Method and device for rocket engine power supply

FIELD: engines and pumps.

SUBSTANCE: invention relates to a device for feeding the chambers of rocket engines (100) with the first and second propellant components. The first power supply circuit (16) of the thrust-generating chamber (10) includes a turbo pump (22) having at least one pump (22a) for pumping the first propellant component from a first tank (12) and a turbine (22b) mechanically connected to said pump (22a). The first power circuit connects the pump output to turbine input through a heat exchanger (24) configured for heating the first rocket fuel component by heat created by thrust-generating chamber in order to operate the turbine. In accordance with the present invention, the second power supply circuit (18) is arranged to supply the second propellant component creating the thrust chamber by the second propellant component from the second tank (14) which is configured to maintain high pressure. The present invention also provides a method for feeding the thrust-generating chamber of the rocket engine chamber with the first and second propellant components.

EFFECT: increased pressure in tanks containing rocket fuel components.

13 cl, 2 dwg

 



 

Same patents:

FIELD: engines and pumps.

SUBSTANCE: Liquid-propellant rocket engine comprises combustion chamber integrated with auxiliary chamber operated at excess of one of fuel components. In compliance with this invention turbopump unit with inlet connected with auxiliary chamber. Besides, said engine incorporates extra gas generator operated at excess second fuel component, its output being connected with combustion chamber nozzle head.

EFFECT: higher power output.

2 cl, 2 dwg

FIELD: aircraft engineering.

SUBSTANCE: proposed system comprises the unit of three cascades of interconnected pumps with independent drives, tank with oxygen and oxygen consumer. System inlet is connected with fuel tank while its outlet is connected with oxygen consumer. In compliance with this invention, this system is provided with high-pressure gas source with valve, mixer and gas consumer. Gas source is connected via valve with third cascade pump drive composed by the turbine. Third cascade turbine outlet is connected with gas consumer and first and second cascade pump drive gas inlets. Said cascades are composed of axial turbines arranged coaxially with appropriate pumps and secured therewith. First and second cascade turbine gas outlets are connected via mixer with first cascade pump liquid oxygen outlet. Note here that oxygen feed channels in first and second cascade pumps extend in diagonal and have axial inlets and outlets while third cascade pump is a rotary design. Method of liquid oxygen feed from tank to consumer consists in that oxygen is fed from tank to first cascade pump. Therefrom, oxygen is fed to second cascade pump. Therefrom, oxygen is fed to third cascade pump. Oxygen from third cascade pump is fed to consumer. Note here that first cascade pump oxygen pressure is increases with due allowance for second cascade pump cavitation-free operation. Liquid oxygen pressure in second cascade pump is increased to supercritical level. Maximum permissible rpm is set for third cascade pump. High pressure gas is fed from compressed gas source to third cascade turbine inlet. Gas energy is converted in third cascade turbine with pressure decreased into mechanical power. At third cascade turbine outlet gas is fed to consumer and to inlet of first and second cascade turbines. Gas energy is converted in first and second cascade turbines into mechanical power at decreased pressure to release this gas into mixer. Wherein this gas is mixed with oxygen flow forced from first cascade pump. Oxygen pressure upstream of third cascade pump is set higher than oxygen critical pressure by not over 10%. Third cascade pump rotor rpm is selected proceeding from rotor intensity parameter B defined by relationship B=Npn2, where Np is pump output, n is rotor rpm. Second cascade pump rotor rpm is set higher than that of first cascade pump rotor.

EFFECT: higher efficiency, decreased weight, higher reliability.

6 cl, 1 dwg

FIELD: engines and pumps.

SUBSTANCE: this turbo pump assy comprises oxidizer pump impeller fitted on the shaft, propellant pump and turbine wheel arranged in turbo pump housing. Note here that it comprises electrical generator with stator and rotor with shaft coupled with turbo pump unit shaft. Note also that magnetic clutch is fitted between turbo pump unit shaft and electrical generator clutch. Said magnetic clutch can be arranged between oxidizer pump and propellant pump. Magnetic clutch can be arranged between propellant pump and extra propellant pump.

EFFECT: ruled out turbo pump explosion at start or in flight.

4 cl, 4 dwg

FIELD: engines and pumps.

