Liquid rocket engine

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Atlas V liquid rocket with payload to Mars

Liquid rocket engines are reaction drives that are mainly used in space travel today.

In contrast to solid drives , in which in the combustion chamber a finished, befindliches in a solid state mixture of fuel and oxidant burns, be in liquid rocket a ( Monergol ) or several ( Diergole , Triergole ) liquid chemical components entrained in (separate) tanks and actual in the Engine promoted. There a continuous chemical reaction occurs (catalytic decomposition of a monergol, combustion of fuel and oxidizer). The gas masses created by the increase in volume flow out of a nozzle as a support mass and thereby generate thrust in the opposite direction. Since the oxidizer is carried in the rocket, the combustion of the fuel can take place without the presence of atmospheric oxygen, e.g. B. in the high atmosphere or in space. In diergolen liquid rockets, the fuel and oxidizer are first mixed in the combustion chamber ; the delivery to the combustion chamber takes place in separate piping systems.

Typical parameters of such a rocket engine are the thrust (the actual propulsion force, usually given in kilonewtons (kN) , often differentiated into ground or take-off thrust and vacuum thrust ) as well as the specific impulse as a key figure for the efficiency of the engine regardless of its size.


1st stage of the Saturn V rocket with a total of five F-1 liquid engines, in the picture Wernher von Braun
Walter rocket motor of the Me 163b

Early theoretical approaches to the use of liquid rockets were published in 1903 by the Russian space pioneer and thought leader Konstantin Eduardowitsch Ziolkowski under the title Exploring Space Using Reaction Apparatuses in the Russian journal Wissenschaftliche Rundschau . On March 16, 1926, the American researcher Robert Goddard succeeded in launching a liquid rocket for the first time (2.5 s flight duration, 14 m height, 50 m flight range). In October 1930, a rocket Goddard had already reached 800 km / h and 610 m altitude. At almost the same time in Germany, from 1930 onwards, test launches with liquid rockets were carried out at the Berlin rocket airfield by the Space Agency. The German research efforts finally led - after the military had taken over the missile program - via the test models A1 , A2 and A3 to the first large rocket with liquid propulsion, the unit 4 (A4) , which was mainly sold under its propaganda name, "Retaliation Weapon 2", short V2, should become known. With the fuel combination of 75% ethanol and oxygen, this crossed the boundary to space for the first time. At the same time during the Second World War smaller monergole were ( "cold") and diergole hydrogen peroxide rocket engines ( H 2 O 2 / Petroleum or N 2 H 4 ) as a starting aid for aircraft, or directly to the drive of interceptors (z. B Me 163 ) used . After the collapse of the German Reich and the withdrawal of scientists and technologies, development was mainly continued by the victorious USA and the Soviet Union, who both made use of captured documents and German developers. During the Cold War , the need for ever more powerful ICBMs pushed ahead engine development - at that time mostly with liquid propulsion. Ultimately, some of these developments could also be used as launch vehicles for space travel purposes (e.g. the R-7 variants for the important Sputnik 1 and Vostok 1 flights with Yuri Gagarin , the first person in space, or the American Titan II Gemini ). The development reached a climax in the late 1960s with the giant F-1 engines of the Saturn V moon rocket . Recent developments are e.g. B. the main engine of the space shuttle or the RD-170 of the Energija rocket, which can be reused. Since the requirements for military missiles have changed (mobility, stationing on submarines as SLBM , permanent and immediate readiness for launch), solid rockets, which are easier to handle, have replaced liquid rockets in this area .

As the history of rocket technology and the fate of some rocket pioneers shows, the development of liquid rockets was initially associated with greater dangers and technical hurdles than that of solid rockets. The reasons are varied: the risk of leaks, evaporation and explosions , damage to pumps and other units , air bubbles or insufficient mixing in the combustion chamber , variable weight distribution during combustion.


A liquid rocket engine essentially consists of a combustion chamber, a nozzle, a pumping device for the propellants (see the section on construction ) and, if necessary, an ignition device. Supplementary components are the thrust frame, which transfers the thrust to the rocket structure, smaller tanks for auxiliary media (e.g. compressed gas, coolant, lubricant, pump and starting fuel) as well as more or less complex pipelines, valves and flow regulators for the operating and auxiliary media. Control elements such as hydraulic cylinders or servomotors for swiveling the combustion chamber or nozzle unit (see also thrust vector control ) can also be part of the engine.

