Cryogenic rocket engines report

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  • 1. Semester VII Branch: Mechanical Engineering Seminar Title: Cryogenic Rocket Engines INTRODUCTION Cryogenics originated from two Greek words “kyros” which means cold orfreezing and “genes” which means born or produced. Cryogenics is the study of verylow temperatures or the production of the same. Liquefied gases like liquid nitrogen andliquid oxygen are used in many cryogenic applications. Liquid nitrogen is the mostcommonly used element in cryogenics and is legally purchasable around the world.Liquid helium is also commonly used and allows for the lowest temperatures to bereached. These gases can be stored on large tanks called Dewar tanks, named afterJames Dewar, who first liquefied hydrogen, or in giant tanks used for commercialapplications. The field of cryogenics advanced when during world war two, when metals werefrozen to low temperatures showed more wear resistance. In 1966, a company wasformed, called Cyro-Tech, which experimented with the possibility of using cryogenictempering instead of Heat Treating, for increasing the life of metal tools. The theory wasbased on the existing theory of heat treating, which was lowering the temperatures toroom temperatures from high temperatures and supposing that further descent wouldallow more strength for further strength increase. Unfortunately for the newly-bornindustry the results were unstable as the components sometimes experienced thermalshock when cooled too fast. Luckily with the use of applied research and the with thearrival of the modern computer this field has improved significantly, creating more stableresults. Another use of cryogenics is cryogenic fuels. Cryogenic fuels, mainly oxygen andnitrogen have been used as rocket fuels. The Indian Space Research Organization(ISRO) is set to flight-test the indigenously developed cryogenic engine by early 2006,after the engine passed a 1000 second endurance test in 2003. It will form the finalstage of the GSLV for putting it into orbit 36,000 km from earth. Cryogenic Engines are rocket motors designed for liquid fuels that have to beheld at very low "cryogenic" temperatures to be liquid - they would otherwise be gas atnormal temperatures. 1 Toc H Institute Of Science And Technology, Arakkunnam, Ernakulam
  • 2. Semester VII Branch: Mechanical Engineering Seminar Title: Cryogenic Rocket Engines The engine components are also cooled so the fuel doesnt boil to a gas in thelines that feed the engine. The thrust comes from the rapid expansion from liquid to gaswith the gas emerging from the motor at very high speed. The energy needed to heatthe fuels comes from burning them, once they are gasses. Cryogenic engines are thehighest performing rocket motors. One disadvantage is that the fuel tanks tend to bebulky and require heavy insulation to store the propellant. Their high fuel efficiency,however, outweighs this disadvantage. The Space Shuttles main engines used for liftoff are cryogenic engines. TheShuttles smaller thrusters for orbital maneuvering use non-cryogenic hypergolic fuels,which are compact and are stored at warm temperatures. Currently, only the UnitedStates, Russia, China, France, Japan and India have mastered cryogenic rockettechnology. All the current Rockets run on Liquid-propellant rockets. The first operationalcryogenic rocket engine was the 1961 NASA design the RL-10 LOX LH2 rocket engine,which was used in the Saturn 1 rocket employed in the early stages of the Apollo moonlanding program. The major components of a cryogenic rocket engine are:  the thrust chamber or combustion chamber  pyrotechnic igniter  fuel injector  fuel turbo-pumps  gas turbine  cryo valves  Regulators  The fuel tanks  rocket engine  nozzle Among them, the combustion chamber & the nozzle are the main components ofthe rocket engine. 2 Toc H Institute Of Science And Technology, Arakkunnam, Ernakulam
