Liquid propellant rocket. RU patent 2506444.
 Device and method for motorisation of rocket engine pump by means of inertia wheel / 2480608Motorisation device of pump (2), which feeds a rocket engine of a space vehicle, includes inertia wheel (1) and device for rotation transfer from an inertia wheel to a pump. As per a preferable version of the proposed invention, it includes measurement instruments of the wheel rotation speed and means (21) for detachment of wheel (1) and pump (2) at the rotation speed that is lower than the specified speed that is lower than nominal wheel rotation speed. This invention especially relates to a space vehicle containing a rocket engine, the feed system of which includes at least one pump driven with the device corresponding to this invention, as well as actuation means of the above device when the aircraft is in flight. |
 Device and method for driving rocket engine pump by internal combustion engine / 2477382Spacecraft rocket engine pump drive comprises aerobic internal combustion engine 1a, 1b running on the mix of fuel oxidiser, air-type fuel and hydrocarbon. Fuel oxidiser and fuel are fed to said engine by tanks and circuit separated from tanks 3 of propergols for rocket engine 16. Invention may be used in rocket engine supply device composed of, at least, two pump incorporating above described drive and ICE pump drive control means 8, 9 adapted to independently vary ICE operating parameters to allow independent control over pump rpm. |
 Lpre pump unit / 2406859Gearbox housing 8 incorporates casing 15 to embrace said housing to make peripheral space 16. Gearbox housing 8 has holes 17 to release cooling fluid from inner space of reduction gearbox 9 into peripheral space 16. Drain pipeline 12 serves to remove cooling fluid from reduction gearbox space 9 and has its ends connected to reduction gearbox 15 and intake branch pipe 18 of booster pump 1 to communicate reduction gearbox peripheral space 16 with booster pump intake branch pipe space 19. Described cooling circuit allows operating in reduction gearbox conditions, not in gear pump conditions, at minimum heating of pumped fluid. |
 Method of experimental confirmation of liquid-propellant rocket engine amplitude-phase characteristics and liquid-propellant rocket engine (versions) / 2406858Proposed method is based on measuring engine parametres responses to perturbation and comparing measured responses. In compliance with this invention, perturbation is produced by generating surges of pressure or fuel component flow rate in at least one lines of engine. Note here that said surge causes damped natural oscillations of the system, components and engine on the whole. Transient processes in engine parametres are noted by measurement system, while responses are compared with those of transient processes in perturbation. Pressure surge is generated at oxidiser or fuel pump inlet or outlet. Surge of flow rate is created in the line feeding oxidiser or fuel into gas generator. Surge of flow rate is created in the line feeding gas to turbo pump unit turbine. In compliance with first version, throttling valve can be arranged at oxidiser or fuel pump inlet or outlet with its time of operation smaller that period of oscillation in hydraulic system. In compliance with second version, engine incorporates at least one line to bypass fuel components at gas generator inlet. Said line incorporates throttling valve with its time of operation smaller that period of oscillation in hydraulic system. In compliance with third version, engine incorporates line to bypass gas around turbine of turbo pump unit with throttling valve with its time of operation smaller that period of oscillation in hydraulic system. |
 Multi-stage carrier rocket, method to launch it and three-component rocket engine / 2385274Invention relates to rocketry. Proposed carrier rocket comprises 1st- and 2nd-stages connected in parallel and having oxidizer and fuel tanks. Second-stage unit accommodates 2nd-fuel tank. Second-stage engines comprise combustion chamber and turbo pump unit arranged in parallel with combustion chamber or at an angle to it. Turbo pump unit comprises turbine, three-component gas generator, oxidizer pump, 2nd-fuel pump, extra 2nd-fuel pump, 1st-fuel pump and extra 1st-fuel pump. 2nd-fuel pump is mounted directly under oxidizer pump. Three-component gas generator outlet is communicated, via gas duct, with combustion chamber. Rocket engine comprises combustion chamber with jet nozzle incorporating regenerative cooling system, has generator, turbo pump unit consisting of turbine gas generator, oxidizer pump and fuel pumps. Turbo pump unit comprises two fuel pumps and two extra fuel pumps intended for consecutive operation on 1st and 2nd fuels without changing oxidizer. 2nd-fuel pump and extra pump are arranged directly under oxidizer pump and are connected, via cut-off valves, with gas generator. Launching method comprises simultaneous starting the engines of 1sr and 2nd stages operating on oxidizers and 1st-fuel, cutting off 1st-stage engines and detachment of1st-stage units. Thereafter, every engine of 2nd-stage is switched over to feeding by second fuel. |
