Low-temperature gas injection thermal protection method for recycling process of reusable rocket and application of low-temperature gas injection thermal protection method

Abstract

The invention belongs to the technical field of aerospace, and particularly relates to a low-temperature gas injection engine thermal protection method for the recycling process of a reusable rocket and application. According to the thermal protection method disclosed by the invention, the low-temperature gas film cooling distribution can be uniform, and meanwhile, the reusability of the rocket engine is improved. A cooling system is installed in the engine and comprises a gas cylinder supercharging device, an engine working device and a low-temperature gas injection device. Wherein the low-temperature gas injection device comprises a valve and a pipeline which are connected with the gas cylinder supercharging device, and a rotational flow runner and a buffer cavity directly connected with the rotational flow runner are arranged on the outer wall surface of the combustion chamber. When the rocket lands, low-temperature gas is injected into the combustion chamber through the rotational flow channel, forms a rotating gas film to be attached to the inner wall face of the combustion chamber under the action of centrifugal force and forms a vortex to be attached to the outer wall face of the engine under the action of reverse incoming flow, and ablation of high-temperature gas to the wall face of the engine is reduced.

Description

Technical Field

[0001] The present invention belongs to the field of aerospace technology, and in particular relates to a thermal protection method and application of a cryogenic gas injection engine used in a reusable rocket recovery process. Background Art

[0002] Reusability is one of the core features of the next generation of liquid launch vehicles, which can effectively reduce the cost of space transportation and support the construction of flight-based space transportation systems. In liquid rocket engines, the nozzle is the main component that generates thrust. The high-temperature and high-pressure gas expands and accelerates in the nozzle, and is discharged from the nozzle outlet to generate thrust. At the same time, when the liquid rocket is recovered and landed, part of the high-temperature gas will flow into the nozzle due to the reaction force of the reverse flow, and the other part will be compressed and expand toward the outer wall of the nozzle, which will produce a severe scouring and ablation effect on the inner and outer walls of the nozzle. At present, the main cooling method of liquid rocket engines is mainly aimed at the inner wall of the combustion chamber and the nozzle, but there is little research on the thermal protection of the outer wall of the engine during rocket landing. In order to avoid the wall material being ablated by the high-temperature and high-speed airflow and realize the reusability of the engine, it is urgent to adopt effective thermal protection means to cool the nozzle and combustion chamber wall structure. Summary of the invention

[0003] The present invention provides a low-temperature gas injection thermal protection method for a reusable rocket recovery process, the purpose of which is to use a low-temperature gas swirl injection into a combustion chamber to avoid the problem of backflow ablation of the inner and outer walls of the nozzle.

[0004] The present invention provides a low-temperature gas injection thermal protection method for a reusable rocket recovery process, wherein a swirl flow channel is arranged circumferentially around the combustion chamber of the engine, and the low-temperature gas is injected into the combustion chamber of the engine through the swirl flow channel after pressure regulation to form a rotating film; liquid fuel and liquid oxidant react with each other in the inner cavity of the combustion chamber to generate a mainstream of fuel gas, and the low-temperature gas adheres to the combustion chamber and the nozzle wall surface and flows in the direction of the mainstream of fuel gas, thereby avoiding direct contact between the high-temperature fuel gas and the wall surface, reducing the scouring and ablation of the engine wall surface by the high-temperature fuel gas, and improving the reusability of the engine.