Combustion chamber

Cut-open RD-107 engine unit (center), above: cylindrical combustion chamber, below: conical nozzle bell

The combustion chamber is a container made of metal in which the fuel is mixed with the oxidizer and burns continuously. As a rule, combustion chambers are made cylindrical for manufacturing reasons. The injection head or an injector plate are arranged on the front side of the combustion chamber opposite the nozzle opening. These have the task of intensively and finely mixing the fuel components brought in separate pipes during the injection in order to ensure complete and complete combustion. The throughput can be several hundred liters per second with large engines (up to 155 tons per minute for the F-1). The length of the combustion chamber must be dimensioned so that the injected components can react completely with one another, on the other hand, the combustion chamber must be as compact as possible in order to avoid undesired heat transfer to the walls. The pressure in the combustion chamber resulting from the combustion can, depending on the design of the engine, reach from under 30 bar to well over 100 bar (currently 205 bar for the SSME and over 245 bar for the RD-170/171).

In order to prevent the combustion chamber from melting and burning through or exploding due to the immense combustion temperatures and pressures inside, it must be cooled. Common methods for this are active or regenerative cooling, in which part of the fuel or the oxidizer flows through in the form of liquid cooling between the double-walled combustion chamber walls before it is injected. If the fuel component is not fed into the combustion process after passing through the cooling jacket, but is released into the environment, this is known as dump cooling . Further measures are shot and fog cooling , in which specifically a local excess fuel is generated in the combustion zone close to the wall or directly to the walls by a particular arrangement of the injection holes to the combustion temperatures lower there, and to utilize the latent heat of evaporation of the fuel; furthermore, the wall is thus also protected from the reaction with the oxidizer. Coatings of the inner walls with heat-resistant, insulating materials (ceramic coatings, mineral fibers such as asbestos ) or ablative materials are also used, which create a heat-insulating boundary layer to the wall through their phase transition when melting . These measures are used for smaller engines with short combustion times, as is the manufacture of the combustion chambers from high-temperature-resistant niobium or tantalum alloys; in these cases, one speaks of passive cooling .

The design of the combustion chamber as well as the injection head or injector plate is a challenge during construction and testing, since malfunctions can lead to discontinuous combustion and even resonant combustion oscillations , which can endanger the entire spacecraft via the reaction via the liquid columns in the fuel lines and the mechanical structure (see pogo effect ).

Thrust nozzle

Rocket nozzle of a Pratt & Whitney RL-10 B of a Delta IV upper stage, the orange and the upper dark part are fixed, the lower dark part is brought into its working position after the stage separation by means of the threaded spindles.
Rocket nozzles made up of individual cooling tubes (XLR-87 of a Titan I)

The exhaust nozzle in the form of a Laval nozzle connects directly to the combustion chamber . This consists of a constriction to increase the speed of the gas, the so-called nozzle neck, which in turn merges into a bell-shaped or conical part in which the thrust is generated by the expansion of the gases. The aerospike engines under development should do without such a thrust nozzle in the conventional sense.

Like the combustion chamber, the nozzle is exposed to high thermal loads, which require cooling measures. Both active and passive cooling processes are used. In the active process, the fuel component diverted for cooling is not only fed into the double wall of the combustion chamber, but also through the double-walled nozzle bell; passive cooling processes are carried out in the same way as with the combustion chamber. A special form of nozzle cooling is the ring-shaped introduction of the relatively cool working gas of the turbo pumps in the bypass flow method into the nozzle bell about halfway between nozzle throat and mouth, which was used in the F-1 engines of the Saturn 5 rocket. Occasionally, especially if an internal curtain or film cooling system is used at the same time, active cooling of the nozzle bell is dispensed with, as is the case with the Viking engine of the Ariane 4 . Here the material heated up to red heat during operation .

Often the combustion chamber and nozzle are manufactured in one part. In order to obtain the coolant channels required for cooling, the basic structure of the combustion chamber or nozzle units of larger engines often consists of bundles of nickel steel tubes (e.g. made of Inconel X-750), which are bent into the shape of the workpieces be brazed . These structures are then reinforced by stiffening rings and massive jackets as well as assembly and connection fittings. During operation, the tubes are flowed through by the cooling medium (fuel or oxidizer), usually in the direction from the nozzle opening to the combustion chamber.

The ratio of the cross-sectional areas of nozzle throat and nozzle mouth is called the relaxation ratio . Depending on the ambient pressure conditions and thus the external pressure "against" which the engine is supposed to work (dense atmosphere on the earth's surface, decreasing pressure with increasing altitude up to the vacuum in space), the expansion ratio in practice is around 10 to 100, a special one The projected European upper stage engine Vinci has a high ratio with 240 in order to achieve a high specific impulse at low ambient pressure. For pure lower stage engines that only work in denser atmospheric layers, smaller expansion ratios are sufficient, upper stage and orbital engines require higher expansion ratios for efficient operation, but the maximum possible and permissible expansion is also limited, cf. for this the Summerfield criterion . In order to circumvent these design problems of the thrust nozzle, research is being carried out on aerospike engines that have an expansion ratio that adapts itself to the ambient pressure.

