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    Rocket Propulsion Basics  The Rocket MotorPump, Combustion Chamber & Nozzle  Rocket Propulsion Principles  The Propellant Pump(s) An essential component of liquid fuelled rocket engines is the means of delivering the propellants (the fuel and the oxidiser) to the combustion chamber. The simplest method used in low thrust rockets is by pressurising the fuel and oxidiser tanks with compressed air or a gas such as nitrogen, but for most liquid fuelled rockets, the high propellant flow rates required are provided by onboard turbopumps. The Injector Plate The injector plate is a passive device which has three purposes. It breaks up the liquid propellants into tiny droplets to aid and speed up combustion, it enables homogeneous mixing of the fuel with the oxidiser and it ensures stable, controlled burning of the fuel, preventing the explosive combustion of the propellants. 帝冢真织免费阅读_帝冢真织最新章节,free xxxx movies hd,中方回应香会主办方期待中国高级代表团与会 6分钟娇喘声 The Nozzle The purpose of the nozzle is to promote the isentropic (constant entropy) expansion of the exhaust gas. As the gas expands, its pressure drops, but since there is no change in total energy, its velocity (kinetic energy) increases to compensate for the reduction in pressure energy. There are thus two factors contributing to the engine thrust, namely, the kinetic energy of the gas particles ejected with high velocity from the exhaust and the pressure difference between the exhaust gas pressure and the ambient pressure of the atmosphere acting across the area of the nozzle exit. The relationship is shown in the following equation.
Engine Thrust F = dm/dt. V_{e} + A_{e}(P_{e}  P_{a})Where dm/dt = Propellant Mass Flow Rate per Second V_{e} = Gas (Exhaust) Velocity at Nozzle Exit A_{e} = Area of Nozzle Exit P_{e} = Gas Pressure at Nozzle Exit P_{a} = Ambient Pressure of the Atmosphere
The first term is known as the momentum thrust and the second term the pressure thrust. Considering the pressure thrust alone, since the ambient pressure decreases with altitude, in the vacuum of free space where the pressure is zero, the rocket thrust will increase to a maximum of 15% to 20% more than the thrust at sea level. (By contrast, the thrust of a jet engine decreases with altitude to zero in free space since it depends for its thrust on air as the oxidiser for the fuel. The rocket on the other hand carries its oxidiser with it.) The momentum of the exhaust gas is however much more effective in creating thrust than the pressure difference at the exhaust exit, so that the more the pressure energy is converted into kinetic energy in the nozzle, the more efficient the nozzle will be. So paradoxically the maximum thrust occurs when the exhaust pressure is equal to the ambient pressure. The effective exhaust velocity V_{e} is a function of the nozzle geometry such as the nozzle expansion ratioA_{e}/A_{t} WhereA_{t} = Area of Nozzle Throat See also Steam Turbine Nozzles  Rockets depend for their action on Newton's Third Law of Motion that: "For every action there is an equal and opposite reaction." In a rocket motor, fuel and oxidiser, collectively called the propellants, are combined in a combustion chamber where they react chemically to form hot gases which expand rapidly and are then accelerated and ejected at high velocity through a nozzle, thereby imparting momentum to the motor in the opposite direction. A rocket can be considered as a large body carrying small units of propellant and travelling with a velocity V. The reaction due to expelling the propellant from the rocket exhaust causes the velocity of the rocket to increase. Assuming no change in ambient pressure, the Conservation of Momentum for the rocket and the expelled propellant gives: (M+dm)V = M(V+dv) + dm(V  V_{e})Where M = The total remaining mass of the rocket and its fuel dm = The mass ejected rearwards through the exhaust nozzle or the change in mass during a given period. V = The initial absolute forward velocity of the rocket just before the ejection of the propellant dv = The increase in forward velocity of the rocket due to the ejection of the exhaust gases V_{e} = The exhaust velocity, relative to the rocket, of the propellant leaving the rocket motor. 帝冢真织免费阅读_帝冢真织最新章节,free xxxx movies hd,中方回应香会主办方期待中国高级代表团与会 6分钟娇喘声 Simplifying we can derive the following: dm.V_{e} = M.dv Dividing by dt to get the instantaneous rates and substituting Newton's Laws a=dv/dt The acceleration of the rocket F=M.a The force or thrust acting on the rocket
This gives dm/dt. V_{e} = M.dv/dt = M.a = F = The Force or ThrustWhere dm/dt = The mass flow rate of the burning fuel mass ejected.
