1.
Vikas (rocket engine)
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The Vikas is a liquid fueled rocket engine built by Indian Space Research Organisation. It was developed at the Liquid Propulsion Systems Center during the 1970s and it is used in the Polar Satellite Launch Vehicle and the Geosynchronous Satellite Launch Vehicle series of expendable launch vehicles for space launch use. The engine is used as the stage of both the PSLV and the GSLV launch vehicles, with four strap-on boosters. The engine is capable of gimballing. The GSLV MK-3 rocket uses two Vikas engines in its L110 core stage, the propellant loading for GSLV Mk-3 vikas engines is 55 tons compared to 40 tons for regular GSLV Mk-2 and PSLV rockets. The engine uses up about 40 metric tons of UDMH as fuel, an up-graded version of the engine has a chamber pressure of 58.5 bar as compared to 52.5 bar in the older version and produces a thrust of 800 kN. The rocket benefited from technological cooperation from the Viking 4A engine built by CNES/SEP of France, the primary difference being that the Vikas is rated for a longer burn time. Class of 1974, Rocket science & reminiscences
2.
Vulcain
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Vulcain is a family of European first stage rocket engines for the Ariane 5. Its development began in 1988 and the first flight was completed in 1996, the updated version of the engine – Vulcain 2 was first successfully flown in 2005. Both members of the family use liquid oxygen/liquid hydrogen cryogenic fuel, as of 2012 no new version of the engine is in development. The development of Vulcain, carried out by a European partnership and it first flew in 1996 powering the ill-fated flight 501 without being the cause of the disaster, and had its first successful flight in 1997. In 2002 the upgraded Vulcain 2 with 20% more thrust first flew on flight 517, the cause was due to flight loads being much higher than expected, as the inquiry board concluded. The first successful flight of the Vulcain 2 occurred in 2005 on flight 521, although different upgrades to the engine have been proposed, there is no current program to develop an uprated version of the engine. If there will ever be one, it is likely that the new engine would be introduced after the PA batch of 30 Ariane 5 ECAs ordered on 10 May 2004 will be expended. On 17 June 2007 Volvo Aero announced that in spring of 2008 it expected to hot-fire test a Vulcain 2 nozzle manufactured with a new sandwich technology, the Vulcain engines are gas-generator cycle cryogenic rocket engines fed with liquid oxygen and liquid hydrogen. The engine operating time is 600 s in both configurations,3 m tall and 1.76 m in diameter, the engine weighs 1686 kg and provides 137 t of thrust in its latest version. The oxygen turbopump rotates at 13600 rpm with a power of 3 MW while the hydrogen turbopump rotates at 34000 rpm with 12 MW of power, the total mass flow rate is 235 kg/s, of which 41.2 kg/s are of hydrogen. The main contractor for the Vulcain engines is Snecma Moteurs, which provides the liquid hydrogen turbopump. The liquid oxygen turbopump is the responsibility of Avio, and the gas turbines power the turbopumps. – Volvo Aero Development of the turbines for the Vulcain 2 turbopumps, – Volvo Aero High cycle fatigue of Vulcain 2 LOx turbine blades. – Volvo Aero An efficient concept design process, – Volvo Aero Vulcain 2 nozzle. – Volvo Aero EADS N. V. – EADS welcomes contract signature for 30 Ariane 5 launchers at ILA2004 in Berlin Three billion Euros contract for 30 Ariane 5 launchers – EADS Astrium
3.
