Liquid propellant rocket engine

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

 

The invention relates to liquid propellant rocket engines rocket engines, working on three components of the fuel, oxidizer and two combustible, and is aimed at improving specific characteristics and reducing the cost of launching missiles on which it is mounted, and a significant improvement of its many characteristics like range, etc. the best option - use as an oxidizer, liquid oxygen, the first fuel - kerosene, the second fuel is liquid hydrogen.

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 chamber when orania. The disadvantage of the 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 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 tract the regenerative cooling the Oia 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 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, MPK7 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 midnihgt, 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, 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 for RF patent for the invention №2232915, publ. 10.09.2003, which contains a camera, 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 staff is ucim. 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 is a complex pneumatic-hydraulic circuit of the engine, the large number of valves and regulators and uniting piping and, as a consequence, a large weight and low reliability and problems at start-up and shutdown of the engine.

The objective of the invention is to improve specific characteristics of LRE, increase its reliability and reduce the cost of launching rockets.

These objectives are achieved in liquid-propellant rocket engine, containing turbopump Assembly, containing in turn mounted on shaft turbine pumps oxidant and fuel, and a combustion chamber having a cylindrical part with jets of oxidant and fuel and the nozzle from the main fuel manifold and system, regenerative cooling, according to the invention of the turbopump Assembly and combustion chamber mounted coaxially, the combustion chamber is made and dual-zone contains the first ring zone to the annular nozzle block and the top of the fuel manifold and a second area with a Central nozzle block having additional fuel nozzles and turbine is installed between the first and second areas of the chamber is gorania. The exit of the oxidizer pump is connected by a pipe containing a valve oxidant, with the combustion chamber. The output of the fuel pump can be connected by pipelines to the main and upper manifolds of fuel. The turbine can be made consisting of a nozzle apparatus, the impeller and directing vanes with a cavity inside, the Central nozzle block is made hollow and its cavity are connected by axial holes through the cavity inside the straightener apparatus with a gap of regenerative cooling of the nozzle and the second zone of the combustion chamber.

The invention is illustrated in the drawings, figure 1...3, where

- figure 1 shows the scheme of LRE,

- figure 2 shows the design of the combustion chamber,

- figure 3 shows the design of the turbine.

Liquid propellant rocket engine is a rocket engine (1...3) contains the camera 1 with the combustion chamber 2 with the nozzle 3, turbopump Assembly TNA 4.

Camera 1 and TNA 4 mounted coaxially and sequentially.

The combustion chamber 2 (Fig 1 and 3) comprises a cylindrical part 5. The nozzle 3 has a tapering portion 6 extending portion 7 and the main fuel manifold 8 in the lower part. The combustion chamber 2 is made dual-zone and contains the first and second zones 9 and 10, respectively. The first zone 9 has an upper fuel manifold 11.

THA 4 contains sequentially mounted on the same shaft 12 from the bottom up: Urbino 13, the oxidizer pump 14, the fuel pump 15. Feature THA 4 is that the turbine 13 is installed in the combustion chamber 2, more precisely at the top of its cylindrical part 5 between the first and second zones, respectively, 9 and 10. The first area 9 of the combustion chamber 2 is made annular and includes an annular head 16 and the annular nozzle unit 17. The annular nozzle block 17 has a nozzle oxidizer 18 and 19 fuel.

In the second zone 10 of the combustion chamber 2 has a Central nozzle unit 20 in the form of a hollow cylinder with a side wall 21 and end wall 22 and the inner cavity 23. The end wall 22 installed additional axial fuel nozzles 24 and on the side wall 21 is a radial fuel nozzles 25.

Turbine 13 contains to allow the device 26, the impeller 27, the disc 28 and a rectifying device 29. A rectifying device 29 is made with a cavity. 30 and the holes 31 and 32 for communication with the path of the regenerative cooling 33 and the inner cavity 23.

THA 4 has three pillars 34...36 and the upper seal 37 (Fig 1 and 2).