  • 3. Semester VII Branch: Mechanical Engineering Seminar Title: Cryogenic Rocket Engines HISTORY The only known claim to liquid propellant rocket engine experiments in thenineteenth century was made by a Peruvian scientist named Pedro Paulet. However, hedid not immediately publish his work. In 1927 he wrote a letter to a newspaper in Lima,claiming he had experimented with a liquid rocket engine while he was a student inParis three decades earlier. Historians of early rocketry experiments, among them Max Valier and Willy Ley,have given differing amounts of credence to Paulets report. Paulet described laboratorytests of liquid rocket engines, but did not claim to have flown a liquid rocket. The first flight of a vehicle powered by a liquid-rocket took place on March 16,1926 at Auburn, Massachusetts, when American professor Robert H. Goddardlaunched a rocket which used liquid oxygen and gasoline as propellants. The rocket,which was dubbed "Nell", rose just 41 feet during a 2.5-second flight that ended in acabbage field, but it was an important demonstration that liquid rockets were possible. 3 Toc H Institute Of Science And Technology, Arakkunnam, Ernakulam
  • 4. Semester VII Branch: Mechanical Engineering Seminar Title: Cryogenic Rocket Engines SPACE PROPULSION SYSTEM Spacecraft propulsion is any method used to accelerate spacecraft and artificialsatellites. There are many different methods. Each method has drawbacks andadvantages, and spacecraft propulsion is an active area of research. However, mostspacecraft today are propelled by forcing a gas from the back/rear of the vehicle at veryhigh speed through a supersonic de Laval nozzle. This sort of engine is called a rocketengine. All current spacecraft use chemical rockets (bipropellant or solid-fuel) for launch,though some have used air-breathing engines on their first stage. Most satellites havesimple reliable chemical thrusters. Soviet bloc satellites have used electric propulsionfor decades, and newer Western geo-orbiting spacecraft are starting to use them fornorth-south station keeping. Interplanetary vehicles mostly use chemical rockets aswell, although a few have used ion thrusters to great success.Classification of Space Propulsion System 4 Toc H Institute Of Science And Technology, Arakkunnam, Ernakulam
  • 5. Semester VII Branch: Mechanical Engineering Seminar Title: Cryogenic Rocket Engines ROCKET ENGINE POWER CYCLESGas pressure feed system A simple pressurized feed system is shown schematically below. It consists of ahigh-pressure gas tank, a gas starting valve, a pressure regulator, propellant tanks,propellant valves, and feed lines. Additional components, such as filling and drainingprovisions, check valves, filters, flexible elastic bladders for separating the liquid fromthe pressurizing gas, and pressure sensors or gauges, are also often incorporated. Afterall tanks are filled, the high-pressure gas valve is remotely actuated and admits gasthrough the pressure regulator at a constant pressure to the propellant tanks. The checkvalves prevent mixing of the oxidizer with the fuel when the unit is not in an rightposition. The propellants are fed to the thrust chamber by opening valves. When thepropellants are completely consumed, the pressurizing gas can also scavenge andclean lines and valves of much of the liquid propellant residue. The variations in thissystem, such as the combination of several valves into one or the elimination andaddition of certain components, depend to a large extent on the application. If a unit isto be used over and over, such as space-maneuver rocket, it will include severaladditional features such as, possibly, a thrust-regulating device and a tank level gauge. 5 Toc H Institute Of Science And Technology, Arakkunnam, Ernakulam
  • 6. Semester VII Branch: Mechanical Engineering Seminar Title: Cryogenic Rocket EnginesGas-Generator Cycle The gas-generator cycle taps off a small amount of fuel and oxidizer from themain flow to feed a burner called a gas generator. The hot gas from this generatorpasses through a turbine to generate power for the pumps that send propellants to thecombustion chamber. The hot gas is then either dumped overboard or sent into themain nozzle downstream. Increasing the flow of propellants into the gas generatorincreases the speed of the turbine, which increases the flow of propellants into the maincombustion chamber (and hence, the amount of thrust produced). The gas generatormust burn propellants at a less-than-optimal mixture ratio to keep the temperature lowfor the turbine blades. Thus, the cycle is appropriate for moderate power requirementsbut not high-power systems, which would have to divert a large portion of the main flowto the less efficient gas-generator flow. 6 Toc H Institute Of Science And Technology, Arakkunnam, Ernakulam
