Liquid-propellant rocket

Liquid propellant rocket engines are reaction drives which are currently used primarily in the aerospace industry.

In contrast to solid drives in which the combustion chamber burns a finished, befindliches in the solid state mixture of fuel and oxidizer are in liquid rocket a ( Monergol ) or more ( Diergole, Triergole ) entrained liquid chemical components in ( separate ) tanks and in the actual sponsored engine. There is a continuous chemical reaction (catalytic decomposition of a Monergols, combustion of fuel and oxidant ). The resulting increase in volume of gas by the mass flow of the support material from a nozzle, thereby generating a thrust in the opposite direction. Since the oxidant is carried in the rocket, which combustion of the fuel can be used without the presence of atmospheric oxygen to take place, for example, high in the atmosphere or in space. The mixture of fuel and oxidizer takes place at diergolen liquid rocket only in the combustion chamber, the promotion to the combustion chamber takes place in separate piping systems.

Typical parameters of such a rocket engine is the thrust (the actual driving force, usually expressed in kN, often differentiated in soil or takeoff thrust and vacuum thrust ) and the specific impulse as a measure of the effectiveness of the engine regardless of its size.

  • 3.1 Compressed gas production
  • 3.2 pumps Promotion 3.2.1 In addition to current methods
  • 3.2.2 main stream method
  • 3.2.3 Expander method

History

Early theoretical approaches on the use of liquid rockets were published in 1903 by the Russian space pioneer Konstantin and vordenker Eduardo Ziolkowski under the title space exploration by means of chemical reactors in the Russian journal Scientific Review. On March 16, 1926 the U.S. researcher Robert Goddard ( 2.5 s flight duration, 14 m high, 50 m flight distance ) scored the first launch of a liquid rocket. In October 1930 a rocket Goddard had already reached 800 km / h and 610 m height. Were almost the same in Germany from 1930 on the rocket launching site in Berlin by the Spaceflight Club launches attempt carried out with liquid rockets. The German research efforts eventually led - after the military had pulled the missile program itself - over the experimental models A1, A2 and A3 to the first large rocket with liquid drive, the unit 4 (A4). This exceeded the fuel combination of 75 % ethanol and oxygen for the first time the edge of space. At the same time the Second World War smaller monergole ("cold " ) and diergole hydrogen peroxide rocket engines ( H2O2/Petroleum or N2H4 ) to help launch aircraft or directly to the drive of interceptors (eg, the Me-163 ) were used. After the collapse of the German Reich, development was mainly pursued by the victorious powers, the U.S. and the Soviet Union, both of captured German documents and developers made ​​use of. In the Cold War, the engine development was driven by the need for ever more powerful ICBMs - then mostly with fluid drives. Ultimately, some of these developments could also be used as launch vehicles for space applications (eg, the R -7 variants of the major flights Sputnik 1 and Vostok 1 with Yuri Gagarin, the first man in space or the American Titan II Gemini). Reached a high point in the development of the late 1960s with the huge F-1 engines of the Saturn V moon rocket. More recent developments are, for example, the main engine of the space shuttle or the RD- 170 of the Energia rocket, which can be reused. Since the requirements have changed to military missiles ( mobility deployments on submarines as SLBM, permanent and immediate launch readiness ) have replaced liquid-fueled rocket in this area, the easier -to-handle solid fuel rockets.

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

Components

A liquid rocket engine consisting essentially of a combustion chamber, a nozzle, a pumping apparatus for the propellant ( see Section types ) and optionally an igniter. Additional components, the thrust frame, which transmits the thrust of the rocket structure, smaller tanks for auxiliary media (such as compressed gas, coolant, lubricant, pumps and start-up fuel) as well as more or less sophisticated piping, valves and flow controllers for operating and auxiliary media. Similarly, controls, such as hydraulic cylinders or servo motors for rotating the combustion chamber or nozzle unit ( see also thrust vector control) can be part of the engine.

Combustion chamber

The combustion chamber is a metal-made vessel in which the fuel is mixed with the oxidant and burned continuously. In general, combustion chambers for manufacturing reasons are cylindrical. At the front end, opposite the nozzle orifice of the combustion chamber of the injection head or Injektorenplatte are arranged. Their task is to mix the brought up in separate piping components in the fuel injection intense and delicate, to ensure a thorough and exhaustive complete combustion. The throughput may be from large engines several hundred liters per second amount ( the F -1 a total of up to 155 tons per minute). The length of the combustion chamber has to be dimensioned so that the injected components can completely react with each other, on the other hand, the combustion chamber has to be as compact as possible in order to avoid undesired heat transfer to the walls. The resulting from combustion pressure in the combustion chamber may, depending on the design of the engine below 30 bar up to 100 bar far beyond reach ( currently at 205 bar and 245 bar for the SSME RD-170/171 ).