SUBSTANCE: proposed engine comprises control system with on-board computer, turbo pump unit and gas generator connected via gas duct with chamber and igniters at combustion chamber and gas generator. Spark plugs are arranged at combustion chamber and gas generator. Electrical generator is fitted on turbo pump unit shaft. Activator of gas generating mix is arranged inside aforesaid gas duct. Compressed air cylinder is connected to starting turbine. Gas generating mix activator can include two electrodes connected by HV conductors with HV unit connected with electrical generator. This engine can comprise central swivel made at gas duct on chamber axis. Central swivel can be shaped to cylinder. Central swivel can be shaped to sphere. This engine can incorporate turbo pump unit shaft rpm transducer connected with on-board computer.

EFFECT: higher specific thrust and reiterated starting.

11 cl, 17 dwg

FIELD: engines and pumps.

SUBSTANCE: proposed rocket engine comprises chamber, gas-generators, fuel pumps and two-stage turbine fed by power gas. First stage outlet is communicated with the chamber nozzle head In compliance with this invention, turbine second stage outlet is communicated with booster pump inlet to turbine housing to feed one of fuel components. Outlet of said pump is communicated with the engine inlet or with ambient atmosphere.

EFFECT: higher power efficiency owing to complete application of gas discharged after second stage.

2 cl, 3 dwg

FIELD: engines and pumps.

SUBSTANCE: this turbopump unit comprises turbine, oxidiser pump and fuel pump with impellers. In compliance with this invention, this turbine represents a birotatory design with two impellers. The latter have no nozzle diaphragms and run in opposite directions, each being communicated with oxidiser pump impeller and fuel pump impeller. Turbopump unit can incorporated an extra fuel pump, impellers of extra fuel pump and fuel pump being fitted on one shaft.

EFFECT: decreased centrifugal loads at turbine rotor.

2 cl, 2 dwg

FIELD: engines and pumps.

SUBSTANCE: proposed turbopump unit comprises oxidiser pump, turbine running on gas fuel, turbine bearing, system of seals separating said oxidiser pump from said turbine. Gas drainage between the system of seals and turbine, seal being fitted on turbine side Turbine bearing is arranged in chamber between said seal and turbine cavity.

EFFECT: decreased losses of separating gas in drainage, better dynamics of rotor.

4 cl, 4 dwg

FIELD: engines and pumps.

SUBSTANCE: invention relates to rocket engineering. Liquid propellant rocket engine comprises a turbopump set, which contains in its turn the following components installed on a shaft: a turbine, oxidant and fuel pumps, and a combustion chamber having a cylindrical part with oxidant and fuel nozzles and a nozzle with the main header of fuel and a system of regenerative cooling, according to the invention, the turbopump set and the combustion chamber are installed coaxially, the combustion chamber is made as two-zone and comprises the first circular zone with a circular nozzle block and an upper fuel header, and the second zone with a central nozzle block, having additional fuel nozzles, and the turbine is installed between the first and second zones of the combustion chamber. The outlet from the oxidant pump is connected by the pipeline comprising an oxidant valve with a combustion chamber. The outlet from the fuel pump is connected by pipelines with the main and upper fuel headers. The turbine is made as consisting of the nozzle block, an impeller and a straightening unit with a cavity inside of it, a central nozzle block is made as hollow, and its cavity is connected by axial holes via the cavity inside the straightening unit with a gap of regenerative cooling of a nozzle and the second zone of the combustion chamber.

EFFECT: invention provides for improvement of specific characteristics of an engine and also its increased reliability.

5 cl, 3 dwg

FIELD: engines and pumps.

SUBSTANCE: liquid propellant rocket engine comprising a turbopump set, which contains the following components installed on a shaft: a turbine, oxidant and fuel pumps, and a combustion chamber having a cylindrical part with oxidant and fuel nozzles and a nozzle with the main header of fuel and a system of regenerative cooling, according to the invention, the turbopump set and the combustion chamber are installed coaxially, the combustion chamber is made as two-zone and comprises the first circular zone with a circular nozzle block and an upper fuel header, and the second zone with a central nozzle block made in the form of a hollow cylinder, having axial additional fuel nozzles, and the turbine is installed between the first and second zones of the combustion chamber, the turbine is made as consisting of the nozzle block, an impeller and a straightening unit with a cavity inside of it, a central nozzle block is made as hollow, and its cavity is connected by axial holes via the cavity inside the straightening unit with a gap of regenerative cooling of a nozzle and the second zone of the combustion chamber, and the cavity inside the straightening unit with slot holes is connected to the second zone. The outlet from the oxidant pump is connected by the pipeline comprising an oxidant valve with a combustion chamber. The outlet from the fuel pump is connected by pipelines with the main and upper fuel headers.

