A composite cooled nozzle and rocket engine

Abstract

The application discloses a composite cooling nozzle and a rocket engine, and relates to the technical field of aerospace engineering.The composite cooling nozzle comprises a first pipe body, which is a single-wall pipe; a second pipe body is fixedly connected with the first pipe body; a first cooling channel and a second cooling channel are arranged in the pipe wall of the second pipe body in an adjacent mode, and the first cooling channel and the second cooling channel are not communicated; wherein, the first cooling channel is arranged on the side of the second pipe body close to the first pipe body, and the first cooling channel is communicated with the inner wall of the first pipe body; a first gas injection assembly is connected with the first cooling channel and is used for cooling the inner wall of the first pipe body; and a second gas injection assembly is connected with the second cooling channel and is used for cooling the second pipe body.The area of the non-cooling area between the two pipe bodies is reduced, and the structural thermal protection reliability is improved.The nozzle has a compact structure, high adaptability and is favorable for controlling the overall weight of the thrust chamber.

Description

Technical Field

[0001] This application belongs to the field of aerospace engineering technology, specifically relating to a composite cooling nozzle and a rocket engine. Background Technology

[0002] The nozzle is a crucial component in the thrust chamber of a liquid rocket engine that generates thrust. During operation, it is subjected to high-temperature combustion gases, requiring reliable thermal protection of its structure.

[0003] In related technologies, there is a technical problem that nozzle cooling has areas that cannot be cooled, and these areas are at risk of thermal protection failure when the engine is under high operating conditions. Summary of the Invention

[0004] This application provides a composite cooling nozzle and a rocket engine, aiming to at least partially solve the technical problem that current nozzle cooling systems have areas that cannot be cooled, and these areas pose a risk of thermal protection failure when the engine is under high operating conditions.

[0005] To solve the above-mentioned technical problems, this application adopts the following technical solution:

[0006] A first aspect of this application provides a composite cooling nozzle, used as a carrier for expanding and accelerating airflow within the thrust chamber of a liquid rocket engine to be ejected outward at high speed. The composite cooling nozzle includes: a first tube body; a second tube body communicating with the first tube body, wherein the second tube body has a first cooling channel and a second cooling channel within its wall; wherein the first cooling channel is located on the side of the second tube body near the first tube body and communicates with the inner wall of the first tube body; a first gas injection assembly connected to the first cooling channel, the first gas injection assembly being used to inject a film cooling medium into the first cooling channel to cool the inner wall of the first tube body; and a second gas injection assembly connected to the second cooling channel, the second gas injection assembly being used to inject and discharge a regenerated cooling medium into the second cooling channel to cool the second tube body.

[0007] In some embodiments, a plurality of acceleration grooves are provided in the first cooling channel. After the gas film cooling working medium enters the acceleration grooves, it forms a continuous supersonic gas film to perform gas film cooling on the inner wall of the first tube.

[0008] In some embodiments, the cross-sectional radius of the acceleration groove gradually narrows from the inlet end to the middle, and the cross-sectional radius of the acceleration groove gradually expands from the middle to the end. After the gas film cooling working fluid enters the acceleration groove, it first contracts and then expands to form a continuous supersonic gas film.

[0009] In some embodiments, the second cooling channel is provided with a plurality of spaced-apart barrier blocks.

[0010] In some embodiments, the barrier block is rectangular.

[0011] In some embodiments, the second pipe body includes an inner wall and an outer wall, the second cooling channel is disposed between the inner wall and the outer wall, and the second pipe body is disposed on the side of the inner wall near the inside of the pipe.

[0012] In some embodiments, the first gas injection assembly includes: a first gas collector fixed to the outer wall, the first gas collector being connected to the first cooling channel through a first air inlet; and a first air inlet flange fixed to the first gas collector, the gas film cooling medium entering the first gas collector through the first air inlet flange and then entering the first cooling channel through the first air inlet.

[0013] In some embodiments, the second air injection assembly includes: a second air intake manifold fixed to the outer wall, the second air intake manifold being connected to one end of the second cooling channel through a second air intake hole; a second air intake flange fixed to the second air intake manifold, the regenerated cooling medium entering the second air intake manifold through the second air intake flange and then entering the second cooling channel through the second air intake hole; a second air outlet manifold fixed to the outer wall, the second air outlet manifold being connected to the other end of the second cooling channel through a second air outlet hole; and a second air outlet flange fixed to the second air outlet manifold, the regenerated cooling medium cooled in the second cooling channel being discharged into the second air outlet manifold and then discharged through the second air outlet flange.

