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.

Claims

1. 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.

 

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