EFFECT: increased specific characteristics of a liquid propellant rocket engine and its increased reliability.

3 cl, 3 dwg

FIELD: motors and pumps.

SUBSTANCE: liquid fuel rocket motor containing turbopump unit with the turbine installed on the shaft, oxidant and fuel pumps, and combustion chamber comprising a cylindrical part with oxidant and fuel injectors, and nozzle with the main fuel manifold and regenerative cooling system, according to the invention turbopump unit and combustion chamber are installed coaxially, the combustion chamber is designed as two-zonal and contains the first annular zone with ring-type spray unit and upper fuel manifold and the second zone with the central spray unit comprising additional fuel injectors, and the turbine is installed between the first and second zones of the combustion chamber. The oxidant pump output is connected with the combustion chamber through the pipeline with the oxidant valve. The fuel pump output is connected through pipelines with the main and upper fuel manifolds. The turbine is designed comprising the nozzle device, impeller and straightener blades with a cavity inside, central spraying unit is designed hollow and its cavity is connected with axial openings through a cavity inside straightener blades with the gap of regenerative cooling nozzle and the second zone of the combustion chamber.

EFFECT: better LFRM specific characteristics and higher reliability.

5 cl, 2 dwg

FIELD: liquid-propellant rocket engines.

SUBSTANCE: proposed system includes lines of different propellant components, gas lines and gas generator. Mounted in front of main turbo-pump unit with multi-stage propellant component pump and drive gas turbine is booster turbo-pump unit with propellant component pump and drive single-stage hydraulic turbine fed from "n" stage of main pump. Working passage of hydraulic turbine is located in connecting line between outlet of "n" stage and inlet of "(n+1)" stage of main pump.

EFFECT: simplified construction; reduced mass of liquid propellant.

3 cl, 1 dwg

FIELD: rocketry, in particular, liquid-propellant rocket engines using helium as a cooler of the engine chamber body.

SUBSTANCE: the liquid-propellant rocket engine has an engine chamber consisting of a combustion chamber and a nozzle, having regenerative-cooling ducts, turbopump assembly including centrifugal pumps of oxidizer, fuel and helium, neutral gas generator fed from the pumps of oxidizer and fuel, and the outlet of the helium pump is coupled to the regenerative cooling passage of the combustion chamber, whose outlet is coupled to the mentioned gas generator, the outlet of the gas generator is coupled to the turbine of the turbopump assembly, whose outlet is coupled to the oxidizer supply line to the combustion chamber mixing head. Besides, cooling of the chamber nozzle is effected by fuel, which, having passed through the regenerative cooling ducts, is supplied to the mixing head. The combustion chamber of the engine chamber and the gas generator operate at a stoichiometric relation of the fuel components. Introduction of the helium additive to the combustion products of the main fuel components to the neutral gas generator and further to the engine combustion chamber makes it possible to enhance the engine specific thrust pulse approximately by 20S, and, with regard to denial of screen cooling, approximately to 30S and more.

EFFECT: enhanced engine specific thrust pulse.

1 cl, 1 dwg

FIELD: rocket and space engineering.

SUBSTANCE: proposed liquid-propellant rocket engine has chamber. Bypass main line with flow rate regulator passing part of fuel into chamber by-passing cooling duct is installed in parallel to line of fuel delivery for cooling.

EFFECT: increased service life of engine chamber owing to reduction of thermal stresses in inner wall.

1 dwg

FIELD: classic and return launch vehicles.

SUBSTANCE: proposed low-thrust cryogenic propulsion module contains main cryogenic engine 10, two auxiliary engine 21, 22 to control position in space, cryogenic ergol supply tanks 31, 32, 33, 34, device for periodically building pressure in tanks 31, 32, 33, 34 and device to generate explosive pulses of main cryogenic engine at pulse mode during period of pressure building in tanks 31, 32, 33, 34. Device to periodically build pressure in tanks 31, 32, 33, 34 has heat exchange system connected with heat accumulator 61, 62 and device 71,72 to excite circulation of preliminary set amount of ergol in heat exchange system. Module contains additionally device for heating heat accumulator 61, 62 in period between two sequential explosive pulses.