[0005] The present invention further provides an application of a low-temperature gas injection thermal protection method for a reusable rocket recovery process, wherein a cooling system is arranged inside the engine, including: a high-pressure gas cylinder, a liquid oxidizer tank, a liquid fuel tank, an oxidizer pump, a fuel pump, a gas generator, a turbine, and a buffer chamber; after the first outlet of the high-pressure gas cylinder is connected to the pressure regulating device through a first gas pipeline, it is respectively connected to the inlet of the liquid oxidizer tank and the liquid fuel tank through two branches of the first gas pipeline; a one-way valve that only allows gas injection is respectively arranged on the inlet branches of the liquid oxidizer tank and the liquid fuel tank. ; The second outlet of the high-pressure gas cylinder is connected to the pressure regulating device through the second gas transmission pipeline, and then connected to the buffer chamber through a one-way valve that only allows gas injection; a swirl flow channel is circumferentially distributed at the front end of the combustion chamber; the buffer chamber pipeline is connected to the swirl flow channel; the liquid oxidant storage tank and the liquid fuel storage tank are respectively connected to the combustion chamber through a liquid transmission pipeline equipped with an oxidant pump and a fuel pump; the inlet of the gas generator is respectively connected to the outlet of the oxidant pump and the fuel pump through a liquid transmission pipeline, the gas chamber outlet of the gas generator is connected to the input end of the turbine, and the output end of the turbine is connected to the oxidant pump and the fuel pump.

[0006] Furthermore, the low-temperature gas stored in the high-pressure gas cylinder is an inert gas below 300K; the gas includes nitrogen, helium, argon, etc.; the pressure regulating device includes a pressure switch, a pressurizing valve, an isolation valve, a pressure reducer, a pressure sensor, and a safety valve installed in sequence on the first gas pipeline and the second gas pipeline.

[0007] Furthermore, the angle between the injection hole of the swirl flow channel and the outer wall of the combustion chamber is 30° to 90°.

[0008] Furthermore, valves for controlling the flow rate of liquid passing through are respectively provided on the first gas pipelines connected to the combustion chamber.

[0009] Furthermore, the liquid oxidant tank and the liquid fuel tank are provided with exhaust valves; and the turbine outlet is connected to a turbine exhaust pipe.

[0010] Furthermore, when the liquid rocket starts the engine to work, the one-way valve is opened to allow part of the low-temperature gas in the high-pressure gas cylinder to be injected into the liquid oxidizer tank and the liquid fuel tank through the first gas pipeline; according to the pressure difference relationship between the high-pressure gas cylinder and the liquid oxidizer tank and the liquid fuel tank, the pressure regulating device is used to regulate the high-pressure gas cylinder so that the pressure of the high-pressure gas cylinder is 10-30 MPa, and the liquid oxidizer tank and the liquid fuel tank are pressurized to 1-3 MPa; part of the liquid oxidizer and liquid fuel in the liquid oxidizer tank and the liquid fuel tank are injected into the combustion chamber under the action of the oxidizer pump and the fuel pump, and a combustion reaction is carried out in the combustion chamber to generate combustion gas, which is then discharged through the nozzle to generate combustion gas. The engine; the other part of the liquid oxidizer and liquid fuel burns in the gas generator, driving the turbine to do work, providing energy for the operation of the oxidizer pump and the fuel pump, and realizing a cyclic working process; when the rocket begins to land, the gas in the high-pressure gas cylinder begins to be injected into the buffer chamber after pressure regulation through the second gas pipeline, and the gas in the buffer chamber is directly injected into the combustion chamber through the swirl flow channel to form a layer of rotating air film, which flows toward the nozzle outlet under the action of the mainstream of the gas. The low-temperature air film adheres to the inner wall of the combustion chamber and the nozzle, reducing the temperature of the inner wall, and is discharged from the nozzle with the mainstream of the gas. Due to the action of the reverse flow, a vortex is formed, which adheres to the outer wall of the engine and reduces the ablation effect of the high-temperature gas on the engine.

[0011] Furthermore, 20% to 40% of the low-temperature gas in the high-pressure gas cylinder is respectively injected into the liquid oxidant tank and the liquid fuel tank through the first gas pipeline; 60% to 80% of the low-temperature gas in the high-pressure gas cylinder is injected into the combustion chamber through the second gas pipeline connected to the swirl flow channel; and 10% to 20% of the fuel in the fuel tank is burned in the gas generator.

