Liquid-propellant rocket engine and the method of its starting

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

 

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, 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 the engine is the deterioration of the cavitation properties of the pump when the bypass condensate.

Also known way to control the traction and component ratio of the rocket engines at the request of the Russian Federation for invention №93032897, publ. 27.07.1996, for implementation of which required an extremely complicated pneumohydraulic the Skye scheme and the electronic unit.

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 is the component feeding fuel into the combustion chamber of the engine, the 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. The engine contains a set of sequentially before the pump is one of the components of the main fuel turbopump Assembly 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 connected with the main turbines the booster turbopump units, the outputs are connected to the mixer.

The disadvantage of this scheme is that the heat removed in cooling the combustion chamber, may be insufficient 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 be weeks is sufficient for cooling the combustion chamber and drive the main turbine.

Furthermore, it is not designed launch system rocket engine, ignition system components and 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, (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 pump connected to the fuel 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 pump fuel made compatible with the STN with a controlled valve preliminary stages.

The disadvantages of this scheme Vlada:

- large system weight engine start,

- low efficiency due to the fact that complete combustion of the fuel components is never previsti 97...98%,

- the complexity of a pneumatic circuit, namely the existence of a large number of valves and regulators and uniting piping,

low reliability of the engine,

- afterburning of fuel when the engine is switched off.

Objectives of the invention: improving its efficiency, simplification of a pneumatic circuit, increased reliability, increased power and reduced weight of the engine, improving start-up and control the traction motor and provision of cleaning from the remaining fuel after shutdown.

These tasks are achieved due to the fact that liquid propellant rocket engine comprises a combustion chamber with a tract of regenerative cooling, turbopump Assembly with the turbine having an input and output line, and a pump oxidizer and fuel, which exits the fuel pump is connected through the valve of the fuel with the combustion chamber, and exits the pump oxidizer through the oxidizer valve connected to the gas generator, with turbopump Assembly contains an additional fuel pump, the inlet of which is connected to the output of the fuel pump, and the output is connected to the gas generator tubing you what the pressure will be about, where there is a high pressure valve and the flow regulator, in line turbine installed traction control, connected to side line and a ground line with a check valve and connector. Per additional pump fuel through the multiplier is connected to the shaft turbopump Assembly. The combustion chamber and the gas generator is equipped with pilot devices connected electrical connections with a power management system.

These tasks have been achieved in the way of starting liquid-propellant rocket engine, including the promotion of a turbopump Assembly and the opening of clonanav oxidizer, fuel, fuel in a high-pressure line, wherein the promotion turbines operate with compressed air from a ground tank and drive the turbine at the work carried out from the side of the container.

The invention is illustrated in the drawing.

Liquid propellant rocket engine comprises a combustion chamber 1 and turbopump unit 2 - TNA. Turbopump Assembly 2, in turn, contains the oxidizer pump 3, a fuel pump 4, a turbine 5, additional fuel pump 6, which shaft 7 is connected by a multiplier 8, accommodated in the housing of the multiplier 9, the shaft 10 turbopump Assembly 2. The gas generator 11 is installed directly above the combustion chamber 1, coaxially with it. The camera is orania 1 contains the nozzle 12, made of two shells with a gap "A" between them, and the head of the combustion chamber 13 within which is made of the outer plate 14 and the inner plate 15 with the cavity "B" between them. Inside the head of the combustion chamber 13 is installed in the oxidizer injector 16 and the fuel injector 17. Injector oxidizer 16 tell cavity "In" with the internal cavity of the combustion chamber D, and fuel nozzles 17 are informed of the cavity "B" with the internal cavity of the combustion chamber 1 - cavity "D". On the outer surface of the combustion chamber 1 has a fuel manifold 18, which runs the fuel lines 19 to the bottom of the nozzle 12. To the fuel manifold 18 is connected to the output of the fuel valve 20, the entrance to which the pipe 21 is connected to the output of the fuel pump 4. The output of the additional fuel pump 6 a high pressure pipeline 22 through the flow regulator 23, equipped with actuator 24, and the high pressure valve 25 is connected to the collector 26 of the gas generator 11 is placed in its lower part, and then through the cooling channel of the gas generator with a cavity "E". The head design of the gas generator 11 and the nozzle is accomplished similarly to the combustion chamber 1. The exit of the oxidizer pump 3 pipeline oxidant 27 through the oxidizer valve 28 is also connected to the gas generator 11, specifically with its cavity "W". On the head 13 of the combustion chamber 1 a of the firing device is TBA 29 (electrosupply or pilosellae), and the gas generator 11 - ignition device 30. Igniting devices 29 and 30 can be applied by one or a few pieces on the combustion chamber 1 and the gas generator 11.