 Liquid propellant rocket engine (versions) / 2301352Invention relates to liquid propellant rocket engine operating according to generatorless circuit. Liquid propellant rocket engine contains chamber, turbopump set to deliver propellant components (fuel and oxidizer) into mixing head of combustion chamber, control and regulation sets, and pipelines. According to invention, engine of first design version is furnished additionally with turbopump set to deliver third auxiliary component. Outlet space of pump is connected with cooling chamber duct, output of cooling duct is connected with blade space of turbines of auxiliary and main turbopump set, and through special nozzles with ambient medium or supersonic part of nozzle to create additional thrust. According to second design version, working medium of auxiliary component is directed into space of mixing head of combustion chamber. According to third design version, before-turbine part of auxiliary component, after getting out of duct of cooling chamber, is connected through restrictor with combustion chamber, and other part is directed through special nozzle into ambient medium or into supersonic part of nozzle of main chamber. Invention provides engine operating on hydrocarbon fuel with high value of specific thrust impulse (close of that of engine with after burning of generator gas). |
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Liquid propellant rocket engine (versions) / 2301352 Invention relates to liquid propellant rocket engine operating according to generatorless circuit. Liquid propellant rocket engine contains chamber, turbopump set to deliver propellant components (fuel and oxidizer) into mixing head of combustion chamber, control and regulation sets, and pipelines. According to invention, engine of first design version is furnished additionally with turbopump set to deliver third auxiliary component. Outlet space of pump is connected with cooling chamber duct, output of cooling duct is connected with blade space of turbines of auxiliary and main turbopump set, and through special nozzles with ambient medium or supersonic part of nozzle to create additional thrust. According to second design version, working medium of auxiliary component is directed into space of mixing head of combustion chamber. According to third design version, before-turbine part of auxiliary component, after getting out of duct of cooling chamber, is connected through restrictor with combustion chamber, and other part is directed through special nozzle into ambient medium or into supersonic part of nozzle of main chamber. Invention provides engine operating on hydrocarbon fuel with high value of specific thrust impulse (close of that of engine with after burning of generator gas). |
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Multi-stage carrier rocket, method to launch it and three-component rocket engine / 2385274 Invention relates to rocketry. Proposed carrier rocket comprises 1st- and 2nd-stages connected in parallel and having oxidizer and fuel tanks. Second-stage unit accommodates 2nd-fuel tank. Second-stage engines comprise combustion chamber and turbo pump unit arranged in parallel with combustion chamber or at an angle to it. Turbo pump unit comprises turbine, three-component gas generator, oxidizer pump, 2nd-fuel pump, extra 2nd-fuel pump, 1st-fuel pump and extra 1st-fuel pump. 2nd-fuel pump is mounted directly under oxidizer pump. Three-component gas generator outlet is communicated, via gas duct, with combustion chamber. Rocket engine comprises combustion chamber with jet nozzle incorporating regenerative cooling system, has generator, turbo pump unit consisting of turbine gas generator, oxidizer pump and fuel pumps. Turbo pump unit comprises two fuel pumps and two extra fuel pumps intended for consecutive operation on 1st and 2nd fuels without changing oxidizer. 2nd-fuel pump and extra pump are arranged directly under oxidizer pump and are connected, via cut-off valves, with gas generator. Launching method comprises simultaneous starting the engines of 1sr and 2nd stages operating on oxidizers and 1st-fuel, cutting off 1st-stage engines and detachment of1st-stage units. Thereafter, every engine of 2nd-stage is switched over to feeding by second fuel. |
