 engine-flights. The Rocketdyne J-2 engine, used in a cluster of five in the Saturn V second stage, and singly in the Saturn IB second stage and Saturn V third stage, had no catastrophic failures in 86 engine-flights.[notes 2]

The Space Shuttle Solid Rocket Booster, used in pairs, caused one notable catastrophic failure in 270 engine-flights.

The RS-25, used in a cluster of three, flew in 46 refurbished engine units. These made a total of 405 engine-flights with no catastrophic in-flight failures. A single in-flight RS-25 engine failure occurred during Space Shuttle Challenger's STS-51-F mission.[32] This failure had no effect on mission objectives or duration.[33]

The solar thermal rocket would make use of solar power to directly heat reaction mass, and therefore does not require an electrical generator as most other forms of solar-powered propulsion do. A solar thermal rocket only has to carry the means of capturing solar energy, such as concentrators and mirrors. The heated propellant is fed through a conventional rocket nozzle to produce thrust. The engine thrust is directly related to the surface area of the solar collector and to the local intensity of the solar radiation and inversely proportional to the I_sp.

Nuclear propulsion includes a wide variety of propulsion methods that use some form of nuclear reaction as their primary power source. Various types of nuclear propulsion have been proposed, and some of them tested, for spacecraft applications:

Slow development of this technology continued up to the later 19th century, when Russian Konstantin Tsiolkovsky first wrote about liquid-fueled rocket engines. He was the first to develop the Tsiolkovsky rocket equation, though it was not published widely for some years.

The modern solid- and liquid-fueled engines became realities early in the 20th century, thanks to the American physicist Robert Goddard. Goddard was the first to use a De Laval nozzle on a solid-propellant (gunpowder) rocket engine, doubling the thrust and increasing the efficiency by a factor of about twenty-five. This was the birth of the modern rocket engine. He calculated from his independently derived rocket equation that a reasonably sized rocket, using solid fuel, could place a one-pound payload on the Moon.

Goddard began to use liquid propellants in 1921, and in 1926 became the first to launch a liquid-propellant rocket. Goddard pioneered the use of the De Laval nozzle, lightweight propellant tanks, small light turbopumps, thrust vectoring, the smoothly-throttled liquid fuel engine, regenerative cooling, and curtain cooling.[9]^:247–266

During the late 1930s, German scientists, such as Wernher von Braun and Hellmuth Walter, investigated installing liquid-fueled rockets in military aircraft (Heinkel He 112, He 111, He 176 and Messerschmitt Me 163).[45]

The turbopump was employed by German scientists in World War II. Until then cooling the nozzle had been problematic, and the A4 ballistic missile used dilute alcohol for the fuel, which reduced the combustion temperature sufficiently.

Staged combustion (Замкнутая схема) was first proposed by Alexey Isaev in 1949. The first staged combustion engine was the S1.5400 used in the Soviet planetary rocket, designed by Melnikov, a former assistant to Isaev.[9] 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.

In the West, the first laboratory staged-combustion test engine was built in Germany in 1963, by Ludwig Boelkow.

Hydrogen peroxide / kerosene fueled engines such as the British Gamma of the 1950s used a closed-cycle process (arguably not staged combustion, but that's mostly a question of semantics) by catalytically decomposing the peroxide to drive turbines before combustion with the kerosene in the combustion chamber proper. This gave the efficiency advantages of staged combustion, whilst avoiding the major engineering problems.

Liquid hydrogen engines were first successfully developed in America, the RL-10 engine first flew in 1962. Hydrogen engines were used as part of the Apollo program; the liquid hydrogen fuel giving a rather lower stage mass and thus reducing the overall size and cost of the vehicle.

Most engines on one rocket flight was 44 set by NASA in 2016 on a Black Brant.[46]