[0014] In some embodiments, the composite cooling nozzle further includes a recovery unit connected to the second outlet flange, the recovery unit being used to guide the cooled regenerated cooling medium back to the thrust chamber to participate in combustion.

[0015] A second aspect of this application provides a rocket engine in which the composite cooling nozzle is mounted in the thrust chamber.

[0016] As can be seen from the above technical solution, this application has at least the following advantages and positive effects:

[0017] This application discloses a composite cooling nozzle, which connects a second tube to a first tube and simultaneously cools both tubes by incorporating a first cooling channel and a second cooling channel within the second tube. This reduces the area of ​​the uncooled region between the two tubes and improves the structural thermal protection reliability of the composite cooling nozzle. The first and second cooling channels are integrally machined within the second tube, ensuring nozzle surface continuity and reducing product molding difficulty. This nozzle has a compact structure, strong adaptability, and facilitates control of the overall weight of the thrust chamber.

[0018] This application discloses a rocket engine that connects a second tube to a first tube by setting a composite cooling nozzle. A first cooling channel and a second cooling channel are provided within the second tube to simultaneously cool both the first and second tubes, reducing the area of ​​the uncooled region between the two tubes and improving the structural thermal protection reliability of the composite cooling nozzle. Both the first and second cooling channels are integrally machined within the second tube, ensuring the continuity of the nozzle profile and reducing product molding difficulty. This nozzle has a compact structure, strong adaptability, and is beneficial for controlling the overall weight of the thrust chamber. Attached Figure Description

[0019] To more clearly illustrate the technical solutions in the embodiments of the present invention, the accompanying drawings used in the description of the embodiments will be briefly introduced below. Obviously, the accompanying drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0020] Figure 1 This is a schematic diagram of the structure of a composite cooling nozzle in an embodiment of this application;

[0021] Figure 2 This is a partial structural schematic diagram of a composite cooling nozzle in an embodiment of this application;

[0022] Figure 3 This is a schematic diagram of the structure of the first cooling channel in the embodiment of this application.

[0023] The reference numerals in the attached drawings are explained as follows: 100, first pipe body; 200, second pipe body; 210, first cooling channel; 211, acceleration groove; 220, second cooling channel; 221, barrier block; 230, inner wall; 240, outer wall; 300, first gas injection assembly; 310, first gas collector; 311, first air inlet; 320, first air inlet flange; 400, second gas injection assembly; 410, second air inlet collector; 411, second air inlet; 420, second air inlet flange; 430, second air outlet collector; 431, second air outlet; 440, second air outlet flange. Detailed Implementation

[0024] To enable those skilled in the art to more clearly understand this application, the technical solutions in the embodiments of this application will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only a part of the embodiments of this application, and not all of the embodiments. Based on the embodiments of this application, all other embodiments obtained by those of ordinary skill in the art without creative effort are within the scope of protection of this application.

[0025] Regenerative cooling is a widely used thermal protection technology for the extension section of liquid rocket engine nozzles. Its principle involves introducing a cryogenic working fluid into the nozzle jacket to cool the inner wall, and then guiding it back to the thrust chamber head for combustion. During this process, the cryogenic working fluid continuously heats up within the nozzle jacket. When the nozzle area ratio is large or the length is excessive, the cooling capacity of the cryogenic working fluid near the nozzle exit decreases due to the rising temperature, potentially preventing it from continuing to cool the inner wall. For large area ratio nozzle extension sections, a film-cooled single-wall nozzle can be added after the regenerative cooling nozzle. This involves accelerating the turbine exhaust gas driving the turbopump or the working fluid that has cooled the regenerative cooling nozzle into a supersonic film in the collector, cooling the downstream single-wall nozzle.

[0026] For large-area nozzle extensions that require both regenerative cooling and film cooling, the regenerative cooling section and the film cooling section are typically machined separately and connected by flanges. However, this connection structure has the following drawbacks: First, to ensure the structural strength of the two nozzle sections and the space constraints of the collector, an uncooled area will form on the inner wall of the area connecting the two nozzle sections. Under high engine operating conditions, this area is at risk of thermal protection failure. Second, the separate machining of the two nozzle sections requires ensuring that the outlet diameter of the inner wall of the regenerative cooling section is not larger than the inlet diameter of the inner wall of the film cooling section after assembly. Otherwise, the high-temperature combustion gas will stagnate at the inverted step formed, causing structural ablation. This is especially critical for large-size nozzle extensions, where the machining accuracy and welding deformation control requirements are extremely high. Third, flange connections typically increase the structural weight significantly, making their application uneconomical in upper-stage rocket engines.