Thus the instantaneous thrust on the rocket is directly proportional to the fuel mass flow or burn rate. For the change in velocity over a longer period we must take into account the reduction dM in the total mass of the rocket as its fuel is consumed and integrate the velocity over time for the duration of the period. From the above: The mass expelled = The reduction in mass of the rocket and its propellant load or dm =  dM and ∫dv = V_{e} ∫dm / Mso that ∫dv =  V_{e} ∫dM / MThus V_{f}V_{i} =  V_{e}(ln M)_{i}^{f} =  V_{e}(lnM_{f}  lnM_{i}) = V_{e} ln (M_{i} / M_{f})Where V_{i} = The initial velocity of the rocket V_{f} = The final velocity of the rocket ln = The natural logarithmic function M_{i} = The initial mass of the rocket including its payload all its propellant M_{f} = The final mass of the rocket and its payload including its remaining propellant M_{i} / M_{f} is known as the rocket's Mass Ratio
This is known as Tsiolkovsky's Equation Note that although a greater initial mass (of propellant) which increases the Mass Ratio, will create a greater increase in velocity, the relationship is not linear and the increase in velocity due to the increased available fuel becomes proportionally less as the initial mass M_{i} increases. This is because some of the extra propellant must be used to accelerate the mass of the extra fuel itself. See also Missile Ballistics, Orbits and Aerodynamics Multistage Rockets, another of Tsiolkovsy's ideas, separate the propulsion into more than one stage, each stage with its independent rocket motor, propellant tanks and pumps or pressurisation systems. The stages may be "stacked" as in the Apollo space vehicle which took the astronauts to the moon, or "piggy backed" as in the Space shuttle. As the propellant in the first stage is used up, the stage is jettisoned and the propulsion taken over by the subsequent stage so that the later stages do not have to waste energy accelerating the useless mass of the jettisoned stages. In this way higher velocity and range can be achieved with the same initial vehicle weight, payload weight and propellant capacity or alternatively a greater payload can be carried with a smaller initial weight.  Impulse, Thrust and Fuel Performance  Rocket Power and Dynamic Conversion Efficiency  Impulse For a constant Thrust F, the Impulse I provided by a motor or a propellant over a specific Period t is defined as; I = F.t TheSpecific Impulse I_{s} is the ratio of the of thrust produced to the weight flow of the propellants (fuel plus oxidiser). It is a measure of the potential effectiveness of a particular fuel and oxidiser combination in converting its chemical energy into useful work and is thus a convenient way of comparing fuel efficiencies. It is defined (in Imperial units) as: Thrust (lbs) Propellant Consumption (lbs/s) 
Specific Impulse = Or I_{s} = F / (dw/dt)Where I_{s} = The specific impulse is expressed in units of time (seconds) F = The thrust w = The combined weight of the fuel and oxidiser dw/dt = The propellant consumption per second
In international SI or MKS units this relationship becomes: I_{s} = F / (dm/dt).g_{0} Rearranging, this becomes: F = I_{s}(dm/dt).g_{0}Where F = The thrust in Newtons m = The mass of propellant in Kg g_{0} = The standard acceleration due to gravity at sea level (32.2 ft/s/s)
Thus increased thrust can be achieved by using propellants with a higher specific impulse and also by increasing the fuel burn rate. From the equations of motion opposite, the exhaust velocity V_{e} is given by V_{e} = F / dm/dtThus V_{e}=I_{s}.g_{0}The exhaust velocity, relative to the motor, is therefore directly proportional to the specific impulse. This is a simple way of determining the exhaust velocity from the specific impulse of the fuel / oxidiser combination. Note: Due to the affect of the ambient air pressure, the specific impulse may be 15% to 20% lower at sea level than in the vacuum of space. (See the thrust equation in the diagram above) Propellant DensityFuel effectiveness also depends on its density as well as the density of its associated oxidiser. High density propellants, can be accommodated in smaller tanks and they can use smaller pumps for feeding the propellants to engine. This allows smaller lighter vehicle structures with less aerodynamic drag. Taking density into account the effective specific impulse is given by: I_{d} = ρ_{av}I_{s}Where: I_{d} = The Density Specific Impulse(Kg.secs/litre) ρ_{av} = The average density of the fuel and the propellant mixture (kg/litre)
