Aerial rudder active cooling system of hypersonic aircraft

Abstract

The invention relates to the technical field of aviation, and particularly discloses an air rudder active cooling system of a hypersonic aircraft, which comprises a leading edge skin and a trailing edge skin, a rudder shaft connected with the leading edge skin and the trailing edge skin is arranged between the leading edge skin and the trailing edge skin, at least two leading edge skin channels are arranged in the leading edge skin, and the leading edge skin channels are communicated with the leading edge skin. All the leading edge skin channels are communicated in the 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 front edge skin channel, the working medium output channel is communicated with at least one rear edge skin channel, and the front edge skin channel which is not communicated with the working medium input channel is communicated with the rear edge skin channel which is not communicated with the working medium output channel. And the air rudder can be cooled through an active thermal protection technology.

Description

Technical Field

[0001] The present invention relates to the field of aviation technology, and in particular to an active cooling system for an air rudder of a hypersonic aircraft. Background Art

[0002] Aircraft play an extremely important role in modern society and are widely used in many fields of human life. Aircraft usually have air rudders, which are important actuators for controlling the flight attitude of the aircraft. They generate lateral control force by changing the air flow, thereby achieving the stability and operability of the aircraft. The full-motion rudder is a type of air rudder. The full-motion rudder refers to a component that combines the horizontal stabilizer and the elevator. It is connected to the fuselage through the rudder shaft. The pilot can control the deflection of the entire horizontal tail and achieve precise control of the flight attitude and direction through the overall deflection. The full-motion rudder is an advanced control component used in aerospace vehicles and high-performance aircraft.

[0003] The fully movable rudder is subjected to extremely high thermodynamic and aerodynamic loads when flying with the aircraft at high speed. For this reason, the fully movable rudder needs to be thermally protected. At present, the industry usually adopts the following technologies to perform thermal protection on the air rudder: The first method is passive thermal protection technology, which relies on the natural properties of the material to protect the aircraft, including heat absorption, radiation heat dissipation and insulation of heat transfer. However, the protection ability of passive thermal protection is limited by the thermal physical properties of the material, especially under high heat flux density, the temperature bearing capacity of the material is insufficient, so it is usually suitable for application scenarios with low and medium heat loads; the second method is semi-passive thermal protection technology, which is a thermal protection method that combines the characteristics of passive and active cooling technologies, and uses the thermal insulation properties and reflection of the material. However, semi-passive thermal protection will face the problem of working fluid material exhaustion in a high heat flux environment, and the expansion or phase change of the working fluid may cause changes in the shape of the aircraft, thereby affecting the aerodynamic performance and control stability. Therefore, it is usually suitable for application scenarios with medium heat loads; the third method is active thermal protection technology. Active thermal protection technology is a technology that absorbs heat through fluid convection heat transfer to cool high-temperature components of the aircraft. It has good thermal protection effects and is suitable for long flight time and reusable advantages. However, active thermal protection technology is currently mainly used for cooling aircraft engines and is less used in full-moving rudders. Applying active thermal protection technology to full-moving rudders has become a key technology in 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 provides an active cooling system for an air rudder of a hypersonic aircraft for cooling the air rudder by using an active thermal protection technology.

[0005] In order to solve the above technical problems, the present invention provides the following technical solutions:

[0006] An air rudder active cooling system for a hypersonic aircraft described in the present invention comprises a leading edge skin and a trailing edge skin, a rudder shaft connected to the leading edge skin and the trailing edge skin is arranged between the two, at least two leading edge skin channels are arranged in the leading edge skin, and all leading edge skin channels are connected in the leading edge region; at least two trailing edge skin channels are arranged in the trailing edge skin, and all trailing edge skin channels are connected in the trailing edge region; the rudder shaft comprises 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, the working fluid output channel is connected to at least one trailing edge skin channel, and the leading edge skin channel not connected to the working fluid input channel is connected to the trailing edge skin channel not connected to the working fluid output channel.

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

[0008] Furthermore, three leading edge skin channels are provided in the leading edge skin, 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; three trailing edge skin channels are provided in the trailing edge skin, 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, a plurality of first skin pipelines are arranged at intervals in the leading edge skin channel, and all the first skin pipelines are connected in the leading edge region; a plurality of second skin pipelines are arranged at intervals in the trailing edge skin channel, and all the second skin pipelines are connected in the trailing edge region.