To prevent melting and burning or explosion of the combustion chamber due to the immense combustion temperatures and pressures in its interior, it must be cooled. Used method for this purpose are the active, or regenerative cooling, in which a part of the fuel or the oxidizer in the form of a liquid-cooled combustion chamber walls between the double-walled produced passes before it is injected. Will the fuel component after the cooling jacket happen not supplied to the combustion, but discharged into the environment, we speak of loss of cooling ( english dump cooling ). Other measures include film and curtain cooling, where in the combustion zone near the wall or directly on the walls by a particular arrangement of the injection holes targeted a local excess fuel is produced to lower the combustion temperatures there and take advantage of the latent heat of evaporation of the fuel; Further, the wall is protected prior to the reaction with the oxidizer. Also coating the inner walls with a heat-resistant insulating materials (ceramic coatings, mineral fibers such as asbestos ), or ablative materials are used that provide a heat insulating barrier to the wall during the melting process due to their phase transition. These measures come with smaller engines with short burn times for use, as well as the production of the combustion chambers of high- temperature-resistant niobium or tantalum alloys, we speak in these cases of passive cooling.

The design of the combustion chamber and injection head or Injektorenplatte is a challenge in the design and testing, as malfunction may result in an unsteady combustion to resonant combustion oscillations, which may jeopardize the entire spacecraft above the retroactive effect of the liquid column in the fuel lines and the mechanical structure (see Pogoeffekt ).

Exhaust nozzle

Directly to the combustion chamber closes the exhaust nozzle in the form of a Laval nozzle. This consists of a constriction of the gas to increase in speed, known as the throat, which in turn merges into a bell-shaped or cone-shaped part is produced in the expansion of the gases of the thrust. The under development aerospike engines are provided without such an exhaust nozzle in the conventional sense.

As well as the combustor, the nozzle is exposed to high thermal loads, which requires measures to cool. There are just as active and passive cooling methods used. The active method, the branched-off fuel component for cooling is guided not only in the double wall of the combustion chamber, but also by the double-walled nozzle cone; passive cooling methods are also carried out as in the combustion chamber. A special form of nozzle cooling the annular introducing the relatively cool working gas, the turbo pump during bypass mode in the nozzle cone to about halfway between the nozzle throat and mouth, which was used in the F-1 engines of the Saturn 5 rocket. Occasionally, especially with simultaneous use of an internal haze or film cooling, need for active cooling of the nozzle cone, such as the Viking engine of the Ariane 4 Here the material is heated to red heat in operation.

Often combustion chamber and nozzle are manufactured in one piece. In order to obtain the necessary cooling coolant channels, the basic structure of the combustion chamber and nozzle units of larger engines often consists of all bundles of stainless steel tube (for example made ​​of Inconel X-750 ), which is bent into the shape of the workpieces and be brazed. These structures are then reinforced by stiffening rings and massive shells and mounting and connection fittings. The tubes are in operation, mostly in the direction of the nozzle mouth down to the combustion chamber, flows through the cooling medium (fuel or oxidizer ).

The ratio of the cross sectional areas of the nozzle throat and nozzle outlet is referred to as an expansion ratio. Depending on the ambient pressure conditions and thus the external pressure, " against" when the engine needs work ( dense atmosphere at the surface, decreasing pressure with increasing height in the universe up to the vacuum) is the expansion ratio in practice about 10 to 100, a particularly high ratio, the projected European upper stage drive Vinci with 240 to achieve a high specific impulse at low ambient pressure. For pure sublevels engines that operate only in the denser layers of the atmosphere, rich minor relaxation ratios, upper stages and orbital engines require higher relaxation conditions for efficient work, however, the maximum possible and permissible expansion is also limited, cf the Summerfield criterion. To circumvent these problems of interpretation of the thruster being researched aerospike engines that have a self-adapting to the ambient pressure expansion ratio.

