A thrust device for a liquid rocket engine with a composite cooling form

Abstract

A thrust device of a liquid rocket engine with a composite cooling form. Fuel and oxidizer are mixed and burned through an annular gap from different directions. The fuel is introduced as a coolant from the throttling inlet at the top of the injection panel into the coolant channel by using the effusion cooling technology, and the heat of the inner wall of the thrust chamber is carried away through convective heat transfer. In the throat expansion section, film cooling and radiation cooling are coupled. The coolant droplets discharged from the coolant discharge port spread on the wall of the thrust chamber to form a protective liquid film. After evaporation and heat absorption, liquid film steam is formed. The high-speed steam can carry away the heat of the gas through external radiation and form a protective barrier at the tail of the nozzle. Low-temperature fuel is introduced at the regenerative coolant inlet at the end of the nozzle. The fuel coolant flows in the regenerative coolant channel and absorbs the heat of the wall of the thrust chamber, and is transported to the regenerative coolant outlet at the top of the injection panel. The preheated fuel can enter the fuel storage cavity again to increase the inlet fuel temperature, enhance the combustion performance of the propellant, and at the same time take into account the cooling of the wall of the thrust chamber.

Description

Technical Field

[0001] The invention belongs to the technical field of liquid rocket engines, and in particular relates to a liquid rocket engine thrust device in a composite cooling form for thermal protection of the liquid rocket engine. Background Art

[0002] During the engineering process of liquid rocket engines, the temperature in the thrust chamber is as high as 3500 K, the gas velocity at the nozzle outlet is as high as Ma 6, the gas pressure can reach 20 MPa, and the heat flux density near the throat is as high as 160 MW / m 2 The intense high-speed convection and radiation heat exchange between the gas and the wall puts the engineering materials of the thrust chamber wall under great test. The melting point of general high-temperature resistant alloy materials, such as tungsten, molybdenum, and niobium alloy materials, is between 2700 K and 3600 K. In the high-temperature, high-speed, and high-pressure gas environment, the thrust chamber wall is prone to overheating, oxidation, corrosion, and burning. Therefore, certain thermal protection measures must be taken to protect the rocket engine thrust chamber.

[0003] At present, there are four main types of cooling technologies for rocket engine thrust chambers:

[0004] 1) Ablation cooling technology. This method is to cover the inner wall of the thrust chamber with an anti-ablation coating. The coating material is a high-temperature resistant carbon composite material. The high-temperature fuel gas consumes the coating material to protect the inner wall of the thrust chamber. This method is widely used in solid rocket thrusters.

[0005] 2) Discharge cooling technology: This method is to inject liquid coolant from the cooling channel opening at a certain position on the head or body of the thrust chamber. The coolant removes the heat from the thrust chamber wall through convection exchange in the channel and is discharged from the channel outlet at the tail.

[0006] 3) Film cooling technology: This method is to spray the coolant liquid film directly onto the inner wall of the thrust chamber. The coolant droplets evaporate and absorb heat, thereby reducing the temperature of the thrust chamber wall.

[0007] 4) Radiative cooling technology: This method uses the coolant vapor’s ability to absorb heat through external radiation, and can cool small areas with relatively low heat flux, such as the nozzle or tail of a small engine.

[0008] In addition, the regenerative cooling technology of introducing low-temperature fuel through a dedicated channel at the tail of the nozzle, absorbing heat from the wall by convection heat transfer, and then re-entering the injection panel at the head for injection combustion has also received widespread attention in recent years. According to the temperature distribution inside the thrust chamber of a liquid rocket and the characteristics of different cooling methods, designing a thrust chamber with multiple cooling forms is an important issue in the field of rocket engine thermal protection. Summary of the invention

[0009] In view of the deficiencies of the existing single cooling or composite cooling technologies for liquid rocket thrust chambers, the present invention proposes a thrust device for a liquid rocket engine with a composite cooling form that adopts multiple cooling methods and can effectively reduce the internal temperature of the thrust chamber.

