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CRYOGENIC ROCKET ENGINES

The use of liquid fuel for rocket engines was first considered as early as the beginning of our century. The Russian K. E. Ziolkowsky, the American Robert H. Goddard and the German-Romanian H. Oberth worked independently on the problems of spaceflight and discovered soon that in order to succeed, rocket engines with a high mass flow were mandatory. Already then the combustion of liquid fuels was seen to be the most promising method for generating the necessary thrust.

However, it was not until some decades after these pioneers made their early attempts, when the first big liquid fuel powered rocket, the German A-4 became reality at the beginning of the forties. This rocket was ingloriously successful as the V-2 weapon. Liquid oxygen was used as oxydizer and 75 percent proof ethyl alcohol as fuel which gave the rocket more than 300 kN thrust. Its range was about 300 km.

As the development of rocket engines continued, higher thrust levels were achieved when liquid oxygen and liquid hydrocarbon were used as fuel. This allowed for the construction of the first intercontinental rocket with a range of more than 10,000 km.

The fuel combination liquid oxygen and RP-1, a kerosene-like hydrocarbon compound, was later the basis for the propulsion of the American intercontinental rockets Atlas and Titan I as well as for the boosters for the first stage of the Saturn family used in the Apollo Moon Program. The Russians also used this fuel combination for far-reaching rocket weapons and launchers. Combinations of oxygen or fluoride as oxydizer and hydrogen or methane as fuel make especially attractive fuel mixtures for rockets. They are far more effective than the above mentioned oxygen-kerosene combination. The real disadvantage, however, is their low density.

Under normal atmospheric conditions, (room temperature: 300 -oK, pressure: 1 bar) these substances are in a gaseous state. One cannot remedy the low density by increasing the pressure, because the required tank structures would end up being too heavy. The answer is to liquefy the fuels by cooling them down. This is why these kinds of fuels are also called cryogenic fuels. The resulting tank volumes, i.e. tank masses correspond with the storable fuel combinations.

Due to handling difficulties the cryogenic fuels are not considered suitable for rocket-based weapons. That is why these fuels are only used for civilian spaceflight, while missiles and weapons use mainly solid fuel or storable liquid fuels.

Among the cryogenic fuels the combination of liquid oxygen (LOX) as oxydizer and liquid hydrogen (LH2) as fuel deserves a special mention. Both components are easily and cheaply available. The fuel is environmentally friendly, non-corrosive and has the highest efficiency of all non-toxic combinations. To liquefy, hydrogen has to be cooled to a temperature of minus 253 -oC. Its boiling point is 20 -oK, only just above absolute zero on the temperature scale. During this process its density increases to above 70 kg/m3. Liquefaction of oxygen takes place at a temperature of minus 184 -oC, its density then is 1,140kg/m3.

The rocket pioneers mentioned earlier knew about the advantageous performance of this fuel combination for rocket technology. However, since there were difficulties in handling liquid hydrogen, their knowledge could not be put into practice back then. However, from the beginning liquid oxygen was used as oxydizer. Goddard and Oberth utilized it in their early trials.

In the sixties, the steadily increasing payload weights and the corresponding demand for more thrust of the launcher lead to the use of liquid hydrogen for the Centaur upper stage with RL-10 engines. The Apollo-Program followed with the J-2 engines. At the peak of this development so far is the US Space Shuttle Main Engine SSME. Europe, Japan, the former Soviet Union and finally China were to follow several years later with engines using this fuel combination.

In principle, cryogenic rocket engines generate thrust like all other rocket engines - by accelerating an impulse carrier to very high speeds.

In conventional aircraft engines the surrounding air is the main impulse carrier and fuel is the energy carrier. This is why such an engine does not only need the atmosphere to burn the fuel, but mainly to generate the required impulse - the thrust.

However, in rocket engines energy and impulse carrier are identical and are present as fuel in the launcher. This type of engine is therefore ideal for space flight. The chemical energy, which is stored in the fuel, is changed into kinetic energy by burning it in the thrust chamber and subsequent expansion of the hot gases in the nozzle. In the process the pressure, which is forced onto the nozzle wall, generates the reaction force or to be precise, the thrust.

In order to compare a variety of fuel combinations and rockets, a quantity is used which is independent from the dimension of the rocket in question. The most important quantity is the specific impulse, which determines the thrust per kilogram of emitted fuel per second. It is recorded in seconds and indicates the deciding quality factor. Unlike the fuel combinations liquid oxygen-kerosene or nitrogen tetroxide- unsymmetrical dimethyhydrazine (UDMH), which is frequently used with launch rockets and generate specific impulses between 300 s and 340 s in a vacuum, the cryogenic combination liquid oxygen-liquid hydrogen can reach up to 450s. This high specific impulse makes this fuel combination very attractive and explains why engines using this fuel are frequently used for high performance missions in spaceflight.