As the tapering portion 6, and the expanding portion 7 of the nozzle 3 is made with the possibility of regenerative cooling (figure 1 and 2) and contain two walls: an inner wall 38 and outer wall 39 a tract of regenerative cooling 33 between them for the passage of cooling gas and the cooling nozzle. Tract regenerative what about the cooling 33 communicates with the cavity 40 of the main fuel manifold 8. Inside the combustion chamber 2 (Fig 1 and 2) is made from the upper plate 41 and the bottom plate 42 with a gap (cavity) between them 43. Above the upper plate 41 is made of the cavity 44. Inside the annular head 16 of the combustion chamber 2, as mentioned earlier, the installed injector oxidizer 18 and fuel nozzles 19. Injector oxidizer 18 report of the cavity 44 with an internal cavity 45 of the combustion chamber 2. Fuel nozzles 19 are informed of the cavity 43 with an internal cavity 45. The ring head 16 of the combustion chamber 2 installed ignition device 46. Ring head 16 of the combustion chamber 2 is attached by multiple pipelines, 47 containing the oxidizer valves 48, TNA 4. Ring head 16 of the combustion chamber 2 and the oxidizer pump 14 is connected intermediate casing 49.

To the main fuel manifold 8 is attached to the pipe 50 with the valve 51. The other end of the pipe 50 is connected to the output of the fuel pump 15.

The upper manifold 11 by the pipe 52, containing the flow regulator 53 and valve 54 also is connected to the output of the fuel pump 15. Additional fuel pump is missing. This simplifies the design of the rocket engine and reduces its weight. Coordination of hydraulic resistances tract regenerative cooling 33 and the small value of the cooling channel of the first zone 9 of the combustion chamber 2 leads to the fact that a large part of the fuel consumption goes through the OS is new additional fuel nozzles 24 and radial fuel nozzles 25, as they carry out its injection into the second zone 10 with a significantly lower pressure than the pressure in the first zone 9. This is due to the pressure loss in the turbine 13.

The supports 34 and 35 are installed in the intermediate casing 49, which is made between TNA 4 and the combustion chamber 2 and contains a protective sleeve 55, is made within the first zone 9 of the combustion chamber 2. Inside the protective sleeve 55 in the lower seal 56. (figure 2). The lower seal 56 prevents the exit of the cooling support 36 fuel into the combustion chamber 2.

For reliable cooling of the bearings 36 applied system cooling fuel, which contains made inside the shaft 12 in the axial hole 57, the input radial hole 58, the intermediate radial bore 59 extending into the gap 60 between the protective sleeve 55 and the shaft 12, and the output of the radial hole 61 made in a protective sleeve 55 and telling the gap 60 with a cavity 43 for use of cooling of the fuel by burning it in the combustion chamber 2.

The engine contains a purging system cylinder 62 with the inert gas, the pipe 63 and valve 64. The pipe 63 is connected to the main fuel manifold 8.

Start LRE is as follows.

Open the oxidizer valves 48 and valves 51 and 54.

The oxidizer and fuel enter the combustion chamber 2, and more precisely in its first zone 9, where ignite Zap lname devices 46 and burn at relatively low temperatures of 500 to 700°C. Through axial additional fuel nozzles 24 and radial fuel nozzles 25 a large part of the fuel fed to the second zone 10, where it is burned at a temperature of from 3500 to 4000°C for maximum specific impulse rocket engine.

Regulation mode LRE is a flow regulator 53. It modifies the flow of combustion products through the turbine 13 and the temperature of the combustion products on the entrance.

Off LRE is in the reverse order. After closing all valves oxidant 48 and fuel 51 and 54 open the purge valve 64, and the inert gas from the cylinder 62 through the pipe 63 is supplied to the main fuel manifold 8 equalising the combustion chamber 2 of the engine for cleaning residue from the fuel.

The use of the invention will:

To reduce the transverse dimensions of the rocket engine by placing the combustion chamber and TNA sequentially in a single line along one common axis.

To reduce the weight of the engine due to the lack of fillers and an additional fuel pump. The function generator performs the first zone of the combustion chamber.

- To simplify the design of the engine by reducing the number of pipelines.

- To lessen the mutual influence of hot and cold (cryogenic propellant components) components and assemblies.

To reduce the imbalance of the shaft TNA due to the use of t is ex supports and schemes of their placement so in order to bring to the rotating parts of the rotor.

To improve the cooling of the bearings TNA due to the use of cooling support, placed directly around the impeller 27 of the turbine 13 inside the combustion chamber 2, i.e. in the zone of high temperatures.