[0010] To achieve the above object, the technical solution adopted by the present invention is: it includes a thrust chamber, a propellant injection system and a thrust chamber cooling system arranged at the head of the thrust chamber;

[0011] The propellant injection system of the thrust chamber includes an injection panel and an injector housing connected to the injection panel. A needle valve and a needle bolt top cap communicated with the oxidizer injection port are arranged at the center of the injector housing. An oxidizer storage cavity is formed between the inner surface of the needle valve and the outer surface of the needle bolt top cap. A fuel injection port is arranged on the outer side surface of the injector housing. A fuel storage cavity communicated with the fuel injection port is formed among the injector housing, the outer page of the needle valve and the injection panel;

[0012] An anti-ablation coating is provided on the inner wall of the thrust chamber from the injection panel to the throat converging section;

[0013] The thrust chamber cooling system includes a coolant throttle valve and a regenerative coolant outlet opened on the injector housing, a coolant channel with a coolant discharge port at the end connected to the coolant throttle valve and extending from the injection panel to the throat expansion area on the inner wall surface of the thrust chamber, a regenerative coolant channel with a regenerative coolant inlet at the lower end on the outer wall of the coolant channel and connected to the regenerative coolant outlet. The regenerative coolant inlet is arranged at the end of the nozzle. The regenerative coolant outlet is connected to the fuel injection port, and the preheated fuel can enter the fuel storage cavity again.

[0014] An annular fuel injection gap with a spacing of 0.05 mm - 0.1 mm is formed between the injection panel and the outer surface of the needle valve.

[0015] An annular oxidizer injection gap with a spacing of 0.1 mm - 1 mm is formed between the inner surface of the needle valve and the needle bolt top cap.

[0016] A regenerative coolant circulation pump is also provided at the inlet of the regenerative coolant channel.

[0017] The coolant throttle valve includes a valve seat. A baffle with a through hole is arranged in the valve seat. The baffle divides the valve seat into upper and lower channels. A valve core is installed in the lower channel. The valve core includes a coil, a fixed iron core and a moving iron core arranged in the coil. A spring is arranged between the fixed iron core and the moving iron core. The front end of the moving iron core is a conical structure adapted to the through hole.

[0018] The injector housing and the injection panel are connected by bolts.

[0019] The injector housing and the needle valve are connected by studs.

[0020] The coolant channels are milled on the inner wall of the thrust chamber, and the regenerative coolant channels are milled on the outer wall of the thrust chamber. The outer sides of the coolant channels and the regenerative coolant channels are sealed by electroplating.

[0021] The coolant channels are in a spiral structure or a vertical tube bundle structure. The cross-section of the pipeline is a rectangle with a length of 6 mm and a width of 4 mm, and the wall thickness of the pipe is 0.6 mm.

[0022] The regenerative coolant channels are in the form of a vertical tube bundle. The cross-section of the pipeline is a rectangle with a length of 10 mm and a width of 8 mm, and the wall thickness of the pipe is 0.6 mm.

[0023] In the present invention, the fuel and the oxidant are mixed and burned from the annular gaps in different directions. The anti-ablation coating in the area from the injection panel to the wall surface of the throat converging section uses the ablation cooling technology to reduce the heat transferred from the combustion gas to the upstream of the thrust chamber. The fuel is used as the coolant and introduced into the coolant channels from the throttling inlet at the top of the injection panel by using the effusion cooling technology, and the heat on the inner wall of the thrust chamber is taken away through the convective heat transfer. In the throat expansion section, film cooling and radiation cooling are coupled. The coolant droplets discharged from the coolant discharge ports spread on the wall surface of the thrust chamber to form a protective liquid film. After evaporation and heat absorption, liquid film steam is formed. The high-speed steam can radiate externally to take away the heat of the combustion gas and form a protective barrier at the tail of the nozzle. Finally, low-temperature fuel is introduced into the regenerative coolant inlet at the end of the nozzle. The fuel coolant flows in the regenerative coolant channels and absorbs the heat on the wall surface of the thrust chamber, and is transported to the regenerative coolant outlet at the top of the injection panel. The preheated fuel can enter the fuel storage cavity again to increase the inlet fuel temperature, enhance the combustion performance of the propellant, and at the same time take into account the cooling of the wall surface of the thrust chamber. Brief Description of the Drawings

[0024] Figure 1 is the overall structural schematic diagram of the present invention.