  • 7. Semester VII Branch: Mechanical Engineering Seminar Title: Cryogenic Rocket EnginesStaged Combustion Cycle In a staged combustion cycle, the propellants are burned in stages. Like the gas-generator cycle, this cycle also has a burner, called a preburner, to generate gas for aturbine. The pre-burner taps off and burn a small amount of one propellant and a largeamount of the other, producing an oxidizer-rich or fuel-rich hot gas mixture that is mostlyunburned vaporized propellant. This hot gas is then passed through the turbine, injectedinto the main chamber, and burned again with the remaining propellants. Theadvantage over the gas-generator cycle is that all of the propellants are burned at theoptimal mixture ratio in the main chamber and no flow is dumped overboard. The stagedcombustion cycle is often used for high-power applications. The higher the chamberpressure, the smaller and lighter the engine can be to produce the same thrust.Development cost for this cycle is higher because the high pressures complicate thedevelopment process. 7 Toc H Institute Of Science And Technology, Arakkunnam, Ernakulam
  • 8. Semester VII Branch: Mechanical Engineering Seminar Title: Cryogenic Rocket Engines COMBUSTION IN THRUST CHAMBER The thrust chamber is the key subassembly of a rocket engine. Here the liquidpropellants are metered, injected, atomized, vaporized, mixed, and burned to form hotreaction gas products, which in turn are accelerated and ejected at high velocity. Arocket thrust chamber assembly has an injector, a combustion chamber, a supersonicnozzle, and mounting provisions. All have to withstand the extreme heat of combustionand the various forces, including the transmission of the thrust force to the vehicle.There also is an ignition system if non-spontaneously ignitable propellants are used. 8 Toc H Institute Of Science And Technology, Arakkunnam, Ernakulam
  • 9. Semester VII Branch: Mechanical Engineering Seminar Title: Cryogenic Rocket Engines FUEL INJECTION The functions of the injector are similar to those of a carburetor of an internalcombustion engine. The injector has to introduce and meter the flow of liquid propellantsto the combustion chamber, cause the liquids to be broken up into small droplets (aprocess called atomization), and distribute and mix the propellants in such a mannerthat a correctly proportioned mixture of fuel and oxidizer will result, with uniformpropellant mass flow and composition over the chamber cross section. This has beenaccomplished with different types of injector designs and elements. The injection hole pattern on the face of the injector is closely related to theinternal manifolds or feed passages within the injector. These provide for the distributionof the propellant from the injector inlet to all the injection holes. A large complexmanifold volume allows low passage velocities and good distribution of flow over thecross section of the chamber. A small manifold volume allows for a lighter weightinjector and reduces the amount of "dribble" flow after the main valves are shut. Thehigher passage velocities cause a more uneven flow through different identical injectionholes and thus a poorer distribution and wider local gas composition variation. Dribbling results in afterburning, which is an inefficient irregular combustion thatgives a little "cutoff" thrust after valve closing. For applications with very accurateterminal vehicle velocity requirements, the cutoff impulse has to be very small andreproducible and often valves are built into the injector to minimize passage volume. Impinging-stream-type, multiple-hole injectors are commonly used with oxygen-hydrocarbon and storable propellants. For unlike doublet patterns the propellants areinjected through a number of separate small holes in such a manner that the fuel andoxidizer streams impinge upon each other. Impingement forms thin liquid fans and aidsatomization of the liquids into droplets, also aiding distribution. The two liquid streamsthen form a fan which breaks up into droplets. 9 Toc H Institute Of Science And Technology, Arakkunnam, Ernakulam