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Method of experimental confirmation of liquid-propellant rocket engine amplitude-phase characteristics and liquid-propellant rocket engine (versions) / 2406858 Proposed method is based on measuring engine parametres responses to perturbation and comparing measured responses. In compliance with this invention, perturbation is produced by generating surges of pressure or fuel component flow rate in at least one lines of engine. Note here that said surge causes damped natural oscillations of the system, components and engine on the whole. Transient processes in engine parametres are noted by measurement system, while responses are compared with those of transient processes in perturbation. Pressure surge is generated at oxidiser or fuel pump inlet or outlet. Surge of flow rate is created in the line feeding oxidiser or fuel into gas generator. Surge of flow rate is created in the line feeding gas to turbo pump unit turbine. In compliance with first version, throttling valve can be arranged at oxidiser or fuel pump inlet or outlet with its time of operation smaller that period of oscillation in hydraulic system. In compliance with second version, engine incorporates at least one line to bypass fuel components at gas generator inlet. Said line incorporates throttling valve with its time of operation smaller that period of oscillation in hydraulic system. In compliance with third version, engine incorporates line to bypass gas around turbine of turbo pump unit with throttling valve with its time of operation smaller that period of oscillation in hydraulic system. |
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Lpre pump unit / 2406859 Gearbox housing 8 incorporates casing 15 to embrace said housing to make peripheral space 16. Gearbox housing 8 has holes 17 to release cooling fluid from inner space of reduction gearbox 9 into peripheral space 16. Drain pipeline 12 serves to remove cooling fluid from reduction gearbox space 9 and has its ends connected to reduction gearbox 15 and intake branch pipe 18 of booster pump 1 to communicate reduction gearbox peripheral space 16 with booster pump intake branch pipe space 19. Described cooling circuit allows operating in reduction gearbox conditions, not in gear pump conditions, at minimum heating of pumped fluid. |
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Device and method for driving rocket engine pump by internal combustion engine / 2477382 Spacecraft rocket engine pump drive comprises aerobic internal combustion engine 1a, 1b running on the mix of fuel oxidiser, air-type fuel and hydrocarbon. Fuel oxidiser and fuel are fed to said engine by tanks and circuit separated from tanks 3 of propergols for rocket engine 16. Invention may be used in rocket engine supply device composed of, at least, two pump incorporating above described drive and ICE pump drive control means 8, 9 adapted to independently vary ICE operating parameters to allow independent control over pump rpm. |
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Device and method for motorisation of rocket engine pump by means of inertia wheel / 2480608 Motorisation device of pump (2), which feeds a rocket engine of a space vehicle, includes inertia wheel (1) and device for rotation transfer from an inertia wheel to a pump. As per a preferable version of the proposed invention, it includes measurement instruments of the wheel rotation speed and means (21) for detachment of wheel (1) and pump (2) at the rotation speed that is lower than the specified speed that is lower than nominal wheel rotation speed. This invention especially relates to a space vehicle containing a rocket engine, the feed system of which includes at least one pump driven with the device corresponding to this invention, as well as actuation means of the above device when the aircraft is in flight. |
 Liquid propellant rocket / 2506444Invention relates to jet machinery. Liquid-propellant engine comprises engine chamber, turbine, fuel pump and jet pre-pump and differs from known designs in that jet pre-pump injection nozzle communicates with turbine inlet or outlet or with chamber cooling circuit. |
FIELD: engines and pumps.
SUBSTANCE: invention relates to jet machinery. Liquid-propellant engine comprises engine chamber, turbine, fuel pump and jet pre-pump and differs from known designs in that jet pre-pump injection nozzle communicates with turbine inlet or outlet or with chamber cooling circuit.
EFFECT: higher efficiency of jet pre-pumps.
3 dwg
The invention relates to the rocket engine and can be used in the development of liquid rocket engines (LPRE).
One of the key challenges when designing LPRE, is the creation, maybe, of simple construction, combined with high power characteristics.
