Turbopump set of rocket engine

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

 

The invention relates to rocket technology, particularly to a liquid-propellant rocket engines rocket engines, operating at cryogenic oxidizer and on hydrocarbon fuel.

Known liquid-propellant rocket engine for RF patent for the invention №2095607 designed for use in space boosters, speed boosters, and as the main engine of the spacecraft, includes a combustion chamber regenerative cooling channel and turbopump Assembly - TNA. TNA contains the pumps components - fuel and oxidant with the turbine on the same shaft, which introduced the capacitor. The output capacitor on the refrigerant line connected to the inlet into the combustion chamber and the entrance to the path of the regenerative cooling of the combustion chamber. The output of the capacitor to the line of fluid connected to the inlet to pump one of the components. The output of the pump of the same component in communication with the inlet of the condenser in the refrigerant line. The second input capacitor is in communication with the outlet of the turbine. The pump outlet of another component in communication with the entrance into the combustion chamber.

The disadvantage of TNA engine is the deterioration of the cavitation properties of the pump when the bypass condensate.

The known method of operation of rocket engines and liquid-propellant rocket engine for RF patent for the invention №2187684. The method of operation of a liquid-propellant rocket engine, the conclusion is to be in the flow of propellants into the combustion chamber of the engine, gasification of one of the components in the path of the cooling of the combustion chamber, supplying it on the turbopump turbine unit with subsequent discharge to the nozzle head of the combustion chamber. Part of the expense of one of the components of the fuel is directed into the combustion chamber, and the remaining part is gasified and is directed to the turbine of the turbopump assemblies. Exhaust turbines on the gaseous component is mixed with a liquid component flowing into the engine at a pressure higher than the vapour pressure of the resulting mixture. Liquid propellant rocket engine comprises a combustion chamber with a tract of regenerative cooling, pumps supply fuel components and turbine. Pumps and turbines arranged in two TNA: primary and booster. The engine contains a set of sequentially before the pump is one of the components of the main fuel turbopump Assembly, the booster pump turbopump Assembly and the mixer. The output of the main pump turbopump Assembly is connected with the nozzle head of the combustion chamber and the path of the regenerative cooling of the combustion chamber. Tract regenerative cooling, in turn, is associated with the turbines of the main and booster turbopump units whose outputs are connected to the mixer.

The disadvantage of this scheme is that the heat, remove the CR is the cooling of the combustion chamber, may not be enough to drive the turbopump Assembly of the engine of great power.

Known LRE by RF patent for the invention №2190114, IPC 7 F02K 9/48, publ. 27.09.2002, This rocket engine includes a combustion chamber with a tract of regenerative cooling, turbopump Assembly TNA pumps oxidant and fuel output line which is connected to the cylinder combustion chamber, the main turbine and the drive circuit of the main turbine. In the circuit drive the main turbine are connected in series between a fuel pump and a tract of regenerative cooling of the combustion chamber, which is connected to the input of the main turbine. The output from the turbine TNA is connected to the input of the second stage pump fuel.

This engine has a significant drawback. Bypass heated in the path of the regenerative cooling of the combustion chamber fuel input to the second stage pump fuel will lead to cavitation. Most LRE use such components in the fuel, the oxidant flow is almost always more fuel consumption. Therefore, for a powerful rocket engine having great traction and a lot of pressure in the combustion chamber, this scheme is not acceptable, because the fuel consumption will not be enough to cool the combustion chamber and drive the main turbine.

Furthermore, it is not designed launch system rocket engine, the ignition system is of komponentov fuel cut-off system rocket engine and clean the rest of the fuel in the path of the regenerative cooling of the combustion chamber.