[0025] In the drawings: 11, oxidant injection port; 12, injector housing; 13, fuel injection port; 14, injection panel; 15, needle valve; 16, needle bolt top cap; 17, fuel storage cavity; 18, oxidant storage cavity; 19, thrust chamber; 2, anti-ablation coating; 31, coolant throttling inlet; 32, coolant channels; 33, coolant discharge ports; 41, regenerative coolant inlet; 42, regenerative coolant circulation pump; 43, regenerative coolant channels; 44, regenerative coolant outlet; 51, connecting bolts; 52, connecting studs.

[0026] Figure 2 is the device diagram of the coolant throttling inlet.

[0027] In the attached drawings: 311, valve seat; 312, coil; 313, stationary iron core; 314, spring; 315, moving iron core; 316, valve core; 317, through hole; 318, baffle plate. Detailed implementation mode

[0028] The present invention will be further described below in conjunction with the embodiments shown in the attached drawings.

[0029] See Figure 1 , the present invention includes a thrust chamber 19, a propellant injection system and a thrust chamber cooling system arranged at the head of the thrust chamber;

[0030] The propellant injection system of the thrust chamber includes an injection panel 14 and an injector housing 12 connected to the injection panel 14 by bolts 51. A needle valve 15 and a needle bolt top cap 16 communicating with an oxidizer injection port 11 are arranged at the center of the injector housing 12. The injector housing 12 and the needle valve 15 are connected by studs 52. An oxidizer storage cavity 18 is formed between the inner surface of the needle valve 15 and the outer surface of the needle bolt top cap 16. An annular oxidizer injection gap is formed between the inner surface of the needle valve 15 and the needle bolt top cap 16, and the spacing is 0.1 mm - 1 mm. A fuel injection port 13 is arranged on the outer side surface of the injector housing 12. The injector housing 12, the needle valve 15 and the injection panel 14 form a fuel storage cavity 17 communicating with the fuel injection port 13; an annular fuel injection gap is formed between the injection panel 14 and the outer surface of the needle valve 15, and the spacing is 0.05 mm - 0.1 mm.

[0031] An anti-ablation coating 2 is provided on the inner wall of the convergent section of the thrust chamber 19 from the injection panel 14 to the throat.

[0032] The thrust chamber cooling system includes a coolant throttle valve 31 and a regenerative coolant outlet 44 opened on the outer side surface of the injector housing 12, a coolant channel 32 connected to the coolant throttle valve 31 and having a coolant discharge port 33 at the end extending from the injection panel 14 to the throat expansion area on the outer wall surface of the thrust chamber 19, and a regenerative coolant channel 43 provided at the lower end of the outer wall of the coolant channel 32 and having a regenerative coolant inlet 41 and a regenerative coolant circulation pump 42 and connected to the regenerative coolant outlet 44.

[0033] The coolant channel 32 of the present invention is machined by milling on the inner wall of the thrust chamber 19, and the regenerative coolant channel 43 is machined by milling on the outer wall of the thrust chamber. The outer sides of the coolant channel 32 and the regenerative coolant channel 43 are sealed by electroplating. This processing method can reduce the material stress and deformation during the manufacturing process and can withstand a higher thermal load during the thermal operation. The length of the coolant channel can be extended to the convergent section in front of the throat or to the expansion section behind the throat according to the specific working conditions.

[0034] The coolant channel 32 is of a spiral structure or a vertical tube bundle structure. The cross-section of the pipe is a rectangle with a length of 6 mm and a width of 4 mm, and the wall thickness of the pipe is 0.6 mm.

[0035] The regenerative coolant channel 43 is in the form of a vertical tube bundle. The cross-section of the pipe is a rectangle with a length of 10 mm and a width of 8 mm, and the wall thickness of the pipe is 0.6 mm, reducing the flow resistance of the regenerative coolant.