[0027] This application provides a composite cooling nozzle and a rocket engine, aiming to at least partially solve the technical problem that current nozzle cooling systems have areas that cannot be cooled, and these areas pose a risk of thermal protection failure when the engine is under high operating conditions.

[0028] Please see Figure 1 and Figure 2 .

[0029] Figure 1 This is a schematic diagram of the structure of a composite cooling nozzle according to an embodiment of this application. Figure 2 This is a partial structural schematic diagram of a composite cooling nozzle according to an embodiment of this application; as shown... Figure 1 and Figure 2As shown, this composite cooling nozzle serves as a carrier for expanding and accelerating airflow within the thrust chamber of a liquid rocket engine to propel it outward at high speed. The composite cooling nozzle includes: a first tube body 100, which is a single-walled tube; and a second tube body 200 connected to the first tube body 100. The second tube body 200 has adjacent first cooling channels 210 and second cooling channels 220 within its wall. The first cooling channels 210 and second cooling channels 220 are not connected. The first cooling channel 210 is located near the first cooling channel 220 within the second tube body. On one side of the tube body 100, the first cooling channel 210 is connected to the inner wall of the first tube body 100; the first gas injection assembly 300 is connected to the first cooling channel 210, and the first gas injection assembly 300 is used to inject a gas film cooling medium into the first cooling channel 210 to cool the inner wall 230 of the first tube body 100; the second gas injection assembly 400 is connected to the second cooling channel 220, and the second gas injection assembly 400 is used to inject a regenerated cooling medium into the second cooling channel 220 and discharge it to cool the second tube body. By connecting the second tube 200 to the first tube 100, and by setting a first cooling channel 210 and a second cooling channel 220 within the second tube 200 to simultaneously cool the first tube 100 and the second tube 200, the area of ​​the uncooled region between the two tubes is reduced, thereby improving the structural thermal protection reliability of the composite cooling nozzle. The first cooling channel 210 and the second cooling channel 220 can be integrally machined within the second tube 200, ensuring the continuity of the nozzle profile and reducing the difficulty of product molding. This nozzle has a compact structure, strong adaptability, and is beneficial for controlling the overall weight of the thrust chamber.

[0030] like Figure 1 As shown, in some embodiments, the first tube 100 and the second tube 200 can be coaxially arranged. The radius of the first tube 100 gradually decreases from the first end to the second end, and the radius of the second tube 200 gradually decreases from the first end to the second end. The second end of the first tube 100 is welded and fixed to the first end of the second tube 200. The structure is compact, highly adaptable, and conducive to controlling the overall weight of the thrust chamber.

[0031] In some embodiments, the air outlet of the first cooling channel 210 is flush with the welding point of the first tube 100 and the second tube 200, and the air film cooling medium is directly blown onto the inner wall of the first tube.

[0032] Please see Figure 3 , Figure 3 This is a schematic diagram of the structure of the first cooling channel 210 in the embodiment of this application;

[0033] like Figure 3As shown, in some embodiments, the first cooling channel 210 is provided with a plurality of acceleration grooves 211. After the gas film cooling working medium enters the acceleration grooves 211, it forms a continuous supersonic gas film to perform gas film cooling on the inner wall 230 of the first tube 100.

[0034] like Figure 3 As shown, in some embodiments, the cross-section of the acceleration groove 211 can be a Venturi tube, that is, its radius gradually narrows from the air inlet end to the middle, and the cross-sectional radius of the acceleration groove 211 gradually expands from the middle to the end. After the gas film cooling working medium enters the acceleration groove 211, it first contracts and then expands to form a continuous supersonic gas film. The continuous supersonic gas film passes through the inner wall 230 of the first tube 100 to perform gas film cooling on the first tube 100.

[0035] Please see Figure 2 In some embodiments, a plurality of spaced-apart baffles 221 are provided within the second cooling channel 220. By providing baffles 221, the travel distance of the regenerated cooling medium is increased, thereby improving the cooling effect.

[0036] In some embodiments, the barrier block 221 is rectangular.

[0037] In other embodiments, the blocking block 221 may also be other shapes, such as trapezoidal, arc-shaped, triangular, or other shapes.