To the turbine 5 docked inlet line 31 with traction control 32 with the actuator of the throttle 33.

The drive system of the turbine contains side highway 34 and the starting line 35 connected to the supply line 31. In the side of the highway 34 has side valve 36 and the side cylinder 37 with compressed air (gas). Starting line 35 contains successively installed check valve 38, the connector 39 and ground cylinder 40 with compressed air (gas).

To the fuel manifold 18 is connected pipeline flushing 41 with the purge valve 42.

Liquid rocket engine has a control unit 43. The control unit 43 is connected to electrical connections with pilot devices 29 and 30, the fuel valve 20, the valve oxidant 27, the actuator control valve 24, the high pressure valve 25, a working valve 32.

When you start LPRE with control unit 43 electrical signals are fed to the draught controller 32. The compressed gas acting on the starting line 35 from the surface of the cylinder 40, spins TNA 2. The pressure of the oxidant and fuel at the outlet of the oxidizer pumps 3, pump 4 fuel and additional fuel pump 6 age is AET. The signal for opening the valves 20, 25 and 27. The oxidant and fuel flows into the combustion chamber 1 and the gas generator 11. A signal to the ignition device 29 and 30, the fuel mixture in the combustion chamber 1 and the gas generator 11 is ignited. The engine started. Flow regulator 23 carry out the regulation of the ratio of components of fuel for its simultaneous consumption of the oxidizer and fuel tanks. The thrust of the engine is changed by the draught controller 32 by changing the flow rate of compressed air through the turbine 5.

After separation of the missile from the ground connector 39 is disconnected, the check valve 38 closes and opens side valve 36. Compressed air (gas) from the onboard tank 37 is supplied through the draught controller 32 to the turbine 5.

When shutting down the engine control unit 43 a signal to the damper 32, the valves 20, 25 and 27, which are closed. Then the signal to open the purge valve 39 and an inert gas, for example nitrogen, pipeline flushing 41 through the purge valve 42 is supplied to the fuel manifold 18 and forth in the cavity "A" to remove excess fuel.

The application of the invention allowed:

1. To simplify the pneumatic-hydraulic circuit of the engine.

2. To increase the reliability of the engine by simplifying the control circuit.

3. To provide for the regulation of engine thrust in widely the range of modes.

4. To increase the capacity and improve specific characteristics LRE due to more complete combustion of the components of rocket fuel, which is its two-stage combustion: first, in the gas generator at a sub-optimal ratio of components, with an excess of oxidizing agent, and then in the combustion chamber at the optimum value.

6. To improve the startup and shutdown of the engine due to the use of land can of compressed air (gas) to run.

7. To prevent high-frequency and low-frequency oscillations in the combustion chamber by placing the gas generator coaxially with the combustion chamber and directly above it.

8. To provide clearance from combustible residues of the cooling jacket of the combustion chamber after switching off the engine.

1. Liquid propellant rocket engine containing a combustion chamber with a tract of regenerative cooling, turbopump Assembly with the turbine having an input and output line, and a pump oxidizer and fuel, which exits the fuel pump is connected through the valve of the fuel with the combustion chamber, and exits the pump oxidizer through the oxidizer valve connected to the gas generator, wherein the turbopump Assembly contains an additional fuel pump, the inlet of which is connected to the output of the fuel pump, and the output is connected to the gas generator tubing high Yes the population, where there is a high pressure valve and the flow regulator, in line turbine installed traction control, connected to the side line and the starting line with a check valve and connector.

2. Liquid propellant rocket engine according to claim 1, characterized in that the shaft of the additional pump fuel through the multiplier is connected to the shaft turbopump Assembly.

3. Liquid propellant rocket engine according to claim 1 or 2, characterized in that the combustion chamber and the gas generator is equipped with pilot devices connected electrical connections with a power management system.

4. Start liquid-propellant rocket engine, including the promotion of a turbopump Assembly and the valve opening of the oxidant, fuel, fuel in a high-pressure line, wherein the promotion turbines operate with compressed air from a ground tank and drive the turbine at the work carried out from the side of the container.



 

Same patents:

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.

EFFECT: simplified construction; reduced mass of liquid propellant.

3 cl, 1 dwg

The invention relates to liquid propellant rocket engines (LPRE), particularly to a rocket engine turbopump with fuel consisting of separately stored oxidizer and fuel

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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