 Power Rocket Engine power P = The maximum available kinetic energy delivered to the exhaust gas stream per second. P = 1/2 dm/dt V_{e}^{2}Vehicle Motive PowerP_{m} = The power transmitted to the vehicle to drive it forwards P_{m}= FVThis implies that the rocket power at any instant is dependent on its velocity and is zero when the forward velocity is zero as it would be at liftoff. Once the rocket starts moving, the available kinetic energy and power are split between the exhaust stream and the rocket vehicle. Thus P = P_{m} + P_{e}So that P_{e} = 1/2 dm/dt (V_{e}  V)^{2}WhereP_{e} is the remaining power in the exhaust stream EfficiencyIgnoring parasitic efficiency losses such as propellant pumping power, frictional losses and nozzle design efficiency, the conversion efficiency of translating the energy in the exhaust gas flow into forward motion of the rocket is given by, η = P_{m}/ PWhereη = The conversion efficiency Thus η = FV / ( FV + dm/dt (VV_{e})^{2}/2)Note that the efficiency is dependent on the rocket's velocity and is maximum when V = V_{e}, that is when the forward velocity of the rocket is equal to the rocket's exhaust velocity. SubstitutingF / V_{e}fordm/dt the above equation simplifies to: η = 2 (V/ V_{e}) / (1+(V / V_{e})^{2})This provides a measure of the rocket's efficiency in terms of velocity alone. Ullage motors are rocket motors used to provide artificial gravity in multistage liquid fuelled rockets by momentarily accelerating the second stage forwards after the first stage burnout. This moment of forward thrust is required in the weightless environment of outer space to make certain that the second stage liquid propellant is in the proper position to be drawn into the pumps and that the gaseous zone above the liquid in the tank is not next to the pump input prior to starting the second stage engines. The extra thrust also helps to make a clean separation between stages. "Ullage" is an old brewers term meaning the air space above beer in a vat. Rocket Fuels and Oxidisers 
Liquid Fuels and Oxidisers Liquid propellants pioneered in 1926 by Robert Goddard are relatively safe and easy to control and easy to start and stop. However they need a complex pumping system, pressure controls, valves and a feed system to deliver the propellants to the combustion chamber all of which reduce the mass ratio and hence the efficiency of the system. Cryogenic Fuels and Oxidisers Some of the highest energy liquid propellants have very low boiling points. Liquid Hydrogen (LH_{2}) fuel for example has a boiling point of 252.9°C and an oxidiser such as Liquid Oxygen (LOX) boils at 183°C. Using these high energy density propellants in gaseous form is impractical since the enormous onboard storage tanks and pumping systems they would require would be too big and heavy. Even in liquid form there are difficulties in using these propellants since the storage tanks may need to be insulated and the pumps must work at very low temperatures with a very high temperature gradient across the body of the pump. Safety, handling and storage are also issues of concern. Nevertheless, cryogenic propellants are used when controllable, maximum thrust is a priority. Solid Fuels and Oxidisers Solid propellant motors contain both the fuel and the oxidiser in a charge called the grain which is stored within the combustion chamber. Invented by the Chinese in 1150, the motors are compact and light weight and do not need pumps, valves or feed systems so they have a very high mass ratio and thrust per unit volume, but for the same reason they are difficult to control. Once the burn starts, it is difficult, if not impossible, to stop until all the fuel is consumed. Hypergolic Propellants Hypergolic propellants are fuel and oxidiser combinations, liquid at room temperature, which ignite spontaneously on contact with eachother. They are easy to control, start, stop and restart. Some combinations are extremely toxic and corrosive. Suitable for engines which must be ignited in space or reoperated numerous times. Elimination of the igniter removes a significant source of unreliability.  Example  Saturn V S1C Engine Performance  Example  Saturn V Fuel Choices  Rocketdyne F1 Engine used in Saturn V S1C Engine dimensions Dry mass: 18,500 lbs Length: 19 ft Maximum diameter: 12ft 4in
