Launch vehicle powerplant

Abstract

The application provides a launch vehicle power device suitable for a launch vehicle, comprising a main shell, a cover body and a coated propellant column, the main shell comprises a combustion shell and a nozzle shell which are arranged in sequence and integrally formed; the front end of the combustion shell is open; the cover body and the coated propellant column are fixedly connected and installed in the cavity of the combustion shell, the cover body is detachably connected with the inner wall at the front end opening of the combustion shell, and the coated propellant column is coaxial with the combustion shell. The launch vehicle power device can prevent the increase of the thrust line deflection amount of the launch vehicle power device caused by the assembly error between the combustion shell and the nozzle shell, effectively reduce the thrust line deflection amount of the launch vehicle power device, effectively correct the thrust line deflection of the launch vehicle power device, improve the flight accuracy of the launch vehicle, effectively prevent the coated propellant column from moving, make the coated propellant column stable on the same axis as the combustion shell, and further reduce the thrust line deflection amount of the launch vehicle power device.

Description

Technical Field

[0001] This application relates to the field of launch vehicle technology, and more specifically, to a launch vehicle propulsion device. Background Technology

[0002] A launch vehicle's propulsion system includes a nozzle and a combustion chamber. Thrust line deviation is generally defined as the angle between the combustion chamber axis and the nozzle axis. Although thrust line deviation is an inherent characteristic of launch vehicle propulsion systems, excessive deviation can severely impact the launch vehicle's flight accuracy.

[0003] Currently, in the production and manufacturing process of launch vehicle propulsion systems, the nozzle and combustion chamber are usually produced separately, and then the nozzle and combustion chamber are connected and fixed together using threads, bolts, and snap rings.

[0004] However, the nozzle and combustion chamber of a launch vehicle's propulsion system are separate structures. From their separate production to final assembly, process fluctuations at each stage of processing can easily cause excessive skewness in the thrust line of the launch vehicle's propulsion system. Summary of the Invention

[0005] This application addresses the shortcomings of existing methods by proposing a launch vehicle propulsion device to solve the technical problem of excessive thrust line skew in launch vehicle propulsion devices in related technologies.

[0006] In a first aspect, embodiments of this application provide a launch vehicle propulsion system, comprising: The main housing includes a combustion housing and a nozzle housing that are arranged sequentially and integrally formed; the combustion housing has an opening at its front end; The cover and the coated propellant are fixedly connected and installed in the cavity of the combustion shell. The cover is detachably connected to the inner wall of the front opening of the combustion shell, and the coated propellant is coaxial with the combustion shell.

[0007] Optionally, the coated drug column includes a first drug segment and a second drug segment, wherein the first drug segment and the second drug segment are arranged along the axial direction of the coated drug column; The cover is provided with a receiving groove, which gradually narrows in the direction away from the second drug segment, and the bottom surface of the receiving groove is a curved surface with an opening facing the covered drug column; At least a portion of the first drug segment is adapted to the shape of the receiving groove and is fixedly connected to the inner wall of the receiving groove.

[0008] Optionally, along the axial direction of the main housing, the outer wall of the cover is threadedly connected to the inner wall of the combustion housing; at least a portion of the first propellant segment is bonded to the inner wall of the receiving groove.

[0009] Optionally, a first heat insulation layer is provided in the cavity of the combustion shell. The cavity of the combustion shell includes a first cavity and a second cavity. The first cavity and the second cavity are arranged sequentially away from the nozzle shell along the axial direction of the combustion shell. The radial dimension of the first cavity is smaller than the radial dimension of the second cavity. The second propellant segment is clearance-fitted with the rear end portion of the combustion shell.

[0010] Optionally, the first insulation layer is connected to a plurality of insulation strips located in the second cavity, and the plurality of insulation strips are arranged in a ring-shaped interval along the outer peripheral wall of the second drug segment. Optionally, the first heat insulation layer is connected to a stop member located in the first cavity. The stop member is annular and located between the coated propellant and the nozzle housing along the axial direction of the combustion shell.

[0011] Optionally, the second drug segment is columnar, and the insulation strip is arc-shaped and adapted to the outer wall of the second drug segment.

[0012] Optionally, the outer wall of the cover is provided with an annular groove, and a sealing ring is provided in the annular groove.

[0013] Optionally, the thickness of the covering sleeve of the coated explosive column is greater than the thickness of the insulation strip.

[0014] Optionally, at least one of the cover and the main housing is made of aluminum alloy, the inner wall of the nozzle housing is provided with a second heat insulation layer, the inner wall of the receiving groove is provided with a third heat insulation layer, the first heat insulation layer and the third heat insulation layer are both made of EPDM rubber, and the second heat insulation layer is made of high silica glass fiber.