The major components of a liquid fuel cryogenic engine are the thrust chamber, the fuel pumps with its valves and regulators and the tanks. The fuel and oxydizer pump system is the main component and can be divided into two different principles. The most simple way is to increase the pressure of the tank with inertial gases to pressurize the tanks against the pressure in the combustion chamber. In this type of engine the fuel and gas tanks are very heavy, which explains why this design principle is only used for smaller rockets with shorter burning times. For the above mentioned liquid hydrogen and oxygen cryogenic fuel, it is not an option at all because of the hydrogen's low density and the corresponding large fuel tank size.

The alternative is to use turbopumps. These can be differentiated into a bypass or a main flow configuration. In the bypass configuration, the fuel flow is split, the main part is used via the combustion chamber to generate thrust, while a small amount of the fuel is used to drive the pump through the turbine and is subsequently emitted. In the main flow design, the entire fuel is fed through the turbines, which drive the pumps, and then further to the combustion chamber.

The main advantage of the turbopumps is that it allows the optimization of the operating conditions by controlling the pump's rpm along with the lower volume and mass requirement as compared to the pressurization version.

The combustion chamber also is a critical component of the engine, because of the high output and accordingly high pressures. High pressures of over 200 bar and temperatures of more than 3,000degrees Fahrenheit cause a great heat strain on the combustion chamber and call for effective cooling. Copper combustion chambers are used, in the outside of which cooling channels are milled that are galvanoplasticly closed. This technology was also developed very early in Germany and is used in all engines of this kind today.

The list of all hydrogen-oxygen cryogenic engines, which have been developed up until today, is relatively long. The best-known is the above mentioned SSME. This engine has a vacuum thrust of 2,090 kN and a specific vacuum impulse of 452 s. The fuel- mass stream of 470kg/s, (which is needed when the engine is operating), is supplied by two turbopumps with an output of 100,000 hp.

The three engines, which are needed for the main stage of the Space Shuttle system, have a combined output of more than 37 million hp or 27,380 Megawatts, which corresponds to about 30 conventional nuclear power plants.

The Vulcan engine of Ariane 5, the European launch rocket, is a further example of successful engine technology using cryogenic fuels. It supplies a vacuum thrust of 1,125kN and has a specific vacuum impulse of 425s. During its burn time of 570s, the engine uses a fuel mass stream of 265kg/s which is supplied by the fuel turbo pumps with an output of 21,000 hp.

So-called Aerospike Engines are a new application of cryogenic rocket engines. With aerospike engines, the combustion chamber is fitted around a cone shaped central body, the Aerospike. This is main difference to conventional rocket engines.

In traditional engines the gas stream, which is generated in the combustion chamber, is expanded within the engine in a laval nozzle and is simultaneously accelerated to several times supersonic speed. The name originates from its inventor deLaval, who developed a nozzle with a convergent and a divergent part. The gas stream is accelerated to the speed of sound in the converging part with the narrowest cross-section, while further acceleration to supersonic speeds takes place in the nozzle's divergent part. The nozzle can be bell or cone-shaped.

The problem of this conventional nozzle shape is that it is only optimized for a fixed operational point. Since the specific impulse increases with an increase of the nozzle size, the engineers are trying to realise the biggest nozzle possible. On the other hand there is the disadvantage that the performance on the ground will suffer - an overexpanded flow can cause a separation of the gases from the nozzle wall. Since the atmospheric pressure decreases constantly while the rocket is climbing but the flow conditions within the nozzle remain the same, the engine's performance during flight changes accordingly. In practise the nozzle is designed to perform best in a vacuum, and the dip in performance during launch can be minimised.

Engines with aerospike nozzles try to avoid this compromise. With aerospike engines, the hot gases are not guided along the inside of the engine nozzle, but stream along the outer surface of a nozzle ramp which has an open side to the atmosphere. This design is supposed to avoid an over expansion on the ground and allows a variety of nozzles with big area proportions and appropriate specific impulses in the vacuum. In order not to make the construction too long, the central cone is actually cut off quite short. This hardly compromises the performance. Instead of a circular combustion chamber with a narrow slot shaped nozzle neck, which is almost impossible to construct, a line of small single combustors is used. They form a kind of circular combustion chamber. Each of these single combustion chambers has to be supplied with fuel via a central turbopump.

The so called linear aerospike is a special case of this type of engine. Here the combustors are arranged in a linear configuration, enabling a better integration with the vehicle and reducing the tail drag. Engines like these were examined and tested in detail in the late sixties and early seventies but, the work was discontinued.

The new American X-33/VentureStar vehicles will be the first application for linear aerospike engines. Boeing Rocketdyne is developing and building the XRS-2200 for the X-33 and planning a full-scale RS-2200 for the future VentureStar orbiter. Both engine variants are using liquid oxygen and liquid hydrogen for the highest possible performance, placing them in the group of modern cryogenic high performance engines.

From page 92 of FLUG REVUE 6/99