Known liquid propellant rocket engine and how to run it on a Russian patent for invention №2232915, publ. 10.09.2003 g (prototype), which contains the combustion chamber, turbopump Assembly, generator, system startup, means for igniting the fuel components and fuel line. The output of the oxidizer pump is connected to the input of the generator. The output of the first stage fuel pump connected to the channels of regenerative cooling chamber and from the mixing head. The output of the second stage of the fuel pump (fuel pump) connected to the flow regulator with electric drive. The other input of the regulator is connected with a starting tank with regular fuel. The output of the regulator is connected to the gas generator. The output of the gas generator connected to the turbine inlet turbopump unit, the output of which is connected to a mixing head. The flow regulator is equipped with a hydraulic drive preliminary stages, through which the cavitating jet, hydrocele connected with starting a tank with regular fuel. Hydrocele connected with the second stage of the fuel pump. The choke installed on the output of the first stage fuel pump, done in conjunction with a controlled valve preliminary stages.

The disadvantage of this scheme is the low reliability of the seal between the primary and secondary fuel pumps and is behind a large differential pressure: 300...400 kgf/cm 2for modern rocket engines.

Objectives of the invention: improved reliability TNA due to full tightness of the seal between the fuel pump and additional fuel pump.

The solution of the stated problem is achieved due to the fact that the turbopump Assembly of the rocket engine, containing mounted on the shaft part of the rotor turbopump Assembly impeller of the oxidizer pump, the impeller of the fuel pump and the turbine wheel, is placed in the case turbopump unit, and an additional fuel pump with the shaft and impeller additional fuel pump, between the shaft turbopump Assembly and shaft of additional fuel pump made magnetic coupling. On the shaft turbopump Assembly has a driving disk magnetic coupling, and on the shaft of the additional fuel pump installed the clutch disc magnetic coupling. Between the master and the slave disk, a magnetic coupling is made partition of a nonmagnetic material. Partition of a nonmagnetic material combined with the optional pump fuel. Master and slave drives, and partition can be made spherical and/or fins.

Conducted patent studies have shown that the proposed solution has novelty, inventive step and industrial applicability. Novi is confirmed on the results of the patent research inventive step - achieve a new effect - an absolute tightness MGN.

Industrial applicability is due to the fact that all the elements included in the layout TNA, known from the prior art and are widely used in engine.

The invention is illustrated in figure 1...3, where figure 1 shows a diagram of the TNA; figure 2 and 3 versions of the magnetic coupling.

Turbopump Assembly of liquid-propellant rocket engine TNA 1 includes a shaft TNA 2, on which components are installed rotor TNA: impeller pump oxidizer 3, the impeller of the fuel pump 4 and the turbine wheel 5. All parts of the rotor TNA is located inside the housing TNA 6. Additional fuel pump 7 having the impeller additional fuel pump 8 and the shaft of the additional fuel pump 9, is made coaxial with the TNA 1 and installed on the side opposite the impeller of the turbine 5. The impeller additional fuel pump 8 is installed in the housing of additional fuel pump 10, the cavity of which "B" is sealed relative to the cavity TNA "And". Between the impeller of the fuel pump 4 and an additional fuel pump 7 in the housing TNA 6 installed magnetic clutch 11. The magnetic clutch 11 is composed of a master disk for magnetic coupling 12, which is installed in the housing TNA 6 and the slave drive magnetic clutch 13, which is installed in building the CoE additional fuel pump 10. Between the disks magnetic coupling 12 and 13 is a partition of a nonmagnetic material 14, for example from non-magnetic steel. Thus, the cavity of the fuel pump "A" and the cavity of the additional fuel pump "B" is completely sealed relative to each other.

The gas generator 15 is installed coaxially with TNA over 1 turbine nozzle device 16. The gas generator 15 includes a cylinder core 17, within which the outer plate 18 and the inner plate 19 with the cavity "B" above them and the cavity "D" between them. Inside the head of the gas generator 17 is installed in the oxidizer injector 20 and the fuel injector 21. Injector oxidizer 20 tell cavity "In" with the internal cavity of the generator D, and fuel nozzles 21 are informed of the cavity "D" with the internal cavity of the gas generator "D". On the outer surface of the core 15 has a fuel manifold 22, which fits the high pressure fuel line 23 from the secondary fuel pump 7. In line high pressure pipeline 23 is set high pressure valve 24 and the flow regulator 25 to the actuator of the flow regulator 26. The exit of the impeller of the fuel pump 4 is connected by a pipe 27 to the input of additional fuel pump 7 and the combustion chamber (combustion chamber 1 is not shown).

