An active cooling system for an air-foil of a hypersonic vehicle

Abstract

The application relates to the field of aviation technology, and particularly discloses an active cooling system for an air rudder of a hypersonic aircraft, which comprises a leading edge skin and a trailing edge skin, a rudder shaft is arranged between the leading edge skin and the trailing edge skin and connected with the two, at least two leading edge skin channels are arranged in the leading edge skin, and all the leading edge skin channels are communicated in a leading edge area; at least two trailing edge skin channels are arranged in the trailing edge skin, and all the trailing edge skin channels are communicated in a trailing edge area; the rudder shaft comprises a working medium input channel and a working medium output channel, the working medium input channel is communicated with at least one leading edge skin channel, the working medium output channel is communicated with at least one trailing edge skin channel, and the leading edge skin channels not communicated with the working medium input channel and the trailing edge skin channels not communicated with the working medium output channel are communicated. The active thermal protection technology can be used to cool the air rudder.

Description

Technical Field

[0001] This invention relates to the field of aviation technology, and in particular to an active cooling system for the air rudder of a hypersonic vehicle. Background Technology

[0002] Aircraft play a vital role in modern society, with widespread applications in various aspects of human life. Aircraft typically possess aerodynamic control surfaces, which are crucial actuators for controlling the aircraft's attitude. By altering airflow, they generate lateral control forces, thereby achieving stability and maneuverability. The all-moving control surface is a type of aerodynamic control surface, referring to a component that combines the horizontal stabilizer and elevator into one unit. It is integrated with the fuselage via a control shaft, allowing the pilot to control the entire horizontal stabilizer's deflection. This overall deflection enables precise control of flight attitude and direction. The all-moving control surface is an advanced control component used in spacecraft and high-performance aircraft.

[0003] All-moving control surfaces experience extremely high thermodynamic and aerodynamic loads during high-speed flight, necessitating thermal protection. Currently, the industry typically employs the following technologies for thermal protection of aerodynamic control surfaces: The first method is passive thermal protection, which relies on the natural properties of materials to protect the aircraft, including heat absorption, radiative dissipation, and heat transfer isolation. However, the protective capability of passive thermal protection is limited by the thermophysical properties of the materials, especially under high heat flux densities, where the materials' temperature resistance becomes insufficient. Therefore, it is generally suitable for applications with low to medium heat loads. The second method is semi-passive thermal protection, a method that combines the characteristics of passive and active cooling technologies, utilizing the thermal insulation and anti-thermal properties of materials. The first approach is active thermal protection. Active thermal protection uses airflow or a working fluid to remove some heat. However, in high-heat-fluid environments, it faces the problem of working fluid material depletion. Furthermore, the expansion or phase change of the working fluid may alter the aircraft's shape, thus affecting aerodynamic performance and handling stability. Therefore, it is generally suitable for applications with medium heat loads. The third approach is active thermal protection technology. Active thermal protection technology is a technology that uses fluid convection to absorb heat and cool high-temperature components of an aircraft. It has good thermal protection effects and is suitable for long-endurance and reusability. However, active thermal protection technology is currently mainly used for cooling aircraft engines and is less commonly used in all-moving rudders. Applying active thermal protection technology to all-moving rudders has become a key technology for the design and development of high-speed aircraft. Summary of the Invention

[0004] The present invention aims to solve the technical problems existing in the prior art and provide an active cooling system for the air rudder of a hypersonic vehicle that uses active thermal protection technology to cool the air rudder.

[0005] To solve the above-mentioned technical problems, the present invention provides the following technical solutions:

[0006] The present invention discloses an active cooling system for the air rudder of a hypersonic vehicle, comprising a leading-edge skin and a trailing-edge skin, with a rudder shaft connected to both the leading-edge and trailing-edge skins. The leading-edge skin has at least two leading-edge skin channels, all of which are interconnected in the leading-edge region. The trailing-edge skin also has at least two trailing-edge skin channels, all of which are interconnected in the trailing-edge region. The rudder shaft includes a propellant input channel and a propellant output channel. The propellant input channel is connected to at least one leading-edge skin channel, and the propellant output channel is connected to at least one trailing-edge skin channel. Leading-edge skin channels not connected to the propellant input channel and trailing-edge skin channels not connected to the propellant output channel are also interconnected.

[0007] The present invention discloses an active cooling system for the air rudder of a hypersonic vehicle. During system operation, the working propellant enters the leading-edge skin channel connected to the rudder shaft through the working propellant input channel, flows along the channel to the leading-edge region of the air rudder, then flows out through the leading-edge skin channel not connected to the working propellant input channel to the trailing-edge skin channel not connected to the working propellant output channel, and flows along the channel to the trailing-edge region of the air rudder. Finally, it flows out through the trailing-edge skin channel connected to the working propellant output channel to the working propellant output channel, and is then discharged from the working propellant output channel of the rudder shaft, completing the cooling cycle for the air rudder. During the flow process, the working propellant can exchange heat and cool the leading-edge region, leading-edge skin, trailing-edge skin, and trailing-edge region of the air rudder, effectively removing the heat generated in the air rudder and effectively ensuring the thermal protection performance of the air rudder.

