Vulcain Engine Pogo

This humorously named phenomenon is nevertheless a serious problem. The pressure at the inlet to the combustion chamber should be constant for a steady flow of propellant to the combustion chamber and hence a steady thrust. As mentioned above, the acceleration of the rocket raises the pressure at the pump inlet, and it is possible to develop a feedback loop between the instantaneous thrust and the pump inlet pressure. If this happens, then a small natural fluctuation in thrust will result in a fluctuation in flow rate to the combustion chamber. The fluctuation will make itself felt, with a slight delay (the time taken for the propellant to flow from the pump to the chamber), as a further fluctuation in thrust. This, in turn, changes the inlet pressure at the pump, which causes another thrust fluctuation, and so on. The time delay is usually in the 10-ms region, and the reinforcement mechanism can result in the build-up of an oscillation in thrust with a period of about 100 Hz. This is very damaging to the rocket and the payload, as a small fluctuation in a mega-Newton of thrust is a large force. For this reason, pogo correction systems are fitted to liquid-propellant rocket engines.

The basic principle is to introduce some capacitance to the system in order to smooth out fluctuations in inlet pressure. A small sealed volume is connected to the propellant line, adjacent to the combustion chamber inlet, and is filled with propellant. It is pressurised using gas from the tank pressurisation system. If the line pressure falls momentarily, additional propellant is very quickly injected from the storage volume, to raise the pressure to its original value. If the pressure rises then some of the excess propellant flows into the sealed volume, again restoring the line pressure to normal. It is usually only necessary to fit pogo correction to one propellant line, which in most cases is the oxidant line. This system can be passive, or it can be actively controlled, to deal with, for example, the much greater pressure fluctuations which occur when an engine is shut down or is started. In such cases the pogo correction system can also protect the turbine from cavitation. For a single-use engine, damage to the turbines at shut-down is of no consequence, but for a reusable engine a turbine damaged due to 'racing' in a heavily cavitated fluid is a serious matter.

3.4 COOLING OF LIQUID-FUELLED ROCKET ENGINES

Before considering examples of actual rocket engines it is convenient to consider the cooling of the combustion chamber and nozzle. High combustion temperature produces a high exhaust velocity. A typical temperature is 3,000 K, but the melting point of most metals is below 2,000 K and so the combustion chamber and nozzle must be cooled. This is done by allowing part of the cool unburnt propellant to carry away the heat conducted and radiated to the walls of the chamber and nozzle. This can be done in a number of ways.

Technically, the simplest method is film cooling. Part of the liquid propellant is caused to flow along the inside surface of the combustion chamber and down the inside surface of the nozzle. The evaporation of this liquid film has a certain cooling effect, and results in a layer of cool gas between the wall and the hot gases passing from the chamber and through the nozzle. The cooling film is introduced through part of the injector next to the wall. This type of cooling works best with lower combustion temperatures such as are encountered in storable propellant engines. The Ariane 4 Viking engine uses film cooling, which results in the simplest configuration of the combustion chamber and nozzle. In this engine the injector is mounted on the cylindrical wall of the chamber rather than at the end, and it is therefore simple to inject part of the UDMH parallel to the wall. This method is suitable for cooling the combustion chamber and throat, because the efficiency of cooling decreases with distance from the injector. The nozzle is less well cooled and may glow red hot, cooling by radiation. The use of a refractory cobalt alloy enables the nozzle to retain its structural strength at this high temperature.

Cryogenic propellants, liquid hydrogen and liquid oxygen generate much higher combustion temperatures, and the cold liquid lends itself to efficient cooling. In such cases the walls of the combustion chamber and the nozzle are made hollow, and one of the propellants—usually hydrogen—is passed through the cavity. This cools the chamber and nozzle walls effectively, at the expense of additional complication and cost in construction. The gas resulting from the waste heat carried away from the walls can be used in various ways. The simplest approach is to exhaust the gas through many small nozzles round the rim of the main nozzle, to generate a little additional thrust. This is called dump cooling, and it can be used to cool the long nozzle of an engine designed for use in a vacuum, where it may be inconvenient to pipe the gas back into the top of the engine. As mentioned above, the gas may also be used to drive the turbine or to pressurise the propellant tanks. The most efficient way of using this gas is to feed it back into the combustion chamber and burn it to contribute to the main thrust. This has two advantages: the chemical energy of the gas—part of the propellant load of the rocket—is not wasted, and the waste heat conducted and radiated out of the combustion chamber is returned to the main combustion. This latter point is very important. Fundamental thermodynamics tells us that extraction of energy from a hot gas depends on the temperature difference between the source and the sink. After cooling the walls, the temperature of the propellant is far below that in the combustion chamber, so not much energy or thrust can be extracted from it. On the other hand, if it is passed into the combustion chamber and heated to the combustion temperature, then much more of the energy acquired during cooling is released. This technique is called regenerative cooling, and results in the most efficient engines. Of course, it leads to further complications and results in a heavier engine, and as always there must be a correct balance between extra thrust and extra weight.

