Reusable airspace engine and thermal management sleeve system

Abstract

The invention provides a reusable airspace engine and thermal management sleeve system, which comprises an engine and a thermal management sleeve system, and is characterized in that the engine comprises an injector and a spray pipe; the injector is connected with the heat management sleeve system, and the spray pipe is located in a cavity of the heat management sleeve system; the engine jet pipe is embedded in the beveled sleeve of the thermal management sleeve system, and the thermal management sleeve system of a composite structure is matched, so that heat insulation can be achieved, surrounding parts can be protected, heat can be absorbed, the jet pipe is prevented from being overheated, and the contradiction between the compact space of a cabin and the thermal protection requirement is solved; the system integrally solves the contradiction between the compact space of the cabin and the thermal protection requirement, can reduce pneumatic interference and thermal influence when the cabin enters the atmosphere, and also solves the problems of detectability and safety in the detection process due to repeated use of an engine.

Description

TECHNICAL FIELD

[0001] The present application belongs to the technical field of reusable spacecraft engine, and particularly relates to a reusable space engine and a thermal management sleeve system. BACKGROUND

[0002] The cargo space shuttle capable of going to and returning from the earth repeatedly becomes one of the feasible technical solutions to meet the demand of the low-cost cargo transportation system of the space station. The cargo space shuttle will be subjected to aerodynamic force heat when going to and returning from the earth, so the attitude and orbit control engine needs to be embedded in the cabin body. However, the temperature of the nozzle is very high when the engine is ignited, which will inevitably have a great thermal radiation effect on the surrounding components in the cabin body, and even directly cause the failure of the surrounding components. At the same time, due to the compact structure in the cabin body, if the nozzle has limited heat dissipation space, the temperature of the nozzle itself will rise by more than 200℃, and when the temperature exceeds the upper limit of the temperature resistance of the nozzle material, the nozzle will be burned and fail. On the other hand, after the engine returns to the ground, the test procedures such as leakage of the highly toxic and strongly corrosive propellant, sealing detection and the like need to be completed on the spacecraft, so that the spacecraft has the ability of rapid reflight. The existing engine products cannot meet the use requirements of the reusable spacecraft with limited space.

[0003] At present, in the related prior art:

[0004] A cooling jacket of a liquid space engine with good cooling effect (patent document CN112282973A) discloses a cooling jacket of a liquid space engine with good cooling effect, comprising a nozzle body, the outer wall of the top of the nozzle body is fixed with symmetrically divided metal pipes, one end of the metal pipe is fixed with a disc, a plurality of uniformly divided air inlet holes are opened in the top of the disc, and the inner wall of the air inlet hole is connected with the metal pipe, and the nozzle body is sequentially provided with an air inlet groove and a heat conduction groove from the outside to the inside. By using the nozzle body, the metal pipe, the disc, the air inlet hole, the air inlet groove and the air outlet valve, the outside cold air can be introduced into the air inlet groove and circulated to cool the nozzle body when the space engine is running, so that the cooling efficiency of the cooling jacket is improved. By using the radial small hole, the heat conduction groove, the metal pipe, the inlet flange, the outlet flange and the turboprop, part of the heat generated can be introduced into the heat conduction groove through the radial small hole, and the contact area between the hot gas and the cooling hydrogen is increased, so that the cooling effect of the device is better. However, this technical solution relies on external cold air medium, and the heat exchange efficiency between the gas in the groove and the solid wall surface is low, which is difficult to provide thermal protection for the space engine generating 3000K high-temperature gas in a vacuum environment without gas medium. SUMMARY

[0005] In view of the defects in the prior art, the purpose of the present application is to provide a reusable space engine and a thermal management sleeve system.

[0006] The reusable aerospace engine and thermal management sleeve system provided by the application comprises:

[0007] The engine and the thermal management sleeve system,

[0008] The engine comprises an injector and a nozzle.

[0009] The injector is connected with the thermal management sleeve system, and the nozzle is located in the cavity of the thermal management sleeve system.