[0015] Optionally, the launch vehicle propulsion system further includes a throat liner; The throat liner is disposed on the inner peripheral wall of the nozzle housing; the throat diameter and outlet inner diameter of the throat liner are obtained by machining with the axis of the nozzle housing as a reference after the throat liner is disposed on the inner peripheral wall of the nozzle housing.

[0016] The beneficial technical effects of the technical solutions provided in this application include: In this embodiment, the combustion shell and nozzle shell of the launch vehicle propulsion system are integrally formed, facilitating coaxiality between them. This prevents assembly errors from increasing the thrust line deviation of the launch vehicle propulsion system, effectively reducing the thrust line deviation and improving flight accuracy. Furthermore, the cover is stably connected to the front opening of the combustion shell, and the propellant grain is fixed to the cover, effectively preventing movement and ensuring the propellant grain remains stable on the same axis as the combustion shell, further reducing the thrust line deviation.

[0017] Additional aspects and advantages of this application will be set forth in part in the description which follows, and will become apparent from the description or may be learned by practice of this application. Attached Figure Description

[0018] The above and / or additional aspects and advantages of this application will become apparent and readily understood from the following description of the embodiments taken in conjunction with the accompanying drawings, wherein: Figure 1 This is a cross-sectional structural schematic diagram of a launch vehicle propulsion device provided in an embodiment of this application; Figure 2 A perspective structural schematic diagram of a launch vehicle propulsion device provided in an embodiment of this application; Figure 3 This is a schematic diagram of the structure of the cover provided in an embodiment of this application; Figure 4 This is a schematic diagram of the structure of the coated drug column provided in the embodiments of this application; Figure 5 This is a schematic diagram of the structure after the cap and the coated drug cartridge are connected, as provided in an embodiment of this application. Figure 6 This is a schematic diagram of the structure of the combustion shell and nozzle shell provided in the embodiments of this application; Figure 7 A schematic diagram of the combustion shell and nozzle shell after the installation of the first heat insulation layer and throat liner provided in the embodiments of this application; Figure 8 This is a schematic diagram of the structure of the insulation strip provided in an embodiment of this application; Figure 9 for Figure 1 A schematic diagram of a local structure in the image; Figure 10 This is a schematic diagram of the structure of the laryngeal liner and larynx provided in the embodiments of this application.

[0019] Explanation of reference numerals in the attached figures 100-main shell; 110 - Combustion shell; 111 - First cavity; 112 - Second cavity; 120 - Nozzle housing; 121 - Limiting ring; 122-Laryngeal liner; 123 - Throat; 124 - Connector; 1241 - Third step surface; 1242 - Fourth step surface; 200 - Cover; 210 - Sealing ring; 220 - Annular groove; 230 - Receiving groove; 300 - Coated drug column; 310 - First drug segment; 320 - Second drug segment; 330 - Coating sleeve; 400 - Insulation strip; 410 - First insulation section; 420 - Second insulation section; 500 - First insulation layer; 600-stop component; 700 - Second insulation layer; 710 - First stepped surface; 720 - Second stepped surface; 800 - Third insulation layer. Detailed Implementation

[0020] The embodiments of this application are described below with reference to the accompanying drawings. It should be understood that the embodiments described below with reference to the accompanying drawings are exemplary descriptions for explaining the technical solutions of the embodiments of this application, and do not constitute a limitation on the technical solutions of the embodiments of this application.

[0021] Those skilled in the art will understand that, unless specifically stated otherwise, the terms "described" and "the" as used herein may also include plural forms. It should be further understood that the term "comprising" as used in this application's specification means the presence of described, integers, steps, operations, elements, and / or components, but does not exclude other features, information, data, steps, operations, elements, components, and / or combinations thereof supported by this art. The term "and / or" as used herein refers to at least one of the items defined by the term; for example, "A and / or B" can be implemented as "A," or as "B," or as "A and B."

[0022] To make the objectives, technical solutions, and advantages of this application clearer, the embodiments of this application will be described in further detail below with reference to the accompanying drawings.

[0023] In related technologies, during the production and manufacturing process of launch vehicle propulsion devices, the nozzle and combustion chamber are usually produced separately, and then the nozzle and combustion chamber are connected and fixed by threads, bolts, clasps, etc., which helps to reduce costs and realize the mass production of launch vehicle propulsion devices.