If hot spots on the chamber and nozzle walls are to be avoided, the propellant must be in contact with the wall everywhere, and the flow must be smooth and continuous. Moreover, there is a large quantity of heat to carry away. Most engines therefore have the nozzle and lower part of the combustion chamber made from metal tubes welded together, wall to wall, to form a continuous surface. The propellant flows through this multiplicity of tubes freely and is, at the same time, constrained to cover the entire inner wall. In some cases the tubes are parallel to the axis of the thrust chamber, and in others a spiral form is used to produce a longer flow path. The two may be combined, with the spiral form being used on the nozzle and the axial form of the combustion chamber. The design of such a complicated structure is very demanding both on the materials and on the function. The operating temperature and pressure are very high, and any interruption of the flow during operation would be fatal. Nevertheless the advantage to be gained in terms of exhaust velocity is significant. The Saturn V engine developed an exhaust velocity of around 4,200 ms"1, while the SSME develops a velocity of 4,550 ms"1.

As mentioned above, these apparently small gains have a major impact on the performance of the rocket, in terms of payload and achievable velocity increment.

3.5 EXAMPLES OF ROCKET ENGINE PROPELLANT FLOW

For most modern launchers, gas-pressure-fed systems are not sufficiently powerful for use in first or second stages. This is just a matter of the required thrust, as pressure-fed systems cannot deliver propellant at a very high flow rate without prohibitively high tank pressures. Pressure-fed systems are advantageous for upper stages, because the reduction in weight helps to produce a high mass ratio, and the thrust and propellant flow requirements are less demanding. Before considering examples of gas generator and turbo-pump systems, a modern pressure-fed system used on the Ariane 5 upper stage will be described.

3.5.1 The Aestus engine on Ariane 5

This is the restartable engine used on the upper stage of the Ariane 5 rocket (Figure 3.5). The propellants are hypergolic: monomethyl-hydrazine (MMH) and nitrogen tetroxide. Both of these are liquid at normal temperature and pressure (NTP) and can be stored safely. Ignition of the rocket results simply from the chemical reaction that occurs spontaneously when the propellants meet in the combustion chamber. The propellant delivery scheme is shown in Plate 1.

There is a single combustion chamber gimballed to allow ±6° of thrust vector control through two actuators. The nozzle is bell-shaped with an expansion ratio of 30 to develop an exhaust velocity of 3,240 ms-1 in vacuo; as an upper stage, it operates only in vacuo. Regenerative cooling is employed for the combustion chamber walls and the inboard part of the nozzle, for which the MMH is used. It flows from the tank into the lower part of the hollow walls, and having extracted heat it enters the combustion chamber through the injector. This is a multi-element coaxial injector with which the swirl technique is used to mix the MMH with the nitrogen tetroxide. While the combustion chamber and the inboard part of the nozzle are regeneratively cooled the nozzle extension is not; it is allowed to glow redhot in use, dissipating heat by radiation.

There are two fuel tanks and two oxidiser tanks. The fuel (MMH) tanks are spherical, while the oxidiser tanks are slightly elongated, reflecting the differing volumes of fuel and oxidant. The oxidant-fuel ratio is 2.05. Both types of tank are made of aluminium alloy. The spherical shape uses the minimum volume of aluminium to contain the propellant, and also produces the minimum wall thickness to safely contain a given pressure. Thus the propellant tanks are optimised for a pressure delivery system. This can be employed for an upper stage in which the quantity of propellant is relatively modest, but the huge amounts of propellant needed for the first stage cannot be contained in spherical tanks. This approach for an upper stage also minimises the length, and hence the structural mass required. The tanks are pressurised with helium from a pair of high-pressure tanks;

Figure 3.5. The Aestus engine on Ariane 5. The high-expansion ratio nozzle and two of the four propellant tanks can be seen. Upper-stage engines should be short to reduce the overall length of the vehicle; here the propellant tanks cluster round the engine.