Fuel: Kerosene (RP1), delivered at 1,754 lb/s (dm_{f}/dt) Oxidizer: Liquid oxygen (LOX), delivered at 3,982 lb/s(dm_{o}/dt) Total Propellant Flow (dm/dt): 5,736 lb/s Mixture mass ratio (r): 2.27:1 oxidiser to fuel The mixture mass ratio is the ratio of oxidiser to fuel for optimum combustion. This is not the same as the rocket's mass ratio which represents the efficiency of the mechanical construction of the rocket compared with its fuel load. Turbopump: 5,550 rpm, 41,000 kW single turbine, powered by a gas generator requiring 1,694 lb/s propellants, driving fuel and oxidiser pumps on the same shaft with a total flow rate of 2,542 litres/sec (1,565 l/s of LOX and 976 l/s of RP1) Thrust (F): 1,522,000 lbs at Sea Level Specific Impulse (I_{s}): F /(dm/dt) = 265.3 secs at Sea Level, 305 secs in vacuum. Exhaust Velocity (V_{e}):(I_{s}*g_{0}) = 8543 ft/s (5825 mph) Expansion ratio: 16:1 with nozzle extension, 10:1 without Combustion chamber pressure: 70 bars Combustion chamber temperature:3,300^{o}C Burn time: rated at 165 seconds  Fuel for Saturn V Main Engines  Hydrogen (LH_{2}) versus Kerosene (RP1)The thrust provided by rocket fuel is proportional to the energy density of the fuel and its propellant and the rate at which the fuel is burned. While liquid hydrogen (LH_{2}) has the highest energy density (energy per unit mass) of all fuels, over 30% more than kerosene, it also has the lowest physical density (mass per unit volume), only one twelfth the density of RP1. Thus RP1 has a greater energy content per unit volume than LH_{2}, while LH_{2} has a greater energy per unit mass. This means that to provide the same energy content as RP1, the fuel tanks, pipes and pumps and the structures needed to contain and transport the less physically dense LH_{2} will be disproportionately large compared with those needed for the kerosene fuel supply. This increases the final, (nonfuel) mass of the rocket, thus decreasing its mass ratio and hence its conversion efficiency. Minimising this nonfuel mass at lift off is particularly important when maximum thrust is required which is why RP1 is considered as an alternative. For lower thrust levels however, the relatively high mass of the fuel supply system needed to supply the liquid hydrogen is less significant compared with the gains made by using the more energy dense hydrogen fuel and there is a crossover point which occurs as the required thrust decreases when the higher energy, though less physically dense, hydrogen becomes the more energy efficient option. This is because the volume of the fuel system needed to contain the less dense hydrogen increases as the cube of the linear dimensions, but the weight of its fuel containers and pipes, which depends roughly on their surface area, only increases as the square of the linear dimensions. For very high, long duration thrusts such as those required from the S1C first stage of the Saturn V launch vehicle to get the heavy Apollo Space Vehicle off the ground, using the lighter hydrogen as the fuel would require an impractically large and heavy on board fuel supply system. For this reason kerosene with its lighter, more compact fuel supply system components was used to power the F1 rocket engines used in the S1C. Once the heavy stage 1 has been jettisoned and the rocket is operating in much reduced gravity, the required thrust is reduced and hydrogen becomes the most efficient option for fuelling the J2 engines powering the lighter stage 2 (S11) and stage 3 (S1VB) of the Saturn V. Details of the specialised propellants chosen for the the various other thrusters used on Saturn V and the rocket motors used on the Apollo 11 spacecraft are given in the table about Apollo11 Rocket Motorsbelow. 