[0024] However, the nozzle and combustion chamber of a launch vehicle's propulsion system are separate structures. From their separate production to final assembly, process fluctuations in each stage of manufacturing can easily cause excessive skewness in the thrust line of the launch vehicle's propulsion system, affecting the launch vehicle's flight accuracy.

[0025] To solve the above technical problems, refer to Figures 1-10 This application provides a launch vehicle propulsion device, including a main shell 100, a cover 200, and a propellant charge 300.

[0026] The main housing 100 includes a combustion housing 110 and a nozzle housing 120 arranged sequentially and integrally formed; the combustion housing 110 has a front end opening; The cover 200 and the coated propellant 300 are fixedly connected and installed in the cavity of the combustion shell 110. The cover 200 is detachably connected to the inner wall of the front opening of the combustion shell 110, and the coated propellant 300 is coaxial with the combustion shell 110.

[0027] Optionally, in this embodiment of the application, the front end of the combustion housing 110 refers to the end away from the nozzle housing 120 along the central axis of the combustion housing 110; the rear end of the combustion housing 110 refers to the end close to the nozzle housing 120 along the central axis of the combustion housing 110.

[0028] The diameter of the combustion shell 110 should not be too small, otherwise the propellant grain 300 installed inside the combustion shell 110 will be too small, failing to meet the launch vehicle's endurance requirements. Conversely, the diameter of the combustion shell 110 should not be too large, otherwise it will increase the volume and weight of the combustion shell 110, affecting the launch vehicle's flight. Optionally, the diameter of the combustion shell 110 is 80mm-200mm. Specifically, the diameter of the combustion shell 110 can be 80mm, 100mm, 120mm, or 200mm, etc. Optionally, the length of the launch vehicle propulsion device in this embodiment is 240mm.

[0029] By integrally molding the combustion shell 110 and the nozzle shell 120, it is easier to make the combustion shell 110 and the nozzle shell 120 coaxial. This prevents an increase in the thrust line deviation of the launch vehicle propulsion system due to assembly errors between the combustion shell 110 and the nozzle shell 120, thereby effectively reducing the thrust line deviation of the launch vehicle propulsion system and effectively correcting the thrust line deviation, thus improving the flight accuracy of the launch vehicle. The cover 200 can be stably connected to the front opening of the combustion shell 110, and the propellant grain 300 is fixed to the cover 200, which can effectively prevent the propellant grain 300 from moving around and keep the propellant grain 300 stable on the same axis as the combustion shell 110, which can further reduce the thrust line deviation of the launch vehicle propulsion system.

[0030] Optionally, such as Figure 1 and Figure 3 As shown, the coated drug column 300 includes a first drug segment 310 and a second drug segment 320, which are arranged along the axial direction of the coated drug column 300. The cap 200 is provided with a receiving groove 230, which gradually narrows away from the second drug segment 320, and the bottom surface of the receiving groove 230 is a curved surface with an opening facing the coated drug column 300. At least a portion of the first drug segment 310 is adapted to the shape of the receiving groove 230 and is fixedly connected to the inner wall of the receiving groove 230.

[0031] It should be noted that the coated propellant column 300 can be composed of a coating sleeve 330 and a propellant column body, with the propellant column body disposed within the coating sleeve 330. The first propellant segment 310 has a part of the coating sleeve 330 and a part of the propellant column body; the second propellant segment 320 has another part of the coating sleeve 330 and another part of the propellant column body.

[0032] In this embodiment, a portion of the first drug segment 310 can be adapted to the shape of the receiving tank 230 and fixedly connected to the inner wall of the receiving tank 230, while the entire first drug segment 310 can also be adapted to the shape of the receiving tank 230 and fixedly connected to the inner wall of the receiving tank 230.

[0033] The gradually contracting shape of the receiving groove 230 allows the first propellant segment 310 to move along the axial direction of the covered propellant grain 300 to connect with the inner wall of the receiving groove 230. At this time, the first propellant segment 310 of the covered propellant grain 300 will generate an interaction force with the inner wall of the receiving groove 230. That is, an interaction force will be generated between the first propellant segment 310 and the bottom surface of the receiving groove 230, and at the same time, an interaction force will be generated between the first propellant segment 310 and the inner wall of the receiving groove 230. This can improve the fit between the first propellant segment 310 and the inner wall of the receiving groove 230, making the first propellant segment 310 fit more tightly with the inner wall of the receiving groove 230, reducing the radial movement range of the covered propellant grain 300, and further reducing the thrust line deviation of the launch vehicle propulsion device.