[0012] Furthermore, when the pressure in the high-pressure gas cylinder is less than 10MPa, the pressure regulating device pressurizes the high-pressure gas cylinder and slowly opens the pressurizing valve; when the target pressure is reached, the pressurizing valve is kept at an appropriate opening to maintain the pressure in the gas cylinder and the pipeline; then the one-way valve is opened to allow the low-temperature gas to be injected into the liquid oxidizer tank, liquid fuel tank and combustion chamber respectively.

[0013] Furthermore, when the pressure in the high-pressure gas cylinder is greater than 30MPa, the pressure regulating device pressurizes the high-pressure gas cylinder, slowly opens the isolation valve, allows the low-temperature gas to enter the pressure reducer, and gradually adjusts the pressure reducer to reduce the pressure of the low-temperature gas to reach the target pressure; during the decompression process, the flow of the low-temperature gas can be controlled by a safety valve. If there is overpressure, the safety valve will automatically exhaust the gas; then open the one-way valve to allow the low-temperature gas to be injected into the liquid oxidizer tank, liquid fuel tank and combustion chamber.

[0014] The beneficial effects of the present invention are:

[0015] The low-temperature gas injection protection method for the reusable rocket recovery process of the present invention is that liquid oxidant and liquid fuel react with each other in the inner cavity of the combustion chamber to generate a mainstream of combustion gas, and the low-temperature gas adheres to the combustion chamber and nozzle wall and flows in the direction of the mainstream of combustion gas, thereby avoiding direct contact between the high-temperature combustion gas and the wall, reducing the scouring and ablation of the engine wall by the high-temperature combustion gas, and improving the reusability of the engine.

[0016] The present invention is installed in the cooling system inside the engine. The low-temperature gas is injected into the combustion chamber through the swirl flow channel and attached to the inner wall of the combustion chamber under the action of centrifugal force to form a rotating air film, thereby achieving the purpose of uniform distribution of the air film.

[0017] The cooling system of the present invention is installed inside the engine, and compared with the traditional liquid engine, only some valves and pipeline structures are added, and the weight burden is relatively small.

[0018] The present invention is installed in the cooling system inside the engine, and the air film rotation can also expand the cooling area, so as to cool the combustion chamber and the nozzle more effectively.

[0019] The present invention is installed in the cooling system inside the engine. When the rocket lands, the low-temperature air film adheres to the combustion chamber and the inner wall surface of the nozzle, reduces the temperature of the inner wall surface, and is discharged from the nozzle along with the mainstream of the fuel gas. Due to the action of the reverse flow, a vortex is formed and adheres to the outer wall surface of the engine, thereby improving the reusability of the engine. The inner wall surface of the combustion chamber and the nozzle and the outer wall surface of the engine can synergistically reduce the temperature, so that the engine can be better protected when the rocket returns to the atmosphere. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] Figure 1 It is an example diagram of the engine cooling system of a liquid rocket engine with cryogenic helium injection;

[0021] Figure 2 This is an example diagram of low-temperature helium being injected into the swirl flow channel of the combustion chamber;

[0022] Figure 3 This is a flow field diagram showing the principle of low-temperature helium cooling the inner and outer walls of the engine.

[0023] The meanings of the accompanying drawings are as follows:

[0024] 1-high-pressure gas cylinder, 2-liquid oxidant tank, 3-liquid fuel tank, 4-pressure switch, 5-pressurization valve, 6-pressure reducer, 7-safety valve, 8-pressure sensor, 9-check valve, 10-isolating valve, 11-exhaust valve, 12-oxidant and fuel pump, 13-liquid valve, 14-gas generator, 15-turbine, 16-turbine exhaust nozzle, 17-buffer cavity, 18-combustion chamber, 19-nozzle, 20-swirl flow channel, 21-combustion chamber inner cavity, 22-nozzle wall, 23-engine wall, 24-cooling gas flow direction, 25-main gas flow direction, 26-recirculation zone, 27-flow separation point, 28-bow shock wave, 29-lateral inflow. DETAILED DESCRIPTION

[0025] The present invention is further described below in conjunction with the accompanying drawings.