Courtesy ESA.

Figure 3.5. The Aestus engine on Ariane 5. The high-expansion ratio nozzle and two of the four propellant tanks can be seen. Upper-stage engines should be short to reduce the overall length of the vehicle; here the propellant tanks cluster round the engine.

Courtesy ESA.

the gas pressure being moderated by a reducing valve to around 18 bar to pressurise the propellant tanks. The propellant is delivered to the engine at 17.8 bar, and the combustion chamber itself operates at 11 bar. There is a considerable pressure reduction across the injector. The passive anti-pogo system is fitted to the oxidant line.

Before the engine is started, the system is purged with helium to remove propellant residues from test firings. The oxidiser valve is then opened, followed, after a short delay, by the fuel valve. The full thrust of 29 kilo-Newton is developed 0.3 seconds after the start signal. Shutdown is initiated by closing the MMH valve, followed shortly by the closure of the oxidiser valve. The engine is then purged with helium to prepare it for the restart. The total burn time of the engine is 1,100 seconds, and the vacuum exhaust velocity is 3,240 ms-1. This engine has been used successfully for the upper stage of Ariane 5 since 1999. The restart capability has been demonstrated for an improved range of orbit options. A pump-fed version has been tested for higher thrust applications (Figure 3.6).

3.5.2 The Ariane Viking engines

This series of rocket engines is used to power the first and second stages of the Ariane 4 launch vehicle. There are three variants. The short nozzle version—Viking 5C—is used in groups of four to power the first stage; the Viking 6—more or less identical to

Figure 3.6. The pump-fed variant Aestus engine firing. In this test the long nozzle extension has been removed. Courtesy ESA.

the 4C—is used for the strap-on boosters; and the Viking 4B powers the second stage and has a long nozzle to produce greater efficiency at high altitude. There is little difference in the propellant delivery systems. The general scheme is shown in Plate 2.

The Viking engine uses the storable hypergolic propellants nitrogen tetroxide and UDMH25 (unsymmetrical dimethyl hydrazine with a 25% admixture of hydrazine hydrate). There is no ignition system because the propellants ignite on contact, which, as mentioned before, is convenient for restartable engines and is also a very reliable system even when the engine is not restartable. In addition to the tanks of propellant, water is also carried to act as a combustion coolant, and high-pressure nitrogen to operate the valves. (High flow rates demand large-diameter pipes and large valves, which are difficult to operate purely electrically). There are two valves to control the flow of the individual propellants to the turbo-pump. This is a single turbine, developing 2,500 kW at 10,000 rpm and driving two pumps on the same shaft; the different flow rates are accommodated by having different sized pumps. A separate pump, driven through a reduction gear, distributes the water. Part of the propellant flow (about 0.5%) is diverted to the gas generator, where the propellants react to produce the hot gas which powers the turbine. Water is injected to cool the combustion products. The hot gases pass to the turbine and then to the turbine exhaust, which is nozzle shaped to add to the thrust. Part of the hot gas is diverted to pressurise the propellant tanks. This static pressure is quite high—about 6 bar—and is enough to prevent cavitation at the pump blades with these room-temperature liquids.

The thrust is stabilised by two control loops. One controls the temperature of the hot gases from the gas generator by varying the amount of injected water, and the other uses the combustion chamber pressure to control the flow of propellant into the gas generator and thus the turbo-pump speed. In this way the thrust is kept constant. A third balancing system controls the relative pressures of the two propellants at the injector to keep the mixture ratio correct. The pogo corrector is a small cylindrical chamber surrounding the main oxidiser pipe and linked to it by small holes; it is pressurised from the nitrogen supply used to operate the valves. The combustion chamber and nozzle are cooled with a film of UDMH from the lower part of the injector. The 5C develops 678 kN of thrust at sea level, with an exhaust velocity of 2,780 ms"'. The high altitude 4B variant develops 805 kN of thrust with a higher exhaust velocity of 2,950ms"1.

This engine has been used successfully for the upper stage of Ariane 5 since 1999. The restart capability has been demonstrated for an improved range of orbit options. A pump-fed version has been tested for higher thrust applications.