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

[0011] Alternatively, it also includes a working fluid output liquid collecting component; the working fluid output liquid collecting component includes a working fluid output liquid collecting cavity, and the working fluid output liquid collecting cavity is respectively connected to the working fluid output channel on the rudder shaft and at least one trailing edge skin channel.

[0012] Furthermore, the working medium input liquid collecting assembly also includes an inlet manifold assembly, the inlet manifold assembly includes an inlet manifold connector, a plurality of inlet pipelines are arranged at intervals in the inlet manifold connector, the inlet manifold connector is connected to at least one leading edge skin channel through the inlet pipeline; an inlet connecting nozzle is also arranged on the inlet manifold connector, the inlet connecting nozzle is connected to the working medium input liquid collecting cavity;

[0013] Alternatively, the working medium output liquid collecting assembly further comprises an outlet manifold assembly, the outlet manifold assembly comprises an outlet manifold connector, a plurality of outlet pipelines are arranged at intervals in the outlet manifold connector, the outlet manifold connector is connected to at least one trailing edge skin channel through the outlet pipeline; an outlet connecting nozzle is further arranged on the outlet manifold connector, the outlet connecting nozzle is connected to the working medium output liquid collecting cavity;

[0014] Alternatively, the rudder shaft includes an inner tube and an outer tube, the outer tube is arranged on the periphery of the inner tube, and the outer tube and the inner tube are arranged at an interval, and the two are connected by a second lattice 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 lattice structure is arranged in the working fluid output channel.

[0015] Furthermore, the length of the outer tube is shorter than that of the inner tube, one end of the inner tube passes through 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 passes through the outside of the outer tube passes through the working fluid output liquid collecting chamber and is connected with the working fluid input liquid collecting chamber, and the end of the outer tube through which the inner tube passes is connected with the working fluid output liquid collecting chamber.

[0016] Furthermore, a plurality of reinforcing ribs are arranged at intervals on the inner wall of the leading edge skin located at the leading edge area. The plurality of reinforcing ribs are arranged along the shape of the leading edge area, and the sizes of the plurality of reinforcing ribs gradually decrease from the middle to both sides.

[0017] Furthermore, the working fluid is injected into the leading edge core layer channel connected thereto from the working fluid input channel of the rudder shaft at high speed, and is ejected from the leading edge core layer channel at high speed toward the reinforcing rib located in the middle.

[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, and a gap is formed in the leading edge skin and the trailing edge skin connected by integral molding at a position opposite to the rudder shaft, and a sealing plate is arranged at the gap.

[0020] Furthermore, it also includes a fuel tank, a valve and a high-pressure pump, which are connected in series, and the whole formed by the fuel tank, the valve and the high-pressure pump is connected to the working fluid input channel of the rudder shaft through a pipeline, and the working fluid output channel of the rudder shaft is connected to the fuel pipeline of the engine of the hypersonic aircraft through a pipeline. BRIEF DESCRIPTION OF THE DRAWINGS

[0021] The above and other purposes, features and advantages of the present invention will become more clear by more specific description of the preferred embodiments of the present invention shown in the accompanying drawings. The same reference numerals indicate the same parts in all the accompanying drawings, and the accompanying drawings are not deliberately scaled to the actual size, and the focus is on illustrating the main purpose of the present invention.

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

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

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

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

[0026] Figure 5 This is a schematic diagram of the structure of the working fluid input liquid collecting component and the working fluid output liquid collecting component.

[0027] Figure 6 and Figure 7 Schematic diagram of the rudder shaft structure in different directions.

[0028] Figure 8 Schematic diagram of the reinforcing rib structure in the leading edge area.

[0029] Fig. 9 The figure is a schematic diagram of the connection structure of the present invention on a hypersonic aircraft.

[0030] Among them, the leading edge skin 1; the trailing edge skin 2; the rudder shaft 3; the leading edge area 4; the trailing edge area 5; the working fluid input channel 6; the working fluid output channel 7; the full-moving rudder 8; the leading edge core layer channel 9; the first leading edge skin channel 10; the second leading edge skin channel 11; the first trailing edge skin channel 12; the second trailing edge skin channel 13; the trailing edge core layer channel 14; the first skin pipeline 15; the second skin pipeline 16; the working fluid input liquid collecting cavity 17; the working fluid output liquid collecting cavity 18; the inlet manifold connector 19; the inlet connecting nozzle 20; the outlet manifold connector 21; the outlet connecting nozzle 22; the first lattice sandwich structure 23; the reinforcing rib 24; the fuel storage tank 25; the valve 26; the high-pressure pump 27. DETAILED DESCRIPTION

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

[0032] It should be noted that when an element is considered to be "connected" to another element, it may be directly connected to another element and integrated therewith, or there may be an intermediate element at the same time. The terms "installed", "one end", "the other end" and similar expressions used herein are for illustrative purposes only.