  • 10. Semester VII Branch: Mechanical Engineering Seminar Title: Cryogenic Rocket Engines Unlike doublets work best when the hole size (more exactly, the volume flow) ofthe fuel is about equal to that of the oxidizer and the ignition delay is long enough toallow the formation of fans. For uneven volume flow the triplet pattern seems to be moreeffective. The non-impinging or shower head injector employs non-impinging streams ofpropellant usually emerging normal to the face of the injector. It relies on turbulence anddiffusion to achieve mixing. The German World War II V-2 rocket used this type ofinjector. This type is now not used, because it requires a large chamber volume forgood combustion. Sheet or spray-type injectors give cylindrical, conical, or other types of spraysheets; these sprays generally intersect and thereby promote mixing and atomization.By varying the width of the sheet (through an axially moveable sleeve) it is possible tothrottle the propellant flow over a wide range without excessive reduction in injector 10 Toc H Institute Of Science And Technology, Arakkunnam, Ernakulam
  • 11. Semester VII Branch: Mechanical Engineering Seminar Title: Cryogenic Rocket Enginespressure drop. This type of variable area concentric tube injector was used on thedescent engine of the Lunar Excursion Module and throttled over a 10:1 range of flowwith only a very small change in mixture ratio. The coaxial hollow post injector has been used for liquid oxygen and gaseoushydrogen injectors by most domestic and foreign rocket designers. It works well whenthe liquid hydrogen has absorbed heat from cooling jackets and has been gasified. Thisgasified hydrogen flows at high speed (typically 330 m/sec or 1000 ft/sec); the liquidoxygen flows far more slowly (usually at less than 33 m/sec or 100 ft/sec) and thedifferential velocity causes a shear action, which helps to break up the oxygen streaminto small droplets. The injector has a multiplicity of these coaxial posts on its face. 11 Toc H Institute Of Science And Technology, Arakkunnam, Ernakulam
  • 12. Semester VII Branch: Mechanical Engineering Seminar Title: Cryogenic Rocket Engines PHASES OF COMBUSTION IN THRUST CHAMBERRapid Combustion Zone In this zone intensive and rapid chemical reactions occur at increasingly highertemperature; any remaining liquid droplets are vaporized by convective heating and gaspockets of fuel-rich and fuel-lean gases are mixed. The mixing is aided by localturbulence and diffusion of the gas species. The further breakdown of the propellantchemicals into intermediate fractions and smaller, simpler chemicals and the oxidationof fuel fractions occur rapidly in this zone. The rate of heat release increases greatlyand this causes the specific volume of the gas mixture to increase and the local axialvelocity to increase by a factor of 100 or more. The rapid expansion of the heated gases also forces a series of local transversegas flows from hot high-burning-rate sites to colder low-burning-rate sites. The liquiddroplets that may still persist in the upstream portion of this zone do not follow the gasflow quickly and are difficult to move in a transverse direction. Therefore, zones of fuel-rich or oxidizer-rich gases will persist according to the orifice spray pattern in theupstream injection zone. The gas composition and mixture ratio across the chambersection become more uniform as the gases move through this zone, but the mixturenever becomes truly uniform. As the reaction product gases are accelerated, they become hotter (due tofurther heat releases) and the lateral velocities become relatively small compared to theincreasing axial velocities. The combustion process is not a steady flow process. Somepeople believe that the combustion is locally so intense that it approches localizedexplosions that create a series of shock waves. When observing any one specificlocation in the chamber, one finds that there are rapid fluctuations in pressure,temperature, density, mixture ratio, and radiation emissions with time. 12 Toc H Institute Of Science And Technology, Arakkunnam, Ernakulam