[0008] Furthermore, the leading edge skin is provided with three leading edge skin channels, namely a first leading edge skin channel, a second leading edge skin channel, and a leading edge core layer channel. The leading edge core layer channel is located between the first leading edge skin channel and the second leading edge skin channel, and the leading edge core layer channel is connected to the working fluid input channel. The trailing edge skin is provided with three trailing edge skin channels, namely a first trailing edge skin channel, a second trailing edge skin channel, and a trailing edge core layer channel. The trailing edge core layer channel is located between the first trailing edge skin channel and the second trailing edge skin channel, and the trailing edge core layer channel is connected to the working fluid output channel. The first leading edge skin channel is connected to the first trailing edge skin channel, and the second leading edge skin channel is connected to the second trailing edge skin channel.

[0009] Furthermore, the leading edge skin channel is provided with multiple first skin pipes at intervals, and all the first skin pipes are connected in the leading edge region; the trailing edge skin channel is provided with multiple second skin pipes at intervals, and all the second skin pipes are connected in the trailing edge region.

[0010] Furthermore, it also includes a working fluid input and collection assembly; the working fluid input and collection assembly includes a working fluid input and collection chamber, which is connected to the working fluid input channel on the rudder shaft and at least one leading edge skin channel respectively;

[0011] Alternatively, it may also include a working fluid output and collection assembly; the working fluid output and collection assembly includes a working fluid output and collection chamber, which is connected to a working fluid output channel on the rudder shaft and at least one trailing edge skin channel.

[0012] Furthermore, the working fluid input collection assembly also includes an inlet manifold assembly, which includes an inlet manifold connector. Multiple inlet pipes are spaced apart within the inlet manifold connector, and the inlet manifold connector is connected to at least one leading edge skin channel through the inlet pipes. An inlet connection nozzle is also provided on the inlet manifold connector, and the inlet connection nozzle is connected to the working fluid input collection chamber.

[0013] Alternatively, the working fluid output collection assembly may further include an outlet manifold assembly, which includes an outlet manifold connector and multiple outlet pipes spaced apart within the outlet manifold connector. The outlet manifold connector is connected to at least one trailing edge skin channel through the outlet pipes. An outlet connection nozzle is also provided on the outlet manifold connector, which is connected to the working fluid output collection chamber.

[0014] Alternatively, the rudder shaft includes an inner tube and an outer tube, the outer tube being disposed around the inner tube and spaced apart from it, and the two being connected by a second dot matrix structure; the inner through hole of the inner tube forms the working fluid input channel; the working fluid output channel is formed between the outer tube and the inner tube, and the second dot matrix structure is disposed in the working fluid output channel.

[0015] Furthermore, the length of the outer tube is shorter than the length of the inner tube, one end of the inner tube extends to the outside of the outer tube, and the other end of the inner tube is flush with the corresponding end of the outer tube; the end of the inner tube extending to the outside of the outer tube passes through the working fluid output collection chamber and is connected to the working fluid input collection chamber, and the end of the outer tube through which the inner tube extends is connected to the working fluid output collection chamber.

[0016] Furthermore, multiple reinforcing ribs are spaced apart on the inner wall of the leading edge skin located in the leading edge region. The multiple reinforcing ribs are arranged along the shape of the leading edge region, and the size of the multiple reinforcing ribs gradually decreases from the middle to both sides.

[0017] Furthermore, the working propellant is injected at high speed from the working propellant input channel of the rudder shaft into the leading edge core channel connected to it, and then impacts the reinforcing rib located in the middle at high speed from the leading edge core channel.

[0018] Furthermore, a first lattice sandwich structure is provided between adjacent leading edge skin channels and between adjacent trailing edge skin channels.

[0019] Furthermore, the leading edge skin and the trailing edge skin are connected by integral molding. The integrally molded leading edge skin and the trailing edge skin form a notch at a position opposite to the rudder shaft, and a sealing plate is provided at the notch.

[0020] Furthermore, it also includes a fuel tank, valves, and a high-pressure pump, which are connected in series. The entire assembly formed by the fuel tank, valves, and high-pressure pump is connected to the working propellant input channel of the rudder shaft through a pipeline, and the working propellant output channel of the rudder shaft is connected to the fuel pipeline of the hypersonic vehicle's engine through a pipeline. Attached Figure Description

[0021] The above and other objects, features, and advantages of the invention will become clearer through a more detailed description of the preferred embodiments illustrated in the accompanying drawings. The same reference numerals denote the same parts throughout the drawings, and the drawings are not intentionally drawn to scale with actual dimensions; the focus is on illustrating the gist of the invention.

[0022] Figure 1 This is a schematic diagram of the structure of the present invention.

[0023] Figure 2 for Figure 1 The front view (showing the direction of working fluid flow).

[0024] Figure 3 This is a partial schematic diagram of the first skin piping structure.

[0025] Figure 4 This is a partial schematic diagram of the second skin piping structure.

[0026] Figure 5 A schematic diagram of the working fluid input and output liquid collection components.