EFFECT: improved mass-and-dimension characteristics of module, reduced time taken for execution of task.

17 cl, 14 dwg

FIELD: rocketry.

SUBSTANCE: according to proposed method of creating thrust of liquid-propellant rocket engine with circulation of heat carrier based on taking of propellant components from tanks, increasing their pressure by pumps driven by turbine and introducing them into gas generator and combustion chamber, combustion of components in gas generator and chamber and creating thrust with ejection of combustion products through nozzle, when introducing component of propellant and products of their gasification into combustion chamber, tangential speed component is imparted to them and part of combustion products is replaced by heat carrier and in process of recirculation it is successively expanded at higher pressure of diverting part of nozzle, cooled, condensed in head exchanger-condenser, pressure is raised by pump and it is then delivered to near-critical part of nozzle to repeat the cycle. Liquid-propellant rocket engine with closed circuit of heat carrier contains chamber with mixing head and regenerative cooling duct, turbopump set with oxidizer and propellant pumps whose output main lines are connected with said mixing head of chamber and gas generator, and said closed circuit of heat carrier is formed with successively interconnected circulating pump, unit to introduce heat carrier to near-critical area of nozzle, heat exchanger-condenser, means to supply condensed component to input of circulating pump. According to invention, closed circuit is provided with section of diverting part of nozzle on which ring ribs made of heat-resistant material are secured over circumference.

EFFECT: reduced cost of launching of useful load into orbit.

5 cl, 3 dwg

FIELD: liquid propellant rocket power plants with turbopump units.

SUBSTANCE: the liquid propellant rocket power plant having liquid-hydrogen and liquid-oxygen tanks with booster pumps and main turbopump units uses also an electrochemical generator with an oxygen inlet and outlet and a hydrogen inlet and outlet, oxygen ejector, hydrogen ejector and two electric motors, one of which is connected to the shaft of the oxygen booster pump, and the other-to the shaft of the hydrogen booster pump, the oxygen inlet of the electrochemical generator is connected through a pipe to the gas cushion of tank with liquid oxygen, and the outlet-to the inlet of oxygen ejector, whose outlet is connected to the gaseous oxygen supply pipe to the reaction chamber: the hydrogen inlet of the electrochemical generator is connected through a pipe to the gas, cushion of the tank with liquid hydrogen, and the outlet is connected to the inlet of the hydrogen ejector, whose outlet is connected to the gaseous hydrogen supply line to the reaction chamber.

EFFECT: provided reliable multiple starting of the liquid propellant rocket power plant.

1 dwg

FIELD: rocket engineering; production of the booster turbo-pump aggregates with an axial pumps used in them.

SUBSTANCE: the invention is pertaining to the field of rocket engineering, in particular, to production of the booster turbo-pump aggregates with the axial pumps used in the liquid rocket engines (LRE). The booster turbo-pump aggregate consists of the body (1), in which there is the pump (3) axial wheel fixed on the shaft (2) and the wheel of the hydraulic turbine (4). The wheel of the hydraulic turbine (4) is connected to the axial wheel of the pump (3) by soldering along its outer diameter. The shaft (2) rests on the fixed bearing (5) and on the movable bearing (6). The axial stops (7) and (8) of the body (1) eliminate the possibility of the axial motion of the fixed bearing (5) with respect to the body (1), and consequently, the motion of the shaft (2). The movable bearing (6) may have the shift in the axial direction concerning the body (1) because of the difference of the axial power and thermal deformations of the body (1) and the shaft (2). From the side of the axial intake in the body there is the axial stop (9). Between the axial stop of the body (1) and the movable bearing (6) the axial spring (10) is installed. The support ring (11) is mounted between the axial spring (10) and the movable bearing (6). The axial spring (10) is made in the form of the resilient conical ring. The invention ensures the increased service life of the fixed bearing (5) and the service life of the whole aggregate.

EFFECT: the invention ensures the increased service life of the fixed bearing (5) and the service life of the whole aggregate.

3 cl, 1 dwg

FIELD: rocket engineering; production of the devices for the liquid propellant rocket engines.

