The method of operation of a liquid-propellant rocket engine with vapor-circuit in the system turbopump fuel

 

The method of operation of a liquid-propellant rocket engine with vapor-circuit in the system turbopump fuel involves the use of heat products of combustion to convert the working fluid from liquid substances in pairs with supercritical temperature and pressure. After completion of the work pairs again turn into a liquid, using the spent cooling steam in the heat exchanger-condenser Hadarezer rocket fuel, returning when the condensate in the corresponding pump. Portion of the liquid working fluid at the outlet from the pump is introduced into the exhaust steam, through its partial condensation with the formation of wet steam at the inlet to the working path of the heat exchanger-condenser. The invention will improve the weight and dimensions of the heat exchanger-condenser. 4 C.p. f-crystals, 2 Il.

The invention relates to liquid propellant rocket engines (LPRE), particularly to a rocket engine turbopump with fuel consisting of separately stored oxidizer and fuel, at least one of these fuel components (usually oxygen oxidizer) is cryogenic.

The known method of operation of rocket engines with vapor-circuit in the system turbopump productlogo liquid substances in pairs with supercritical temperature and pressure, after doing the work I again draw in fluid, using Hadarezer rocket fuel in the spent cooling steam in the heat exchanger-condenser, and the condensate returned to the appropriate pump (see U.S. Pat. EN 2155273 C1, 18.08.1999 prototype of the invention).

The first necessary condition for the functioning of LPRE with vapor-circuit in the system turbopump fuel supply is providing the energy balance of this system, i.e. equality between available power turbines and a total capacity of pumps. To obtain high values of specific impulse rocket engine (Iyyou need to create high pressure in the chamber (pto). In this case, for balancing power is necessary to heat sufficient propellant mass of the turbine (e.g., ammonia) to a high temperature and make the resulting steam (or rather - gas) at high pressure drop. At the end of the working cycle of the supply system should be given the residual heat of the exhaust steam flowing into the LRE cold fuel to cool the steam to complete his treatment in the condensate.

Usually the main Hadarezer fuel is concentrated in the cryogenic oxygen oxidizer having at the outlet of the pump temperatureIn the General case, this unit contains three working areas cooling (e.g., ammonia) vapor to the saturation temperature, the plot of wet steam (actually condensation) and the cooling section of the condensate to ensure cavitation free operation of the pump. At high value of ptothe implementation of the workflow in the heat exchanger-condenser requires highly developed surface of the unit, and its weight is excessively large, which prevents the implementation of the prototype method.

The invention solves the technical problem of improving mass and size parameters of the heat exchanger-condenser.

The technical problem is solved in that in the method of operation of a rocket engine with vapor-circuit in the system turbopump fuel supply, including the use of heat products of combustion to convert the working fluid from the source of liquid substances in pairs with supercritical temperature and pressure, after which doing work once again turn into a liquid using Hadarezer rocket fuel in the spent cooling steam in the heat exchanger-condenser, and the condensate returned to the appropriate pump according to the invention a portion of the liquid rnam wet steam at the entrance to the working channel Teploobmennik condenser.

In cases of carrying out the invention:

- enter in the exhaust vapor of the liquid is pre-cooled by heat exchange with fuel rocket fuel, which can use part consumed through the engine mass fuel;

- enter in the exhaust vapor of the liquid is pre-cooled by heat exchange with the oxidizer propellant, which can use part consumed through the engine weight of the oxidizer.

When carrying out the invention are expected technical result, which coincides with the essence of the problem being solved.

The invention is explained using Fig.1 and 2, which presents a functional diagram for a liquid propellant rocket engine on the proposed method.

According Fig.1, LRE contains creating traction camera 1 with the nozzle head 1A, the combustion chamber 1b and a supersonic jet nozzle 1C; the camera body is formed of two coaxial shells (external and internal), forming a path for 1d flow cooler. For supplying fuel to the engine is provided turbopump Assembly (TNA), which includes a pump cryogenic oxidizer (usually liquid oxygen) 2, pump fuel (e.g. kerosene) 3, a pump 4 for feeding the condensed working fluid of the turbine (for example the 7 pump 3 is connected to a jet head 1a. She communicated with the pump 2 through the high-pressure line 8 with the heat exchanger-condenser 9. It is designed for cooling with condensation of the exhaust steam of the turbine, which is supplied through line 10; the resulting condensate is returned through the pipe 11 into the pump 4.

In the line 10 between the turbine and the heat exchanger-condenser posted by the mixing device 12 for inputting the exhaust steam supplied through a pipeline 13 after pre-cooling in the heat exchanger 7 the mass of the condensate taken from the line 14, which connects the output of pump 4 to the input of the cooling tract 1d camera. The output reported by the pipeline 15 to the input of the turbine 5. The pump 4 together with the turbine 5, the heat exchanger-condenser 9, the mixing device 12, a cooling path 1d and connective expenditure highways form a closed circuit for circulation of the working fluid undergoing phase transformations. To the specified circuit connected to the circulation loop (4-7-13-12-9-11-4).

Described LPRE works as follows.