Higher expansion ratios require larger and therefore heavier nozzle bells, which, due to their overall length, can also have an unfavorable effect on the overall design of the rocket (longer stage adapters are required to accommodate the nozzles), which is why some upper stage engines have an extendable nozzle for after stage separation and before ignition the lower extension part of the nozzle bell is extended telescopically over the part of the bell that is firmly connected to the combustion chamber (projected for the Vinci , implemented for the RL10B-2 in the upper stage of Delta IV ).

Types of fuel delivery

Every liquid rocket engine has a combustion chamber with an adjoining thrust nozzle as a central component. The main differences between the different designs lie in the way in which the fuel gets from the tanks into the combustion chamber and in what way, in the case of engines with turbo pumps, the working medium of the turbines (the hot gas) as well as the fuels and oxidizers are conveyed.

Compressed gas delivery

Scheme of the pressurized gas propulsion of the Apollo spacecraft (CSM)

The compressed gas feed (English Pressure-fed cycle ) is the simplest embodiment, it completely avoids mechanical pumps and promotes the fuels by the tanks with an inert gas (usually helium ), which in separate pressure cylinders is carried pressurized, and pressurized be . The liquids are pressed into the combustion chambers by the tank pressure via simple pipes. The limits of this design, which is simple and relatively reliable due to the small number of components, are that the tanks must be made relatively stable and heavy as pressure vessels in order to withstand the pressure of the conveying gas, and the achievable combustion chamber pressure is also limited by the maximum permissible overpressure in the tanks. The use is therefore limited to smaller and weaker thrust applications, for example control and maneuver thrusters for spacecraft or apogee engines . Practical examples are the ascent and descent engines of the Apollo lunar module or the main engine of the command / service module of the Apollo spacecraft . The use of hypergolic components made it possible to build very simple, reliable engines with very few mechanical components that could be reliably ignited even after missions lasting several days or that were designed to be re-ignited many times, like the main engine of the Apollo-CSM.

Pump delivery

Cut open turbo pump of an A4 rocket

More powerful engines, on the other hand, use mechanical pumps to transport the fuels from the tanks, which are only under very little overpressure, into the combustion chamber ("active fuel delivery"). Since the drive power requirement for this pumping work is very high (up to several dozen megawatts per engine, with the Mark 10 pump each of the five F- 1s of the Saturn moon rocket over 40 megawatts, the equivalent of 55,000 wave horsepower, 190 megawatts for the Russian RD- 170 ) only compact centrifugal pumps driven by gas turbines come into consideration, the working gas of which is generated with the rocket fuels carried, regardless of the ambient atmosphere. Such a turbo pump usually consists of a device for generating the working gas, the working turbine itself and one or more single or multi-stage radial pumps (one each for fuel and oxidizer) that are mechanically driven by the turbine. Often at least the turbine and the pump assemblies are combined in a housing and arranged on a common shaft. As a rule, the turbo pumps are mounted on an equipment rack on the engine in the immediate vicinity of the combustion chamber. There are also arrangements in which a central turbo pump supplies several individual combustion chambers at the same time, as in the RD-170 with one pump for four combustion chambers.

Depending on the type of hot gas generation and the flow pattern of the various media, hot gas and fuels, different variants of active fuel delivery have developed over time. The basic variants mentioned can often be divided into sub-variants.

Sidestream process

In the bypass flow process ( gas generator cycle or open cycle ), part of the fuel and oxidizer pumped to the combustion chamber is diverted and burned in a separate combustion chamber. A non-stoichiometric combustion (fuel or oxidizer excess) is aimed for in order to reduce the hot gas temperatures to a level that is tolerable for the turbine materials (400 to 700 K ). After the hot gas flow in the turbine has performed its work, the expanded hot gas is either used to cool the nozzle or released into the environment via an exhaust pipe next to the thrust nozzle. In this engine variant, there are at least two streams (main stream to the main combustion chamber and the fuel to the gas generator combustion chamber in the secondary stream; possibly a third stream for nozzle and combustion chamber cooling). Around five percent of the total fuel in a stage is used to drive the pump due to incomplete combustion and is no longer available for the actual thrust generation of the rocket engine; on the other hand, it is a tried and tested, proven and manageable technology. The sidestream process is the oldest and most common variant. Many larger rocket engines work according to this principle, including the F-1 of the Saturn sub-stage S1C . A sub-variant is the use of a separate fuel for the turbopump gas generator as in the V2 / A4 rocket or the RD-107 of the Soviet Soyuz / R7 rocket , both of which use the catalytic decomposition of hydrogen peroxide to generate the pump working gas.