3.5.3 The Ariane HM7 B engine

The HM7 B liquid hydrogen-liquid oxygen engine is used to power the third stage of the Ariane 4 series of launchers and a version is presently used as a cryogenic upper stage for the Ariane 5 while the Vinci engine is being developed. The schematic is shown in Plate 3. It uses a single gas generator and turbine driving two pumps on different shafts. The high-speed pump driven directly by the turbine at 60,000 rpm delivers the liquid hydrogen at 55 bar, while the low speed pump driven through a gear chain at 13,000 rpm delivers the liquid oxygen. The static pressure in the gas lines is raised by coaxial impellers to a level sufficient to prevent cavitation The gas from the turbine is exhausted through a shaped nozzle to generate additional thrust. The nozzle throat and combustion chamber are cooled regeneratively by passing most of the hydrogen through 128 axial tubes forming the wall, before it enters the combustion chamber itself. The rest of the nozzle is dump cooled by routing a fraction of the hydrogen through 242 spiral tubes and then through micro-nozzles at the end of the main nozzle. The gas generator is fed a hydrogen rich mixture, which keeps the temperature down and reduces oxidation of the turbine blades. The gas generation rate—and therefore the propellant flow rate—is stabilised by controlling the oxygen flow into the gas generator. The valves which control the flow of propellant are operated by helium at high pressure, switched by electromagnetic valves. A pogo corrector is fitted to the liquid oxygen line, pressurised by helium.

A particular requirement of cryogenic engines is to purge the system before ignition, and to deal with the boil-off of the cryogenic propellants. Neither liquid oxygen nor liquid hydrogen can remain liquid under achievable pressures, and so the tanks have to vent continuously to the atmosphere until a few minutes before launch. The need for purging is twofold. Firstly, all the components—the valves, pumps and combustion chambers—need to be brought down to the temperature of the propellants to avoid localised boiling of the cryogen. This would generate back pressure and interrupt flow. Secondly, the entire system must be freed of atmospheric gases which would freeze and block the system on coming into contact with the cryogenic liquids. For this reason, purging valves are provided to enable a free flow of cold gas from the boiling cryogens through the system before the main valves are opened.

The pre-launch sequence includes the chilling and purging of the system. The gas generator is then started—in this case by a pyrotechnic igniter. When the turbines are delivering full power, the main propellant valves are opened and the main combustion chamber is started by another pyrotechnic igniter.

3.5.4 The Vinci cryogenic upper-stage engine for Ariane 5

Further increase in the payload mass to geostationary transfer orbit (GTO) with the Ariane 5 makes use of a cryogenic upper stage, to replace the Aestus storable propellant engine. This engine, called Vinci will power the upper stage from 2006 (Figure 3.7). For an upper stage, mass ratio is very important, and the system does not use a gas generator to power the turbo-pumps, instead the turbines are driven by hot hydrogen emerging from the cooling channels of the combustion chamber and upper nozzle. This expander cycle can be used when the propellant delivery rates and chamber pressure are not too high. The two turbines are connected, in series, on the hot hydrogen line, the gas being routed first to the hydrogen turbine. On emerging from the oxygen turbine, the gas enters the combustion chamber; all the hydrogen follows this route while the oxygen is delivered in liquid form to the combustion chamber straight from the turbo-pump. The exhaust velocity is 4,650m/s, thanks to this efficient regenerative cooling and an expansion ratio of 240 (achieved by a deployable nozzle extension). The thrust is 180 kN. These values are to be compared with the thrust of the Aestus, 29 kN, and its exhaust velocity, 3,240 m/s. This is an example of the modern trend to reduce the complexity of rocket engines, and to address all the factors that make the vehicle efficient. This engine only weighs 550 kg, which helps to keep the mass ratio of the upper stage high.

3.5.5 The Ariane 5 Vulcain cryogenic engine

The Vulcain cryogenic engine used for the main propulsion stage of Ariane 5 develops 1.13 MN of thrust and operates at 110 bar combustion chamber pressure. It is similar to the HM7 B in design, but uses full regenerative cooling of the combustion chamber and nozzle. The single gas generator drives two separate turbo-pumps, with nozzle exhausts. The propellants enter through 516 coaxial

Figure 3.7. The Vinci cryogenic upper-stage engine. Note the very long nozzle extension to give the high exhaust velocity; it is deployed after separation of the main stage. This new cryogenic engine is specified for the updated Ariane 5, which will have a 10 tonne capacity to GTO.