[0036] See Figure 2 , the coolant throttle valve 31 of the present invention includes a valve seat 311. A baffle 318 with a through hole 317 is arranged in the valve seat 311. The baffle 318 divides the valve seat 311 into upper and lower channels. A valve core 316 is installed in the lower channel. The valve core 316 includes a coil 312, a fixed iron core 313 and a movable iron core 315 arranged in the coil 312. A spring 314 is arranged between the fixed iron core 313 and the movable iron core 315. The front end of the movable iron core 315 is a conical structure adapted to the through hole 317.

[0037] Its working process is as follows: fuel enters from the fuel injection port 13 opened on the side of the injector housing 12, fills the entire fuel storage chamber 17, and then enters the thrust chamber head through the annular gap formed by the outer side of the needle valve 15 and the injection panel 14; oxidant enters from the oxidant injection port 11 opened on the top surface of the injector housing 12, fills the entire oxidant storage chamber 18, and then enters the thrust chamber head through the annular gap formed by the inner side of the needle valve 15 and the needle plug top cap 16; two jets of fuel and oxidant in different directions are atomized and mixed at the thrust chamber head, and after combustion, high-temperature and high-pressure gas is generated, and the heat load upstream of the thrust chamber is relatively large. During cooling, ablation cooling is first used, and an anti-ablation coating 2 is set on the wall surface of the area from the injection panel to the throat convergence section. The hot combustion gas quickly reacts with the high-temperature resistant carbon composite material in the anti-ablation coating 2 to generate carbon solid particles that adhere to the inner wall of the thrust chamber, isolating the combustion gas from transferring heat to the inner wall of the thrust chamber; secondly, part of the fuel enters the coolant channel 32 from the coolant throttling inlet 31 as a coolant, takes away the heat of the inner wall of the thrust chamber through convection exchange, and is discharged from the coolant discharge port 33 to complete the discharge cooling. The amount of fuel coolant is controlled by a throttling device. When the coil 312 is energized, the fixed iron core 313 attracts the moving iron core 315 to move rightward, thereby compressing the spring 314, and the valve core 316 connected to the moving iron core 315 moves rightward, and the flow cross-sectional area becomes larger, thereby increasing the fuel flow rate entering the coolant channel 32, and vice versa. The coolant droplets discharged from the coolant discharge port 33 can spread on the wall to form a protective liquid film if the evaporation temperature has not been reached, and form liquid film steam after evaporation and heat absorption. The high-speed steam radiates heat to the outside, which can take away the heat of the gas and form a protective barrier at the tail of the nozzle, playing a combined role of film cooling and radiation cooling. Finally, in order to increase the inlet fuel temperature, enhance the combustion performance of the propellant, and take into account the cooling of the inner wall of the thrust chamber, a low-temperature fuel is introduced into the regenerative coolant inlet 41 at the end of the nozzle. Through the action of the regenerative coolant circulation pump 42, the coolant flows in the regenerative coolant channel 43 and absorbs the heat of the thrust chamber wall, and is transported to the regenerative coolant outlet 44 above the injection panel 14. The regenerated coolant outlet 44 is connected to the fuel injection port 13 , and the preheated fuel can enter the fuel storage chamber 17 again.

[0038] The present invention improves the existing single cooling method for the thrust chamber wall of a liquid rocket engine. According to the heat distribution characteristics inside the thrust chamber, a composite design is made for ablation cooling and emission cooling in the upstream area from the injection panel to the throat where the heat concentration is relatively large; the heat in the downstream area of ​​the nozzle is relatively low, so film cooling and radiation cooling can be coupled. Finally, regenerative cooling is introduced at the end of the nozzle to absorb the heat from the wall and preheat the fuel. The design of the present invention couples multiple cooling methods, which can effectively reduce the temperature inside the thrust chamber of a liquid rocket engine.