The exit of the impeller of the pump oxidizer 3 pipeline oxidizer 28 through klapa the oxidant 29 is connected to the cavity "B" gas generator 15. In the gas generator 15 is set to one or more of the firing device 30. The control unit 31 is connected to electrical connections with the firing device 30, the high pressure valve 24, valve oxidizer 29 and the actuator of the flow regulator 26.

Master and slave drives magnetic coupling 12 and 13 can be made with fins 32 (figure 2 and 3) to increase the strength of the shell TNA 6 and the body extra fuel pump 10 To reduce the thickness of the walls of non-magnetic material, it is, as well as the coupling halves 12 and 13 can be made spherical and/or fins 32 (1...3).

When you start LPRE with control unit 31 electrical signals are fed to the valves 24 and 29 and pilot (pilot) device 30. The oxidizer and fuel from the impellers of the pumps 3, 4 and 7 by gravity in the gas generator 15, which is ignited, the combustion cycle turbine wheel 5, the pressure at the outlet of the impellers of the pumps 3 and 4 increases. Part of fuel (about 10%) is supplied in additional fuel pump 7, where its pressure is greatly increased. Additional fuel pump 7 is driven by the slave disk 13 of the magnetic clutch 11. Consequently, due to the lack of seal on the shaft of the additional fuel pump 9 its reliability increases. When the pressure at the inlet to the impeller of the fuel pump 3 is of the order of R12...4 kgf/cm 2at the exit of the impeller pumps fuel 4 R2=300 kgf/cm2and when the outlet pressure of the additional fuel pump 7 is approximately R3=700 kgf/cm2arising between them a pressure drop of approximately 400 kgf/cm2perceived wall of non-magnetic material 14. The seal used in the prototype is missing, but on the impeller additional fuel pump 8 is transmitted torque from the shaft TNA 2.

The application of the invention allowed:

1. To increase the reliability TNA due to the lack of seal on the shaft of the additional fuel pump and its full integrity through the use of a magnetic coupling.

2. To simplify the kinematic scheme TNA due to the refusal of the gearbox.

3. To reduce the overall weight TNA due to elimination of the gearbox and its body.

4. To reduce pressure and leak fuel drain gearbox and from the cavity between the pumps of oxidizer and fuel.

5. To improve fire safety TNA due to:

- reduce the probability of contact of the oxidant and fuel in the cavity between the pumps of oxidizer and fuel,

- exceptions to design the cooling system gearbox flammable component of rocket fuel is flammable.

1. Turbopump Assembly of the rocket engine, containing mounted on the shaft part of the rotor turbopump Assembly: krill ATCU the oxidizer pump, the impeller of the fuel pump and the turbine wheel, is placed in the case turbopump unit, and an additional fuel pump with the shaft and impeller additional fuel pump, characterized in that between the shaft turbopump Assembly and shaft of additional fuel pump made magnetic coupling.

2. Turbopump Assembly of the rocket engine according to claim 1, characterized in that the shaft has a driving disk magnetic coupling, and on the shaft of the additional fuel pump installed a slave drive magnetic clutch.

3. Turbopump Assembly of the rocket engine according to claim 2, characterized in that between the leading and trailing drive magnetic clutch made partition of a nonmagnetic material.

4. Turbopump Assembly of the rocket engine according to claim 3, characterized in that the wall of non-magnetic material combined with the optional pump fuel.

5. Turbopump Assembly of the rocket engine according to claim 4, characterized in that the partition of a nonmagnetic material, master and slave drives are made spherical and/or fins.



 

Same patents:

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: 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 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: 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: 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: 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: 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: 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: 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.

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3 cl, 1 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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