[0027] Figure 6 and Figure 7 This is a schematic diagram of the rudder shaft structure in different directions.

[0028] Figure 8 A schematic diagram showing the installation of reinforcing ribs in the leading edge region.

[0029] Figure 9 This is a simplified schematic diagram of the connection structure of the present invention on a hypersonic vehicle.

[0030] The components include: leading edge skin 1; trailing edge skin 2; rudder shaft 3; leading edge region 4; trailing edge region 5; working fluid input channel 6; working fluid output channel 7; all-moving rudder 8; leading edge core layer channel 9; first leading edge skin channel 10; second leading edge skin channel 11; first trailing edge skin channel 12; second trailing edge skin channel 13; trailing edge core layer channel 14; first skin pipeline 15; second skin pipeline 16; working fluid input collection chamber 17; working fluid output collection chamber 18; inlet manifold connector 19; inlet connector 20; outlet manifold connector 21; outlet connector 22; first lattice sandwich structure 23; reinforcing rib 24; fuel tank 25; valve 26; and high-pressure pump 27. Detailed Implementation

[0031] To facilitate understanding of the present invention, a more comprehensive description will be given below with reference to the accompanying drawings.

[0032] It should be noted that when a component is considered to be "connected" to another component, it can be directly connected to and integrated with the other component, or there may be an intervening component present. The terms "mounted," "one end," "the other end," and similar expressions used in this document are for illustrative purposes only.

[0033] Unless otherwise defined, all technical and scientific terms used herein have the same meaning as commonly understood by one of ordinary skill in the art. The terminology used in this specification is for the purpose of describing particular embodiments only and is not intended to be limiting of the invention. The term "and / or" as used herein includes any and all combinations of one or more of the associated listed items.

[0034] In this application, unless otherwise expressly specified and limited, "above" or "below" the second feature can mean that the first feature is in direct contact with the second feature, or that the first feature is in indirect contact with the second feature through an intermediate medium. Furthermore, "above," "on top of," and "over" the second feature can mean that the first feature is directly above or diagonally above the second feature, or simply that the first feature is at a higher horizontal level than the second feature. "Below," "below," and "under" the second feature can mean that the first feature is directly below or diagonally below the second feature, or simply that the first feature is at a lower horizontal level than the second feature.

[0035] In the description of this specification, the references to terms such as "preferred embodiment," "another embodiment," "other embodiment," or "specific example," etc., indicate that a specific feature, structure, material, or characteristic described in connection with that embodiment or example is included in at least one embodiment or example of this application. In this specification, the illustrative expressions of the above terms do not necessarily refer to the same embodiment or example. Furthermore, the specific features, structures, materials, or characteristics described may be combined in a suitable manner in any one or more embodiments or examples. Moreover, without contradiction, those skilled in the art can combine and integrate the different embodiments or examples described in this specification, as well as the features of different embodiments or examples.

[0036] This invention provides a specific implementation of an active cooling system for the air rudders of a hypersonic vehicle, see [link to relevant documentation]. Figure 1The system includes a leading edge skin 1 and a trailing edge skin 2. A rudder shaft 3 is provided between the leading edge skin 1 and the trailing edge skin 2 and connected to both. The leading edge skin 1 has at least two leading edge skin channels, all of which are connected in the leading edge region 4. Specifically, the term "leading edge" is a prior art term referring to the intersection of the upper and lower surfaces of the leading edge skin 1. There are at least two leading edge skin channels. When there are only two channels, one channel is located on the upper surface of the leading edge skin 1, and the other is located on the lower surface. The lengths of both channels extend to the upper surface of the leading edge skin 1. The intersection of the upper and lower surfaces, i.e., the leading edge region 4, is connected within this region; the trailing edge skin 2 has at least two trailing edge skin channels, all of which are connected in the trailing edge region 5. Specifically, the "tailing edge" is also a prior art term, referring to the intersection of the upper and lower surfaces of the trailing edge skin 2. There are at least two trailing edge skin channels. When there are only two trailing edge skin channels, one is located on the upper surface of the trailing edge skin 2, and the other is located on the lower surface. The length of both trailing edge skin channels extends to the intersection of the upper and lower surfaces of the trailing edge skin 2, i.e. There are 5 trailing edge regions, and they are connected within this region; the rudder shaft 3 includes a working propellant input channel 6 and a working propellant output channel 7. The working propellant input channel 6 is connected to at least one leading edge skin channel. When there are only two leading edge skin channels, the working propellant input channel 6 is connected to only one of the leading edge skin channels. The working propellant output channel 7 is connected to at least one trailing edge skin channel. When there are only two trailing edge skin channels, the working propellant output channel 7 is connected to only one of the trailing edge skin channels. The leading edge skin channels not connected to the working propellant input channel 6 and the trailing edge skin channels not connected to the working propellant output channel 7 are connected. The working propellant can be supercritical aircraft fuel (e.g., Fluids such as aerospace kerosene or carbon dioxide are selected according to the actual situation. In this embodiment, aerospace kerosene is preferred as the working medium. Aerospace kerosene can operate efficiently in a stable state, ensuring the safety and stability of the aircraft. It has a high specific heat capacity and can absorb more heat in a small volume, thereby effectively preventing overheating problems. Furthermore, the aerospace kerosene used by the aircraft itself can be used as a working medium to cool the air rudder. That is, the aircraft's own fuel has the dual role of fuel and cooling, eliminating the need for an additional cooling working medium, reducing the overall weight of the aircraft, and the high-temperature fuel discharged by the air rudder has higher combustion efficiency.