Liquid-propellant rocket
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A liquid-propellant rocket or liquid rocket is a rocket engine that uses liquid propellants. This permits the use of low-mass propellant tanks, resulting in a mass ratio for the rocket. An inert gas stored in a tank at a pressure is sometimes used instead of pumps in simpler small engines to force the propellants into the combustion chamber. These engines may have a mass ratio, but are usually more reliable. And are therefore used widely in satellites for orbit maintenance, some designs are throttleable for variable thrust operation and some may be restarted after a previous in-space shutdown. Liquid propellants are used in hybrid rockets, in which a liquid oxidizer is generally combined with a solid fuel. This seminal treatise on astronautics was published in 1903, but was not distributed outside Russia until years later, german V-2 inventor Wernher von Braun considered Paulet one of the fathers of aeronautics. The National Air & Space Museum in Washington, D. C. has a plaque honoring the memory of Paulet. He was the only known developer of liquid-propellant rocket engine experiments in the 19th century, however, he did not publish his work. In 1927 he wrote a letter to a newspaper in Lima, historians of early rocketry experiments, among them Max Valier, Willy Ley, and John D. Clark, have given differing amounts of credence to Paulets report. Paulet described laboratory tests of, but did not claim to have launched a liquid rocket. Wernher von Braun, in his book World History of Aeronautics states, Pedro Paulet was in Paris in those years, experimenting with his tiny two-and-a-half kilogram motor, by this act, Paulet should be considered the pioneer of the liquid fuel propulsion motor. Further, in his History of Rocketry and Space Travel, von Braun recognizes that by his efforts, Paulet helped man reach the Moon. The rocket, which was dubbed Nell, rose just 41 feet during a 2. 5-second flight that ended in a cabbage field, goddard proposed liquid propellants about fifteen years earlier and began to seriously experiment with them in 1921. In Germany, engineers and scientists became enthralled with liquid-fuel rockets and this amateur rocket group, the VfR, included Wernher von Braun, who became the head of the army research station that secretly built the V-2 rocket weapon for the Nazis. The German-Romanian Hermann Oberth published a book in 1922 suggesting the use of liquid propellants, after World War II the American government and military finally seriously considered liquid-propellant rockets as weapons and began to fund work on them. The Soviet Union did likewise, and thus began the Space Race, bipropellant liquid rockets generally use a liquid fuel, such as liquid hydrogen or a hydrocarbon fuel such as RP-1, and a liquid oxidizer, such as liquid oxygen. The engine may be a rocket engine, where the fuel and oxidizer
4.
Dinitrogen tetroxide
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Dinitrogen tetroxide, commonly referred to as nitrogen tetroxide, is the chemical compound N2O4. It is a reagent in chemical synthesis. It forms a mixture with nitrogen dioxide. Dinitrogen tetroxide is an oxidizer that is hypergolic upon contact with various forms of hydrazine. Dinitrogen tetroxide could be regarded as two nitro groups bonded together and it forms an equilibrium mixture with nitrogen dioxide. The molecule is planar with an N-N bond distance of 1.78 Å, the N-N distance corresponds to a weak bond, since it is significantly longer than the average N-N single bond length of 1.45 Å. Unlike NO2, N2O4 is diamagnetic since it has no unpaired electrons, inevitably, some dinitrogen tetroxide is a component of smog containing nitrogen dioxide. Nitrogen tetroxide is made by the oxidation of ammonia, steam is used as a diluent to reduce the combustion temperature. The gas is essentially pure nitrogen dioxide, which is condensed into dinitrogen tetroxide in a brine-cooled liquefier, dinitrogen tetroxide can also be made through the reaction of concentrated nitric acid and metallic copper. This synthesis is more practical in a setting and is commonly used as a demonstration or experiment in undergraduate chemistry labs. The unstable species further react to form nitrogen dioxide which is purified and condensed to form dinitrogen tetroxide. Nitrogen tetroxide is used as an oxidizer in one of the more important rocket propellants because it can be stored as a liquid at room temperature and it became the storable oxidizer of choice for many rockets in both the United States and USSR by the late 1950s. It is a propellant in combination with a hydrazine-based rocket fuel. One of the earliest uses of this combination was on the Titan family of rockets used originally as ICBMs and then as launch vehicles for many spacecraft. Used on the U. S. Gemini and Apollo spacecraft and also on the Space Shuttle, it continues to be used as stationkeeping propellant on most geo-stationary satellites, and many deep-space probes. It now seems likely that NASA will continue to use this oxidizer in the next-generation crew-vehicles which will replace the shuttle and it is also the primary oxidizer for Russias Proton rocket. When used as a propellant, dinitrogen tetroxide is usually referred to simply as Nitrogen Tetroxide, most spacecraft now use MON instead of NTO, for example, the Space Shuttle reaction control system used MON3. On 24 July 1975, NTO poisoning affected the three U. S. astronauts on board the Apollo-Soyuz Test Project during its final descent, one crew member lost consciousness during descent
5.