[0038] Please see Figure 2 In some embodiments, the second tube body 200 includes an inner wall 230 and an outer wall 240, and the second cooling channel 220 is disposed around the inner wall 230 and the outer wall 240. The second tube body 200 is disposed on the side of the inner wall 230 closer to the inside of the pipe. By simultaneously machining the first cooling channel 210 and the second cooling channel 220 on the inner wall 230, the area of ​​the uncooled region between the two cooling structures is reduced, and the reliability of the structural thermal protection is improved.

[0039] Please see Figure 1 In some embodiments, the first gas injection assembly 300 includes: a first gas collector 310, fixed to the outer wall 240, the first gas collector 310 being connected to the first cooling channel 210 through a first air inlet 311; and a first air inlet flange 320, fixed to the first gas collector 310, the gas film cooling medium entering the first gas collector 310 through the first air inlet flange 320 and then entering the first cooling channel 210 through the first air inlet 311.

[0040] In some embodiments, the first air inlet 311 is a circular hole. In other embodiments, the first air inlet 311 may also be elliptical, square, or other shapes, and there is no limitation on this.

[0041] In some embodiments, the second air injection assembly 400 includes: a second air intake manifold 410 fixed to the outer wall 240, the second air intake manifold 410 being connected to one end of the second cooling channel 220 through a second air intake hole 411; a second air intake flange 420 fixed to the second air intake manifold 410, the regenerated cooling medium entering the second air intake manifold 410 through the second air intake flange 420 and then entering the second cooling channel 220 through the second air intake hole 411; a second air outlet manifold 430 fixed to the outer wall 240, the second air outlet manifold 430 being connected to the other end of the second cooling channel 220 through a second air outlet hole 431; and a second air outlet flange 440 fixed to the second air outlet manifold 430, the regenerated cooling medium cooled in the second cooling channel 220 being discharged into the second air outlet manifold 430 and then discharged through the second air outlet flange 440.

[0042] In some embodiments, both the second air inlet 411 and the second air outlet 431 can be rectangular holes. In other embodiments, the second air inlet 411 can also be elliptical, square, or other shapes, which are not limited here.

[0043] In some embodiments, multiple sets of the first air intake assembly and the second air intake assembly may be provided and are circumferentially fixed to the outer wall 240.

[0044] In some embodiments, the composite cooling nozzle further includes a recovery unit connected to the second outlet flange 440, which is used to guide the cooled regenerated cooling medium back to the thrust chamber to participate in combustion.

[0045] A second aspect of this application provides a rocket engine in which the composite cooling nozzle is installed within the thrust chamber. By using the composite cooling nozzle to connect the second tube 200 to the first tube 100, and by providing a first cooling channel 210 and a second cooling channel 220 within the second tube 200 to simultaneously cool both the first tube 100 and the second tube 200, the area of ​​the uncooled region between the two tubes is reduced, improving the structural thermal protection reliability of the composite cooling nozzle. The first cooling channel 210 and the second cooling channel 220 are integrally machined within the second tube 200, ensuring the continuity of the nozzle profile and reducing the difficulty of product molding. This nozzle has a compact structure, strong adaptability, and is beneficial for controlling the overall weight of the thrust chamber.

[0046] As can be seen from the above technical solution, this application has at least the following advantages and positive effects:

[0047] This application discloses a composite cooling nozzle, which connects a second tube to a first tube and simultaneously cools both tubes by incorporating a first cooling channel and a second cooling channel within the second tube. This reduces the area of ​​the uncooled region between the two tubes and improves the structural thermal protection reliability of the composite cooling nozzle. The first and second cooling channels are integrally machined within the second tube, ensuring nozzle surface continuity and reducing product molding difficulty. This nozzle has a compact structure, strong adaptability, and facilitates control of the overall weight of the thrust chamber.

[0048] This application discloses a rocket engine that connects a second tube to a first tube by setting a composite cooling nozzle. A first cooling channel and a second cooling channel are provided within the second tube to simultaneously cool both the first and second tubes, reducing the area of ​​the uncooled region between the two tubes and improving the structural thermal protection reliability of the composite cooling nozzle. Both the first and second cooling channels are integrally machined within the second tube, ensuring the continuity of the nozzle profile and reducing product molding difficulty. This nozzle has a compact structure, strong adaptability, and is beneficial for controlling the overall weight of the thrust chamber.

[0049] In this application, unless otherwise expressly specified and limited, "above" or "below" the second feature can include direct contact between the first and second features, or contact between the first and second features through another feature between them. Furthermore, "above," "over," and "on top" of the second feature includes the first feature being directly above or diagonally above the second feature, or simply indicates that the first feature is at a higher horizontal level than the second feature. "Below," "below," and "under" the second feature includes the first feature being directly below or diagonally below the second feature, or simply indicates that the first feature is at a lower horizontal level than the second feature.