[0010] Preferably, the nozzle is a Laval nozzle, the profile is first contracted and then expanded, and an annular groove is arranged on the outlet end surface of the nozzle; and the highest working temperature of the nozzle without the thermal management sleeve system is ≤1200℃.

[0011] Preferably, the thermal management sleeve system comprises a bevel sleeve, an inner layer anti-oxidation coating, an outer layer metal film and multi-layer thermal insulation materials.

[0012] The nozzle is arranged in the thermal management sleeve system, and the outlet end surface is shorter than the short side of the bevel sleeve.

[0013] The inner wall surface of the bevel sleeve is provided with the inner layer anti-oxidation coating, and the outer wall surface is provided with the outer layer metal film, and the outermost layer is entirely coated with the multi-layer thermal insulation materials.

[0014] Preferably, the bevel sleeve is provided with an inlet boss, a cylindrical segment, an outlet flange, bolt mounting holes and threaded holes.

[0015] The bolt mounting holes are uniformly arranged on the inlet boss, and the threaded holes are uniformly arranged on the outlet flange.

[0016] Preferably, the inner profile of the bevel sleeve is a cylindrical structure, the outlet end surface is beveled, and an angle exists between the cutting surface and the axis.

[0017] Preferably, the material of the nozzle is niobium hafnium or niobium tungsten alloy, the material of the bevel sleeve is C / C composite material, C / SiC composite material or GH3128 high-temperature alloy, the multi-layer thermal insulation materials are composed of several layers of double-sided aluminum-coated polyimide films, the inner layer anti-oxidation coating is a composite coating of SiC / Si-B4C system, which can maintain good anti-oxidation performance below 1500℃, the outer wall surface is provided with the outer layer metal film, the outer layer metal film is a low-emissivity film, and the emissivity of the film is ≤0.4.

[0018] The bevel sleeve is connected and fixed with the engine through the plurality of bolts; and the distance between the inner wall of the cylindrical segment of the bevel sleeve and the outer wall of the nozzle is 1-5mm.

[0019] Preferably, the outlet flange thickness is greater than or equal to 5 mm; the bolt mounting hole diameter is greater than or equal to 4 mm, and the number is greater than or equal to 3; the threaded hole diameter is greater than or equal to 4 mm, and the number is greater than or equal to 4; the wall thickness of the beveled sleeve cylindrical segment is greater than or equal to 2.5 mm.

[0020] Preferably, the angle between the beveled sleeve outlet end face and the axis is 40° to 80°.

[0021] A method for using a reusable aerospace engine and a thermal management sleeve system provided by the application, comprising:

[0022] Step S1: After the engine returns to the ground with the spacecraft, an O-ring made of perfluoroether material is placed in the annular groove at the outlet end face of the nozzle;

[0023] Step S2: The sealing tool is fastened to the sleeve flange by bolts, and the O-ring is pressed to close the nozzle;

[0024] Step S3: The residual propellant is discharged along the nozzle through the tool center flow channel;

[0025] Step S4: The outlet of the plugging tool is sealed, and the engine is subjected to a sealing test.

[0026] Compared with the prior art, the application has the following beneficial effects:

[0027] 1. The engine nozzle is embedded in the beveled sleeve, and a thermal management sleeve system with a composite structure is arranged, which can not only protect the surrounding components from heat, but also absorb heat to prevent the nozzle from overheating, solving the contradiction between the compact space of the cabin and the heat protection demand;

[0028] 2. The outlet end face of the beveled sleeve protrudes from the nozzle outlet and is beveled at a certain angle to adapt to the cabin profile, which can reduce the aerodynamic interference and thermal influence when entering the atmosphere;

[0029] 3. The annular groove is machined at the outlet end face of the nozzle to install a sealing O-ring, and the threaded hole of the beveled sleeve flange is machined to connect the sealing tool, forming a closed discharge flow channel, which meets the cleaning and testing requirements of the engine on the spacecraft, greatly shortens the maintenance time compared with the traditional engine, and solves the problem of detectability and safety during the detection process of the reusable engine.