Courtesy ESA.

Figure 3.7. The Vinci cryogenic upper-stage engine. Note the very long nozzle extension to give the high exhaust velocity; it is deployed after separation of the main stage. This new cryogenic engine is specified for the updated Ariane 5, which will have a 10 tonne capacity to GTO.

Courtesy ESA.

injectors and generate an exhaust velocity of 4,300 ms"1. The schematic is shown in Plate 5.

The propellants are stored in a cylindrical tank 24 metres long, which also provides the main structural element of the stage. Combining the functions of fuel tank and rocket structure reduces the dead weight. The 25.5 tonnes of liquid hydrogen occupies most of the volume of the tank, the 130 tonnes of oxygen being stored in the upper portion, separated by a hemispherical bulkhead. The density of liquid oxygen is much higher than that of liquid hydrogen. The hydrogen tank is pressurised by gaseous hydrogen produced by the regenerative cooling circuit—that is, heated by the combustion chamber. The oxygen tank is pressurised by helium stored in a spherical tank containing 140 kg of liquid helium. The helium is heated by the turbo-pump exhaust. A separate gaseous helium supply is used to operate the propellant valves and the pogo corrector, and to pressurise the liquid helium tank. This is stored in separate spherical tanks.

The gas generator and the combustion chamber are both fitted with pyrotechnic igniters. A separate solid propellant cartridge provides the gas pressure to start the turbo-pumps. The hydrogen and oxygen then enter the gas generator and the combustion chamber and are ignited. The engine is started 8 seconds before firing the boosters. This allows it to be checked out before the irrevocable booster ignition. The engine is stopped by closing the propellant valves.

The Vulcain 2 engine (Figure 3.8) specified for Ariane 5 launchers after 2002, to give an additional tonne of payload into GTO, is an updated version of the Vulcain engine used before 2002. The new engine incorporates a number of improvements, the most notable being an increase of 10% in the mass of propellant available, as a result of changing the fuel-oxidiser ratio of the engine in favour of more oxygen; the ratio was changed from 5.3 to 6.15. Because of the higher density of liquid oxygen, this can be accomplished without increasing the total volume of the propellant tanks. More oxygen increases both the mass ratio and the thrust. Normally, this would be expected to decrease the exhaust velocity because the mean molecular weight of the

Figure 3.8. The Vulcain 2 under test. This is the new version of the Vulcain specified for Ariane 5. It uses a more oxygen-rich mixture to improve the mass ratio, and a longer nozzle to restore the exhaust velocity. Courtesy ESA.

exhaust increases, however other improvements mitigate this effect and in fact the exhaust velocity is some 30m/s faster. The exhaust velocity is maintained by a higher expansion ratio—60 compared with 45. The cooling of the longer nozzle is accomplished by routing the turbo-pump exhaust into the nozzle extension to create a film of cooler gas, protecting the walls from the hot exhaust. The quantity of oxygen carried is increased by 23% and a re-designed two-stage turbopump for the oxygen line gives a 40% higher delivery rate. This combined with an increase in throat area gives a higher thrust of 1,350 kN, compared with 1,140 kN for the Vulcain.

3.5.6 The Space Shuttle main engine

The SSME uses the same cryogenic propellants as the Ariane engines, but is different in concept. It is intended to be reused many times, and to be highly efficient. It uses the staged combustion system to drive the turbo-pumps, and has full regenerative cooling. The vacuum exhaust velocity is 4,550 ms"', and the thrust is controllable from 67% to 109% of nominal. The propellant distribution system is shown in Plate 4.

The propellants are stored in the external tank. The hydrogen tank is pressurised by gas from the regenerative cooling of the combustion chamber, and the oxygen tank by gas resulting from regenerative cooling of the oxidiser gas generator. The propellants are delivered to the combustion chamber by separate turbo-pumps, with individual gas generators. These are called 'pre-burners' because the exhaust from the turbo-pumps passes to the combustion chamber for further burning. The propellants are raised from tank pressure to combustion chamber pressure in two stages, using separate low-pressure and high-pressure turbo-pumps.

The most important aspect of the SSME design, for our purposes, is the fact that all the exhaust from the fuel delivery system passes into the combustion chamber so that all the energy stored in the exhaust contributes to the thrust. This recovery of energy is much more efficient if enabled at high temperature in the combustion chamber than by venting the gas at the turbine exhaust temperature as in, for example, the Vulcain engine. Since the propellant flow is rather complicated, we shall examine each propellant system in turn.