  • 13. Semester VII Branch: Mechanical Engineering Seminar Title: Cryogenic Rocket EnginesInjection/Atomization Zone Two different liquids are injected with storable propellants and with liquidoxygen/hydrocarbon combinations. They are injected through orifices at velocitiestypically between 7 and 60 m/sec or about 20 to 200 ft/sec. The injector design has aprofound influence on the combustion behavior and some seemingly minor designchanges can have a major effect on instability. The pattern, sizes, number, distribution,and types of orifices influence the combustion behavior, as do the pressure drop,manifold geometry, or surface roughness in the injection orifice walls. The individual jets, streams, or sheets break up into droplets by impingement ofone jet with another (or with a surface), by the inherent instabilities of liquid sprays, orby the interaction with gases at a different velocity and temperature. In this first zone theliquids are atomized into a large number of small droplets. Heat is transferred to thedroplets by radiation from the very hot rapid combustion zone and by convection frommoderately hot gases in the first zone. The droplets evaporate and create local regionsrich either in fuel vapor or oxidizer vapor. This first zone is heterogeneous; it contains liquids and vaporized propellant aswell as some burning hot gases. With the liquid being located at discrete sites, there arelarge gradients in all directions with respect to fuel and oxidizer mass fluxes, mixtureratio, size and dispersion of droplets, or properties of the gaseous medium. Chemicalreactions occur in this zone, but the rate of heat generation is relatively low, in partbecause the liquids and the gases are still relatively cold and in part becausevaporization near the droplets causes fuel-rich and fuel-lean regions which do not burnas quickly. Some hot gases from the combustion zone are re-circulated back from therapid combustion zone, and they can create local gas velocities that flow across theinjector face. The hot gases, which can flow in unsteady vortexes or turbulence patterns, areessential to the initial evaporation of the liquids. The injection, atomization andvaporization processes are different if one of the propellants is a gas. For example, thisoccurs in liquid oxygen with gaseous hydrogen propellant in thrust chambers or pre-combustion chambers, where liquid hydrogen has absorbed heat from cooling jackets 13 Toc H Institute Of Science And Technology, Arakkunnam, Ernakulam
  • 14. Semester VII Branch: Mechanical Engineering Seminar Title: Cryogenic Rocket Enginesand has been gasified. Hydrogen gas has no droplets and does not evaporate. The gasusually has a much higher injection velocity (above 120 m/sec) than the liquidpropellant. This cause shear forces to be imposed on the liquid jets, with more rapid dropletformation and gasification. The preferred injector design for gaseous hydrogen andliquid oxygen is different from the individual jet streams used with storable propellants.Stream Tube Combustion Zone In this zone oxidation reactions continue, but at a lower rate, and some additionalheat is released. However, chemical reactions continue because the mixture tends to bedriven toward an equilibrium composition. Since axial velocities are high (200 to 600m/sec) the transverse convective flow velocities become relatively small. Streamlinesare formed and there is relatively little turbulent mixing across streamline boundaries.Locally the flow velocity and the pressure fluctuate somewhat. The residence time inthis zone is very short compared to the residence time in the other two zones. Thestreamline type, inviscid flow, and the chemical reactions toward achieving chemicalequilibrium persist not only throughout the remainder of the combustion chamber, butare also extended into the nozzle. Actually, the major processes do not take placestrictly sequentially, but several seem to occur simultaneously in several parts of thechamber. The flame front is not a simple plane surface across the combustion chamber There is turbulence in the gas flow in all parts of the combustion chamber. Theresidence time of the propellant material in the combustion chamber is very short,usually less than 10 milliseconds. Combustion in a liquid rocket engine is very dynamic,with the volumetric heat release being approximately 370 MJ/m3-sec, which is muchhigher than in turbojets. Further, the higher temperature in a rocket causes chemicalreaction rates to be several times faster (increasing exponentially with temperature)than in turbojet. The four phases of combustion in the thrust chamber are 1. Primary Ignition 14 Toc H Institute Of Science And Technology, Arakkunnam, Ernakulam