[0033] Unless otherwise defined, all technical and scientific terms used herein have the same meaning as those generally understood by those skilled in the art. The terms used in the specification herein are only for the purpose of describing specific embodiments and are not intended to limit the present invention. The term "and / or" used herein includes any and all combinations of one or more related listed items.

[0034] In the present application, unless otherwise clearly specified and limited, a first feature being “above” or “below” a second feature may mean that the first and second features are in direct contact, or the first and second features are in indirect contact through an intermediate medium. Moreover, a first feature being “above”, “above”, and “above” a second feature may mean that the first feature is directly above or obliquely above the second feature, or simply means that the first feature is higher in level than the second feature. A first feature being “below”, “below”, and “below” a second feature may mean that the first feature is directly below or obliquely below the second feature, or simply means that the first feature is lower in level than the second feature.

[0035] In the description of this specification, the description with reference to the terms "preferred embodiment", "further embodiment", "other embodiments" or "specific example" etc. means that the specific features, structures, materials or characteristics described in conjunction with the embodiment or example are included in at least one embodiment or example of the present application. In this specification, the schematic representations of the above terms do not necessarily refer to the same embodiment or example. Moreover, the specific features, structures, materials or characteristics described may be combined in any one or more embodiments or examples in a suitable manner. In addition, those skilled in the art may combine and combine the different embodiments or examples described in this specification and the features of the different embodiments or examples, unless they are contradictory.

[0036] The present invention provides a specific implementation of an active cooling system for an air rudder of a hypersonic aircraft, see Figure 1, including a leading edge skin 1 and a trailing edge skin 2, a rudder shaft 3 connected to the leading edge skin 1 and the trailing edge skin 2 is arranged between the two, at least two leading edge skin channels are arranged in the leading edge skin 1, and all leading edge skin channels are connected in the leading edge area 4. Specifically, the so-called leading edge is a prior art term, that is, the intersection position of the upper surface and the lower surface of the leading edge skin 1, and there are at least two leading edge skin channels. When there are only two leading edge skin channels, one leading edge skin channel is arranged in the upper surface of the leading edge skin 1, and the other leading edge skin channel is arranged in the lower surface of the leading edge skin 1, and the lengths of the two leading edge skin channels extend to the upper surface of the leading edge skin 1. and the lower surface, that is, the leading edge area 4, and are connected in this area; at least two trailing edge skin channels are arranged in the trailing edge skin 2, and all trailing edge skin channels are connected in the trailing edge area 5. Specifically, the so-called trailing edge is also a term in the prior art, that is, the intersection of the upper surface and the lower surface of the trailing edge skin 2, and there are at least two trailing edge skin channels. When there are only two trailing edge skin channels, one trailing edge skin channel is arranged in the upper surface of the trailing edge skin 2, and the other trailing edge skin channel is arranged in the lower surface of the trailing edge skin 2, and the lengths of the two trailing edge skin channels extend to the intersection of the upper surface and the lower surface of the trailing edge skin 2, that is The rudder shaft 3 is located at the trailing edge region 5 and is connected in this region; the rudder shaft 3 includes a working fluid input channel 6 and a working fluid output channel 7, the working fluid input channel 6 is connected to at least one leading edge skin channel, when there are only two leading edge skin channels, the working fluid input channel 6 is only connected to one of the leading edge skin channels; the working fluid output channel 7 is connected to at least one trailing edge skin channel, when there are only two trailing edge skin channels, the working fluid output channel 7 is only connected to one of the trailing edge skin channels; the leading edge skin channel not connected to the working fluid input channel 6 is connected to the trailing edge skin channel not connected to the working fluid output channel 7; wherein the working fluid may be an aircraft fuel in a supercritical state (for example Aviation kerosene) or carbon dioxide and other fluids, select a suitable working fluid according to actual conditions, and in this embodiment, aviation kerosene is preferably used as the working fluid. Aviation kerosene can operate efficiently in a stable state to ensure the safety and stability of the aircraft. It has a high specific heat capacity and can absorb more heat in a smaller volume, thereby effectively preventing overheating problems. It is also possible to consider using the aviation kerosene used by the aircraft itself as a working fluid to cool the air rudder, that is, the aircraft's own fuel has the dual roles of fuel and cooling, and no additional cooling working fluid is needed, thereby reducing the overall weight of the aircraft, and the high-temperature fuel discharged from the air rudder has a higher combustion efficiency.