[0026] The present invention discloses a low-temperature gas injection thermal protection method for a reusable rocket recovery process, wherein a swirl flow channel 20 is arranged circumferentially around a combustion chamber 18 of an engine, and low-temperature helium is injected into the combustion chamber 18 of the engine through the swirl flow channel after pressure regulation to form a rotating film; liquid fuel and liquid oxidant react with each other in an inner cavity 21 of the combustion chamber to generate a mainstream of fuel gas, and the low-temperature helium adheres to the combustion chamber 18 and the nozzle wall 22 along the flow direction of the mainstream of fuel gas, thereby avoiding direct contact between the high-temperature fuel gas and the wall, reducing the scouring and ablation of the engine wall by the high-temperature fuel gas, and improving the reusability of the engine.

[0027] The present invention discloses an application of a thermal protection method for a cryogenic gas injection engine in a reusable rocket recovery process, such as Figure 1 As shown, a cooling system is installed inside the engine, including a high-pressure gas storage device, an engine working device, a low-temperature helium injection device and various pipelines. The high-pressure gas storage device includes a high-pressure gas cylinder 1 for storing high-pressure low-temperature helium, a pressure switch 4 for outputting a control signal for the helium pressure, a pressurizing valve 5 for increasing the pressure of the helium in the gas cylinder, a pressure reducer 6 for reducing the pressure of the helium in the pipeline, a safety valve 7 for preventing the pressure in the pressurized equipment from exceeding the design allowable value, a pressure sensor 8 for monitoring the pressure of the helium, a one-way valve 9 for preventing the backflow of the helium, and an isolation valve 10 for cutting off the flow of the helium.

[0028] The engine working device comprises a liquid oxidant tank 2 and a liquid fuel tank 3. The high-pressure gas storage device is directly connected to the liquid fuel and oxidant tanks to increase or maintain the gas pressure in the tanks to ensure normal operation. The exhaust valve 11 can discharge the excess helium in the tank to maintain a stable pressure, the oxidant pump and the fuel pump 12 drive the fuel and oxidant to flow to the combustion chamber, the liquid valve 13 controls the flow rate, the gas generator 14 provides energy for the turbine 15, the combustion chamber 18 is the main place for the chemical reaction between the oxidant and the fuel, and the nozzle 19 is a device for converting the heat released by combustion into the kinetic energy of the gas.

[0029] The cryogenic helium injection device includes various pipes and valves similar to the high-pressure gas storage device, and also includes a buffer chamber 17 and a swirl flow channel 20. The buffer chamber is used to make the flow of each swirl flow channel 20 injected into the combustion chamber 18 consistent to avoid the engine from generating lateral force. The purpose of the high-pressure gas storage device is to maintain the pressure of the liquid fuel tank 3 and the liquid oxidant tank 2 to ensure normal operation, while the cryogenic helium injection device directly injects the cryogenic helium into the combustion chamber 18 through the swirl flow channel 20. Figure 2 As an example of a swirl flow channel, the swirl flow channels 20 are distributed circumferentially at the front end of the combustion chamber 18, and there are six of them. The injection holes are tangent to the outer wall of the combustion chamber 18 at an angle of 90°.

[0030] Cooling systems installed inside the engine, such as Figure 1 , 2 As shown; the first outlet of the high-pressure gas cylinder 1 is connected to the pressure regulating device through the first gas pipeline, and then connected to the inlets of the liquid oxidant tank 2 and the liquid fuel tank 3 through two branches of the first gas pipeline; a one-way valve 9 and an isolation valve 10 are respectively provided on the inlet branches of the liquid oxidant tank 2 and the liquid fuel tank 3; the second outlet of the high-pressure gas cylinder 1 is connected to the pressure regulating device through the second gas pipeline, and then connected to the buffer chamber 17 through the one-way valve 9 that only allows gas injection; a swirl flow channel 20 is circumferentially distributed at the front end of the combustion chamber 18; the pipeline of the buffer chamber 17 is connected to the swirl flow channel 20; the liquid oxidant tank 2 and the liquid fuel tank 3 are respectively connected to the combustion chamber 18 through a liquid delivery pipeline equipped with an oxidant pump and a fuel pump; a valve 13 for controlling the flow rate of liquid passing through is respectively provided on the first gas delivery pipeline connecting the combustion chamber 18; the inlet of the gas generator 14 is respectively connected to the outlets of the oxidant pump and the fuel pump through a liquid delivery pipeline, the gas chamber outlet of the gas generator 14 is connected to the input end of the turbine 15, and the output end of the turbine 15 is connected to the oxidant pump and the fuel pump. The pressure regulating device comprises a pressure switch 4, a helium pressurizing valve 5, an isolation valve 10, a pressure reducer 6, a pressure sensor 8, and a safety valve 7 which are sequentially installed on the first gas pipeline and the second gas pipeline.