[0037] When the active cooling system of the air rudder is working, especially during hypersonic flight, the leading edge region 4 of the air rudder (such as the all-moving rudder 8) is subjected to severe aerodynamic heating. This region often has the highest temperature, so effective cooling of the leading edge region 4 is necessary to ensure the thermal protection performance of the air rudder. In this invention, the working propellant enters the leading edge skin channel connected to the working propellant input channel 6 of the rudder shaft 3, and flows along the channel to the leading edge region 4 of the air rudder. Then, it flows out from the leading edge skin channel not connected to the working propellant input channel 6 to the trailing edge skin channel not connected to the working propellant output channel 7, and flows along the channel to the trailing edge region 5 of the air rudder. Finally, it flows out from the trailing edge skin channel connected to the working propellant output channel 7 to the working propellant output channel 7, and is then discharged from the working propellant output channel 7, completing the cooling cycle of the air rudder. During its flow, the working fluid can exchange heat and cool the leading edge skin 1, the leading edge region 4 of the air rudder, the trailing edge skin 2, and the trailing edge region 5 of the air rudder, effectively removing the heat generated in the air rudder and ensuring its thermal protection performance. Furthermore, gaps typically exist at the connection points between the air rudder and the aircraft body, resulting in heat flow at these points. Since the rudder shaft 3 has both a working fluid input and output channel, and connects the air rudder to the aircraft body via the rudder shaft 3, the working fluid flowing within the rudder shaft 3 can exchange heat with the high-temperature walls at the connection gaps, removing aerodynamic heating heat and further ensuring the air rudder's thermal protection performance.

[0038] In the preferred embodiment, see Figure 1The leading edge skin 1 has three leading edge skin channels: a first leading edge skin channel 10, a second leading edge skin channel 11, and a leading edge core layer channel 9. The leading edge core layer channel 9 is located between the first leading edge skin channel 10 and the second leading edge skin channel 11, and it is connected to the working fluid input channel 6. Specifically, the first leading edge skin channel 10 is located in the upper surface of the leading edge skin 1, and the second leading edge skin channel 11 is located in the lower surface of the leading edge skin 1. A sandwich leading edge layer is provided in the leading edge skin 1, and the leading edge core layer channel 9 is located in the sandwich leading edge layer. The sandwich leading edge layer can be connected to the leading edge skin 1 through an integral molding technology. This method has good sealing performance and strength. However, the two can also be formed separately and then connected by fasteners or welding. In the case of separate configuration, attention needs to be paid to the sealing performance and strength of the connection between the two. The trailing edge skin 2 has three trailing edge skins. The channels are a first trailing edge skin channel 12, a second trailing edge skin channel 13, and a trailing edge core layer channel 14. The trailing edge core layer channel 14 is located between the first trailing edge skin channel 12 and the second trailing edge skin channel 13, and it is connected to the working fluid output channel 7. Specifically, the first trailing edge skin channel 12 is disposed in the upper surface of the trailing edge skin 2, and the second trailing edge skin channel 13 is disposed in the lower surface of the trailing edge skin 2. A sandwiched trailing edge layer is disposed in the trailing edge skin 2, and the trailing edge core layer channel 14 is disposed in the sandwiched trailing edge layer. The sandwiched trailing edge layer can be connected to the trailing edge skin 2 through an integral molding technology. This method has good sealing performance and strength. However, the two can also be formed separately and then connected by fasteners or welding. In the case of separate configuration, attention needs to be paid to the sealing performance and strength of the connection between the two. It should be noted that the "upper surface" and "lower surface" mentioned above are based on Figure 1 The diagrams illustrating the directional states are for ease of understanding and should not be considered as limitations on the technical solution. Figure 1When the orientation of the diagram changes, the directional terminology will also change accordingly. The first leading edge skin channel 10 is connected to the first trailing edge skin channel 12, and the second leading edge skin channel 11 is connected to the second trailing edge skin channel 13. In addition, the outlet of the leading edge core layer channel 9 is directly opposite the center of the leading edge region 4, while the inlet of the trailing edge core layer channel 14 is directly opposite the center of the trailing edge region 5. During the operation of the aircraft system, the working propellant is input at high speed from the working propellant input channel 6 of the rudder shaft 3 and enters the leading edge core layer channel 9. It is then jetted at high speed into the leading edge region 4. Since the outlet of the leading edge core layer channel 9 is directly opposite the center of the leading edge region 4, the high-speed flowing working propellant directly impacts the center of the leading edge region 4 upon ejection. This impact jet carries away heat from the leading edge region 4, achieving localized cooling. The impact effect of the high-speed jet flow enhances the cooling effect on the inner wall of the leading edge region 4. Furthermore, based on the impact cooling principle combined with the aforementioned structure, the cooling effect is further improved, and the pressure drop of the working propellant in the leading edge skin channel is avoided from affecting the cooling effect. After the working fluid ejected from the core layer channel 9 enters the leading edge region 4, it can flow out of the leading edge skin 1 through the first leading edge skin channel 10 and the second leading edge skin channel 11, simultaneously exchanging heat between the upper and lower surfaces of the leading edge skin 1. It then flows into the trailing edge skin 2 through the first trailing edge skin channel 12 and the second trailing edge skin channel 13. Since the heat generated by the trailing edge skin 2 is lower than that generated by the leading edge skin 1, the working fluid that cools the leading edge skin 1 can also effectively cool the trailing edge skin 2. The working fluid flowing out from the first trailing edge skin channel 12 and the second trailing edge skin channel 13 merges into the trailing edge region 5 and flows out through the trailing edge core layer channel 14. Subsequently, it is discharged from the air rudder through the working fluid output channel 7 of the rudder shaft 3, completing the entire cooling cycle.