Claims

1. A thrust device for a liquid rocket engine with a composite cooling form, characterized in that: It includes a thrust chamber (19), a propellant injection system and a thrust chamber cooling system provided at the head of the thrust chamber; The propellant injection system of the thrust chamber includes an injection panel (14) and an injector housing (12) connected to the injection panel (14). A needle valve (15) and a needle bolt top cap (16) communicating with an oxidizer injection port (11) are provided at the center of the injector housing (12). An oxidizer storage cavity (18) is formed between the inner surface of the needle valve (15) and the outer surface of the needle bolt top cap (16). A fuel injection port (13) is provided on the outer side surface of the injector housing (12). A fuel storage cavity (17) communicating with the fuel injection port (13) is formed between the outer surface of the injector housing (12), the needle valve (15) and the injection panel (14); An ablation-resistant coating (2) is provided on the inner wall of the thrust chamber (19) from the injection panel (14) to the throat converging section; The thrust chamber cooling system includes a coolant throttle valve (31) and a regenerative coolant outlet (44) provided on the injector housing (12), a coolant channel (32) located on the inner wall surface of the thrust chamber (19) extending from the injection panel (14) to the throat expansion area and connected to the coolant throttle valve (31) with a coolant discharge port (33) at the end, a regenerative coolant channel (43) located on the outer wall of the coolant channel (32) with a regenerative coolant inlet (41) at the lower end and connected to the regenerative coolant outlet (44). The regenerative coolant inlet (41) is provided at the end of the nozzle. The regenerative coolant outlet (44) is connected to the fuel injection port (13), and the preheated fuel can enter the fuel storage cavity (17) again.

2. The thrust device of the liquid rocket engine with a composite cooling form according to claim 1, characterized in that: An annular fuel injection gap is formed between the injection panel (14) and the outer surface of the needle valve (15), and the spacing is 0.05 mm - 0.1 mm.

3. The thrust device of the liquid rocket engine with a composite cooling form according to claim 1, characterized in that: An annular oxidizer injection gap is formed between the inner surface of the needle valve (15) and the needle bolt top cap (16), and the spacing is 0.1 mm - 1 mm.

4. The thrust device of the liquid rocket engine with a composite cooling form according to claim 1, characterized in that: A regenerative coolant circulation pump (42) is also provided at the inlet of the regenerative coolant channel (43).

5. The thrust device of the liquid rocket engine with a composite cooling form according to claim 1, characterized in that: The coolant throttle valve (31) includes a valve seat (311). A baffle (318) with a through hole (317) is provided in the valve seat (311). The baffle 318 divides the valve seat (311) into upper and lower channels. A valve core (316) is installed in the lower channel. The valve core 316 includes a coil (312), a fixed iron core (313) and a moving iron core (315) provided in the coil (312). A spring (314) is provided between the fixed iron core (313) and the moving iron core (315). The front end of the moving iron core (315) is a conical structure adapted to the through hole (317).

6. The thrust device of the liquid rocket engine with a composite cooling form according to claim 1, characterized in that: The injector housing (12) and the injection panel (14) are connected by bolts (51).

7. The thrust device of the liquid rocket engine with a composite cooling form according to claim 1, characterized in that: The injector housing (12) and the needle valve (15) are connected by studs (52).

8. The thrust device of the liquid rocket engine with a composite cooling form according to claim 1, characterized in that: The described coolant channel (32) is milled on the inner wall of the thrust chamber (19), and the regenerative coolant channel (43) is milled on the outer wall of the thrust chamber. The outer sides of the coolant channel (32) and the regenerative coolant channel (43) are sealed by electroplating.

9. The thrust device of the liquid rocket engine with a composite cooling form according to claim 1, characterized in that: The described coolant channel (32) is in a spiral structure or a vertical tube bundle structure. The cross-section of the pipe is a rectangle with a length of 6 mm and a width of 4 mm, and the wall thickness of the pipe is 0.6 mm.

10. The thrust device of the liquid rocket engine with a composite cooling form according to claim 1, characterized in that: The described regenerative coolant channel (43) is in the form of a vertical tube bundle. The cross-section of the pipe is a rectangle with a length of 10 mm and a width of 8 mm, and the wall thickness of the pipe is 0.6 mm.

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