SUBSTANCE: the invention is pertaining to the field of rocket engineering and may be used in the liquid propellant rocket engines (LPRE). The device for separation of the pump and the turbine of the booster turbo-pump aggregate of the LPRE consists of the pump (2), the turbine (3), the separating cavity (1) located between the pump (2) and the turbine (3) and the external intake tract (4). The separating cavity (1) is limited from the side of the pump (2) by the shaft gasket (5), which diameter is made smaller than the diameter of the shaft (10) in the area of the seat of the bearing of the turbine (11), and from the side of the turbine (3) - by the unloading disk (6) aligned with the turbine impeller (7). On the turbine impeller (7) there is the gasket of the unloading disk (8). The axial impeller of the pump (9) and the turbine impeller (7) are fixed on the shaft (10). From the direction of the turbine (3) the shaft (10) rests on the turbine bearing (11), which is brought out beyond the bounds of the separating cavity (1) and is installed from the direction of the pump (2). The cavity of the turbine bearing (12) which is adjoining the shaft gasket (5) is connected by the delivery channels (13) with the pump outlet (14). The offered device ensures the minimum losses of the power used for separation of the pump and the turbine, and also the effective refrigeration of the bearings by the liquid monophase hydrogen.

EFFECT: the invention ensures the minimum losses of the power used for separation of the pump and the turbine, the effective refrigeration of the bearings by the liquid monophase hydrogen.

2 cl, 2 dwg

FIELD: rocket engineering; production of the liquid propellant rocket engines.

SUBSTANCE: the invention is pertaining to the field of rocket engineering, in particular, to production of the liquid propellant rocket engines powered by the cryogenic oxidant and the hydrocarbon propellant. The liquid propellant rocket engine contains the combustion chamber with the tract of the regenerative cooling, the turbo-pumping aggregate with the turbine having the inlet and outlet trunks, and the pumps of the oxidant and the propellant, for which the outlet of the propellant p[ump is connected through the propellant valve to the combustion chamber, and the outlet of the oxidant pump through the oxidant valve is connected to the gas generator. At that the turbo-pump aggregate contains the additional propellant pump, which inlet is connected to the outlet of the propellant pump, and the outlet is connected to the gas generator through the high pressure pipeline, in which there is the high-pressure valve and the consumption regulator. In the trunk of the turbine there is the thrust regulator, to which the on-board trunk and the starting trunk with the return valve and the connector are connected. The method of the liquid propellant rocket engine starting provides for the spinning-up of the turbo-pump aggregate and opening of valves of the oxidant, the propellant, the propellant in the high-pressure trunk, run-up of the turbine conduct a compressed air from a land bulb, and the turbine spinning-up is exercised by the compressed air from the ground pressure vessel and the turbine drive at operation is exercised from the on-board vessel. The invention ensures simplification of the pneumatic-hydraulic circuit, the increased reliability, the increase of the power and the specific characteristics of the liquid propellant rocket engine, the decreased mass of the engines, the improved engine starting and cutoff and provision of the engine cleansing from the leavings of the propellant after the engine cutoff.

EFFECT: the invention ensures simplification of the liquid propellant rocket engine pneumatic-hydraulic circuit, the increased its reliability, power and specific characteristics, the decreased mass of the engine, the improved the engine starting, cutoff and cleansing from the leavings of the propellant after its cutoff.

4 cl, 1 dwg

FIELD: rocketry.

SUBSTANCE: invention relates to liquid-propellant rocket engines operating on cryogenic oxidizer and on hydrocarbon fuel. Proposed turbopump set of rocket engine contains the following parts of rotor of turbopump set mounted on shaft: oxidizer pump impeller, fuel pump impeller and turbine wheel arranged in housing of turbopump set and additional fuel pump with shaft and impeller of additional fuel pump. Design peculiarity of turbopump set is that magnetic clutch is installed between rotor of turbopump set and rotor of additional fuel pump. Driving disk of magnetic clutch is installed on shaft of turbopump set, and driven disk is mounted on shaft of additional fuel pump. Partition made on nonmagnetic material is found between driving and driven disks of magnetic clutch. Said partition is aligned with housing of additional fuel pump. Partition, driving and driven disks can be made spherical and/or provided with ribbing.

EFFECT: increased reliability of turbopump set.

5 cl, 3 dwg

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