The unique aspect of the SSME is that nearly all of the hydrogen from the fuel tank passes through the pre-burners or gas generators, and only a small fraction passes directly to the main combustion chamber after driving the low-pressure fuel pump; as the exhaust from the pre-burners will eventually enter the combustion chamber, this does not matter. It has the further advantage that a fuel-rich mixture— to keep the pre-burner exhaust temperature low enough for the turbine blades—is automatically achieved.

Liquid hydrogen arrives at the inlet of the low-pressure pump at the static pressure of about 2 bar. The pump raises this to 18 bar. It is powered by hot hydrogen gas emerging from the cooling channels in the combustion chamber. The liquid hydrogen is then pressurised to 440 bar by the high-pressure turbo-pump. It then follows three separate paths. Part of the flow enters the cooling channels in the

Figure 3.9. The SSME on a test stand. Note the long bell-shaped nozzle to extract the maximum exhaust velocity from the hot gas, and the complexity of the propellant feed system above. The hydrogen turbo-pump is visible on the left and the (smaller) oxygen pump on the right of the engine. The pipes to feed liquid hydrogen into the cooling channels of the nozzle are visible. Courtesy NASA.

Figure 3.9. The SSME on a test stand. Note the long bell-shaped nozzle to extract the maximum exhaust velocity from the hot gas, and the complexity of the propellant feed system above. The hydrogen turbo-pump is visible on the left and the (smaller) oxygen pump on the right of the engine. The pipes to feed liquid hydrogen into the cooling channels of the nozzle are visible. Courtesy NASA.

combustion chamber and emerges as hot gas, which is routed to the low-pressure fuel pump turbine to drive it. Emerging (now cool) from the turbine some of it goes to pressurise the fuel tank, and the rest cools the hot gas manifold before entering the combustion chamber. The second path passes through the cooling channels of the nozzle (Figure 3.9) before joining the third path, which routes most of the hydrogen to both of the pre-burners.

The exhaust from the pre-burners is effectively hydrogen-rich steam, at quite a high temperature (850 K). This is the fuel supply for the main combustion chamber. Consequently the hydrogen 'injector' is handling hot gas rather than cold liquid, and is called the 'hot gas manifold'. This takes the exhaust from both turbo-pumps and feeds it into the combustion chamber, where it burns with the liquid oxygen, to generate a thrust of 2 MN.

All of the liquid hydrogen is routed through the chamber or nozzle cooling channels, and afterwards becomes gaseous. In contrast, most of the oxygen remains in the liquid state right up to the combustion chamber injector. The static pressure in the oxygen tank is higher than in the hydrogen tank—about 6 bar—and the low-pressure oxygen turbo-pump raises this pressure to about 30 bar for the inlet to the high-pressure oxygen turbo-pump. After this turbo-pump the pressure of the liquid oxygen is 300 bar. The flow now divides into four separate paths. The first path carries some of the liquid oxygen to the low-pressure turbo-pump to drive the turbine, and on leaving the turbine it re-enters the main flow to the high-pressure turbo-pump. In the second path the liquid oxygen cools the high-pressure pre-burner and is converted into gas, which is used to pressurise the main oxygen tank and the pogo corrector. The third path carries most of the oxygen to the main combustion chamber injector. The fourth and final path takes liquid oxygen to an additional turbo-pump attached to the main pump shaft, which boosts the pressure to 500 bar for injection into the two pre-burners. This oxygen is burned with part of the hydrogen and forms the hot steam in the pre-burner exhaust, which then enters the combustion chamber. These routes can be followed in Plate 4. The thrust and the mixture ratio are controlled by the fuel and oxidant pre-burner valves which regulate the flow of oxygen to the pre-burners, and hence the turbine speed. Since the mixture is fuel rich, it is only necessary to vary the oxygen flow to the pre-burners to control the speed.

As with all highly developed devices the SSME (Plates 8 and 9) seems complicated in its propellant distribution. The main aim is, however, simple: to run each element of the system at its maximum efficiency, and then convert all the energy released from the burnt propellants into thrust, at a high exhaust velocity. In previous chapters we have seen that high exhaust velocity is the ultimate determinant of the success of a rocket as a launcher.

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