Coming into the engine oxidizer propellant (e.g., liquid oxygen) is served by a pump 2 through line 8 in the nozzle head 1a came). Coming in LRE fuel rocket fuel (e.g. kerosene) served by the pump 3 through line 6 into the nozzle head 1A of the camera. On the way the fuel is heated in the heat exchanger 7 (from the hot product from line 14). In the combustion chamber 1b of the fuel components are burned, and the resulting high-temperature gas is supplied to the jet nozzle 1C, creating cravings camera 1 (and LRE as a whole). Circulating in a closed loop working fluid (e.g., ammonia) to drive the turbine 5 are served by the condensate pump 4 through line 14 into the cooling tract 1d camera. After his passing the working fluid is converted to superheated steam (gas) with supercritical temperature and pressure, serves on the pipeline 15 to the turbine 5, causing the pumps 2, 3, 4 via a common shaft (usually consists of two parts, connected by a spring). The exhaust steam of the turbine serves on line 10 sequentially in the mixer 12 and the heat exchanger-condenser 9. In the first of these units is also served by pipeline 13 condensate from line 14, cooling it on the way in the heat exchanger 7 colder fuel. The cooled condensate is mixed in mixer 12 with the flow of exhaust gas turbine, receiving the result of wet steam coming next work in Proc. of the l feed system repeats.

In Fig.1 dashes shows the heat exchanger 7a to cool the condensate fed to the mixer 12, through the use of glatorians oxidant (not fuel, as in the case of the heat exchanger 7).

In Fig.2 shows a variant of placement in the line between the turbine 5 and the pump 4 not one but two heat exchangers: heat exchanger-cooler heat exchanger 9a and the capacitor 9b. In the first unit carry out a preliminary cooling (oxidant) of the exhaust gas turbine, which is then diluted in the mixer 12A cooled condensate from unit 7 to the pipe 13a. In unit 9b carry out the condensation of the resulting vapor-liquid mixture.

The invention is not limited to the shown in Fig.1 and 2 schemes, for example:

- the presence of the heat exchangers 7 and 7a is not mandatory;

- coming out of the cooling channel 1d pairs before feeding the turbine 5 can be heated additionally by the gas generated by the combustion of part of the fuel in a special gas generator;

the number of impellers in the pumps and the turbine may be different;

- use as a working fluid turbine of the same product that fuel rocket fuel (for example, liquefied methane or natural gas is t from the implementation of the invention will show a specific example: for LRE fuel oxygen kerosene" with a thrust of 1, 2 MN, functioning under the scheme according to Fig.1 without heat exchangers 7 and 7a.

Initial data for calculation of energy balance LRE:

- pressure in the chamber 15 MPa;

the flow of oxidant through the engine 260 kg/s;

- the flow of fuel through the engine 100 kg/s;

- working body of the turbine - ammonia vapor (gas) with supercritical temperature and pressure.

The results of the calculation:

the steam flow through the turbine 36 kg/s;

the steam temperature at the inlet/outlet of the turbine 300/60C;

- steam pressure at the inlet/outlet of the turbine 17/0,8 MPa;

the flow of condensate through a pump 41 kg/s;

- the temperature of ammonia at the inlet/outlet condensate pump 0/12C;

- pressure of ammonia at the inlet/outlet condensate pump 0,6/23 MPa;

- at the output of mixer 12 steam has a pressure of 0.9 MPa and a temperature of 21C and the degree of dryness of 0.96;

the working surface of the heat exchanger-condenser 60 m2;

- weight duralumin construction of the heat exchanger-condenser 170 kg

Thus, in a specific example shows that the invention allows a rocket engine with a high value of pto(which correspond to high values of Iy) at an acceptable size and weight of the heat exchanger-condenser. This technical result is achieved due to the fact that in the ode turbine, and wet steam with dryness 0,96), i.e. the liquid-vapor mixture, which contains condensation nuclei contributing to the process of condensation.

Claims

1. The method of operation of a liquid-propellant rocket engine with vapor-circuit in the system turbopump fuel supply, including the use of heat products of combustion to convert the working fluid from the source of liquid substances in pairs with supercritical temperature and pressure, after which doing work once again turn into a liquid using Hadarezer rocket fuel in the spent cooling steam in the heat exchanger-condenser, and the condensate returned to the respective pump, characterized in that the portion of the liquid working fluid at the outlet from the pump is introduced into the exhaust steam, through its partial condensation with the formation of wet steam at the inlet to the working path of the heat exchanger-condenser.

2. The method of operation of a liquid-propellant rocket engine under item 1, characterized in that an input in the exhaust vapor of the liquid is pre-cooled by heat exchange with fuel rocket fuel.

3. The method of operation of a liquid-propellant rocket engine under item 1, ulele rocket fuel.

4. The method of operation of a liquid-propellant rocket engine under item 2, characterized in that the cooling fluid used part consumed through the engine mass fuel.

5. The method of operation of a liquid-propellant rocket engine under item 3, characterized in that the cooling fluid used part consumed through the engine weight of the oxidizer.

 

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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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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.

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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.

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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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