Higher relaxation ratios require larger and therefore heavier nozzle bells, which can also affect the overall design of the rocket unfavorably by their length (longer stages adapters to accommodate the nozzle required), so some upper stage engines have an extendable nozzle, in which, after stage separation and before the ignition the lower extension of the nozzle cone is extended telescopically over the firmly connected to the combustion chamber of the bell ( configured for the Vinci, realized the RL10B -2 in the upper stage of the Delta IV).

Types of fuel Promotion

Each liquid rocket engine has as its central component a combustion chamber with subsequent thereto exhaust nozzle. The main differences of the different types are in the way, as the fuel from the tank enters the combustion chamber and in what way in engines with turbo pumps, the working medium of the turbine ( hot gas ) and the subsidized fuels and oxidizers are performed.

Pressure gas production

The compressed gas extraction (english Pressure -fed cycle) is the most basic version, they dispensed with entirely mechanical pumps and promotes fuels by the tanks with an inert gas (usually helium), which is carried in separate pressure cylinders, pressurized and pressurized. The liquids are forced through simple piping through the tank pressure in the combustion chambers. The boundaries of these simple due to the small number of components and relatively reliable design are that the tanks need to be designed as a pressure container is relatively stable and difficult to withstand the pressure of the transport gas, as well as the achievable combustion chamber pressure is limited by the maximum permissible excess pressure in the tanks. The use is therefore limited to smaller and weaker shear applications, such as control and maneuvering thrusters for spacecraft or apogee motors. Practical examples include 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 so very simple, reliable engines could be built with very few mechanical components that could be ignited reliably even after several days of emissions or on a plurality of times re-ignition, as the main engine of the Apollo CSM, were designed.

Pumps promotion

More powerful engines, however, use mechanical pumps to transport the fuels of the standing only under very low pressure tank into the combustion chamber ( " Active Treibstofförderung "). Since the drive power requirement is very high for this pump work (up to tens of thousands of kilowatts per engine in the Mark 10 pumps each of the five F-1 of the Saturn moon rocket than 40,000 kW, the equivalent of 55,000 shaft horsepower, 190 MW at the Russian RD- 170) come only compact, driven by gas turbines centrifugal pumps are those whose working gas is generated independent of the ambient atmosphere with entrained solid propellants. Such a turbopump is generally of a device for generation of the working gas, the power turbine itself, and one or more single or multi- stage centrifugal pumps ( one for the fuel and oxidant ), which are mechanically driven by the turbine. Often at least the turbine, the pump modules are combined in a housing and arranged on a common shaft. The turbo pumps are mounted usually in close proximity to the combustion chamber at an equipment rack on the engine. There are also arrangements in which a central turbo pump simultaneously supplies a plurality of individual combustion chambers as in the RD -170 with a pump for four combustion chambers.

Depending on the nature of the hot gas generator and the flow diagram of the various media and the hot gas fuels in the course of time various types of active Treibstofförderung developed. The basic variants mentioned can be often divided into sub-variants.

Bypass mode

When bypass mode (English gas generator cycle or open cycle) is diverted part of the pumped fuel and oxidizer to the combustion chamber and burned in a separate combustion chamber. Here, a non-stoichiometric combustion ( fuel or Oxidatorüberschuss ) is desirable to reduce the temperatures of the hot gas to an acceptable degree for the turbine materials (about 400-700 K). After the hot gas stream has not given his job performance in the turbine, the expanded hot gas is either used for nozzle cooling or discharged via an exhaust pipe next to the exhaust nozzle in the area. In this engine variant, therefore ( to the main combustion chamber and in addition to the main electrical power to the fuel gas generator combustion chamber; possibly a third stream to the nozzle and combustion chamber cooling) there are at least two streams. Approximately five percent of the total fuel one stage be availed by imperfect combustion to drive pumps and are thus no longer the actual generation of thrust of the rocket motor available; on the other hand is a tried and proven and controllable technology. The bypass procedure is the oldest and most popular variant. Many larger rocket engines operate according to this principle, including the F -1 of the Saturn stage S1C. A sub-variant is to use a separate fuel pump for the turbo gas generator as in the V2/A4-Rakete or RD 107 Soviet Sojus/R7-Rakete, both of which use the catalytic decomposition of hydrogen peroxide to produce the pump working gas.