  • 15. Semester VII Branch: Mechanical Engineering Seminar Title: Cryogenic Rocket Engines 2. Flame Propagation 3. Flame Lift off 4. Flame AnchoringPrimary Ignition  begins at the time of deposition of the energy into the shear layer and ends when the flame front has reached the outer limit of the shear layer  starts interaction with the recirculation zone.  phase typically lasts about half a millisecond  it is characterised by a slight but distinct downstream movement of the flame .  The flame velocity more or less depends on the pre-mixedness of the shear layer only.Flame Propagation  This phase corresponds to the time span for the flame reaching the edge of the shear layer, expands into in the recirculation zone and propagates until it has consumed all the premixed propellants.  This period lasts between 0.1 and 2 ms.  It is characterised by an upstream movement of the upstream flame front until it reaches a minimum distance from the injector face plate.  It is accompanied by a strong rise of the flame intensity and by a peak in the combustion chamber pressure. 15 Toc H Institute Of Science And Technology, Arakkunnam, Ernakulam
  • 16. Semester VII Branch: Mechanical Engineering Seminar Title: Cryogenic Rocket Engines  The duration of this phase as well as the pressure and emission behaviour during this phase depend strongly on the global characteristics of the stationary cold flow before ignition.Flame Lift Off  phase starts when the upstream flame front begins to move downstream away from the injector because all premixed propellants in the recirculation zone have been consumed until it reaches a maximum distance.  This period lasts between 1 and 5 ms.  The emission of the flame is less intense showing that the chemical activity has decreased.  The position where the movement of the upstream flame front comes to an end, the characteristic times of convection and flame propagation are balanced. 16 Toc H Institute Of Science And Technology, Arakkunnam, Ernakulam
  • 17. Semester VII Branch: Mechanical Engineering Seminar Title: Cryogenic Rocket EnginesFlame Anchoring.  This period lasts from 20 ms to more than 50 ms, depending on the injection condition.  It begins when the flame starts to move a second time upstream to injector face plate and ends when the flame has reached stationary conditions.  During this phase the flame propagates upstream only in the shear layer .  Same as flame lift-off phase the vaporisation is enhanced by the hot products which are entrained into the shear layer through the recirculation zone.  The flame is stabilised at a position where an equilibrium exists between the local velocity of the flame front and the convective flow velocity.  This local flame velocity is depending on the upstream LOX-evaporation rates, i.e., the available gaseous O2, mixing of O2 and H2, hot products and radicals in the shear layer.  At the end of this phase, combustion chamber pressure and emission intensity are constant. 17 Toc H Institute Of Science And Technology, Arakkunnam, Ernakulam
  • 18. Semester VII Branch: Mechanical Engineering Seminar Title: Cryogenic Rocket EnginesDIFFERENT TYPES OF CRYOGENIC ENGINESHM-7B Rocket Engine HM-7 cryogenic propellant rocket engine has been used as an upper stageengine on all versions of the Ariane launcher. The more powerful HM-7B version wasused on Arianes 2, 3 and 4 and is also used on the ESC-A cryogenic upper stage ofAriane 5. Important principles used in the HM-7 combustion chamber were adopted byNASA under license and it is this technology that formed the basis of todays US spaceshuttle main engines - the first reusable rocket engine in the world. The HM7 engine was built upon the development work of the 40kN HM-4 engine.In 1973, the Ottobrunn team started development of the HM-7 thrust chamber for thethird stage of Ariane 1. Six years later, the HM-7 engine was successfully qualified withthe first launch of Ariane 1 in December 1979. 18 Toc H Institute Of Science And Technology, Arakkunnam, Ernakulam