[0039] In the preferred embodiment, see Figure 3 Multiple first skin channels 15 are spaced apart in the leading edge skin channel, and all first skin channels 15 are connected in the leading edge region 4. Specifically, multiple first skin channels 15 are spaced apart in the first leading edge skin channel 10, the second leading edge skin channel 11, and the leading edge core layer channel 9, and the multiple first skin channels 15 in the first leading edge skin channel 10, the second leading edge skin channel 11, and the leading edge core layer channel 9 are all distributed along the axial direction of the rudder shaft 3. The length of a single first skin channel 15 extends along the intersection of the upper and lower surfaces of the leading edge skin 1, that is, it extends towards the leading edge region 4. See also Figure 4Multiple second skin tubes 16 are spaced apart in the trailing edge skin channel, and all second skin tubes 16 are connected in the trailing edge region 5. Specifically, multiple second skin tubes 16 are spaced apart in the first trailing edge skin channel 12, the second trailing edge channel 13, and the trailing edge core layer channel 14. The multiple second skin tubes 16 in the first trailing edge skin channel 12, the second trailing edge skin channel 13, and the trailing edge core layer channel 14 are all distributed along the axial direction of the rudder shaft 3. The length of a single second skin tube 16 extends along the intersection of the upper and lower surfaces of the trailing edge skin 2, that is, towards the trailing edge region 5. Through the above structure, the working fluid can be evenly distributed in the leading edge skin channel and the trailing edge skin channel. Combined with the impinging jet, pressure drop during the flow process is avoided, which would affect the cooling effect. Furthermore, the diameters of the first skin pipe 15 and the second skin pipe 16 are between 1mm and 2mm, such as 1mm, 1.1mm, 1.2mm, 1.3mm, 1.4mm, 1.5mm, 1.6mm, 1.7mm, 1.8mm, 1.9mm, or 2mm. Setting the pipe diameters between 1mm and 2mm can further ensure uniform distribution of the working fluid to uniformly cool the leading edge skin 1 and the trailing edge skin 2. It also further avoids pressure drop and ensures effective cooling. In addition, it controls the weight of the aerodynamic rudder and the structural load-bearing capacity. While ensuring thermal protection, it avoids the risk of increasing the weight of the aircraft and reducing the structural strength, thereby improving the performance and safety of the aircraft in hypersonic environments. Furthermore, the diameter of the first skin tube 15 in the first leading edge skin channel 10 and the second leading edge skin channel 11 is larger than the diameter of the first skin tube 15 in the leading edge core layer channel 9. The diameter of the first skin tube 15 in the first leading edge skin channel 10 and the second leading edge skin channel 11 can be set to 1.5mm < α ≤ 2mm, while the diameter of the first skin tube 15 in the leading edge core layer channel 9 can be set to 1mm ≤ α < 1.5mm. This is because the temperature difference between the inner and outer surfaces of the leading edge skin is extremely large, and it is a direct cooling area. By controlling the diameter of the first skin tube 15 in the first leading edge skin channel 10 and the second leading edge skin channel 11, the excessive temperature gradient of the structure can be avoided, which would lead to excessive thermal stress. Since the leading edge core layer channel 9 is located inside, there will be no excessive temperature gradient. By minimizing the diameter of the first skin tube 15 in the leading edge core layer channel 9, a better impact cooling effect can be achieved in the leading edge area.

[0040] In the preferred embodiment, see Figure 5It also includes a working fluid input collection assembly; the working fluid input collection assembly includes a working fluid input collection cavity 17, which is connected to the working fluid input channel 6 on the rudder shaft 3 and at least one leading edge skin channel, that is, the working fluid input collection cavity 17 is simultaneously connected to the working fluid input channel 6 and at least one leading edge skin channel. In this embodiment, the working fluid input collection cavity 17 is connected to the leading edge core layer channel 9; it also includes a working fluid output collection assembly; the working fluid output collection assembly includes a working fluid output collection cavity 18, which is connected to the working fluid output channel 7 on the rudder shaft 3 and at least one trailing edge skin channel, that is, the working fluid output collection cavity 18 is simultaneously connected to the working fluid output channel 7 and at least one trailing edge skin channel. In this embodiment, the working fluid output collection cavity 18 is connected to the trailing edge core layer channel 14.

