The method of operation of a liquid rocket engine turbopump feed oxygen-methane fuel
The method of operation of a liquid rocket engine turbopump feed oxygen-methane fuel, in which part of the consumed methane fuel is used as a refrigerant for flow of the cooling chamber. After use, the refrigerant is mixed with the remaining mass of the fuel, increase the pressure of the mixture and serve it in the gas generator for the combustion of part of the oxygen of the oxidant in order to obtain a reducing gas, which is installed on the turbine, and the exhaust gas dorogaya in the chamber with the rest mass of the oxidizer. Before mixing with the remaining mass of fuel used, the refrigerant is cooled by heat exchange with an oxygen oxidizer, and part of the fuel spent on creating Severnogo chamber cooling by the coolant on the inner wall of the chamber provided through the belt holes. The invention will increase the specific impulse of the rocket engine. 1 C.p. f-crystals, 2 Il. The invention relates to liquid propellant rocket engines (LPRE), particularly to a rocket engine turbopump with a supply of two-component fuel, including liquid oxygen (oxidizer) and liquefied methane fuel (including natural gas, consisting of in which methane fuel is used as a refrigerant for flow of the cooling chamber and to obtain a reducing gas during combustion in the gas generator (GG) with part of the oxygen oxidizer, moreover, the resulting gas fired turbine and exhaust gas dorogaya in the chamber with the rest mass of oxidant - see Acta Astronautica, Vol. 41, Nos 4-10, R. 211, fig.2 - analogue of the invention.In the known method-similar to cool cameras use all methane fuel consumed LRE; after cooling chamber heated fuel is fed directly to the combustion in the cities. Thus, the necessary pressure of methane pump includes the total loss of pressure in the duct of the cooling chamber and the turbine. Consequently, when the pressure level in the chamber Rto15 MPa pressure in the cooling path reaches 50 MPa, which leads to the destruction of the mechanical linkage between the inner and outer shells of the camera. To avoid this, there is a reduction of ptothat restricts get the value of the specific impulse rocket engine (Ibeats), i.e. does not allow a sufficient degree of potential chemical energy of oxygen-methane fuel to generate thrust.The known method of operation of rocket engine turbopump with a supply of oxygen-methane fuel, in which part of the consumed methane fuel is used as a refrigerant for protochnoye mixture and serve it in the gas generator for the combustion of part of the oxygen of the oxidant in order to obtain a reducing gas, which work on the turbine, and the exhaust gas dorogaya in the chamber with the rest mass of oxidant - see Acta Astronautica, Vol. 41, Nos 4-10, R. 211, fig.3 - the prototype of the invention.Using the prototype method allows to reduce the pressure in the duct of the cooling chamber to a level acceptable for reasons of structural strength. However, the pre-supply pump stage, providing increased pressure of the mixture of fuel, runs on the hot product, whose density is small, and therefore power consumption (along with the dimensions and weight) of the charge pump stage can reach large values (increasing with increasing pto). This circumstance limits the achievable level of ptothe value of17 MPa, which restricts obtain the value of Ibeats.The present invention solves the technical problem of increasing the specific impulse rocket engine.The technical problem is solved in that in the method of operation of a rocket engine turbopump with a supply of oxygen-methane fuel, in which part of the consumed methane fuel is used as a refrigerant for flow of the cooling chamber, the refrigerant is mixed with underwater oxidant in order to obtain a reducing gas, which work on the turbine, and the exhaust gas dorogaya in the chamber with the rest mass of the oxidant according to the invention before mixing with the remaining mass of fuel used, the refrigerant is cooled by heat exchange with an oxygen oxidant of the fuel is spent on creating Severnogo chamber cooling by the coolant on the inner wall of the chamber through the inclusion of the belt holes. In the particular case of the invention on Zavetnoe chamber cooling consume (4-10)% of the total weight of fuel.When carrying out the invention are expected technical result, which coincides with the essence of the problem being solved.The invention is illustrated by drawings, where - Fig.1 presents the scheme of the rocket engine, functioning according to the invention; - Fig.2 shows the change of the parameters of methane refrigerant depending on the pressure in the chamber.According Fig. 1 LRE contains creating traction camera 1 with the nozzle head 1A and a supersonic jet nozzle 1B, intended for the supply of liquid fuel turbopump Assembly (TNA), which includes coaxially installed and consistently located the pump oxygen oxidizer 2 with booster stage 2A, the pump methane fuel Stom exhaust gazivoda 6 - the said jet head 1A. This head is connected also with the oxidizer pump through the high pressure line 7 defined therein a throttle 8 and the heat exchanger 9. The gas generator 5 is connected by supply line fuel through the high pressure pipeline 10 to the outlet pump stage 3A. Line power oxidant GG connected to the output of the pump stage 2A through high-pressure pipe 11 with the installed controller 12. The camera has a housing with two walls forming the flow path of the cooling 1B. He reported to work (fire) space camera through the belt holes 1G, made in the inner wall of the chamber. Tract 1B informed input through a pipe 13 with the fuel pump 3, and the output is in communication with the pumping stage 3A through pipe 14, which is located above the heat exchanger 9. After he mentioned the pipe 14 is connected (15) to the inlet of the pump stage 3A.Described LPRE works as follows. Liquefied oxygen enters the