The majority of modern rocket engines run engine turbopump feed system with fuel, and to provide the necessary cavitation stocks of major centrifugal pumps are used for additional pumps small that are in front of main pumps ().
In the Russian practice is most often used as axial flow pumps or type with screw cutting blades (augers), which are mounted on the same shaft with key impeller pump and are, thus, in the unit. However, for high-speed modern pumps auger is not enough, and then use free running lopastnogo type, operating from hydraulic and gas turbines or inkjet type liquid-to-liquid - prototype. Inkjet widely used on engines development of 1960-1970 (see «Construction and design of liquid rocket engines», .. and others, Moscow, Mashinostroenie, 1989, .224, 225). The advantages of inkjet (they are sometimes called jet or injectors) is their structural simplicity and reliability in operation. The disadvantages include low efficiency, which requires relatively large costs of high-pressure of the active liquid taken after the main pumps, which will impact negatively on the General power balance of unit. For example, to create pressure in 4-5 ATM component for the fuel supplied to the input of the main pump, nozzle jet need to submit a high-pressure liquid with a pressure of approximately 300 atmospheres in the amount of about 15-20% of the consumption component of the fuel supplied to the combustion chamber. This shortcoming was a barrier to the use of inkjet in the construction of modern rocket engines, differing extremely high levels of specific parameters.
The aim of the invention is to eliminate these shortcomings engines, using inkjet , namely, to increase the efficiency of jet
This goal is achieved by a liquid-propellant rocket engines, containing the camera engine, turbine, fuel pump and upstream towards him inkjet with nozzle injection, according to the invention of the injection nozzle jet reported to the entrance or exit from the turbine or with path of the cooling chamber.
In this case, the efficiency of the jet significantly increases due to higher adiabatic work on gas nozzles compared with adiabatic work of the liquid at the same differential pressure at the nozzle. For example, when triggered, the differential pressure of 300 atmospheres. the jet nozzles to create pressure in a passive stream of 4 ATM. (with inlet pressure of 4 ATM and efficiency, equal to 7%, the operating body - hydrogen) is required:
for liquid-to-liquid flow of the active liquid equal to 19% of the total flow into the pump;
for «gas-to-liquid (gas temperature (300 K) - 3% of the total consumption.
It should be borne in mind that for the normal work of the main pump TNA working gas jet must fully condense in a passive flow, and a rear passive flow caused by the condensation of steam, should not affect the cavitation breakdown of the main pump. This can be achieved provided that the working gas is a pair of a fuel component and with limitation of the ratio of costs active pair and passive liquid at a given temperature and pressure.
As in the majority of practical cases, require little additional pressure created , the aforementioned restrictive conditions must be met for almost all used in the operation easily propellant components: liquid hydrogen, liquefied natural gas, liquid oxygen, nitrogen tetroxide.
The best variant would be used as the active worker of the body clean steam fuel component, capable to complete condensation outlet jet pump. However, a variant of use of products of combustion in large excess of one of the components of fuel (combustion products contain, besides vapor component of fuel and other gases in small quantities, e.g., water vapour, carbon dioxide). In this case, if as a fuel component consider cryogenic product (for example, liquid hydrogen or oxygen), small admixtures will crystallize and not have a significant impact on the operation of the main pump TNA.
The invention in the variant when the injection nozzle jet reported to the exit from the turbine (i.e. when using the exhaust gases), is illustrated by picture engine, shown in figure 1.
Figure 1 presents:
1. Luggage engine.
2. Gas generator.
3. Pump fuel.
4. The oxidizer pump
5. Turbine.
6. Input highway fuel.
7. Inkjet .
8. Nozzle injection.
9. Line selection.
Engine, presented in figure 1, consists of a chamber 1, core 2, pump 3 fuel, oxidizer pump 4, turbine 5. Core 2 communicated with the turbine 5 and next to the camera 1. Pump fuel 3 communicated to the camera 1 and 2. Input highway fuel 6 represents the pipeline, which supplies the fuel pump 3, and the input line oxidant is an inkjet 7, according to which the oxidizer is served in the pump 4. Nozzle injection 8 feed on the highway selection 9, communicating with the output from the turbine 5.