Gas-generator cycle
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The gas-generator cycle is a power cycle of a bipropellant rocket engine. Some of the propellant is burned in a gas generator and the hot gas is used to power the engines pumps. Because something is thrown away this type of engine is known as open cycle. There are several advantages to the cycle over its counterpart. The gas generator turbine does not need to deal with the pressure of injecting the exhaust into the combustion chamber. This simplifies plumbing and turbine design, and results in a less expensive, the main disadvantage is lost efficiency due to discarded propellant. Gas-generator cycles tend to have lower specific impulse than staged combustion cycles, as in most cryogenic rocket engines, some of the fuel in a gas-generator cycle may be used to cool the nozzle and combustion chamber. Without any rocket combustion chamber and nozzle heat mitigation, the engine would fail catastrophically
6.
Thrust-to-weight ratio
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The instantaneous thrust-to-weight ratio of a vehicle varies continually during operation due to progressive consumption of fuel or propellant and in some cases a gravity gradient. The thrust-to-weight ratio based on initial thrust and weight is often published and used as a figure of merit for quantitative comparison of the performance of vehicles. The thrust-to-weight ratio can be calculated by dividing the thrust by the weight of the engine or vehicle, note that the thrust can also be measured in pound-force provided the weight is measured in pounds, the division of these two values still gives the numerically correct thrust-to-weight ratio. For valid comparison of the initial ratio of two or more engines or vehicles, thrust must be measured under controlled conditions. The thrust-to-weight ratio and wing loading are the two most important parameters in determining the performance of an aircraft, for example, the thrust-to-weight ratio of a combat aircraft is a good indicator of the maneuverability of the aircraft. The thrust-to-weight ratio varies continually during a flight, thrust varies with throttle setting, airspeed, altitude and air temperature. Weight varies with fuel burn and changes of payload, for aircraft, the quoted thrust-to-weight ratio is often the maximum static thrust at sea-level divided by the maximum takeoff weight. In cruising flight, the ratio of an aircraft is the inverse of the lift-to-drag ratio because thrust is the inverse of drag. Rockets and rocket-propelled vehicles operate in a range of gravitational environments. The thrust-to-weight ratio is calculated from initial gross weight at sea-level on earth and is sometimes called Thrust-to-Earth-weight ratio. The thrust-to-Earth-weight ratio of a rocket or rocket-propelled vehicle is an indicator of its acceleration expressed in multiples of earth’s gravitational acceleration, the thrust-to-weight ratio for a rocket varies as the propellant is burned. If the thrust is constant, then the ratio is achieved just before the propellant is fully consumed. Each rocket has a characteristic thrust-to-weight curve or acceleration curve, not just a scalar quantity, for a takeoff from the surface of the earth using thrust and no aerodynamic lift, the thrust-to-weight ratio for the whole vehicle must be more than one. In general, the ratio is numerically equal to the g-force that the vehicle can generate. Take-off can occur when the vehicles g-force exceeds local gravity, many factors affect a thrust-to-weight ratio. The instantaneous value typically varies over the flight with the variations of thrust due to speed and altitude along with the due to the remaining propellant. The main factors include air temperature, pressure, density. Depending on the engine or vehicle under consideration, the performance will often be affected by buoyancy
7.