1. Liquid propellant rocket engine, containing turbopump Assembly, containing in turn mounted on shaft turbine pumps oxidant and fuel, and a combustion chamber having a cylindrical part with jets of oxidant and fuel and the nozzle from the main fuel manifold and system, regenerative cooling, characterized in that the turbopump Assembly and combustion chamber mounted coaxially, the combustion chamber is performed by duhaney and contains the first ring zone to the annular nozzle block and the top of the fuel manifold, and a second zone with a Central nozzle unit, made in the form of a hollow cylinder having an axial additional fuel nozzles and radial fuel nozzles, and turbine is installed between the first and second zones of the combustion chamber.

2. Liquid propellant rocket engine according to claim 1, characterized in that the exit of the oxidizer pump is connected by a pipe containing a valve oxidant, with the combustion chamber.

3. Liquid propellant rocket engine according to claim 1 or 2, characterized in that the output of the fuel pump connection is replaced by pipelines to the main and upper manifolds of fuel.

4. Liquid propellant rocket engine according to claim 1 or 2, characterized in that the turbine is made consisting of a nozzle apparatus, the impeller and directing vanes with a cavity inside, the Central nozzle block is made hollow and its cavity are connected by axial holes through the cavity inside the straightener apparatus with a gap of regenerative cooling of the nozzle and the second zone of the combustion chamber.

5. Liquid propellant rocket engine according to claim 3, characterized in that the turbine is made consisting of a nozzle apparatus, the impeller and directing vanes with a cavity inside, the Central nozzle block is made hollow and its cavity are connected by axial holes through the cavity inside the straightener apparatus with a gap of regenerative cooling of the nozzle and the second zone of the combustion chamber.



 

Same patents:

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: engines and pumps.

SUBSTANCE: invention relates to liquid-propellant engines, particularly, to multichamber rocket engines. Proposed engine has at least two chambers incorporating regenerative cooling circuit and mixing heads, turbopump system to feed gas generators and engine chambers. Besides, it incorporates control system with starting-shutoff valves, thrust regulator and fuel components ratio throttle. In compliance with this invention, engine feed turbopump system comprises two turbopump units supplied by two self-contained oxidative gas producers. Note here that first and second turbopump units feature identical power and comprises aligned fuel pump, oxidiser pump and gas turbine fitted on one shaft. Note also that second turbopump unit pump represents a two-stage design. Besides, outlets of first turbopump unit fuel and oxidiser pumps are communicated via pipelines with inputs of second turbopump unit fuel and oxidiser pumps. Second turbopump unit oxidiser pump communicates with mixing heads of said gas producers via pipelines incorporating starting-shutoff valves. Outlet of second turbopump unit oxidiser pump first stage is connected with mixing heads of engine chambers via fuel components ratio throttle, starting-shutoff valves, pipelines and chamber regenerative cooling circuits. Note also that second turbopump unit fuel pump second stage is connected with gas producer mixing heads via pipeline and thrust regulator.

EFFECT: decreased dynamic loads at turbopump unit, higher thrust.

1 dwg

FIELD: engines and pumps.

SUBSTANCE: proposed engine comprises turbo pump unit including turbine, oxidiser and propellant pumps, extra propellant pump, all being fitted on turbine shaft, and combustion chamber. Combustion chamber has cylindrical part with oxidiser and fuel nozzles and nozzle with propellant main manifold. Turbo pump unit and combustion chamber are aligned. Double-area combustion chamber comprises first circular area with circular nozzle unit and second area with central nozzle unit incorporating extra fuel nozzles while turbine being arranged inside combustion first area. Oxidiser pump outlet is communicated via pipeline equipped with oxidiser valves with combustion chamber. Central hollow nozzle unit communicates via axial bore extending inside turbo pump unit shaft with extra fuel pump inlet.

EFFECT: decreased transverse size and weight.

3 cl, 2 dwg

FIELD: engines and pumps.

SUBSTANCE: liquid-propellant rocket engine comprises chamber, gas generator, fuel pump, two-stage turbine fed with generator gas. In compliance with this invention, turbine first stage outlet is communicated with the chamber nozzle head while turbine second stage outlet communicates with pump inlet for one of fuel components or with atmosphere.

EFFECT: higher efficiency of rocket engine with generator gas afterburning.

2 dwg

FIELD: engines and pumps.