[0034] The shape of the first propellant segment 310 is adapted to the shape of the receiving groove 230, which enables more of the first propellant segment 310 to connect with the inner wall of the receiving groove 230, improves the connection stability between the first propellant segment 310 and the inner wall of the receiving groove 230, and further improves the fit between the first propellant segment 310 and the inner wall of the receiving groove 230, ensuring that the coated propellant column 300 and the combustion shell 110 are stably on the same axis.

[0035] The bottom surface of the receiving tank 230 is curved, which can guide the first propellant segment 310 to be coaxial with the cover 200 when the first propellant segment 310 is connected to the inner wall of the receiving tank 230, thereby facilitating the coaxiality of the cover 200, the propellant column 300 and the combustion shell 110.

[0036] The propellant grain 300 of the launch vehicle's propulsion system employs end-face combustion. The combustion reaction generates high-temperature, high-pressure gas. The gradually contracting receiving groove 230, when gas leaks from the periphery of the propellant grain 300 into the space between the propellant grain 300 and the inner wall of the receiving groove 230, distributes the pressure of the gas between the first propellant section 310 and the inner wall of the receiving groove 230, reducing stress concentration and effectively resisting gas pressure. This prevents deformation and damage to the cover 200, thus preventing potential problems. The inner surface of the receiving groove 230 can be curved, further dispersing gas pressure and improving the cover 200's resistance to deformation. Furthermore, the curved bottom surface of the receiving groove 230 further enhances the cover 200's resistance to gas pressure.

[0037] Optionally, such as Figure 1 , Figure 3 and Figure 5 As shown, along the axial direction of the main housing 100, the outer wall of the cover 200 is threadedly connected to the inner wall of the combustion housing 110. At least a portion of the first propellant segment 310 is bonded to the inner wall of the receiving groove 230.

[0038] This facilitates the joint installation of the cover 200 and the propellant charge 300 into the cavity of the combustion shell 110. After the cover 200 is connected to the combustion shell 110, the external thread on the outer wall of the cover 200 and the internal thread of the combustion shell 110 will axially limit the cover 200, thereby limiting the propellant charge 300 fixed to the cover 200 in the axial direction of the combustion shell 110, reducing the possibility of axial movement of the propellant charge 300, and reducing the deviation of the thrust line of the launch vehicle propulsion system. Both the external thread on the outer wall of the cover 200 and the internal thread of the combustion shell 110 are M72 threads, ensuring the strength and stability of the threaded connection between the cover 200 and the combustion shell 110.

[0039] Part of the first drug segment 310 can be bonded to the inner wall of the receiving groove 230, and the entire first drug segment 310 can also be bonded to the inner wall of the receiving groove 230, facilitating the connection between the encapsulated drug cartridge 300 and the cap 200. Furthermore, the gradually tapering shape of the receiving groove 230 makes the bond between the first drug segment 310 and the inner wall of the receiving groove 230 more stable, and the curved inner surface of the receiving groove 230 further enhances the stability of the bond between the first drug segment 310 and the inner wall of the receiving groove 230.

[0040] Optionally, such as Figure 1 and Figure 7As shown, a first heat insulation layer 500 is provided inside the cavity of the combustion shell 110. The cavity of the combustion shell 110 includes a first cavity 111 and a second cavity 112. The first cavity 111 and the second cavity 112 are arranged sequentially away from the nozzle shell along the axial direction of the combustion shell 110. The radial dimension of the first cavity 111 is smaller than the radial dimension of the second cavity 112. The second propellant section 320 is clearance-fitted with the rear end portion of the combustion shell 110.

[0041] The first heat insulation layer 500 prevents the combustion shell 110 from being burned during the combustion of the propellant charge 300. The first heat insulation layer 500 can be adhered to the inner wall of the cavity of the combustion shell 110, facilitating its installation within the cavity of the combustion shell 110. By making the second propellant segment 320 clearance-fitted with the rear end portion of the combustion shell 110, the propellant charge 300 is easily inserted into the cavity of the combustion shell 110. After the second propellant segment 320 clearance-fits with the rear end portion of the combustion shell 110, the inner wall of the first cavity 111 can radially constrain the portion of the second propellant segment 320 away from the first propellant segment 310, ensuring that the propellant charge 300 is stably installed within the cavity of the combustion shell 110 and coaxial with the combustion shell 110.