[0030] 4. The engine itself temperature control and thermal management sleeve heat insulation scheme are used, which can realize heat protection in the vacuum environment in space, and the structure is simple and reliable, does not involve the physical state change of the cooling medium, and does not depend on external cold air medium, so that the heat protection can be provided for the aerospace engine generating 3000K high-temperature gas in the vacuum environment without gas medium. BRIEF DESCRIPTION OF DRAWINGS

[0031] Other features, objects and advantages of the present invention will become more apparent upon reading the detailed description of non-limiting embodiments with reference to the following drawings:

[0032] Figure 1 This is a schematic structural diagram of a reusable aerospace engine and thermal management sleeve system according to an embodiment of the present invention.

[0033] Figure 2 It is a schematic cross-sectional view of the structure of the beveled sleeve of the present invention.

[0034] Figure 3 for Figure 1 Enlarged view of point I in the middle.

[0035] The figure shows:

[0036] DETAILED DESCRIPTION

[0037] The present invention will be described in detail below with reference to specific embodiments. The following examples will help those skilled in the art to further understand the present invention, but are not intended to limit the present invention in any form. It should be noted that, for those skilled in the art, several changes and improvements can be made without departing from the scope of the present invention. These all fall within the scope of protection of the present invention.

[0038] like Figure 1 As shown, an embodiment of the present invention provides a reusable aerospace engine and thermal management sleeve system, comprising: an engine 1 and a thermal management sleeve system 2. The engine 1 includes an injector 11 and a nozzle 12. The injector 11 is connected to the thermal management sleeve system 2, and the nozzle 12 is located within the cavity of the thermal management sleeve system 2. Furthermore, the nozzle 12 is a Laval nozzle, with a profile that first contracts and then expands. The outlet end surface of the nozzle 12 is provided with an annular groove 121. The maximum operating temperature of the nozzle 12 when not equipped with the thermal management sleeve system 2 is ≤1200°C.

[0039] Furthermore, the thermal management sleeve system 2 includes a beveled sleeve 21, an inner anti-oxidation coating 22, an outer metal film 23 and a multi-layer thermal insulation material 24; the nozzle 12 is arranged in the thermal management sleeve system 2, and the outlet end face is slightly shorter than the short side of the beveled sleeve 21; the inner wall surface of the beveled sleeve 21 is prepared with an inner anti-oxidation coating 22, the outer wall surface is prepared with an outer metal film 23, and the outermost layer is entirely covered with a multi-layer thermal insulation material 24; specifically, the inner anti-oxidation coating 22 can be a composite coating of the SiC / Si-B4C system, which can maintain good anti-oxidation performance below 1500°C, and the outer wall surface is prepared with an outer metal film 23, and the outer metal film 23 is a low-emissivity film;

[0040] Further, the bevel sleeve 21 is provided with an inlet boss 211, a cylindrical section 212, an outlet flange 213, bolt mounting holes 214 and threaded holes 215, the inner profile of the bevel sleeve 21 is a cylindrical structure, the outlet end face is beveled, and there is an angle between the bevel surface and the axis; the bolt mounting holes 214 are uniformly arranged on the inlet boss 211, the threaded holes 215 are uniformly arranged on the outlet flange 213, and the cylindrical section 212 is the outlet flange 213 between the inlet boss 211 and the outlet flange 213 in the middle section of the bevel sleeve 21; further, the bevel sleeve 21 is fixedly connected with the engine 1 through a plurality of bolts; the distance between the inner wall of the cylindrical section 212 of the bevel sleeve 21 and the outer wall of the nozzle 12 is 1-5 mm, the angle θ between the outlet end face of the bevel sleeve 21 and the axis is 40°-80°, the thickness of the outlet flange 213 is ≥5 mm, the hole diameter of the bolt mounting hole 214 is ≥4 mm, and the number of the bolt mounting hole 214 is ≥3; the hole diameter of the threaded hole 215 is ≥4 mm, and the number of the threaded hole 215 is ≥4; and the wall thickness of the cylindrical section 212 of the bevel sleeve 21 is ≥2.5 mm.