Some Liquid and Solid Fuel CharacteristicsFuel Type  Fuel  Fuel Density ρ_{f} (g/cm^{3})  Fuel Boiling Point (deg C)  Fuel Specific Impulse (Secs)  Oxidiser  Oxidiser Density (g/cm^{3})  Oxidiser Boiling Point (deg C)  Oxidiser / Fuel Mix Ratio (r)  Density Specific Impulse of Mix Kg.secs/L  Density of Mix (g/cm^{3}) ***  Comments about the Fuel  Liquid Bipropellant Petroleum  Kerosene Paraffin (RP1)  0.820  216.3  265 (Sea level) 305 (Vacuum)  Liquid Oxygen (LOX)  1.14  183.0  2.29  264  1.03  Inexpensive, Practical. As with most liquid fuels, relatively easy to control, start and stop. Stable at room temperature. Complex ignition process. Low explosion hazard. Less energy per unit mass than hydrogen. More energy per unit volume than hydrogen Lower specific impulse than cryogenic fuels, but more than hypergolic propellants. Low temperature oxidiser needs insulation.  Liquid Bipropellant Cryogenic  Liquid Hydrogen (LH_{2})  0.071  252.9  425 (Vacuum)  Liquid Oxygen (LOX)  1.14  183.0  5.0  294  0.29  Very high specific impulse 30% to 40% higher than most other fuels Low temperature means difficult to store and handle. Needs insulated tanks. Very low density fuel needs large storage tanks and pumps.  Hypergolic  Hydrazine  1.004  113.5  286  Nitrogen tetroxide  1.45  21.15  1.08  342   Fuels and oxidizers ignite spontaneously on contact with each other. Easy to start, stop and restart Highly toxic and must be handled with extreme care. Remain liquid at room temperature. Relatively easy to control.  UDMH (Unsymmetrical dimethyl hydrazine)  0.791  63.9  277  Nitrogen tetroxide  1.45  21.15  2.10  316   Aerozine 50 (50/50% mix of Hydrazine with UDMH)    280  Nitrogen tetroxide  1.45  21.15  1.59  326   MMH (Monomethyl hydrazine)  0.866  87.5  280  Nitrogen tetroxide  1.45  21.15  1.73  325   Solid  Aluminium with HTPB (Hydroxy terminated Polybutadiene)    277  Ammonium perchlorate    2..12  474   Fuel contained in the combustion chamber. No tanks or pumps required. Compact, lightweight motor designs with a very high mass ratio. Safe, Easy to store, Quick to start Difficult to control. Needs an ignition system. Low specific impulse, but high thrust per unit volume. Allows lighter, simpler, and more reliable casing / combustion chamber designs.  Aluminium with PBAN (Polybutadiene Acrylonitrile)    277  Ammonium perchlorate    2.33  476  
*** Average Density ρ_{av} is given by: ρ_{av} = ρ_{o}ρ_{fv}(1+r) / (ρ_{f}r+ρ_{o})
Where ρ_{o} = The density of the oxidiser ρ_{f} = The density of the fuel r = The ratio of oxidiser mass to fuel mass 帝冢真织免费阅读_帝冢真织最新章节,free xxxx movies hd,中方回应香会主办方期待中国高级代表团与会 6分钟娇喘声
 Apollo 11 Rocket Motors and their Propellants  Application  Motor  NumberUsed  FuelType  Fuel  Oxidiser  PropellantFeed  SpecificImpulse(Secs)  Thrust(lbs)  GrossWeight(lbs)  Propellant Weightlbs / (%)  BurnTime(Secs)  Comments  Saturn V Launch Vehicle Stage 1 (SIC) (Uprated version)  F1 Propulsion  5  Petroleum  Kerosene RP 1 (Paraffin)  Liquid Oxygen LOX  Turbopump  289  1,530,000  4,792,000 (Stage1)  4,492,000 (93.7%) (Stage 1)  150  5 F1 engines giving the S1C a total thrust of 7,650,000 lbs  Retrorockets  8  Solid  Composite of polysulphides  Ammonium perchlorate  NA  277  87,913  504  278  0.633  Stage 1 2 separation  Saturn V Launch Vehicle Stage 2 (SII)  J2 