Drosselweg oxygen-hydrocarbon liquid propellant rocket engine with afterburning gas recovery
Drosselweg oxygen-hydrocarbon liquid propellant rocket engine with afterburning reducing gas includes a fuel line, a cooled hydrocarbon fuel chamber operating on oxygen and hydrocarbon fuel with an excess of fuel gas generator and fuel pumps driven by gas turbines. The control system of the engine includes installed in highways oxygen oxidizer the throttle and choke control fuel ratio of the components. In the system control is provided by sagastume on throttle mode, the device creating an additional burden for hydrocarbon fuel pump. Device additional burden for hydrocarbon fuel pump is either hydraulic resistance introduced directly into the supply line of the gas generator fuel, or by-pass line with a hydraulic resistance. The invention will ensure that the performance of liquid rocket engines in a wide range of throttling and will prevent the formation of soot in the gas generator. 1 Il., 1 PL.
The invention relates to a liquid of realities and fuel.
Known LRE, including a line of liquid and gaseous bodies, designed for cooling flow of hydrocarbon fuel chamber with the nozzle head and the supersonic jet nozzle operating on the oxygen and hydrocarbon fuel with an excess of fuel gas generator, the fuel pump is driven by a gas turbine, an exhaust pipe connected to the nozzle head camera system control with traction control and throttle control fuel ratio of components - see Acta Astronautica, Vol.41, Nos 4-10, pp.209-217, published by Elsevier Science Ltd, 1997 - the prototype of the invention.
LRE, made by the scheme with afterburning, are widely used in boosters, which are the output of payloads into space. These rocket engines can operate at high pressure in the chamber (pto) that provides a high degree of conversion of chemical energy used two-component liquid fuel to generate thrust. However, the device prototype has a significant drawback. In fact, by throttling known LRE (that is, when the managed reduction of thrust) within the defined flight conditions booster, along with the decline of paramedic the camera. This fact narrows the range of the rocket engine throttling.
The present invention solves the technical problem of providing health LRE in a wide range of throttling.
The technical problem is solved by the fact that in the LRE, including a line of liquid and gaseous bodies, designed for cooling flow of hydrocarbon fuel chamber with the nozzle head and the supersonic jet nozzle operating on the oxygen and hydrocarbon fuel with an excess of fuel gas generator, the fuel pump is driven by a gas turbine, an exhaust pipe connected to the nozzle head camera system control with traction control and throttle control fuel ratio of the components according to the invention in the system control is provided by sagastume on throttle mode, the device creating an additional burden for hydrocarbon fuel pump.
In some cases, of the invention:
mentioned device is a nonadjustable or adjustable hydraulic resistance introduced directly into the supply line of the fuel gas generator;
- mentioned device includes a bypass was magistrala technical result coinciding with the essence of the problem being solved.
The invention is illustrated with the aid of the drawing, which shows a functional diagram of the LRE, arranged according to the invention. LRE contains creating traction camera 1 with the nozzle head 1A, supersonic jet nozzle 1B and intended for the supply of liquid fuel turbopump Assembly (TNA). It includes single-spaced diagram of a two-stage pump oxygen oxidant (e.g., liquid oxygen) 2 with booster stage 2A, the two-stage pump of hydrocarbon fuels (such as liquefied methane) 3 with booster stage 3A and the gas turbine 4. It is connected at the inlet to the gas generator 5 and the output - through exhaust pipe (gazvoda) 6 - mentioned jet head 1A. This head is connected also with the oxidizer pump through the high-pressure line 7 installed in it a motorized throttle 8. The gas generator 5 is intended to provide a working fluid turbine, carried out by combustion of part of the spent rocket engine, two fuel with an excess of fuel (in the case of methane). Jet head 5A of the gas generator is connected to pump fuel through the high energy of obsen managing pipeline 11 by line 12, which in GG enters the oxidizer from the pump stage 2A. In the specified trunk-mounted motorized regulator 13. The camera has a housing with two walls forming the flow path of the cooling 1C. He communicated through the inlet pipeline 14 with the output of the pump stage 3 and indicated by a discharge line 15 to the input of the pump stage 3A.
As shown in the drawing by the dashed lines, is described in LRE through the supply line YY fuel may be provided instead of the controller 10 of the bypass line 16, normally closed valve 17.