SUBSTANCE: invention relates to machine building, particularly to high rpm auger rotor turbo pumps of throttled liquid-propellant rocket engines. Proposed method consists in using booster pumps upstream of the main pumps. Every booster pump is driven by gas or hydraulic turbine. Note here that inlet of every turbine is communicated via pipeline with outlet of one of said pumps in hydraulic turbine and, in gas turbine, with gas line of turbo pump turbine. Note also that at pressure drop at pump inlet below that required for continuous operation of the pumps possible at high throttling head at booster pumps is increased by working fluid feed to extra nozzles with their intake manifolds. The latter are preliminary mounted in said turbines via pipelines with control devices. The latter can be composed by multiple station valves or pressure control valves.

EFFECT: cavitation-free operation of fuel feed turbo pumps at low loads.

3 cl, 1 dwg

FIELD: weapons and ammunition.

SUBSTANCE: proposed missile comprises axially symmetric body with oxidiser and propellant tanks and liquid-propellant rocket with combustion chamber and turbo pump unit, and four radial control nozzles. Combustion two-zone chamber comprises cylindrical part with oxidiser atomisers, nozzle with propellant main manifold, extra propellant atomisers, annular and extra manifolds. Turbo pump unit comprises turbine, oxidiser and propellant pumps, extra propellant pump, top and bottom ball hinges. Ball hinge with inner ball and outer spherical shell is arranged between combustion chamber and oxidiser pump. Said four control nozzles are communicated with combustion chamber first zone via pipes.

EFFECT: higher reliability of launching and specific characteristics of liquid-propellant engine.

6 cl, 7 dwg

FIELD: engines and pumps.

SUBSTANCE: proposed engine comprises turbo pump including, in its turn, turbine, oxidiser and propellant pumps, extra fuel pump and combustion chamber with cylindrical part with oxidiser and propellant nozzles and nozzle with the propellant main header. It differs from known designs in that said turbo pump unit and combustion chamber are aligned. Note also that turbine is arranged inside combustion chamber cylindrical part and composed of two zones and comprises propellant extra nozzles at its cylindrical part below the turbine. Oxidiser pump outlet is communicated via pipeline incorporating oxidiser valves with combustion chamber. Mid header can be arranged at combustion chamber cylindrical part mid part. Intermediate header is arranged below mid header with propellant extra nozzles extending therein.

EFFECT: higher specific characteristics and reliability and lower launching costs.

4 cl, 4 dwg

Nuclear submarine // 2494004

FIELD: transport.

SUBSTANCE: invention relates to ship building. Proposed atomic submarine comprises strong hull embraced by light hull, cisterns arranged there between, solid deckhouse and rescue float chamber arranged inside strong hull under said deckhouse, stern with propeller screw with hub fitted on propeller shaft coupled with motor and, at least one nuclear reactor communicated via circulation pipelines with turbo generator connected via electric cables with storage batteries and motor. The latter is connected via electric cable with storage batteries and motor. Besides it comprises missile compartment. Two streamlined torpedo modules with screw propellers are attached to aforesaid hull on its both sides while rocket module is secured atop it and furnished with fast-release end plug and liquid-propellant rocker engine.

EFFECT: higher speed in attack from surface position, better controllability.

4 cl, 10 dwg

FIELD: transport.

SUBSTANCE: set of invention relates to ship building and may be used in construction of warships. Proposed atomic submarine comprises strong hull embraced by light hull, cisterns arranged there between, solid deckhouse and rescue float chamber arranged inside strong hull under said deckhouse, stern with propeller screw with hub fitted on propeller shaft coupled with motor and, at least one nuclear reactor communicated via circulation pipelines with turbo generator connected via electric cables with storage batteries and motor. Two streamlined combat modules with screw propellers are attached to aforesaid hull on its both sides while rocket modules is secured atop it and furnished with fast-release end plug and liquid-propellant rocker engine. Said liquid-propellant rocker engine comprises bearing frame, combustion chamber hinged thereto and including head, cylindrical section and nozzle, gas generator and a pump unit, in its turn, comprising turbine, oxidiser and propellant pumps, gas duct communicating turbine outlet with combustion chamber head via suspension assy. Gas generator and turbo pump unit are aligned with combustion chamber. Note here that gas generator and turbo pump unit are integrated into single unit.

EFFECT: higher speed in attack from surface position, better controllability.

5 cl, 10 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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