[0037] When the air rudder active cooling system is working, especially during hypersonic flight, the leading edge region 4 of the air rudder (such as the full-moving rudder 8) will be subjected to severe aerodynamic heating, and the temperature in this region is often the highest. Therefore, it is necessary to effectively cool the leading edge region 4, which is crucial to ensure the thermal protection performance of the air rudder. In the present invention, the working fluid enters the leading edge skin channel connected to the working fluid input channel 6 of the rudder shaft 3, and flows along the channel to the leading edge region 4 of the air rudder, and then flows out from the leading edge skin channel not connected to the working fluid input channel 6 to the trailing edge skin channel not connected to the working fluid output channel 7, and flows along the channel to the trailing edge region 5 of the air rudder, and finally flows out from the trailing edge skin channel connected to the working fluid output channel 7 to the working fluid output channel 7, and then is discharged from the working fluid output channel 7, completing the cooling cycle of the air rudder. During the flow process, the working fluid can exchange heat and cool the leading edge skin 1, the leading edge area 4 of the air rudder, the trailing edge skin 2, and the trailing edge area 5 of the air rudder, effectively taking away the heat generated in the air rudder, and effectively ensuring the thermal protection performance of the air rudder. In addition, when the air rudder is connected to the aircraft body, there is usually a gap at the connection, resulting in heat flow there. Since the rudder shaft 3 has a working fluid input channel and a working fluid output channel, and the air rudder is connected to the aircraft body through the rudder shaft 3, the high-temperature wall surface at the connection gap can be heat exchanged by the working fluid flowing in the rudder shaft 3, taking away the aerodynamic heating heat, and further effectively ensuring the thermal protection performance of the air rudder.