Described LPRE works as follows. Liquefied oxygen enters the pump 2, from which the main part of the liquid (80%) through line 7 is fed to the jet head 1A of the camera 1. The remaining portion of the oxidant enters the secondary pumping stage 2A, from which the line 12 is fed in the nozzle head 5A of the gas generator 5. Liquefied methane is fed into the pump 3. Part of the fuel through the pipeline 14 is pumped into the flow path of the cooling 1C of the camera, and the heated refrigerant is discharged through pipe 15 to the input of charge pump stage 3A. She increases the pressure of the entire mass of fuel to supply his labour produces the reducing gas with temperature Tyyabout 500...1000). He goes to the turbine 4, resulting in the rotation of its rotor, and with it the fuel pumps. The exhaust turbine gas flows through gazivoda 6 jet head 1A of the camera and digiguide in fire space with oxidizer received from highway 7. High temperature combustion products are expanded in the jet nozzle 1B, creating thrust rocket engine.
Control operating mode LRE exercise influence on the bodies 8, 10, 13, 17. In this case, rotation of the throttle valve 8 leads to a change in the flow rate of the oxidizer through the engine, thus achieving necessary (for simultaneous production of fuel from tanks booster) changing ratio of the fuel components. The movement of the needle controller 13 changes the flow rate of the oxidant (maboutin line power BIENNIUM, resulting in the changing ratio of the fuel components in GG (Kyyand, therefore, the temperature of the gas generator. As a consequence, changes the power TNA, and LRE is translated into another thrust. With increasing mothe value of Tyyincreases, and the engine is forced, and with decreasing maboutthe value of Tyyis reduced, and the engine throttled.
When throttling is th body of the regulator 10 is moved to the position 10A to cover the orifice. Made this additional hydraulic resistance (R) creates an additional load on the pump hydrocarbon fuel, which is compensated in order to ensure the necessary flow of fuel through the YEARS (mg) - elevated (compared to the rocket engine prototype) speed TNA (n). The relative flow rate of fuel flowing through the pipeline 14 to the cooling chamber (mg, OHL), also increases (compared to the rocket engine prototype). The selection of the appropriate values ofp and n provide the value of mg, OHLsufficient for reliable cooling of the camera on the throttle mode. The same effect can be achieved and the opening of the valve 17, perepuskajutsja part of the flow of fuel from the pipe 9 to the inlet of the pump stage 3A.
The invention is not limited to the above specific LRE. For example, pumps of oxidizer and fuel can operate from its own gas turbines, the number of impellers in pumps and turbines may be different. Select the unit type (valve, regulator, throttle), providing additional load pump hydrocarbon fuel, is determined by the specific technical requirements of LRE. Specified the highway I can get directly between the outlet of the pump (pumping speed) and another spot on the desktop tract LPRE with a lower pressure (for example, the engine inlet).
We show the effectiveness of our invention on the example of a specific project rocket engine turbopump with a supply of two-component fuel oxygen - methane". This rocket engine rated thrust 2 MN at pto=24 MPa and at work in the booster should grossulariata up to Rto, min= 0,4 ptothat corresponds to the same reduction in thrust. Values of other parameters LRE nominal and minimum modes of traction presented in the table below. For comparison, the last column in italics is given the parameter values at the minimum mode for rocket engine prototype. The table contains, along with mentioned in the text, the parameter Ng- power pump methane fuel.
As can be seen from the table, when the rocket engine throttling up to 40% of the nominal pressure in the chamber, the fuel consumption for the cooling is reduced to 36%, i.e. approximately in the same degree that ensures reliable cooling chamber. In contrast, when the throttling rocket engine prototype camera would be burnt, because the value of mg, OHLwould have fallen to 25%, which is unacceptable under the terms of the cooling design.
So, on a specific project shows that the proposed invention solves t technical result of the invention is confirmed.
In some cases it is possible to obtain a very important additional technical result. It is due to the fact that by throttling the proposed LRE is possible to prevent the decrease of Tyyto dangerous levels, which in the generator gas would be formed soot deposited on the structure elements with subsequent breach of the engine. This danger exists for a liquid propellant rocket engine at two oxygen-hydrocarbon fuels, when (as in our case) the drive turbine is reductive gas generator. The invention resolves that danger.
Drosselweg oxygen-hydrocarbon liquid propellant rocket engine with afterburning reducing gas, including line of liquid and gaseous bodies, designed for cooling flow of hydrocarbon fuel chamber with the nozzle head and the supersonic jet nozzle operating on the oxygen and hydrocarbon fuel with an excess of fuel gas generator, the fuel pump is driven by a gas turbine, an exhaust pipe connected to the nozzle head camera control system work with regulyatornogo oxidant, characterized in that the system control is provided by sagastume on throttle mode, the device creating additional load for the pump hydrocarbon fuel, which represents either a hydraulic resistance introduced directly into the supply line of the gas generator fuel, or by-pass line with a hydraulic resistance.
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.
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
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.
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
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