[0041] In a preferred embodiment, the working fluid input collection assembly further includes an inlet manifold assembly, which includes an inlet manifold connector 19. Multiple inlet pipes are spaced apart within the inlet manifold connector 19 and arranged along the length of the inlet manifold connector 19. The inlet manifold connector 19 is connected to at least one leading edge skin channel through the inlet pipes. In this embodiment, the inlet manifold connector 19 is connected to the leading edge core layer channel 9 through the inlet pipes. That is, the number of inlet pipes is the same as the number of first skin pipes 15 in the leading edge core layer channel 9, and the pipe diameter is the same. The inlet pipes are connected to the first skin pipes 15 one by one. An inlet connector 20 is also provided on the inlet manifold connector 19, and the inlet connector 20 is connected to the working fluid input collection chamber 17. The working fluid output collection assembly also includes an outlet manifold assembly, which includes an outlet manifold connector 21. Multiple outlet pipes are spaced apart within the outlet manifold connector 21 and arranged along the length of the outlet manifold connector 21. The outlet manifold connector 21 is connected to at least one trailing edge skin channel through the outlet pipes. In this embodiment, the outlet manifold connector 21 is connected to the trailing edge core layer channel 14 through the outlet pipes. That is, the number of outlet pipes is the same as the number of second skin pipes 16 in the trailing edge core layer channel 14, and the pipe diameter is the same. The outlet pipes are connected to the second skin pipes 16 one by one. An outlet connection nozzle 22 is also provided on the outlet manifold connector 21, and the outlet connection nozzle 22 is connected to the working fluid output collection chamber 18. The working fluid input through the working fluid input channel 6 is first transported to the working fluid input collection chamber 17. The working fluid collected in the working fluid input collection chamber 17 can be evenly distributed into the leading edge core layer channel 9, ensuring that the working fluid effectively enters the leading edge core layer channel 9. The working fluid flowing out from the trailing edge core layer channel 14 is first evenly transported to the working fluid output collection chamber 18, and then flows from the working fluid output collection chamber 18 into the working fluid output channel 7 of the rudder shaft 3, and is discharged from the air rudder.

[0042] In the preferred embodiment, see Figure 6 and 7The rudder shaft 3 includes an inner tube and an outer tube. The outer tube surrounds the inner tube, and the two tubes are spaced apart and connected by a second lattice structure (not shown in the figure). The inner through-hole of the inner tube forms a working fluid input channel 6. A working fluid output channel 7 is formed between the outer tube and the inner tube, and the second lattice structure is disposed in the working fluid output channel 7. By setting the rudder shaft 3 as a double-layer structure, active heat exchange and cooling can be performed on the wall surface at the connection gap between the air rudder and the fuselage during the inflow and outflow of the working fluid, further achieving effective thermal protection for the air rudder. The second lattice sandwich structure not only enhances the load-bearing capacity and overall rigidity of the rudder shaft 3, but also effectively improves the thermal conductivity of the structure, thereby reducing the concentration of thermal stress and preventing structural failure caused by local overheating.

[0043] In the preferred embodiment, see Figure 5-7 The outer tube is shorter than the inner tube. One end of the inner tube extends beyond the outer tube, while the other end of the inner tube is flush with the corresponding end of the outer tube. The end of the inner tube extending beyond the outer tube passes through the working fluid output collecting chamber 18 and connects to the working fluid input collecting chamber 17. The end of the outer tube through which the inner tube exits also connects to the working fluid output collecting chamber 18. This structure achieves effective thermal protection while also simplifying and making the air rudder structure more compact.

[0044] In the preferred embodiment, see Figure 3 and 4 A first lattice sandwich structure 23 is provided between adjacent leading edge skin channels and adjacent trailing edge skin channels. Specifically, the first lattice sandwich structure 23 is provided between the first leading edge skin channel 10 and the leading edge core layer channel 9, between the leading edge core layer channel 9 and the second leading edge channel 11, between the first trailing edge skin channel 12 and the trailing edge core layer channel 14, and between the trailing edge core layer channel 14 and the second trailing edge channel 13. By setting the first lattice sandwich structure 23, heat conduction can be further improved, heat can be effectively dispersed, and the heat flow can be more evenly distributed throughout the entire air rudder, reducing the concentration of local thermal stress. At the same time, due to the lightweight characteristics of the lattice sandwich structure, the structural strength and thermal protection capability of the air rudder can be guaranteed while the weight of the air rudder can be significantly reduced, which helps to improve the overall performance of the hypersonic vehicle.

[0045] The first and second lattice sandwich structures described above are existing technologies, mostly used in the field of aerospace technology. The lattice sandwich structure itself is not the inventive point of this invention.