[0031] Workflow: When the liquid rocket starts the engine, first open the pressure switch and the one-way valve 9 installed at the liquid oxidizer tank 2 and the liquid fuel tank 3, allowing 20% ​​to 40% of the low-temperature helium in the high-pressure gas cylinder to continuously flow into the liquid oxidizer tank 2 and the liquid fuel tank 3; according to the pressure difference between the high-pressure gas cylinder and the tank, ensure that the gas cylinder pressure is 10 to 30MPa when injected into the tank, and perform pressurization, decompression and flow control processes respectively; when pressurizing, slowly open the pressure valve 5, monitor the pressure and flow in the pipeline in real time, and when the target pressure is reached, keep the helium pressure valve 5 at an appropriate opening to maintain the pressure of the gas cylinder and the pipeline. When decompressing, slowly open the isolation valve 10, allow the high-pressure gas to enter the pressure reducer 6, gradually adjust the pressure reducer 6, reduce the helium pressure, and achieve the target decompression effect. The flow rate during the decompression process can be controlled by the safety valve 7. If there is overpressure, the safety valve 7 will automatically exhaust. Finally, the one-way valve 9 is opened to allow helium to be injected into the liquid oxidant tank 2 and the liquid fuel tank 3. The working pressure of the tank is 1-3 MPa, so as to achieve the effect of pressurizing the tank. The liquid oxidant and liquid fuel in the tank are injected into the combustion chamber 18 under the action of the oxidant pump and the fuel pump. 10% to 20% of them are burned in the gas generator 14, driving the turbine 15 to do work, providing energy for the operation of the oxidant pump and the fuel pump, and realizing the cyclic working process. The remaining liquid fuel and liquid oxidant are burned in the combustion chamber 18 to generate combustion gas, which is then discharged from the engine through the nozzle 19. When the rocket returns to the atmosphere, 60% to 80% of the low-temperature helium flows into the second gas pipeline equipped with the pressure regulating device. The operation process of the pressure regulating device is the same as the above-mentioned valve working process. Then, the low-temperature helium cyclone is injected into the combustion chamber 18 to form a rotating gas film, which flows to the outlet of the nozzle 19 under the action of the main flow of the gas. The low-temperature gas film covers the inner wall surface of the nozzle, reduces the temperature of the inner wall surface of the engine to achieve a cooling effect. As the main stream of gas exits the nozzle, the low-temperature gas film forms a vortex due to the action of the reverse flow, and flows along the outer wall of the engine, reducing the ablation effect of the high-temperature gas on the engine.

[0032] Figure 3 This is a flow field diagram of the principle of cryogenic helium cooling the inner and outer walls of the rocket engine during recovery and landing. At this time, liquid fuel and liquid oxidizer react with each other in the inner cavity 21 of the combustion chamber to generate a mainstream gas, and the cryogenic helium flows along the combustion chamber and nozzle wall 23 along the flow direction 25 of the mainstream gas. The flow field at the end of the engine forms a bow shock wave 28 due to the interference of the external lateral flow 29. Then, the mainstream gas 25 begins to flow separate at the flow separation point 27 under the action of the bow shock wave, forming a reflow zone 26 on the outer wall of the engine. The flow direction 24 of the cryogenic helium still flows along the outer wall of the engine to avoid direct contact between the high-temperature gas and the wall, thereby reducing the scouring and ablation of the engine wall by the high-temperature gas and improving the reusability of the engine.