Specific impulse
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Specific impulse is a measure of the efficiency of rocket and jet engines. By definition, it is the total impulse delivered per unit of propellant consumed and is equivalent to the generated thrust divided by the propellant flow rate. If mass is used as the unit of propellant, then specific impulse has units of velocity, if weight is used instead, then specific impulse has units of time. Multiplying flow rate by the gravity before dividing it into the thrust. In rockets, this means the engine is more efficient at gaining altitude, distance and this is much less of a consideration in jet engines that employ wings and outside air for combustion to carry payloads that are much heavier than the propellant. Specific impulse includes the contribution to impulse provided by air that has been used for combustion and is exhausted with the spent propellant. Jet engines use air, and therefore have a much higher specific impulse than rocket engines. The specific impulse in terms of propellant mass spent is in units of distance per time and this is higher than the actual exhaust velocity because the mass of the combustion air is not being accounted for. Actual and effective exhaust velocity are the same in rocket engines not utilizing air, Specific impulse is inversely proportional to specific fuel consumption by the relationship Isp = 1/ for SFC in kg/ and Isp = 3600/SFC for SFC in lb/lbf-hr. The amount of propellant is normally measured either in units of mass or weight, however, if propellant weight is used, an impulse divided by a force turns out to be a unit of time, and so specific impulses are measured in seconds. These two formulations are both used and differ from each other by a factor of g0, the dimensioned constant of gravitational acceleration at the surface of the Earth. Note that the rate of change of momentum of a rocket per unit time is equal to the thrust, the higher the specific impulse, the less propellant is needed to produce a given thrust during a given time. In this regard a propellant is more efficient the greater its specific impulse and this should not be confused with energy efficiency, which can decrease as specific impulse increases, since propulsion systems that give high specific impulse require high energy to do so. It is important that thrust and specific impulse not be confused, the specific impulse is a measure of the impulse produced per unit of propellant expended, while thrust is a measure of the momentary or peak force supplied by a particular engine. In many cases, propulsion systems with high specific impulses—some ion thrusters reach 10,000 seconds—produce low thrusts. When calculating specific impulse, only propellant that is carried with the vehicle use is counted. Air resistance and the inability to keep a high specific impulse at a fast burn rate are why all the propellant is not used as fast as possible. If an engine weighs more in order to gain a specific impulse, it may not be as efficient in gaining altitude, distance
8.
Gimbaled thrust
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Gimbaled thrust is the system of thrust vectoring used in most modern rockets, including the Space Shuttle and the Saturn V lunar rockets. In a gimbaled thrust system, the exhaust nozzle of the rocket can be swiveled from side to side, as the nozzle is moved, the direction of the thrust is changed relative to the center of gravity of the rocket. The middle rocket shows the flight configuration in which the direction of thrust is along the center line of the rocket. On the rocket at the left, the nozzle has been deflected to the left, since the thrust no longer passes through the center of gravity, a torque is generated about the center of gravity and the nose of the rocket turns to the left. If the nozzle is gimballed back along the line, the rocket will move to the left. On the rocket at the right, the nozzle has been deflected to the right and the nose is moved to the right
9.
Ariane 1
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Ariane 1 was the first rocket in the Ariane family of expendable launch systems. It was developed and operated by the European Space Agency, which had formed in 1973. Ariane 1 was the first launcher to be developed with the purpose of sending commercial satellites into geosynchronous orbit. Crucially, it was designed with the ability of sending a pair of satellites into orbit on a single launcher, thus reducing costs. As the size of satellites grew, Ariane 1 quickly gave way to the more powerful Ariane 2 and Ariane 3 launchers, which were heavily based upon the original rocket. The Ariane 4 was the last rocket to heavily draw upon the Ariane 1, as a result, the Europa IIIB proposal was scaled back and soon reemerged as the L3S. In January 1973, Willy Brandt, the Chancellor of Germany, formally agreed to the L3S project following a series of approaches by Georges Pompidou. On 21 September 1973, the agreement for the L3S, was signed. Under this agreement, the Europa III was formally cancelled, while the L3S would be developed as a multinational project. Development of L3S was seen as a crucial test for the ESA, according to author Brian Harvey, L3S was one of the major European engineering projects in the last quarter of the century. Other major companies involved included the French electronics firm Matra, Swedish manufacturer Volvo, development of the third stage was a major focus point for the project - prior to Ariane, only the United States had ever flown a launcher that utilised hydrogen-powered upper stages. In order to accommodate Ariane launches, the Guiana Space Centre at Kourou, the former Europa launch site was re-designated as ELA1 and was rebuilt with a lowered base and elongated tower. While all Ariane launches would take place from French Guiana, rocket construction would be performed at Aérospatiales facility in Les Mureaux, Paris. This was considered a coup for the programme as Intelsat was viewed as heavily committed to using the rival Space Shuttle launcher for a large number of its satellites at that point. One week later, ESA announced its commitment to a run of 10 Ariane 1 launchers. There was considerable pressure for Ariane to perform its maiden flight prior to end of 1979, despite fears that the launch would have to be delayed for a month, it was decided to resume the countdown for a second attempt. CAT was successfully placed into an orbit of 202km by 35, 753km, with lift-off mass of 210,000 kg, Ariane 1 was able to put in geostationary transfer orbit one satellite or two smaller of a maximal weight of 1,850 kg. The cost of program is estimated at 2 billion euros, the Ariane 1 was a four-stage vehicle
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