[0046] In the preferred embodiment, see Figure 8Multiple reinforcing ribs 24 are spaced apart on the inner wall of the leading edge skin 1 located in the leading edge region 4. These ribs 24 are arranged along the shape of the leading edge region 4, and their size gradually decreases from the center to both sides. Simultaneously, the leading edge region 4 is arc-shaped, and the free ends of all reinforcing ribs 24 face the center of the arc-shaped leading edge region 4. Through this structure, combined with the working fluid impingement jet, the convective heat transfer between the working fluid and the inner wall of the skin in the leading edge region 4 can be effectively enhanced. The gradual change in the size of the reinforcing ribs 24 is to enhance heat transfer efficiency while reasonably controlling the pressure drop in the relatively low heat flux region. By controlling the gradual change in the size and the spacing of the reinforcing ribs, it can be ensured that there is no excessive pressure drop in the leading edge skin channel, maximizing cooling performance.

[0047] In a preferred embodiment, the working fluid is injected at high speed from the working fluid input channel 6 of the rudder shaft 3 into the leading edge core layer channel 9 connected thereto, and then impacts the reinforcing rib 24 located in the middle at high speed from the leading edge core layer channel 9. Since the middle of the leading edge region 4 is the most severely heated, the size of the reinforcing rib 24 gradually decreases from the middle to both sides, forming a gradient, which can achieve the best heat transfer effect without causing excessive pressure drop. This not only effectively balances the overall cooling demand and pressure drop loss, but also makes the temperature of the leading edge region 4 more uniform and reduces thermal stress. The spacing between adjacent reinforcing ribs 24 is represented by p, and the height of the reinforcing rib 24 is represented by h. The reinforcing ribs 24 are distributed from the middle to both sides, satisfying the condition that p / h gradually increases. Reinforcing ribs 24 are also arranged along the length of the leading edge region 4, thereby achieving effective heat transfer throughout the entire leading edge region 4.

[0048] In a preferred embodiment, the leading edge skin 1 and the trailing edge skin 2 are connected by integral molding. A notch is formed between the integrally molded leading edge skin 1 and the trailing edge skin 2 at a position opposite to the rudder shaft 3, and a sealing plate is provided at this notch. Specifically, the integral molding method can be achieved using 3D printing technology, resulting in good sealing performance. Of course, the leading edge skin 1 and the trailing edge skin 2 can also be molded separately and then assembled into one piece using connectors, but attention must be paid to the sealing performance at the connection point. The notch is reserved to facilitate the installation of the rudder shaft 3, the working fluid input collection assembly, and the working fluid output collection assembly. Finally, the notch is sealed with a sealing plate.

[0049] It should be noted that, Figure 1-8 The air rudders in this embodiment are all illustrated using a simplified drawing method, which is a conventional simplified drawing method in this field. In addition, in this embodiment, the rudder shaft 3, the working fluid input collection assembly, the working fluid output collection assembly, and other components are surrounded internally by the leading edge skin 1 and the trailing edge skin 2 to protect these components.

[0050] In the preferred embodiment, see Figure 9The system also includes a fuel tank 25, a valve 26, and a high-pressure pump 27. These components are connected in series, and the entire assembly is connected to the working fluid input channel 6 of the rudder shaft 3 via a pipeline. Specifically, in this embodiment, the fuel tank 25, valve 26, and high-pressure pump 27 are connected sequentially, with the high-pressure pump 27 also connected to the working fluid input channel 6 via a pipeline. In another embodiment, the order of the valve 26 and high-pressure pump 27 can be interchanged, meaning the fuel tank 25, high-pressure pump 27, and valve 26 are connected sequentially, with the valve 26 connected to the working fluid input channel 6 via a pipeline. The working fluid output channel 7 of the rudder shaft 3 is connected to the fuel pipeline of the hypersonic vehicle's engine via a pipeline. The fuel on the hypersonic vehicle is first pumped into the working fluid input channel 6 of the rudder shaft 3 to cool the air rudder, and then discharged from the working fluid output channel 7 of the rudder shaft 3 and enters the engine's fuel pipeline, where it is burned in the engine's combustion chamber. The above structure enables hypersonic vehicles to use their own fuel as a working medium to cool the air rudders. In other words, the hypersonic vehicle's own fuel serves the dual role of fuel and cooling, eliminating the need for additional cooling working medium, reducing the overall weight of the hypersonic vehicle, and the high-temperature fuel discharged by the air rudders has higher combustion efficiency.

[0051] Although embodiments of this application have been shown and described above, it is understood that the above embodiments are exemplary and should not be construed as limiting this application. Those skilled in the art can make changes, modifications, substitutions and variations to the above embodiments within the scope of this application.

Claims

1. An active cooling system for the air rudder of a hypersonic vehicle, characterized in that: The system includes a leading edge skin and a trailing edge skin, with a rudder shaft connected to both the leading and trailing edge skins. The leading edge skin has at least two leading edge skin channels, all of which are connected in the leading edge region. The trailing edge skin has at least two trailing edge skin channels, all of which are connected in the trailing edge region. The rudder shaft includes a working fluid input channel and a working fluid output channel. The working fluid input channel is connected to at least one leading edge skin channel, and the working fluid output channel is connected to at least one trailing edge skin channel. The leading edge skin channels not connected to the working fluid input channel and the trailing edge skin channels not connected to the working fluid output channel are also connected.