[0041] Further, the material of the nozzle 12 is niobium hafnium or niobium tungsten alloy, the material of the bevel sleeve 21 is C / C composite material, C / SiC composite material or GH3128 high-temperature alloy, and the multilayer thermal insulation material 24 is composed of a plurality of layers of double-sided aluminum-coated polyimide film.

[0042] In a more specific embodiment, the maximum working temperature of the nozzle 12 without installing the thermal management sleeve system 2 is 1100℃; the distance between the inner wall of the bevel sleeve cylindrical section 212 and the outer wall of the nozzle 12 is 2 mm, the angle θ between the outlet end face of the bevel sleeve 21 and the axis is 50°, the thickness of the outlet flange 213 is 7 mm, the hole diameter of the bolt mounting hole 214 is 6 mm, and the number of the bolt mounting hole 214 is 3; the hole diameter of the threaded hole 215 is 6 mm, and the number of the threaded hole 215 is 6. The material of the nozzle 12 is niobium tungsten alloy, the material of the bevel sleeve 21 is C / C composite material, the inner layer oxidation-resistant coating 22 can be a composite coating of SiC / Si-B4C system, which can maintain good oxidation resistance below 1500℃, the outer wall surface is provided with an outer layer metal film 23, the outer layer metal film 23 is a low-emissivity (emissivity ≤0.4) film, and the multilayer thermal insulation material 24 is composed of a plurality of layers of double-sided aluminum-coated polyimide film.

[0043] The embodiment of the present application also provides a use method of the reusable aerospace engine and the thermal management sleeve system, which comprises the following steps:

[0044] Step S1: After the engine returns to the ground with the spacecraft, an O-shaped ring made of perfluoroether material is placed in the annular groove at the outlet end face of the nozzle;

[0045] Step S2: A sealing tool is fastened to the sleeve flange by bolts, and the O-shaped ring is pressed to seal the nozzle;

[0046] Step S3: discharging the residual propellant along the nozzle through the tooling center flow channel;

[0047] Step S4: blocking the tooling outlet, and performing the sealing detection on the engine.

[0048] In summary, the embodiment of the present application provides a reusable aerospace engine and a thermal management sleeve system, comprising: an engine 1 and a thermal management sleeve system 2, the engine 1 comprising: an injector 11 and a nozzle 12; the injector 11 is connected with the thermal management sleeve system 2, and the nozzle 12 is located in the cavity of the thermal management sleeve system 2; the engine nozzle 12 is embedded in the bevel sleeve 21, and the composite structure of the thermal management sleeve system 2 is arranged in cooperation, which can not only protect the surrounding components, but also absorb heat to prevent the nozzle 12 from overheating, and solve the contradiction between the compact space of the cabin and the thermal protection demand; the outlet end surface of the bevel sleeve 21 exceeds the outlet of the nozzle 12, and is bevelled at a certain angle to adapt to the cabin surface, which can reduce the aerodynamic interference and thermal influence when entering the atmosphere; an annular groove is processed on the outlet end surface of the nozzle 12 to install an O-ring, and a threaded hole 215 is processed on the flange of the bevel sleeve 21 to connect the sealing tooling, form a closed discharge flow channel, meet the cleaning and detection requirements of the engine on the spacecraft, greatly shorten the maintenance time compared with the traditional engine disassembly and cleaning, and solve the problems of detectable and safety in the detection process of the reusable engine.

[0049] In the description of the present application, it should be understood that the terms "upper", "lower", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer" and the like indicate the orientation or positional relationship based on the orientation or positional relationship shown in the drawings, and are only for the convenience of describing the present application and simplifying the description, and do not indicate or imply that the device or element referred to must have a particular orientation, be constructed and operated in a particular orientation, and therefore cannot be understood as a limitation on the present application.

[0050] The specific embodiments of the present application are described above. It should be understood that the present application is not limited to the above specific embodiments, and those skilled in the art can make various changes or modifications within the scope of the claims, which does not affect the essential content of the present application. The embodiments of the present application and the features in the embodiments can be arbitrarily combined with each other without conflict.