[0042] Along the direction perpendicular to the axial direction of the combustion shell 110, the gaps between the portion of the second propellant segment 320 located within the first cavity 111 and the first heat insulation layer 500, and the gaps between the portion of the second propellant segment 320 located within the second cavity 112 and the first heat insulation layer 500, facilitate the smooth flow of the gas exiting from the periphery of the encapsulated propellant column 300 into the nozzle shell 120. Furthermore, the larger radial dimension of the second cavity 112 results in a larger gap between the portion of the second propellant segment 320 located within the second cavity 112 and the first heat insulation layer 500, and a smaller gap between the portion of the second propellant segment 320 located within the first cavity 111 and the first heat insulation layer 500. This increases the flow velocity of the gas flowing into the gap between the portion of the second propellant segment 320 located within the first cavity 111 and the first heat insulation layer 500, enabling the gas to flow rapidly into the nozzle shell 120.

[0043] Optionally, such as Figure 1 , Figure 7 and Figure 8 As shown, the first insulation layer 500 is connected to a plurality of insulation strips 400 located in the second cavity 112, and the plurality of insulation strips 400 are arranged in a ring-shaped interval along the outer peripheral wall of the second drug section 320.

[0044] The multiple insulation strips 400 can further prevent the combustion shell 110 from burning when the propellant charge 300 is burning, and can also further radially constrain the part of the first propellant segment 310 located in the second cavity 112, reducing the skew of the launch vehicle thrust line.

[0045] There is a gap between two adjacent insulating strips 400, so that there is a gap between the part of the propellant column 300 not covered by the insulating strip 400 and the first insulating layer 500, and there is also a gap between the second propellant section 320 and the insulating strip 400. This can prevent problems before they occur. That is, it can allow the gas flowing out from the periphery of the propellant column 300 to flow in a timely manner towards the direction of the launch vehicle propulsion device nozzle shell 120 through the gap, preventing the local combustion rate from being too high and affecting the stability of combustion. It can also allow the gas to carry away the heat from the surface of the propellant column 300 in a timely manner, preventing the high temperature gas from being trapped and causing the inner wall of the combustion shell 110 to overheat.

[0046] Because the gap between the part of the propellant charge 300 not covered by the insulation strip 400 and the first insulation layer 500 is relatively large, when the gas flows from this gap to the gap between the propellant charge 300 and the first insulation layer 500 at the first cavity 111, the gas flow velocity will increase, which will enable the gas to flow more promptly toward the direction of the launch vehicle power unit nozzle housing 120.

[0047] It should be noted that the covering sleeve of the coated propellant grain enables the propellant of the coated propellant grain to burn within a predetermined combustion zone. This ensures that the combustion process of the coated propellant grain begins at the end face of the second propellant segment near the nozzle housing and continues along the axis of the coated propellant grain until the portion of the first propellant segment is away from the nozzle housing. After the coated propellant grain is installed in the cavity of the combustion housing, before combustion, there is a gap between the second propellant segment 320 and the insulating strip, and a gap between the portion of the second propellant segment within the first cavity and the first insulating layer. In other words, the entire second propellant segment 320 of the coated propellant grain is slightly suspended. The adhesion between the first propellant segment and the inner wall of the receiving groove maintains this slight suspension. After the coated propellant grain burns, the slight suspension of the portion of the second propellant segment within the second cavity is also maintained by the adhesion between the first propellant segment and the inner wall of the receiving groove.

[0048] Optionally, the insulation strip 400 includes a first insulation section 410 and a second insulation section 420, the first insulation section 410 and the second insulation section 420 being arranged along the axial direction of the covered propellant column 300, with a plurality of first insulation sections 410 surrounding the second propellant column 320. A plurality of second insulation sections 420 surrounding the portion of the first propellant column 310 near the second propellant column 320.

[0049] Thus, the second insulation section 420 can axially limit the cover 200, connecting the cover 200 with part of the first drug segment 310. The cover 200 and the second drug segment 320 are spaced a certain distance apart, so that there is a large gap between the part of the first drug segment 310 near the second drug segment 320 and the second insulation section 420 and the cover 200. This allows the gas flowing out from the side wall of the first drug segment 310 to flow sequentially through this large gap to the gap between the second drug segment 320 and the first insulation section 410, as well as the gap between the part of the second drug segment 320 located in the first cavity 111 and the first insulation layer 500. This allows the gas to be doubly accelerated and flow into the nozzle housing 120 more promptly, preventing the gas pressure from damaging the cover 200.

[0050] Optionally, such as Figure 1 As shown, the first heat insulation layer 500 is connected to a stop member 600 located within the first cavity 111. The stop member 600 is annular and is located between the propellant grain 300 and the nozzle housing 120 along the axial direction of the combustion shell. In this way, the annular stop member 600 can axially limit the propellant grain 300, further reducing the thrust line deflection of the launch vehicle propulsion system.