2. The active cooling system for the air rudder of a hypersonic vehicle according to claim 1, characterized in that: The leading edge skin has three leading edge skin channels: a first leading edge skin channel, a second leading edge skin channel, and a leading edge core layer channel. The leading edge core layer channel is located between the first and second leading edge skin channels and is connected to the working fluid input channel. The trailing edge skin has three trailing edge skin channels: a first trailing edge skin channel, a second trailing edge skin channel, and a trailing edge core layer channel. The trailing edge core layer channel is located between the first and second trailing edge skin channels and is connected to the working fluid output channel. The first leading edge skin channel is connected to the first trailing edge skin channel, and the second leading edge skin channel is connected to the second trailing edge skin channel.

3. The active cooling system for the air rudder of a hypersonic vehicle according to claim 1 or 2, characterized in that: The leading edge skin channel is provided with multiple first skin pipes at intervals, and all the first skin pipes are connected in the leading edge region; the trailing edge skin channel is provided with multiple second skin pipes at intervals, and all the second skin pipes are connected in the trailing edge region.

4. The active cooling system for the air rudder of a hypersonic vehicle according to claim 1 or 2, characterized in that: It also includes a working fluid input and collection assembly; the working fluid input and collection assembly includes a working fluid input and collection chamber, which is connected to a working fluid input channel on the rudder shaft and at least one leading edge skin channel. Alternatively, it may also include a working fluid output and collection assembly; the working fluid output and collection assembly includes a working fluid output and collection chamber, which is connected to a working fluid output channel on the rudder shaft and at least one trailing edge skin channel.

5. The active cooling system for the air rudder of a hypersonic vehicle according to claim 4, characterized in that: The working fluid input collection assembly also includes an inlet manifold assembly, which includes an inlet manifold connector. Multiple inlet pipes are spaced apart within the inlet manifold connector. The inlet manifold connector is connected to at least one leading edge skin channel through the inlet pipes. An inlet connection nozzle is also provided on the inlet manifold connector, which is connected to the working fluid input collection chamber. Alternatively, the working fluid output collection assembly may further include an outlet manifold assembly, which includes an outlet manifold connector and multiple outlet pipes spaced apart within the outlet manifold connector. The outlet manifold connector is connected to at least one trailing edge skin channel through the outlet pipes. An outlet connection nozzle is also provided on the outlet manifold connector, which is connected to the working fluid output collection chamber. Alternatively, the rudder shaft includes an inner tube and an outer tube, the outer tube being disposed around the inner tube and spaced apart from it, and the two being connected by a second dot matrix structure; the inner through hole of the inner tube forms the working fluid input channel; the working fluid output channel is formed between the outer tube and the inner tube, and the second dot matrix structure is disposed in the working fluid output channel.

6. The active cooling system for the air rudder of a hypersonic vehicle according to claim 5, characterized in that: The length of the outer tube is shorter than the length of the inner tube. One end of the inner tube extends out to the outside of the outer tube, and the other end of the inner tube is flush with the corresponding end of the outer tube. The end of the inner tube that extends out to the outside of the outer tube passes through the working fluid output collection chamber and is connected to the working fluid input collection chamber. The end of the outer tube through which the inner tube extends is connected to the working fluid output collection chamber.

7. The active cooling system for the air rudder of a hypersonic vehicle according to claim 2, characterized in that: Multiple reinforcing ribs are spaced apart on the inner wall of the leading edge skin located in the leading edge region. The multiple reinforcing ribs are arranged along the shape of the leading edge region, and the size of the multiple reinforcing ribs gradually decreases from the middle to both sides.

8. The active cooling system for the air rudder of a hypersonic vehicle according to claim 7, characterized in that: The working propellant is injected at high speed from the working propellant input channel of the rudder shaft into the leading edge core channel connected to it, and then impacts the reinforcing rib located in the middle at high speed from the leading edge core channel.

9. The active cooling system for the air rudder of a hypersonic vehicle according to claim 1 or 2, characterized in that: A first lattice sandwich structure is provided between adjacent leading edge skin channels and between adjacent trailing edge skin channels; Alternatively, the leading edge skin and the trailing edge skin are connected by integral molding, and the integrally molded leading edge skin and the trailing edge skin form a notch at a position opposite to the rudder shaft, and a sealing plate is provided at the notch.

10. The active cooling system for the air rudder of a hypersonic vehicle according to claim 1 or 2, characterized in that: It also includes a fuel tank, valves, and a high-pressure pump, which are connected in series. The entire assembly formed by the fuel tank, valves, and high-pressure pump is connected to the working propellant input channel of the rudder shaft through a pipeline, and the working propellant output channel of the rudder shaft is connected to the fuel pipeline of the hypersonic vehicle's engine through a pipeline.

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