  • 19. Semester VII Branch: Mechanical Engineering Seminar Title: Cryogenic Rocket Engines With the introduction of Ariane 2 and Ariane 3, it became necessary to increasethe performance of the HM-7 engine. This was achieved by raising the combustionchamber pressure from 30 to 35 bar and extending the nozzle, thereby raising thespecific impulse. The burn time was also increased from 570 to 735 seconds. Theupgraded engine was thus designated HM-7B and was qualified in 1983. Whensubsequently used on Ariane 4, the burn time was increased to 780 seconds. In February 2005, the HM-7B successfully powered the new cryogenic upperstage of Ariane 5, designate ESC-A (Etage Superior Cryo-technique A). This flight wasa tribute to the performance and flight proven reliability of an engine first developed 30years ago. With the ESC-A upper stage, the payload performance of Ariane 5 isincreased to 10 tonnes. In order to inherit the proven reliability of the HM-7B enginefrom over one hundred Ariane 4 flights, engine changes were kept to a minimum. Themain change being a 20% increase in burn time from 780 seconds to 950 seconds onAriane 5 ESC-A. Use of HM-7B on Ariane 5 is a first step toward increasing the payloadperformance of Ariane 5. A second step will be the introduction of the new Vinciexpander cycle engine to an ESC-B cryogenic upper stage, increasing the payloadperformance to 12 tonnes 19 Toc H Institute Of Science And Technology, Arakkunnam, Ernakulam
  • 20. Semester VII Branch: Mechanical Engineering Seminar Title: Cryogenic Rocket Engines The HM7B engine is a gas generator liquid oxygen / liquid hydrogen engine thatpowers the Ariane 4 third stage. The HM7 engine built upon the development work ofthe 40 kN thrust HM4. The HM7 development program began in 1973 as part ofEuropes effort to develop an indigenous launch capability. Final qualification of theHM7 engine occurred in 1979 and the engine went on to power the third stage of theAriane 1. SEP continued to perfect and upgrade the engine, increasing the specificimpulse by 4 seconds by increasing chamber pressure and lengthening the nozzle. Thenew engine, the HM7B, powered the third stage of the Ariane 2,3 and 4. As of June 1st,1995, SEP had produced 111 HM7B engines, with a cumulated total of 171,700seconds of operation, including 47,400 in flight.300 N Cryogenic Engine: This 300 N cryogenic propellant engine has a vacuum Isp of 415 seconds - thehighest value ever achieved in Europe for an engine of such small size. Being pressure-fed, the engine assembly is relatively simple and avoids the needfor a turbo-pump. The thrust chamber and throat region of the engine are regenerativecooled using hydrogen propellant. The nozzle extension is radiation cooled. The engine incorporates a splash-plate injector having a star shapedconfiguration. Ignition and subsequent re-ignition is achieved using Tri-ethyl aluminum(TEA) - which is hypergolic with the oxygen propellant. The number of re-ignitions is afunction of the volume of Tri-ethyl aluminum accommodated. The engine nominallyprovides for 1 ignition and 3 re-ignitions using just 1.5 cc of Tri-ethyl aluminum. The useof a chemical ignition system enables a very compact design. The engine needs no pre-cooling prior to ignition. Only the propellant feed linesto the engine propellant valves need be pre-cooled. 20 Toc H Institute Of Science And Technology, Arakkunnam, Ernakulam
  • 21. Semester VII Branch: Mechanical Engineering Seminar Title: Cryogenic Rocket Engines Engine construction materials are mainly stainless steel, Nimonic 75 (Chromium-Nickel Alloy) and copper.Applications The 300 N cryogenic engines enable the simplicity of a pressure fed propulsionsystem whilst offering the performance of a turbo-pump propulsion system. Being pressure fed, the engine does not require an additional turbo-pump, withits associated complexity. 21 Toc H Institute Of Science And Technology, Arakkunnam, Ernakulam
  • 22. Semester VII Branch: Mechanical Engineering Seminar Title: Cryogenic Rocket Engines The 300 N cryogenic engines may be used as a main engine in dedicated stagesfor orbital insertion, orbital transfer, orbital, and interplanetary applications, including: Upper stages Kick stages Vernier stages Transfer stages The 300 N cryogenic engines may also be used as a thruster, or thruster clusterwith existing cryogenic turbo-pump propulsion systems and stages for such applicationsas performance augmentation, upgrades, roll control.Vulcain Rocket Engine Vulcain (also known as HM-60) was the first main engine of the Ariane 5cryogenic first stage (EPC). The development of Vulcain, assured by a Europeancollaboration, began in 1988 with the Ariane 5 rocket program. It first flew in 1996powering the ill-fated flight 501 without being the cause of the disaster, and had its firstsuccessful flight