[0051] Optionally, such as Figure 1 As shown, the second propellant section 320 is cylindrical, and the insulating strip 400 is arc-shaped and adapted to the outer wall of the second propellant section 320. This ensures that the spacing between the corresponding parts of the insulating strip 400 and the second propellant section 320 is equal, guaranteeing a more uniform heat distribution between the two when the combustion gases flow between them. This improves the combustion stability of the propellant column 300 and further reduces the thrust line deflection of the launch vehicle's propulsion system.

[0052] The length of the insulation strip 400 extending circumferentially along the second propellant section 320 should not be too short, otherwise the insulation effect will be reduced. Conversely, the length of the insulation strip 400 extending circumferentially along the second propellant section 320 should not be too long, otherwise the gap for gas flow will be reduced. Optionally, the arc length of the curved insulation strip 400 is 5-10 mm. That is, the length of the insulation strip 400 extending circumferentially along the second propellant section 320 is greater than or equal to 5 mm and less than or equal to 10 mm. Specifically, the arc length of the insulation strip 400 can be 5 mm, 8 mm, or 10 mm.

[0053] Optionally, such as Figure 2 and Figure 3 As shown, the outer wall of the cover 200 is provided with an annular groove 220, and a sealing ring 210 is provided inside the annular groove 220. In this way, the sealing performance between the cover 200 and the combustion shell 110 can be improved, preventing the gas generated by the combustion of the propellant 300 from flowing out between the cover 200 and the combustion shell 110, and ensuring that the gas flows to the nozzle shell 120 and is ejected from the nozzle shell 120.

[0054] To ensure effective heat insulation of the insulation strip 400, it can be made thicker without taking up too much space. The space occupied by the insulation strip 400 should not be too large. That is, when the insulation strip 400 is connected to the second cavity 112, the gap between the insulation strip 400 and the second propellant segment 320 should be 0.5mm-1mm. Within this gap range, the radial constraint of the insulation strip 400 on the second propellant segment 320 can be ensured, and it is also convenient for the propellant column 300 to be installed into the cavity of the combustion shell 110. Optionally, such as Figure 1 As shown, the thickness of the covering sleeve 330 covering the propellant grain 300 is greater than the thickness of the insulation strip 400. Thus, the thicker covering sleeve 330 can ensure that the propellant grain body burns in the preset combustion zone, reducing the possibility of being burned through, and further reducing the possibility of the gas passing through the covering sleeve 330, preventing combustion outside the preset zone from causing the thrust line of the launch vehicle propulsion system to deviate.

[0055] Optionally, such as Figure 1 and Figure 3 As shown, at least one of the cover 200 and the main housing 100 is made of aluminum alloy. The inner wall of the nozzle housing 120 is provided with a second heat insulation layer 700, and the inner wall of the receiving groove 230 is provided with a third heat insulation layer 800. The materials of the first heat insulation layer 500 and the third heat insulation layer 800 both include EPDM rubber, and the material of the second heat insulation layer 700 includes high silica glass fiber.

[0056] The aluminum alloy cover 200 and main shell 100 are lightweight, reducing the load on the launch vehicle and improving its flight performance. The first heat insulation layer 500, second heat insulation layer 700, and third heat insulation layer 800 prevent the cover 200, combustion shell 110, and nozzle shell 120 from burning. The EPDM rubber on the inner wall of the receiving groove 230 and inside the combustion shell 110 effectively resists the high-temperature gases generated by the combustion of the propellant charge 300, and because EPDM rubber is not easily chemically reactive, it reduces the corrosion of the cover 200 and combustion shell 110 by the gases. The second heat insulation layer 700, made of high-silica glass fiber reinforced by a molding process, effectively resists the high-speed flowing gases inside the nozzle shell 120.

[0057] Optionally, such as Figure 1 , Figure 2 , Figure 7 and Figure 9As shown, the launch vehicle propulsion system also includes a throat liner 122. The throat liner 122 is disposed on the inner peripheral wall of the nozzle housing 120. The throat diameter and outlet inner diameter of the throat liner 122 are machined with the axis of the nozzle housing 120 as a reference after the throat liner 122 is disposed on the inner peripheral wall of the nozzle housing 120. This ensures that the throat liner 122, the nozzle housing 120, and the combustion shell 110 are coaxial, further effectively reducing the thrust line skew of the launch vehicle propulsion system. Optionally, the outlet inner diameter of the throat liner 122 can be 32 mm.