[0033] The above is only a preferred embodiment of the present invention. It should be pointed out that for ordinary technicians in this technical field, several improvements and modifications can be made without departing from the technical principles of the present invention. For example, using different forms of air intake structures, different air intake structure placement positions and angles, etc. These improvements and modifications should also be regarded as the protection scope of the present invention. For example, using other types of low-temperature gas injection should also be regarded as the protection scope of the present invention.

Claims

1. A cryogenic gas injection heat protection method for a reusable rocket recovery process, characterized in that: A swirl flow channel (20) is arranged around the combustion chamber of the engine, and the low-temperature gas is injected into the combustion chamber of the engine through the swirl flow channel (20) after pressure regulation to form a swirl film; the liquid fuel and the liquid oxidant react with each other in the inner cavity of the combustion chamber to generate a main flow of the combustion gas, and the low-temperature gas flows along the combustion chamber and the nozzle wall in the direction of the main flow of the combustion gas, thereby avoiding direct contact between the high-temperature combustion gas and the wall, reducing the scouring and ablation of the engine wall by the high-temperature combustion gas, and improving the reusability of the engine.

2. An application of the low-temperature gas injection heat protection method for a reusable rocket recovery process as claimed in claim 1, characterized in that: A cooling system is arranged inside the engine, comprising: a high-pressure gas cylinder (1), a liquid oxidant tank (2), a liquid fuel tank (3), an oxidant pump, a fuel pump, a gas generator (14), a turbine (15), and a buffer chamber (17); After the first outlet of the high-pressure gas cylinder (1) is connected to the pressure regulating device through the first gas pipeline, it is respectively connected to the inlets of the liquid oxidant tank (2) and the liquid fuel tank (3) through two branches of the first gas pipeline; the inlet branches of the liquid oxidant tank (2) and the liquid fuel tank (3) are respectively provided with a one-way valve that only allows gas injection; after the second outlet of the high-pressure gas cylinder (1) is connected to the pressure regulating device through the second gas pipeline, it is connected to the buffer chamber (17) through the one-way valve that only allows gas injection; a swirl flow channel (20) is circumferentially distributed at the front end of the combustion chamber; the pipeline of the buffer chamber (17) is connected to the swirl flow channel (20); The liquid oxidant tank (2) and the liquid fuel tank (3) are connected to the combustion chamber through a liquid delivery pipeline installed with an oxidant pump and a fuel pump respectively; the inlet of the gas generator (14) is connected to the outlet of the oxidant pump and the fuel pump respectively through a liquid delivery pipeline, the gas chamber outlet of the gas generator (14) is connected to the input end of the turbine (15), and the output end of the turbine (15) is connected to the oxidant pump and the fuel pump.

3. The application of the low temperature gas injection heat protection method for the reusable rocket recovery process according to claim 2 is characterized in that: The pressure regulating device comprises a pressure switch, a pressurizing valve, an isolating valve, a pressure reducer, a pressure sensor and a safety valve which are sequentially installed on the first gas pipeline and the second gas pipeline.

4. The application of the low temperature gas injection heat protection method for the reusable rocket recovery process according to claim 2 is characterized in that: The low-temperature gas stored in the high-pressure gas cylinder (1) is an inert gas with a temperature below 300K.

5. The application of the low temperature gas injection heat protection method for the reusable rocket recovery process according to claim 2, characterized in that: The liquid oxidant tank (2) and the liquid fuel tank (3) are provided with exhaust valves (11); the injection hole of the swirl flow channel (20) is at an angle of 30° to 90° to the outer wall of the combustion chamber.

6. The application of the low temperature gas injection heat protection method for the recovery process of a reusable rocket according to claim 2, characterized in that: The turbine (15) outlet is connected to a turbine exhaust pipe (16); and valves for controlling the flow rate of liquid passing through are respectively provided on the first gas pipeline connected to the combustion chamber.