[0038] In the preferred embodiment, see Figure 1Three leading edge skin channels are arranged in the leading edge skin 1, namely the first leading edge skin channel 10, the second leading edge skin channel 11 and the 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 the leading edge core layer channel 9 is connected to the working medium input channel 6. Specifically, the first leading edge skin channel 10 is arranged in the upper surface of the leading edge skin 1, and the second leading edge skin channel 11 is arranged in the lower surface of the leading edge skin 1. A sandwich leading edge layer is arranged in the leading edge skin 1, and the leading edge core layer channel 9 is arranged in the sandwich leading edge layer. The sandwich leading edge layer can be connected to the leading edge skin 1 by an integrated molding technology. This method has better sealing performance and strength. However, the two can also be formed separately and then connected by fasteners or welding. In the case of a separate setting, it is necessary to pay attention to the sealing performance and strength of the connection between the two. Three trailing edge skins are arranged in the trailing edge skin 2 The channels are respectively 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 the trailing edge core layer channel 14 is connected to the working medium output channel 7. Specifically, the first trailing edge skin channel 12 is arranged in the upper surface of the trailing edge skin 2, and the second trailing edge skin channel 13 is arranged in the lower surface of the trailing edge skin 2. A sandwich trailing edge layer is arranged in the trailing edge skin 2, and the trailing edge core layer channel 14 is arranged in the sandwich trailing edge layer. The sandwich trailing edge layer can be connected to the trailing edge skin 2 by an integrated molding technology. This method has better sealing performance and strength. However, the two can also be formed separately and then connected by fasteners or welding. In the case of a separate setting, it is necessary to pay attention 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 directions in the diagram are described for ease of understanding and cannot be used as a limitation of the technical solution. Figure 1When the illustrated direction in FIG. 1 changes, the azimuth terms will also change adaptively. 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 faces the middle position of the leading edge region 4, and the inlet of the trailing edge core layer channel 14 faces the middle position of the trailing edge region 5. During the system operation of the aircraft, the working fluid is input at high speed from the working fluid input channel 6 of the rudder shaft 3 and enters the leading edge core layer channel 9, and is ejected at high speed from the leading edge core layer channel 9 to the leading edge area 4. Since the outlet of the leading edge core layer channel 9 is directly opposite to the middle position of the leading edge area 4, the high-speed flowing working fluid directly impacts the middle position of the leading edge area 4 when ejected, thereby taking away the heat of the leading edge area 4 through the impact jet, thereby achieving local cooling of the leading edge area 4, and the impact effect of the high-speed jet flow is used to enhance the cooling effect on the inner wall of the leading edge area 4. In addition, based on the impact cooling principle combined with the aforementioned structure, the cooling effect is further improved, and the cooling effect is also avoided from being affected by the pressure drop of the working fluid in the leading edge skin channel; After the working fluid ejected from the edge core layer channel 9 enters the leading edge area 4, it can flow out of the leading edge skin 1 from the first leading edge skin channel 10 and the second leading edge skin channel 11, and at the same time, heat is exchanged on the upper and lower surfaces of the leading edge skin 1, and flow 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 converges into the trailing edge area 5, flows out from the trailing edge core layer channel 14, and is then 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 pipelines 15 are arranged at intervals in the leading edge skin channel, and all the first skin pipelines 15 are connected in the leading edge region 4. Specifically, multiple first skin pipelines 15 are arranged at intervals 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 pipelines 15 in the first leading edge skin channel 10, the multiple first skin pipelines 15 in the second leading edge skin channel 11, and the multiple first skin pipelines 15 in the leading edge core layer channel 9 are all distributed along the axial direction of the rudder shaft 3, and the length of a single first skin pipeline 15 extends along the intersection of the upper surface and the lower surface of the leading edge skin 1, that is, extends to the leading edge region 4. See Figure 4, multiple second skin pipelines 16 are arranged at intervals in the trailing edge skin channel, and all the second skin pipelines 16 are connected in the trailing edge region 5. Specifically, multiple second skin pipelines 16 are arranged at intervals in the first trailing edge skin channel 12, the second trailing edge channel 13, and the trailing edge core layer channel 14, and the multiple second skin pipelines 16 in the first trailing edge skin channel 12, the multiple second skin pipelines 16 in the second trailing edge skin channel 13, and the multiple second skin pipelines 16 in the trailing edge core layer channel 14 are all distributed along the axial direction of the rudder shaft 3, and the length of a single second skin pipeline 16 extends along the intersection of the upper surface and the lower surface of the trailing edge skin 2, that is, extends to the trailing edge region 5. Through the above structure, the working medium can be evenly distributed in the leading edge skin channel and the trailing edge skin channel, and combined with the impingement jet, it is avoided that the pressure drop in the flow process affects the cooling effect. In addition, the diameter of the first skin pipeline 15 and the diameter of the second skin pipeline 16 are between 1 mm and 2 mm, such as 1 mm, 1.1 mm, 1.2 mm, 1.3 mm, 1.4 mm, 1.5 mm, 1.6 mm, 1.7 mm, 1.8 mm, 1.9 mm or 2 mm. Setting the diameter at 1 mm to 2 mm can further ensure uniform distribution of the working fluid to uniformly cool the leading edge skin 1 and the trailing edge skin 2, while further avoiding the generation of pressure drop to ensure effective cooling. In addition, the weight and structural bearing performance of the air rudder are also controlled. While ensuring the thermal protection effect, the risk of increasing the weight of the aircraft and reducing the structural strength is avoided, thereby improving the performance and safety of the aircraft in a hypersonic environment. In addition, the diameter of the first skin pipeline 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 pipeline 15 in the leading edge core layer channel 9, and the diameter of the first skin pipeline 15 in the first leading edge skin channel 10 and the second leading edge skin channel 11 can be set at 1.5mm<α≤2mm, while the diameter of the first skin pipeline 15 in the leading edge core layer channel 9 can be set at 1mm≤α<1.5. 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 pipeline 15 in the first leading edge skin channel 10 and the second leading edge skin channel 11, it is possible to avoid excessive thermal stress caused by excessive structural temperature gradient. Since the leading edge core layer channel 9 is located inside, there will be no excessively high temperature gradient. By reducing the diameter of the first skin pipeline 15 in the leading edge core layer channel 9 as much as possible, a better impact cooling effect can be achieved on the leading edge area.