[0058] The throat liner 122 creates a throat 123 inside the nozzle housing 120. Using the axis of the nozzle housing 120 as a reference axis, the coaxiality of the throat 123 is +0.01, the coaxiality of the exit portion of the throat liner 122 is -0.01, and the distance between the center of the throat 123 and the exit portion of the throat liner 122 is 21 mm. Thus, the angle of deviation of the axis of the throat liner 122 relative to the axis of the nozzle housing 120 can be calculated as arctan(0.02 / 21) = 3.9′, which meets the requirement that the angle of deviation of the thrust line of the launch vehicle propulsion system should not exceed 5′.

[0059] The second heat insulation layer 700 can be bonded to the inner peripheral wall of the nozzle housing 120. Optionally, along the axial direction of the nozzle housing 120, a limiting ring 121 is provided at the end of the nozzle housing 120 away from the combustion housing 110. The portion of the second heat insulation layer 700 near the limiting ring 121 is stepped and has a first stepped surface 710 and a second stepped surface 720. The first stepped surface 710 is connected to the surface of the limiting ring 121 near the combustion housing 110, and the second stepped surface 720 is connected to the inner ring surface of the limiting ring 121. The inner peripheral wall of the nozzle housing 120 is provided with... The connector 124 is stepped near the second insulation layer 700 and has a third stepped surface 1241 and a fourth stepped surface 1242. The third stepped surface 1241 is connected to the end face of the second insulation layer 700 near the combustion shell 110, and the fourth stepped surface 1242 is connected to the inner peripheral wall of the second insulation layer 700. The connector 124 is integrally formed with the throat liner 122, and both the inner peripheral wall of the second insulation layer 700 and the connector 124 are connected to the throat liner 122.

[0060] The limiting ring 121 is integrally formed with the nozzle housing 120. The third stepped surface 1241 of the connecting body 124 and the limiting ring 121 can limit the second heat insulation layer 700 in the axial direction of the nozzle housing 120, so that the second heat insulation layer 700 is more firmly connected to the inner peripheral wall of the nozzle housing 120. At the same time, the second heat insulation layer 700 can also limit the connecting body 124 axially, thereby limiting the throat liner 122 axially, ensuring the heat insulation effect of the second heat insulation layer 700 on the nozzle housing 120, and preventing the distance between the center position of the throat 123 and the outlet of the throat liner 122 from changing, reducing the deviation of the thrust line of the launch vehicle propulsion device.

[0061] The beneficial technical effects of the technical solutions provided in this application include: The combustion shell 110 and the nozzle shell 120 are integrally formed, facilitating their coaxiality. This prevents assembly errors between the combustion shell 110 and the nozzle shell 120 from increasing the thrust line deviation of the launch vehicle's propulsion system, thereby effectively reducing the thrust line deviation and improving the launch vehicle's flight accuracy. The cover 200 is stably connected to the front opening of the combustion shell 110, and the propellant grain 300 is fixed to the cover 200, effectively preventing the propellant grain 300 from shifting and ensuring it remains stable on the same axis as the combustion shell 110, further reducing the thrust line deviation of the launch vehicle's propulsion system.

[0062] The gradually contracting shape of the receiving groove 230 allows the first propellant segment 310 to fit more tightly against the inner wall of the receiving groove 230, reducing the radial movement range of the encapsulated propellant column 300 and further reducing the thrust line deviation of the launch vehicle power plant.

[0063] The gradually contracting shape of the receiving groove 230 can also disperse the pressure of the gas between the first propellant segment 310 and the inner wall of the receiving groove 230, reducing stress concentration, effectively resisting the pressure of the gas, preventing deformation and damage to the cover 200, and preventing problems before they occur. The inner surface of the receiving groove 230 can be curved, which can further disperse the pressure of the gas and improve the deformation resistance of the cover 200. The bottom surface of the curved receiving groove 230 can also further enhance the ability of the cover 200 to resist the pressure of the gas.

[0064] The bottom surface of the receiving trough 230 is curved, which can also guide the first propellant segment 310 to be coaxial with the cover 200 when the first propellant segment 310 is connected to the inner wall of the receiving trough 230, thereby facilitating the coaxiality of the cover 200, the covering propellant column 300 and the combustion shell 110.

[0065] The second propellant segment 320 is clearance-fitted with the first cavity 111, facilitating the insertion of the coated propellant grain 300 into the cavity of the combustion shell 110. The inner wall of the first cavity 111 can radially constrain the portion of the second propellant segment 320 that is away from the first propellant segment 310, ensuring that the coated propellant grain 300 is stably installed in the cavity of the combustion shell 110 and coaxial with the combustion shell 110, thereby reducing the thrust line skew of the launch vehicle.