7. The application of the low temperature gas injection heat protection method for the reusable rocket recovery process according to claim 2, characterized in that: When the liquid rocket starts the engine, the one-way valve is opened to allow part of the gas in the high-pressure gas cylinder (1) to be injected into the liquid oxidant tank (2) and the liquid fuel tank (3) through the first gas pipeline; according to the pressure difference between the high-pressure gas cylinder and the liquid oxidant tank (2) and the liquid fuel tank (3), the pressure regulating device is used to regulate the high-pressure gas cylinder (1) so that the pressure of the high-pressure gas cylinder (1) is between 10 and 30 MPa, and the pressure of the liquid oxidant tank (2) and the liquid fuel tank (3) is increased to 1 to 3 MPa; part of the liquid oxidant and liquid fuel in the oxidant tank (2) and the fuel tank (3) are injected into the combustion chamber under the action of the oxidant pump and the fuel pump. , a combustion reaction is carried out in the combustion chamber to generate combustion gas, which is then discharged from the engine through the nozzle; the other part of the liquid oxidant and liquid fuel is burned in the gas generator (14), driving the turbine (15) to do work, providing energy for the operation of the oxidant pump and the fuel pump, thereby realizing a cyclic working process; when the rocket begins to land, the gas in the high-pressure gas cylinder (1) begins to be injected into the buffer chamber (17) after pressure regulation through the second gas transmission pipeline, and the gas in the buffer chamber (17) is directly injected into the combustion chamber through the swirl flow channel (20), forming a layer of rotating gas film, which flows toward the nozzle outlet under the action of the main flow of the gas, and the low-temperature gas film adheres to the combustion chamber and the inner wall surface of the nozzle, thereby reducing the temperature of the inner wall surface; As the main stream of gas exits the nozzle, the low-temperature gas film forms a vortex due to the action of the reverse flow, and flows along the outer wall of the engine, reducing the ablation effect of the high-temperature gas on the engine.

8. The application of the low temperature gas injection heat protection method for the reusable rocket recovery process according to claim 2, characterized in that: 20% to 40% of the low-temperature gas in the high-pressure gas cylinder (1) is respectively injected into the liquid oxidant tank (2) and the liquid fuel tank (3) via the first gas pipeline; 60% to 80% of the low-temperature gas in the high-pressure gas cylinder (1) is injected into the combustion chamber via the second gas pipeline connected to the swirl flow channel (20); and 10% to 20% of the fuel in the fuel tank (3) is burned in the gas generator (14).

9. The application of the low temperature gas injection heat protection method for the reusable rocket recovery process according to claim 2, characterized in that: When the pressure in the high-pressure gas cylinder (1) is less than 10MPa, the pressure regulating device pressurizes the high-pressure gas cylinder (1), slowly opens the pressurizing valve, and monitors the pressure in the pipeline in real time through the pressure sensor; when the target pressure is reached, the pressurizing valve is kept at an appropriate opening to maintain the pressure in the gas cylinder and the pipeline; then the one-way valve is opened to allow the low-temperature gas to be injected into the liquid oxidant tank (2), the liquid fuel tank (3) and the combustion chamber respectively.

10. The application of the low temperature gas injection heat protection method for the reusable rocket recovery process according to claim 2, characterized in that: When the pressure in the high-pressure gas cylinder (1) is greater than 30 MPa, the pressure regulating device pressurizes the high-pressure gas cylinder (1), slowly opens the isolation valve, allows the low-temperature gas to enter the pressure reducer, and gradually adjusts the pressure reducer to reduce the pressure of the low-temperature gas to reach the target pressure; during the pressure reduction process, the flow of the low-temperature gas can be controlled by the safety valve, and if there is overpressure, the safety valve will automatically exhaust; then open the one-way valve to allow the low-temperature gas to be injected into the liquid oxidant tank (2), the liquid fuel tank (3) and the combustion chamber.

Images