[0050] In the description of this application, it should be understood that the terms "center", "longitudinal", "lateral", "length", "width", "thickness", "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", "clockwise", and "counterclockwise" indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are used only for the convenience of describing this application and simplifying the description, and do not indicate or imply that the device or element 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.

[0051] In this application, unless otherwise expressly specified and limited, the terms "connection," "fixed," etc., should be interpreted broadly. For example, "fixed" can mean a fixed connection, a detachable connection, or an integral part; it can mean a mechanical connection or an electrical connection; it can mean a direct connection or an indirect connection through an intermediate medium; it can mean the internal communication of two components or the interaction between two components. Those skilled in the art can understand the specific meaning of the above terms in this application according to the specific circumstances.

[0052] Furthermore, the use of terms such as "first" and "second" in this application is for descriptive purposes only and should not be construed as indicating or implying their relative importance or implicitly specifying the number of technical features indicated. Therefore, features defined with "first" or "second" may explicitly or implicitly include one or more features. In the description of this application, "multiple" means two or more, unless otherwise explicitly specified.

[0053] Although embodiments of this application have been shown and described, those skilled in the art will understand that various changes, modifications, substitutions and variations can be made to these embodiments without departing from the principles and spirit of this application, the scope of which is defined by the claims and their equivalents.

Claims

1. A composite cooling nozzle, used as a carrier in the thrust chamber of a liquid rocket engine to expand and accelerate airflow therein for high-speed outward ejection, characterized in that, The composite cooling nozzle includes: first tube body; The second tube body is connected to the first tube body. The second tube body has a first cooling channel and a second cooling channel inside its tube wall. The first cooling channel is located on the side of the second tube body close to the first tube body and is connected to the inner wall of the first tube body. The first gas injection component is connected to the first cooling channel. The first gas injection component is used to inject a gas film cooling medium into the first cooling channel to cool the inner wall of the first tube. The second gas injection assembly is connected to the second cooling channel. The second gas injection assembly is used to inject regenerated cooling medium into the second cooling channel and discharge it to cool the second pipe body. The second pipe body includes an inner wall and an outer wall, the second cooling channel is disposed between the inner wall and the outer wall, and the second pipe body is disposed on the side of the inner wall near the inside of the pipe; The second gas injection assembly includes: The second air intake manifold is fixed to the outer wall and is connected to one end of the second cooling channel through the second air intake hole. The second air inlet flange is fixed to the second air inlet collector. The regenerated cooling medium enters the second air inlet collector through the second air inlet flange and then enters the second cooling channel through the second air inlet hole. The second air outlet collector is fixed to the outer wall, and the second air outlet collector is connected to the other end of the second cooling channel through the second air outlet hole; The second exhaust flange is fixed to the second exhaust collector. The regenerated cooling medium cooled in the second cooling channel is discharged into the second exhaust collector and then discharged through the second exhaust flange.

2. The composite cooling nozzle according to claim 1, characterized in that, The first cooling channel is provided with multiple acceleration grooves. After the gas film cooling working medium enters the acceleration grooves, it forms a continuous supersonic gas film to cool the inner wall of the first tube.

3. The composite cooling nozzle according to claim 2, characterized in that, The cross-sectional radius of the acceleration groove gradually narrows from the inlet end to the middle, and gradually expands from the middle to the end. After the gas film cooling working medium enters the acceleration groove, it first contracts and then expands to form a continuous supersonic gas film.

4. The composite cooling nozzle according to claim 1, characterized in that, The second cooling channel is equipped with multiple spaced-apart barrier blocks.

5. The composite cooling nozzle according to claim 4, characterized in that, The barrier block is rectangular.

6. The composite cooling nozzle according to claim 1, characterized in that, The first gas injection assembly includes: A first gas collector is fixed to the outer wall, and the first gas collector is connected to the first cooling channel through a first air inlet. The first inlet flange is fixed to the first gas collector. The gas film cooling medium enters the first gas collector through the first inlet flange and then enters the first cooling channel through the first inlet hole.

7. The composite cooling nozzle according to claim 1, characterized in that, The composite cooling nozzle also includes a recovery unit, which is connected to the second outlet flange. The recovery unit is used to guide the cooled regenerated working fluid back to the thrust chamber to participate in combustion.

8. A rocket engine, characterized in that, The rocket engine's thrust chamber is equipped with a composite cooling nozzle as described in any one of claims 1 to 7.

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