[0040] In the preferred embodiment, see Figure 5, and also includes a working fluid input liquid collecting component; the working fluid input liquid collecting component includes a working fluid input liquid collecting chamber 17, which is respectively 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 liquid collecting chamber 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 liquid collecting chamber 17 is connected to the leading edge core layer channel 9; and also includes a working fluid output liquid collecting component; the working fluid output liquid collecting component includes a working fluid output liquid collecting chamber 18, which is respectively 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 liquid collecting chamber 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 liquid collecting chamber 18 is connected to the trailing edge core layer channel 14.

[0041] In a preferred embodiment, the working fluid input liquid collecting assembly also includes an inlet manifold assembly, which includes an inlet manifold connector 19, in which a plurality of inlet pipelines are arranged at intervals, and the plurality of inlet pipelines are arranged along the length direction of the inlet manifold connector 19, and the inlet manifold connector 19 is connected to at least one leading edge skin channel through the inlet pipeline. In this embodiment, the inlet manifold connector 19 is connected to the leading edge core layer channel 9 through the inlet pipeline, that is, the number of inlet pipelines is consistent with the number of first skin pipelines 15 in the leading edge core layer channel 9, the pipe diameters are consistent, and the inlet pipelines are connected one by one with the first skin pipelines 15; an inlet connecting nozzle 20 is also provided on the inlet manifold connector 19, and the inlet connecting nozzle 20 is connected to the working fluid input liquid collecting cavity 17. The working fluid output liquid collection component also includes an outlet manifold component, which includes an outlet manifold connector 21. Multiple outlet pipelines are arranged at intervals in the outlet manifold connector 21, and the multiple outlet pipelines are arranged along the length direction of the outlet manifold connector 21. The outlet manifold connector 21 is connected to at least one trailing edge skin channel through the outlet pipeline. In this embodiment, the outlet manifold connector 21 is connected to the trailing edge core layer channel 14 through the outlet pipeline, that is, the number of outlet pipelines is consistent with the number of second skin pipelines 16 in the trailing edge core layer channel 14, the pipe diameters are consistent, and the outlet pipelines are connected one by one with the second skin pipelines 16; an outlet connecting nozzle 22 is also provided on the outlet manifold connector 21, and the outlet connecting nozzle 22 is connected to the working fluid output liquid collection cavity 18. The working fluid inputted by the working fluid input channel 6 is firstly transported to the working fluid input collecting chamber 17, and the working fluid collected in the working fluid input collecting chamber 17 can be evenly distributed to the leading edge core layer channel 9, so as to ensure that the working fluid effectively enters the leading edge core layer channel 9; the working fluid flowing out of the trailing edge core layer channel 14 is firstly evenly transported to the working fluid output collecting chamber 18, and then flows from the working fluid output collecting chamber 18 into the working fluid output channel 7 of the rudder shaft 3 to be 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 is arranged on the periphery of the inner tube, and the outer tube and the inner tube are arranged at intervals, and the two are 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 arranged in the working fluid output channel 7. By setting the rudder shaft 3 as a double-layer structure, the wall surface at the connection gap between the air rudder and the fuselage can be actively heat exchanged and cooled during the inflow and outflow of the working fluid, and effective thermal protection of the air rudder can be further achieved. By setting the second lattice sandwich structure, not only the bearing capacity and overall stiffness of the rudder shaft 3 are enhanced, but also the thermal conductivity of the structure is effectively improved, 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 length of the outer tube is shorter than that of the inner tube, one end of the inner tube passes through 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 passes through the outer tube passes through the working medium output liquid collecting chamber 18 and is connected to the working medium input liquid collecting chamber 17, and the end of the outer tube through which the inner tube passes is connected to the working medium output liquid collecting chamber 18. The above structure can achieve effective thermal protection while making the air rudder structure simple and more compact.

[0044] In the preferred embodiment, see Figure 3 and 4 The 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. The first lattice sandwich structure 23 can further improve heat conduction, effectively disperse heat, make the heat flow more evenly distributed in the entire air rudder, and reduce 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 ensured while significantly reducing the weight of the air rudder, which helps to improve the overall performance of the hypersonic aircraft.

[0045] The first lattice sandwich structure and the second lattice sandwich structure are both prior arts, and are mostly used in the field of aerospace technology. The lattice sandwich structure itself is not the inventive point of the present invention.