Propulsion  5  Cryogenic  Liquid Hydrogen LH_{2}  Liquid Oxygen LOX  Turbopump  381  225,000  1,037,000  942,000 (90.8%) (Stage 2)  359  5 nonrestartable J2 engines giving the S11 a total thrust of 1,125,000 lbs  Retrorockets  4  Solid  Composite of polysulfides  Ammonium perchlorate  NA  277  34,810 each  377.5 each 帝冢真织免费阅读_帝冢真织最新章节,free xxxx movies hd,中方回应香会主办方期待中国高级代表团与会 6分钟娇喘声  268.2  1.52  Stage 23 separation  Ullage rockets  4  Solid  Flexadyne Polybutadiene (CTPB)  Ammonium perchlorate  NA  277  22,700 each  504 each  336  3.7  Stage 2 ullage  Saturn V Launch Vehicle Stage 3 (SIVB)  J2 Propulsion  1  Cryogenic  Liquid Hydrogen LH_{2}  Liquid Oxygen LOX  Turbopump  381  225,000  262,000  228,000 (87%) (excluding reserves)  480  1 restartable J2 engine 2 burns  Ullage rockets  2  Solid  Composite of polysulfides  Ammonium perchlorate  NA  277  3390   58.8  3.8  Main third stage ullage  Saturn V Launch Vehicle Stage 3 (SIVB) Auxiliary Propulsion System (APS)  Ullage  2  Hypergolic  MMH  Nitrogen Tetroxide  Helium pressurised tanks  280  70   303  50  2 APS in stage 3 1 Ullage motor in each APS Used during third stage restart  Attitude Control  6  Hypergolic  MMH  Nitrogen Tetroxide  Helium pressurised tanks  280  150   303  0.07  2 APS in stage 3 3 Attitude control thrusters in each APS  Apollo Command Module  Reaction Control System (RCS)  12  Hypergolic  UDMH  Nitrogen Tetroxide  Helium pressurised tanks  280  92   270  Variable   Apollo Service Module  Service Module Propulsion  1  Hypergolic  Aerozine 50  Nitrogen Tetroxide  Helium pressurised tanks  311  20,500  55,000  40,974  Variable  Nonthrottleable But can be switched on and off  Reaction Control System (RCS)  16  Hypergolic  MMH  Nitrogen Tetroxide  Helium pressurised tanks  280  100   1,362  Variable  16 used in groups of 4  Apollo LM Descent  LM Descent Motor Propulsion  1  Hypergolic  Aerozine 50  Nitrogen Tetroxide  Helium pressurised tanks  311  10,125max Variable 1,020 to 6,800  25,600*  19,500   Throttleable thrust *Gross weight without crew  Apollo LM Ascent  LM Ascent Motor Propulsion  1  Hypergolic  Aerozine 50  Nitrogen Tetroxide  Helium pressurised tanks  311  3,500  9,900*  5,200   *Gross weight without crew  Reaction Control System (RCS)  16  Hypergolic  MMH  Nitrogen Tetroxide  Helium pressurised tanks  290  100   605  Variable  16 used in groups of 4  Apollo Launch Escape System  Escape Motor  1  Solid  Composite of polysulfides  Ammonium perchlorate  NA  277  147,000  Total 8,910   4  Ejects Command Module from a dangerous launch  Tower Jettison  1  Solid  Composite of polysulfides  Ammonium perchlorate  NA  277  31,500    Jettisons tower after safe launch or when it is no longer required  Launch Vehicle Pitch control  1  帝冢真织免费阅读_帝冢真织最新章节,free xxxx movies hd,中方回应香会主办方期待中国高级代表团与会 6分钟娇喘声 Solid  Composite of polysulfides  Ammonium perchlorate  NA  277  2,400    Provides an initial pitch manoeuvre away from the launch pad toward the Atlantic Ocean in case of an abort  Total Apollo Rockets  87   An Earlier Example  The German WWII V2 Missile  V2Rocket  Motor  1  BioEthanol or Petroleum  Ethyl Alcohol (Ethanol)  Liquid Oxygen LOX  Turbopump  269  56,000  27,500 Including payload  19,301 (70.0%)  65  Reference(German V2 Missile) 

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