pump 2, from which the main part of the liquid (80%) through the pipeline 7 is fed to the jet head 1A of the camera 1. The remaining portion of the oxidant enters po. Liquefied methane or liquefied natural gas) enters the pump 3 from which part of the fuel (mFR) through the pipeline 13 is fed in the path of flow of the cooling 1B of the camera. The rest of the fuel comes in pre-supply pump stage 3A.From the cooling tract 1B small fraction of fuel (mveils) is fed through the belt holes 1G on the inner wall of the chamber in order to provide extra cooling (from the inside). The cooling effect is due to the fact that interior walls, washed by the stream of hot gases, a relatively cold protective layer (the veil) gas-liquid mixture moving in the same direction as the main stream. The cooling effect is due both to absorption of heat by evaporation of the veil, and by the fact that a pair of coolant entering the boundary layer, increasing its thickness and thereby reduce heat transfer into the wall.After the tract 1B, the refrigerant flow rate (mFR-mveils) flows through the pipe 14 through the heat exchanger 9 to the input of charge pump stage 3A previously (15) mixed with fuel supplied from step 3. Step 3A increases the pressure of the fuel and delivers it in a spray head of the gas generator 5. From the combustion of liquid is rbine 4, which causes the rotation of the fuel pump via a common shaft (usually consisting of two parts connected by a spring). The exhaust turbine gas flows through the exhaust gazivoda 6 in the nozzle head camera and digets in the working space of the camera with the oxidant supplied from the pipe 7. High temperature combustion products are expanded in the jet nozzle 1B, creating thrust rocket engine.The changing ratio of the fuel components in LRE effect is achieved by the inductor 8, the change in thrust is achieved by means of the controller 12.The role of the heat exchanger 9 is to be cooled by glatorians liquid oxygen fuel, heated in the path of flow of the cooling chamber. Cooling of the fuel is achieved by increasing its density (up to the value for the original liquid product), which reduces the cost of power to drive the charge pump stage 3A and TNA as a whole; as a result, it becomes possible to raise the working pressure in the chamber and to increase specific impulse. However, the final evaluation of the obtained winning (i.e. the increment of the mass of the payload of the rocket apparatus) must take into account the weighting of the design of the engine accounts for the cooling chamber. First, it reduces the amount of refrigerant flow mFRand/or its speed, and therefore, reduce the hydraulic resistance of the cooling tract (p), and finally to reduce the required capacity of methane pump 3. Secondly, the introduction of Severnogo cooling reduces heating of the fuel in the path of flow of cooling (T), which together with the previously mentioned reduction of mFRcreates favorable conditions for the operation of the heat exchanger. It is obvious that with increasing flow rate, mveilsreduced specific impulse rocket engine. The best results are achieved when mveils=(4-10)% of the flow of fuel through the motor mg(a particular case of the invention).We will demonstrate the effectiveness of the proposed method with the results of calculation of the energy balance for oxygen-methane rocket engine, designed for nominal thrust in vacuum Rp=18 TC at a nominal pto=150 kgf/cm2. The camera of this LPRE us were forced by pressure with a proportional increase in thrust. Accordingly, changing the value of mFR,p andT (accepted mveils= 0,05 mg): see Fig. 2. which which provides a balance of capacity pumps and leading their turbines. The desired value of ptomaxamounted to 250 kgf/cm2when the temperature of the reducing gas before the turbine 1000 K (the ratio of the pressure at the turbine was 2,34). The mass of the heat exchanger is estimated in (15-20) kg when the working surface is not more than 3 m2that is quite acceptable.Thus, for a calculated example of the oxygen-methane rocket engine, the proposed method achieves pto= 250 kgf/cm2(corresponds to deadlift 30 TC) compared to pto=170 kgf/cm2for the prototype method. The corresponding increase of the parameterbeats70 m/s, which is a significant increase. So, the expected technical result was confirmed.The most appropriate scope for the proposed method are the rocket engine with a thrust from several units to several tens of ton-forces achieved the greatest (in quantitative terms) the technical result.In conclusion, we note that the invention is not limited to the above specific scheme LRE illustrating the proposed method: the number of working impellers in the pumps may be different, the camera may include an output to allow the nozzles to radiation or ablation is mediocre with the path of flow of cooling), and so D.
Claims1. The method of operation of a liquid rocket engine turbopump feed oxygen-methane fuel, in which part of the consumed methane fuel is used as a refrigerant for flow of the cooling chamber, the refrigerant is mixed with the remaining mass of the fuel, increase the pressure of the mixture and serve it in the gas generator for the combustion of part of the oxygen of the oxidant in order to obtain a reducing gas, which is installed on the turbine, and the exhaust gas dorogaya in the chamber with the rest mass of the oxidizing agent, characterized in that before mixing with the remaining mass of fuel used, the refrigerant is cooled by heat exchange with an oxygen oxidizer, part of the fuel spent on creating Severnogo chamber cooling by the coolant on the inner wall of the chamber through the inclusion of the belt holes.2. The method of operation of a liquid-propellant rocket engine under item 1, characterized in that Zavetnoe chamber cooling consume (4-10)% of the total weight of 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.
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