[0046] In the preferred embodiment, see Figure 8, multiple reinforcing ribs 24 are arranged at intervals on the inner wall of the leading edge skin 1 located at the leading edge area 4, and the multiple reinforcing ribs 24 are arranged along the shape of the leading edge area 4, and the size of the multiple reinforcing ribs 24 gradually decreases from the middle to both sides. At the same time, the leading edge area 4 is arc-shaped, and the free ends of all reinforcing ribs 24 are facing the center of the arc-shaped leading edge area 4. Through the above structure and combined with the working fluid impact jet, the convective heat exchange between the working fluid and the inner wall of the skin of the leading edge area 4 can be effectively further enhanced. The gradual change of the size of the reinforcing ribs 24 is to enhance the heat exchange efficiency while reasonably controlling the pressure drop in the relatively low heat flow area. By controlling the gradual change of the size of the reinforcing ribs and the interval setting, it can be ensured that there will be no excessive pressure drop in the leading edge skin channel, and the cooling performance can be maximized.

[0047] In a preferred embodiment, the working fluid is injected into the leading edge core layer channel 9 connected thereto from the working fluid input channel 6 of the rudder shaft 3 at high speed, and is ejected from the leading edge core layer channel 9 at high speed toward the reinforcing rib 24 located in the middle. Since the middle part of the leading edge region 4 is most seriously heated, the size of the reinforcing rib 24 gradually decreases from the middle part to both sides, forming a gradient, which can achieve the best heat exchange effect without causing excessive pressure drop. This can not only effectively balance the overall cooling demand and pressure drop loss, but also make the temperature of the leading edge region 4 more uniform and reduce thermal stress. The spacing between adjacent reinforcing ribs 24 is represented by p, and the height of the reinforcing ribs 24 is represented by h. They are distributed from the reinforcing ribs 24 in the middle to both sides thereof, satisfying the gradual increase of p / h. Reinforcing ribs 24 are also arranged in the length direction of the leading edge region 4, so as to achieve effective heat exchange for 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, and the leading edge skin 1 and the trailing edge skin 2 connected by integral molding form a gap at a position opposite to the rudder shaft 3, and a sealing plate is provided at the gap. Specifically, the integral molding method can be realized by 3D printing technology, and the sealing performance is good. Of course, the leading edge skin 1 and the trailing edge skin 2 can also be formed separately and then assembled into one by a connector, but attention should be paid to the sealing performance of the connection; by reserving a gap, it is convenient to install the rudder shaft 3, the working fluid input liquid collection component and the working fluid output liquid collection component, and finally the gap is closed by a sealing plate.

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

[0050] In the preferred embodiment, see Fig. 9, further comprising a fuel tank 25, a valve 26 and a high-pressure pump 27, the fuel tank 25, the valve 26 and the high-pressure pump 27 are connected in series, and the whole formed by the fuel tank 25, the valve 26 and the high-pressure pump 27 in series is connected to the working medium input channel 6 of the rudder shaft 3 through a pipeline. Specifically, in this embodiment, the fuel tank 25, the valve 26 and the high-pressure pump 27 are connected in sequence, and the high-pressure pump 27 is also connected to the working medium input channel 6 through a pipeline. In other embodiments, the position sequence of the valve 26 and the high-pressure pump 27 can be exchanged, that is, the fuel tank 25, the high-pressure pump 27 and the valve 26 are connected in sequence, and the valve 26 is connected to the working medium input channel 6 through a pipeline; the working medium output channel 7 of the rudder shaft 3 is connected to the fuel pipeline of the engine of the hypersonic aircraft through a pipeline. The fuel on the hypersonic aircraft is first pumped into the working medium input channel 6 of the rudder shaft 3 to cool the air rudder, and then discharged from the working medium output channel 7 of the rudder shaft 3 and enters the fuel pipeline of the engine, and enters the combustion chamber of the engine through the fuel pipeline of the engine for combustion. Through the above structure, the hypersonic aircraft can use its own fuel as a working fluid to cool the air rudder, that is, the hypersonic aircraft's own fuel has the dual role of fuel and cooling, and no additional cooling working fluid is needed, thereby reducing the overall weight of the hypersonic aircraft, and the high-temperature fuel discharged by the air rudder has a higher combustion efficiency.

[0051] Although the embodiments of the present application have been shown and described above, it can be understood that the above embodiments are exemplary and cannot be understood as limitations on the present application. Ordinary technicians in this field can change, modify, replace and modify the above embodiments within the scope of the present application.