in 1997 (flight 502). In 2002 the upgraded Vulcain 2 with 20% morethrust first flew on flight 517, although a problem with the engine turned the flight into afailure. The cause was due to flight loads being much higher than expected, as theinquiry board concluded. Subsequently, the nozzle has been redesigned, reinforcing the structure andimproving the thermal situation of the tube wall, enhancing hydrogen coolant flow aswell as applying thermal barrier coating to the flame-facing side of the coolant tubes,reducing heat load. The first successful flight of the (partially redesigned) Vulcain 2occurred in 2005 on flight 521. The Vulcain engines are gas-generator cycle cryogenicrocket engines fed with liquid oxygen and liquid hydrogen. They feature regenerative cooling through a tube wall design, and the Vulcain 2introduced a particular film cooling for the lower part of the nozzle, where exhaust gasfrom the turbine is re-injected in the engine They power the first stage of the Ariane 5 22 Toc H Institute Of Science And Technology, Arakkunnam, Ernakulam
  • 23. Semester VII Branch: Mechanical Engineering Seminar Title: Cryogenic Rocket Engineslauncher, the EPC (Étage Principal Cryo technique, main cryogenic stage) and provide8% of the total lift-off thrust (the rest being provided by the two solid rocket boosters).The engine operating time is 600 s in both configurations. The coaxial injector elements cause the LOX and LH2 propellants to be mixedtogether. LOX is injected at the centre of the injector, around which the LH2 is injected.These propellants are mainly atomized and mixed by shear forces generated by thevelocity differences between LOX and LH2. The final acceleration of hot gases, up tosupersonic velocities, is achieved by gas expansion in the nozzle extension, therebyincreasing the thrust. 23 Toc H Institute Of Science And Technology, Arakkunnam, Ernakulam
  • 24. Semester VII Branch: Mechanical Engineering Seminar Title: Cryogenic Rocket EnginesApplications:  main engine of the Ariane 5 cryogenic first stage (EPC)VINCI Rocket Engine: Vinci is a European Space Agency cryogenic rocket engine currently underdevelopment. It is designed to power the new upper stage of Ariane 5, ESC-B, and willbe the first European re-ignitable cryogenic upper stage engine, raising the launchersGTO performances to 12 t. Vinci is an expander cycle rocket engine fed with liquidhydrogen and liquid oxygen. Its biggest improvement from its predecessor, the HM-7 isthe capability of restarting up to five times. It is also the first European expander cycle 24 Toc H Institute Of Science And Technology, Arakkunnam, Ernakulam
  • 25. Semester VII Branch: Mechanical Engineering Seminar Title: Cryogenic Rocket Enginesengine, removing the need for a gas generator to drive the fuel and oxydizer pumps. Itfeatures a carbon ceramic extendable nozzle in order to have a large, 2.15 m diameternozzle extension with minimum length: the retracted nozzle part is deployed only afterthe upper stage separates from the rest of the rocket; after extension, the enginesoverall length increases from 2.3 m to 4.2 m.Applications:  upper stage of Ariane 5 25 Toc H Institute Of Science And Technology, Arakkunnam, Ernakulam
  • 26. Semester VII Branch: Mechanical Engineering Seminar Title: Cryogenic Rocket Engines CONCLUSION The area of Cryogenics in Cryogenic Rocket Engines is a vast one and it cannotbe described in a few words. As the world progress new developments are being mademore and more new developments are being made in the field of Rocket Engineering.Now a day cryo propelled rocket engines are having a great demand in the field ofspace exploration. Due to the high specific impulse obtained during the ignition of fuelsthey are of much demand. 26 Toc H Institute Of Science And Technology, Arakkunnam, Ernakulam
  • 27. Semester VII Branch: Mechanical Engineering Seminar Title: Cryogenic Rocket Engines REFERENCES  “Rocket propulsion elements” by G. P. Sutton, 7 th edition.  “Advances in propulsion” by K. Ramamurthy.  “Rocket and Spacecraft Propulsion” by M. J. Turner.  “Ignition of cryogenic H2/LOX sprays” by O. Gurliat, V. Schmidt, O.J. Haidn, M. Oschwald.  National Aeronautics and Space Administration, United States Of America  Vikram Sarabhai Space Centre, Thiruvananthapuram 27 Toc H Institute Of Science And Technology, Arakkunnam, Ernakulam