Claims

1. A reusable aerospace engine and thermal management sleeve system, characterized in that: include: Engine (1) and thermal management sleeve system (2), The engine (1) comprises: an injector (11) and a nozzle (12); The injector (11) is connected to the thermal management sleeve system (2), and the nozzle (12) is located in the cavity of the thermal management sleeve system (2).

2. A reusable aerospace engine and thermal management sleeve system according to claim 1, characterized in that: The nozzle (12) is a Laval nozzle, the profile of which contracts first and then expands, and an annular groove (121) is provided on the outlet end face of the nozzle (12); the maximum operating temperature of the nozzle (12) when the thermal management sleeve system (2) is not installed is ≤1200°C.

3. The reusable aerospace engine and thermal management sleeve system according to claim 1, characterized in that: The thermal management sleeve system (2) comprises: a beveled sleeve (21), an inner anti-oxidation coating (22), an outer metal film (23) and a multi-layer thermal insulation material (24); The nozzle (12) is arranged in the thermal management sleeve system (2), and the outlet end surface is shorter than the short side of the beveled sleeve (21); An inner layer of anti-oxidation coating (22) is prepared on the inner wall surface of the beveled sleeve (21), an outer layer of metal film (23) is prepared on the outer wall surface, and the outermost layer is entirely covered with a multi-layer heat insulation material (24).

4. A reusable aerospace engine and thermal management sleeve system according to claim 3, characterized in that: The beveled sleeve (21) is provided with an inlet boss (211), a cylindrical section (212), an outlet flange (213), a bolt mounting hole (214) and a threaded hole (215). The bolt mounting holes (214) are evenly arranged on the inlet boss (211), and the threaded holes (215) are evenly arranged on the outlet flange (213).

5. The reusable aerospace engine and thermal management sleeve system according to claim 3, characterized in that: The inner profile of the beveled sleeve (21) is a cylindrical structure, the outlet end face is beveled, and an angle exists between the cut surface and the axis.

6. The reusable aerospace engine and thermal management sleeve system according to claim 3, characterized in that: The material of the nozzle (12) is niobium hafnium or niobium tungsten alloy, the material of the beveled sleeve (21) is C / C composite material, C / SiC composite material or GH3128 high-temperature alloy, the multi-layer thermal insulation material (24) is composed of several layers of double-sided aluminum-plated polyimide films; the inner anti-oxidation coating (22) is a composite coating of the SiC / Si-B4C system, which can maintain good anti-oxidation performance below 1500°C, and an outer metal film (23) is prepared on the outer wall surface, and the outer metal film (23) is a low-emissivity film, and the film emissivity is ≤0.

4.

7. The reusable aerospace engine and thermal management sleeve system according to claim 4, characterized in that: The beveled sleeve (21) is fixed to the engine (1) via a plurality of bolts; the distance between the inner wall of the cylindrical section (212) of the beveled sleeve (21) and the outer wall of the nozzle (12) is 1 to 5 mm.

8. The reusable aerospace engine and thermal management sleeve system according to claim 4, characterized in that: The thickness of the outlet flange (213) is ≥5mm; the diameter of the bolt mounting hole (214) is ≥4mm, and the number is ≥3; the diameter of the threaded hole (215) is ≥4mm, and the number is ≥4; the wall thickness of the cylindrical section (212) of the beveled sleeve (21) is ≥2.5mm.

9. The reusable aerospace engine and thermal management sleeve system according to claim 5, characterized in that: The angle θ between the outlet end surface of the beveled sleeve (21) and the axis is 40° to 80°.

10. A method for using the reusable aerospace engine and thermal management sleeve system according to any one of claims 1 to 9, comprising: Step S1: After the engine returns to the ground with the spacecraft, an O-ring made of perfluoroether material is placed in the annular groove on the nozzle outlet end face; Step S2: The sealing fixture is fastened to the flange of the sleeve by bolts, and the O-ring is compressed to seal the nozzle; Step S3: discharge the residual propellant along the nozzle through the central flow channel of the tooling; Step S4: Seal the tool outlet and perform a sealing test on the engine.

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