rocket_liquid
DESCRIPTION
A liquid-propellant rocket or a liquid rocket is a rocket engine that uses propellants in liquid form. Liquids are desirable because their reasonably high density ...TRANSCRIPT
History
Rocket technology first became known to Europeans following their use by the Mongols Genghis Khan and Ögedei Khan, when they conquered parts of Russia, Eastern, and Central Europe --12th
gunpowder (75% of saltpeter, 15% of carbon and 10% of sulphur).
The name Rocket comes from the Italian Rocchetta (i.e. little fuse), a name of a small firecracker created by the Italian artificer Muratori in 1379.After that Korea 15th
Between 1529 and 1556 Conrad Haas wrote a book that described the concept of multi-stage rocketsindiaThe first iron cased and metal-cylinder rocket artillery made from iron tubes, were developed by the weapon suppliers of Tipu Sultan an Indian ruler of the Kingdom of Mysore
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Types
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Liquid
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Requirement
For long distance Variable thrust- the amount of fuel and rate of burn can be changed in
flight Liquid-fuel boosters are more easily re-usable Liquid-fueled rockets have higher specific impulse than solid rockets The primary performance advantage of liquid propellants is due to the
oxidizer. Several practical liquid oxidizers (liquid oxygen , nitrogen tetroxide, and hydrogen peroxide) are available which have better specific impulse than the ammonium perchlorate used in most solid rockets
liquid propellants are cheaper than solid propellants
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Oxidizer liquid oxygen (LOX, O2)T-Stoff (80% hydrogen peroxid)nitric acid(HNO3)inhibited red fuming nitric acid
(IRFNA, HNO3 + N2O4)
nitric acid 73% with dinitrogen tetroxide 27% (=AK27)
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Fuelliquid hydrogen(LH2, H2) kerosene or RP-1alcohol (ethanol, C2H5OH)hydrazine(N2H4)Aerozine 50monomethylhydrazine(MMH,
(CH3)HN2H2)
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Propellant
LOX and liquid hydrogen, used in the Space Shuttle orbiter LOX and kerosene (RP-1). Used for the first stages of the Saturn
V, Atlas V and Falcon, the Russian Soyuz, Ukrainian Zenit Nitrogen tetroxide(N2O4) and hydrazine(N2H4), MMH, or UDMH.
Used in military, orbital, and deep space rockets because both liquids are storable for long periods at reasonable temperatures and pressure PSLV
hydrogen peroxide and kerosene hydrazine(N2H4) and red fuming nitric acid – Nike Ajax Antiaircraft
Rocket monomethylhydrazine (MMH, (CH3)HN2H2) and dinitrogen tetroxide
– Space Shuttle orbiter's Orbital maneuvering system (OMS) engines and Reaction control system (RCS) thrusters
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Combustion A liquid Rocket combustion chamber to designed to accommodate And allow to sufficient time for following jobs
Injection ,atomization ,vaporization and even mixing liquid fuel and oxidizer Thermal decomposition of oxidizer to enable chemical reaction with fuel Ignition and flame stabilization and combustion of fuel, oxidizer mixture
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Even dispersion of combustion products towards the nozzle The volume ,length and shape of combustion chamber needs to selected To all the steps. Various fuel and oxidizer combustion provides for characteristic lengthL* for rocket
L* =CC volume/throat aeraThe value of L* found experimently
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Desirable properties of liquid propellant
Low freezing point High specific gravity Good chemical stability during storage High specific heat and high thermal coundtivity , High thermal
decomposition Pumping property – flowablity (under cryogenic condition) Temperature stability of physical property (viscosity ,vapor pressure etc.)Under cryogenic condition.
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Currently, six governments have successfully developed and deployed cryogenic rocket engines
United States
European Space Agency
Russia
China
India CE-7.5 CE-20
Japan
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GSLV-D5 MK II, launched on January 5, 2014
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CE-7.5The CE-7.5 is a cryogenic rocket engine developed by India to power the upper stage of its GSLV Mk-2 launch vehicle
CE-7.5 is a regeneratively cooled, variable thrust, staged combustion cycle engine.
SpecificationsThe specifications of the engine:Operating Cycle - Staged combustionPropellant Combination - LOX / LH2Thrust Nominal (Vacuum) - 75 kNOperating Thrust Range - 73.55 kN to 93.1 kN (To be set at any fix values)Chamber Pressure (Nom) - 58 barEngine Mixture ratio (Oxidizer/Fuel by weight) - 5.05Engine Specific Impulse - 454 ± 3 seconds (4.452 ± 0.029 km/s)Engine Burn Duration (Nom) - 720 secondsPropellant Mass - 12800 kg
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