Engine, presented in figure 1, is as follows. Components of fuel (for example, liquid oxygen as an oxidizer and liquid hydrogen as fuel) are fed into the engine. Oxidizer passes inkjet 7, which receives a small increment pressure. fuel in this scheme is missing. Components of fuel arrive in the main pumps 3 and 4, where the main pressure is created, and then oxidizer is served in full flow in the gas generator 2, where he gasified by the heat generated by the combustion of it a small amount of fuel supplied from the pump. Gas-fired oxidizer, consisting in this case, mainly from the vapor oxygen and a small amount of water vapour enters the turbine 5, bringing it into rotation, and later in the combustion chamber 1, where it reacts combustion of fuel, coming back to the main flow of the fuel pump 3. Part of the combustion products is taken after the turbine and is filed under excess pressure at the nozzle injection 8 jet 7, where the gas when triggered differential pressure accelerates and interacting with a liquid the main thread, it gives its kinetic energy, owing to what the main stream of fluid gets increment head. In this option, when the engine gas, having passed nozzle , condenses, while, with the remnants of water vapor, which crystallizes and in the General flow with the concentration of ice crystals order of 0.5-0.6% to the input of the pump 4 onwards according to diagram 1. Figure 2 shows a diagram of LPRE, where in contrast to the flow chart in figure 1, injection nozzle jet reported to the entrance of the turbine 5 line 9. The composition of the units and their symbols are the same as those in figure 1.
The invention in a use case reported with path of the cooling chamber, is illustrated in the scheme of LPRE, shown in figure 3.
Figure 3 presents:
1. Luggage engine.
3. Pump fuel.
4. Pump oxidant.
5. Turbine.
8. Nozzle injection.
9. Line selection.
10. Input line oxidant.
12. Fitting selection.
13. Inkjet .
14. Cooling tract camera.
15. Line supply.
16. Line-of-way.
Engine presented in figure 3, consists of a chamber 1, pump 3 fuel, oxidizer pump 4, turbine 5. The turbine is connected artery supplying 15 with cooling system camera, the output of a turbine connected to a camera highway drainage 16. Input line oxidant 10 is a pipeline that oxidizer is served in the pump 4 and later in chamber 1. Input highway fuel is an inkjet 13, which delivers fuel at the pump inlet 3. Nozzle injection 8 jet 13 reported with path of the cooling chamber by a selection of 9 and fitting selecting the 12 located in the cooling tract camera.
Engine presented in figure 3, is as follows. Components of fuel (for example, liquid oxygen as an oxidizer and liquid hydrogen as fuel) are fed into the engine. Fuel passes inkjet 13, where he received little additional pressure. Oxidizer comes on the input line is oxidant 10. oxidant in this scheme is missing. Components of fuel arrive in the main pumps 3 and 4, where the main pressure is created, and then oxidizer is served full-flow chamber 1, where it reacts burning gaseous fuel coming out of there main expenditure of fuel pump 3 through cooling tract camera 14, turbine 5 and exhaust line for 16.
Of cooling duct camera fuel in the form of pure vapor, in this case, hydrogen, partially withdrawn through the nozzle selection of 12 and on the trunk selection 9 is filed under excess pressure at the nozzle injection 8 jet 13, where steam when triggered differential pressure accelerates and interacting with a liquid the main thread, it gives its kinetic energy, owing to what the main stream of fluid gets increment head. Contact with the liquid vapor cools, condenses and further joint thread enters the pump fuel 3.
Thus, the use of the invention will allow to improve the internal energy of LPRE, simplify the design and thus increase its operational indicators (resource, profitability and reliability).
Liquid rocket engine, comprising a motor, turbine, fuel pump and upstream towards him inkjet with nozzle injection, wherein the injection nozzle jet reported to the entrance or exit from the turbine, or with path of the cooling chamber.