Mainstream process

RD-170 model, a main flow engine with a central turbo pump for four combustion chambers

In the later developed main flow process (English staged combustion or closed cycle ), the principle of the bypass flow process is varied in such a way that a larger part or the entire flow of a fuel component runs through a gas generator (here called pre- burner ) and with a very small proportion of the other component reacts unstoichiometrically. The result is a hot gas flow that still contains large excess quantities of unreacted fuel or oxidizer, which after driving the power turbine of the turbo pump is fed directly into the main combustion chamber and there takes part in the regular combustion reaction to generate thrust with the remaining components injected there. In contrast to the bypass flow process, no unused fuel components go overboard that do not contribute to the overall momentum of the engine. With the main flow process, the highest combustion chamber pressures and high specific impulses can be achieved; on the other hand, this process places the highest demands on development and production due to the high pressures in the pipelines and the handling of the hot gas flow. Well-known representatives of the main flow process are the SSME , the RD-0120 and again the RD-170 .

Expander process

A variation of the main flow process is the expander cycle . This differs from the main flow process in that no gas generator or preburner is used. Rather, one of the two fuel components is pumped through its cooling jacket to cool the combustion chamber. The liquid evaporates and the expanding stream of superheated steam drives the working turbine of the feed pumps. After passing through the turbine, this flow is directed into the main combustion chamber as in the main flow process. This process only works with substances that do not decompose during evaporation and that are still in the gaseous phase after the expansion in the turbine, such as e.g. B cryogenic oxygen (LOX) or hydrogen or low molecular weight hydrocarbons such as methane , ethane and propane ; Kerosene, for example, would condense again too quickly here. Examples of expander cycle engines are the RL-10 of the Centaur upper stage or the European Vinci . The process was modified in places in such a way that only a small amount of fuel was evaporated in the combustion chamber cooling jacket and, after it was used as a working medium for the turbo pump, was released into the environment ( expander bleed cycle ), e.g. B. the LE-5A of the Japanese HIIA missile.

Advantages and disadvantages


  • In contrast to solid rockets, certain liquid engines can be switched off and re-ignited. This is important for steering thrusters when only short impulses are required or to leave the earth orbit (for example in the S-IVB sequence of the Apollo moon flights).
  • The rocket can be assembled without fuel and transported to the launch site, making it lighter and there is no risk of explosion or fire during assembly and transport. Refueling takes place shortly before the start. However, special facilities must be available at the launch pad .
  • Liquid engines can be checked for their functionality (thrust, pump speed, combustion chamber pressure) between the ignition and the lift-off of the rocket from the launch pad.
  • The thrust can be regulated during operation.
  • Liquid rockets often use the fuel more efficiently than solid rockets and thus achieve higher top speeds with the same amount of fuel.
  • The frequently used fuel combination LOX / LH2 burns to water and is therefore locally ecologically harmless.


  • Liquid rockets and engines are more expensive, more complex and therefore more error-prone than solid rockets.
  • The missile's center of gravity shifts as the fuel is consumed . The missile's stabilization and control system must be able to compensate for this displacement.
  • The pogo effect (vibrations in the engine power due to resonance of the liquid columns in the fuel lines and the mechanical structure of the rocket) can occur.
  • Liquid missiles are more dangerous to explode if they leak because the liquids are more easily flammable.
  • Some fuels (including hydrazine derivatives) are toxic; if released (false starts, burnt-out steps fall back to the ground), environmental damage can occur.
  • Cryogenic fuel components may only be refueled shortly before take-off, otherwise they will evaporate prematurely due to warming, which prevents quick-reacting starts or a longer-lasting readiness for take-off. Some storable liquid fuels are highly caustic or corrosive and over time attack the materials of the rocket structure.


The most energetic fuel mixture used in liquid rockets today is cryogenic oxygen and hydrogen (LOX / LH 2 ).

Depending on the fuel mixture used, temperatures of up to 4200 ° C and pressures of over 25 MPa can occur in the combustion chamber.

Manufacturer (selection)

See also


Individual evidence

  1. ^ Kyrill von Gersdorff, Kurt Grasmann, Helmut Schubert (1995) Aircraft engines and jet engines Bernard & Graefe Verlag. ISBN 3-7637-6107-1 , p. 268 ff.
  2. ^ Picture and description of the Walter 109-509C of the Me 163
  3. a b c d e f Stages to Saturn - Fire, Smoke, and Thunder: The Engines Publication in NASA's history archive about the F-1 engine (English)
  4. Representation of the extendable exit cone on the RL-10B2 in the Encyclopedia Astronautica (English)
  5. Power Cycles - Description of the various pump delivery processes at (English)
  6. Article on the technology of rocket engines on Bernd Leitenberger's website
  7. Wiebke Plenkers, Martin B. Kalinowski: Danger scenarios of the release of plutonium by a successful launch with a missile defense system. (PDF; 1.2 MB) Carl Friedrich von Weizsäcker Center for Science and Peace Research, December 2008, p. 17 , accessed on December 5, 2015 .

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