Claims

1. An active cooling system for an air rudder of a hypersonic vehicle, characterized in that: It includes a leading edge skin and a trailing edge skin, a rudder shaft connected to the leading edge skin and the trailing edge skin is arranged between the two, at least two leading edge skin channels are arranged in the leading edge skin, and all the leading edge skin channels are connected in the 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 connected in the trailing edge area; 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, the working fluid output channel is connected to at least one trailing edge skin channel, the leading edge skin channel not connected to the working fluid input channel is connected to the trailing edge skin channel not connected to the working fluid output channel.

2. The air rudder active cooling system for a hypersonic aircraft according to claim 1, characterized in that: The leading edge skin is provided with three leading edge skin channels, namely the first leading edge skin channel, the second leading edge skin channel and the 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 with the working fluid input channel; the trailing edge skin is provided with three trailing edge skin channels, namely the first trailing edge skin channel, the second trailing edge skin channel and the 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 with the working fluid output channel; the first leading edge skin channel is connected with the first trailing edge skin channel, and the second leading edge skin channel is connected with the second trailing edge skin channel.

3. The air rudder active cooling system for a hypersonic aircraft according to claim 1 or 2, characterized in that: A plurality of first skin pipelines are arranged at intervals in the leading edge skin channel, and all the first skin pipelines are connected in the leading edge region; a plurality of second skin pipelines are arranged at intervals in the trailing edge skin channel, and all the second skin pipelines are connected in the trailing edge region.

4. The air rudder active cooling system for a hypersonic aircraft according to claim 1 or 2, characterized in that: It also includes a working fluid input liquid collecting component; the working fluid input liquid collecting component includes a working fluid input liquid collecting cavity, and the working fluid input liquid collecting cavity is respectively connected to the working fluid input channel on the rudder shaft and at least one leading edge skin channel; Alternatively, it also includes a working fluid output liquid collecting component; the working fluid output liquid collecting component includes a working fluid output liquid collecting cavity, and the working fluid output liquid collecting cavity is respectively connected to the working fluid output channel on the rudder shaft and at least one trailing edge skin channel.

5. The air rudder active cooling system for a hypersonic aircraft according to claim 4, characterized in that: The working medium input liquid collecting assembly also includes an inlet manifold assembly, the inlet manifold assembly includes an inlet manifold connector, a plurality of inlet pipelines are arranged at intervals in the inlet manifold connector, the inlet manifold connector is connected to at least one leading edge skin channel through the inlet pipeline; an inlet connecting nozzle is also arranged on the inlet manifold connector, the inlet connecting nozzle is connected to the working medium input liquid collecting cavity; Alternatively, the working medium output liquid collecting assembly further comprises an outlet manifold assembly, the outlet manifold assembly comprises an outlet manifold connector, a plurality of outlet pipelines are arranged at intervals in the outlet manifold connector, the outlet manifold connector is connected to at least one trailing edge skin channel through the outlet pipeline; an outlet connecting nozzle is further arranged on the outlet manifold connector, the outlet connecting nozzle is connected to the working medium output liquid collecting cavity; Alternatively, the rudder shaft includes an inner tube and an outer tube, the outer tube is arranged on the periphery of the inner tube, and the outer tube and the inner tube are arranged at an interval, and the two are connected by a second lattice 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 lattice structure is arranged in the working fluid output channel.

6. The air rudder active cooling system for a hypersonic aircraft according to claim 5, characterized in that: The length of the outer tube is shorter than that of the inner tube, one end of the inner tube passes through 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 passes through the outside of the outer tube passes through the working fluid output liquid collecting chamber and is connected with the working fluid input liquid collecting chamber, and the end of the outer tube through which the inner tube passes is connected with the working fluid output liquid collecting chamber.

7. The air rudder active cooling system for a hypersonic aircraft according to claim 2, characterized in that: A plurality of reinforcing ribs are arranged at intervals on the inner wall of the leading edge skin located at the leading edge area. The plurality of reinforcing ribs are arranged along the shape of the leading edge area, and the sizes of the plurality of reinforcing ribs gradually decrease from the middle to both sides.

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

9. The air rudder active cooling system for a hypersonic aircraft 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 a gap is formed in the leading edge skin and the trailing edge skin connected by integral molding at a position opposite to the rudder shaft, and a sealing plate is provided at the gap.

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

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