[0066] The throat diameter and outlet inner diameter of the throat liner 122 are machined based on the axis of the nozzle housing 120, which can ensure that the throat liner 122, the nozzle housing 120 and the combustion housing 110 are coaxial, further effectively reducing the thrust line deviation of the launch vehicle power plant.

[0067] In the description of this application, the terms "center," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," and "outer," etc., indicate directions or positional relationships based on the exemplary directions or positional relationships shown in the accompanying drawings. They are used to facilitate the description or simplification of the embodiments of this application and are not intended to indicate or imply that the device or component referred to must have a specific orientation or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this application.

[0068] The terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of technical features indicated. Therefore, a feature defined as "first" or "second" may explicitly or implicitly include one or more of that feature. In the description of this application, unless otherwise stated, "a plurality of" means two or more.

[0069] In the description of this specification, specific features, structures, materials, or characteristics may be combined in any suitable manner in one or more embodiments or examples.

[0070] The above description is only a partial implementation of this application. It should be noted that for those skilled in the art, other similar implementation methods based on the technical concept of this application, without departing from the technical concept of this application, also fall within the protection scope of the embodiments of this application.

Claims

1. A launch vehicle propulsion system, characterized in that, include: The main housing includes a combustion housing and a nozzle housing arranged sequentially and integrally formed, wherein the inner peripheral wall of the nozzle housing is provided with a throat liner; the front end of the combustion housing is open; The cover and the coated propellant are installed in the cavity of the combustion shell. The cover is detachably connected to the inner wall of the front opening of the combustion shell, and the coated propellant is coaxial with the combustion shell. The coated drug column includes a first drug segment and a second drug segment, which are arranged along the axial direction of the coated drug column; The cover is provided with a receiving groove, which gradually narrows in the direction away from the second drug segment, and the bottom surface of the receiving groove is a curved surface with an opening facing the covered drug column; At least a portion of the first drug segment is adapted to the shape of the receiving groove and is fixedly connected to the inner wall of the receiving groove.

2. The launch vehicle propulsion system according to claim 1, characterized in that, Along the axial direction of the main housing, the outer wall of the cover is threadedly connected to the inner wall of the combustion housing; at least a portion of the first propellant segment is bonded to the inner wall of the receiving groove.

3. The launch vehicle propulsion system according to claim 1, characterized in that, The combustion shell has a first heat insulation layer inside its cavity. The cavity of the combustion shell includes a first cavity and a second cavity. The first cavity and the second cavity are arranged sequentially away from the nozzle shell along the axial direction of the combustion shell. The radial dimension of the first cavity is smaller than the radial dimension of the second cavity. The second propellant segment is clearance-fitted with the rear end portion of the combustion shell.

4. The launch vehicle propulsion system according to claim 3, characterized in that, The first insulation layer is connected to a plurality of insulation strips located in the second cavity, and the plurality of insulation strips are arranged in a ring-shaped interval along the outer peripheral wall of the second drug segment.

5. The launch vehicle propulsion system according to claim 3, characterized in that, The first heat insulation layer is connected to a stop member located in the first cavity. The stop member is annular and is located between the coated propellant and the nozzle housing along the axial direction of the combustion shell.

6. The launch vehicle propulsion system according to claim 4, characterized in that, The second drug segment is columnar, and the insulation strip is arc-shaped and adapted to the outer wall of the second drug segment.

7. The launch vehicle propulsion system according to claim 1, characterized in that, The outer wall of the cover is provided with an annular groove, and a sealing ring is provided in the annular groove.

8. The launch vehicle propulsion system according to claim 4, characterized in that, The thickness of the sheath covering the propellant column is greater than the thickness of the insulation strip.

9. The launch vehicle propulsion system according to claim 3, characterized in that, At least one of the cover and the main housing is made of aluminum alloy. The inner wall of the nozzle housing is provided with a second heat insulation layer, and the inner wall of the receiving groove is provided with a third heat insulation layer. The materials of the first heat insulation layer and the third heat insulation layer both include EPDM rubber, and the material of the second heat insulation layer includes high silica glass fiber.

10. The launch vehicle propulsion system according to claim 1, characterized in that, It also includes the throat liner; The throat liner is disposed on the inner peripheral wall of the nozzle housing; the throat diameter and outlet inner diameter of the throat liner are obtained by machining with the axis of the nozzle housing as a reference after the throat liner is disposed on the inner peripheral wall of the nozzle housing.

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