Main stream method

When later developed main stream method (English Staged combustion or closed cycle), the principle of the bypass procedure varies in that a greater part or the whole current of a fuel component a gas generator (called pre-burner, English Preburner ) passes through and with a very small proportion of the other component unstöchiometrisch responding. This results in a hot gas stream, which still contains large amounts of excess of unreacted fuel or oxidizer that is fed directly into the main combustor after driving the power turbine of the turbo pump and participate therein to the regular combustion reaction to the thrust generation with the remaining therein injected components. Thus, in contrast to go bypass mode no unused fuel components on board, which do not contribute to the total momentum of the engine. With the main stream method can be the highest combustion chamber pressures and high specific impulses achieve, on the other hand, this method due to the high pressures in the pipelines and the handling of the hot gas stream the highest demands on the development and manufacturing. Known representatives of the main current method are the SSME, the RD -0120, and again the RD- 170.

Expander process

A variation of the main stream process is the expander process (English expander cycle). This deviates from the main stream method, as no gas producer or pre-burner ( Preburner ) will be used. Rather, one of the two fuel components to cool the combustion chamber is pumped through the cooling jacket. The liquid and the expanding hot vapor stream evaporates driving the power turbine to the feed pump. After passing through the turbine, this stream is passed as the main process stream in the main combustion chamber. This procedure only works with materials that do not decompose during evaporation and after expansion in the turbine is still present in the gaseous phase, such B cryogenic oxygen ( LOX) or hydrogen or low molecular weight hydrocarbons such as methane, ethane, and propane; Kerosene for example, would condense here too quickly. Examples of expander cycle engines are the RL- 10 of the Centaur upper stage or the European Vinci. The process was locally modified in a way that only a small amount of fuel was vaporized in the combustion chamber cooling jacket and was discharged after use as a working medium for the turbopump in the environment ( expander bleed cycle), such as the LE -5A Japanese HIIA rocket.

Pros and Cons

Advantages:

  • In contrast to solid rocket engines, certain liquid can turn off and relight. This is important for control thrusters, if only short pulses are needed or to leave the Earth's orbit (for example, the S- IVB sequence of the Apollo moon flights).
  • The rocket can be mounted without fuel and transported to the launch site, it is lighter and during assembly and transport, there is no explosion or fire hazard. The filling is then only shortly before the start. However, specific to the launch pad facilities must be available for it.
  • Liquid engines can be checked from the launch pad on its function ( boost pump speed, combustion chamber pressure ) between the ignition and liftoff of the rocket.
  • The thrust is adjustable during operation.
  • Liquid rockets often use fuel more efficiently than solid rockets and achieve the same amount of fuel higher top speeds.
  • The fuel combination often used LOX/LH2 burns to water, making it hazardous to the environment.

Cons:

  • And liquid rocket engines Advance are more expensive, more complex and thus error-prone than solid rockets.
  • Through the consumption of fuels, the focus of the rocket moves. The stabilization and control system of the rocket must be able to compensate for this shift.
  • It may be the Pogoeffekt occur ( vibrations of engine power due to resonances of the liquid column in the fuel lines and the mechanical structure of the rocket ).
  • Liquid rockets are dangerous in case of leakages in terms of explosion because the liquids are easier to ignite.
  • Some fuels (including hydrazine derivatives) are toxic if released ( false starts, Rücksturz burned-out stages for earth) may cause environmental damage.
  • Cryogenic fuel components may be fueled only shortly before the start, since they evaporate by heating out prematurely, preventing responsive start-up or a longer lasting start standby. Some storable liquid propellants are highly corrosive or corrosive and engage with the time the materials of the missile structure.

Fuels

See main article: rocket fuels.

  • An example of a monergolen fuel is highly concentrated hydrogen peroxide.
  • As fuel may be used, for example, often kerosene RP -1, hydrazine and its derivatives, or liquefied hydrogen ( LH2).
  • As an oxidizer is usually liquid oxygen (LOX ), but used in hydrazine or its derivatives as fuel nitrogen tetroxide.

The most energy-rich fuel mixture, which is now applied to the liquid rocket, is cryogenic oxygen and hydrogen ( LOX/LH2 ).

Depending on the fuel mixture in use temperatures of up to 4200 ° C and pressures may occur up to more than 25 MPa in the combustion chamber.

Manufacturer (selection)

Solid-fuel rocket Robert H. Goddard Berlin rocket launching site Aggregate (rocket family) V-2 rocket JATO R-7 Semyorka Rocketdyne F-1 Space Shuttle Main Engine RD-107 Bar (unit) Pogo oscillation RL-10 Viking (rocket engine) Vinci (rocket engine) Apollo (spacecraft) Hypergolic propellant RD-0120 Centaur (rocket stage) RP-1 Hybrid rocket Brennschluss
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