A design method of a thermal management system for a high-speed aircraft based on aerodynamic heat utilization technology

Abstract

The application relates to a high-speed aircraft thermal management system design method based on a pneumatic heat utilization technology. The system comprises a thermal control box arranged in a control surface, a cooling channel arranged at the bottom end of the control surface and a cooling assembly arranged in a fuselage. A gap is arranged between the bottom end of the control surface and the fuselage, and the cooling channel is provided with an exhaust port facing the gap. One end of the cooling assembly is communicated with the thermal control box through a pipeline, the pipeline is provided with a pressure control member, and the other end is communicated with the cooling channel. The thermal control box is provided with a cooling working medium, and the cooling working medium absorbs the heat of the control surface to generate phase change. The application can realize the multiple functions of leading edge drag reduction, heat protection, semi-active thermal control of the gap and connecting members and the like, and effectively ensure the safe and stable working temperature of the servo mechanism.

Description

Technical Field

[0001] This application relates to the field of thermal protection technology for high-speed aircraft, and in particular to a design method for a thermal management system for high-speed aircraft based on aerodynamic heat utilization technology. Background Technology

[0002] High-speed aircraft represent the forefront of aerospace technology innovation. With technological advancements, the need for aircraft has gradually shifted from simply being able to fly to "flying faster." Therefore, research and application of higher Mach flight and reusability for high-speed aircraft have become a key focus for countries worldwide. However, during high Mach flight, the nose, wings, and rudders will be subjected to extreme aerodynamic forces and thermal loads, potentially compromising the effectiveness of structural materials and structural integrity, thus threatening the stability and safety of the aircraft.

[0003] Among them, the nose and control wings / rudders of high-speed aircraft are key components exposed to high-speed airflow. The leading edge and gap areas have extremely high heat flow. External aerodynamic heating will be transferred to the drive and actuator mechanisms inside the cabin along the connecting parts. Local high temperature will cause the temperature-sensitive devices inside the cabin to fail, thereby threatening flight stability and safety.

[0004] Conventional control wings / rudders employ passive thermal protection, with a structure of "metal frame + heat shield," which is a passive form of "heat resistance." However, with increasing flight speed and duration, the long-term accumulation of high heat flux leads to shortcomings in the performance of passive thermal protection, such as insufficient performance or excessive compensation. Active thermal management, due to its superior thermal protection performance, has become a more promising design approach. Current research shows that single thermal protection methods often have limitations; for example, single evaporative cooling has weak drag reduction effects, and single reverse jet cooling performs poorly. Furthermore, the aforementioned active thermal protection methods typically require carrying large amounts of coolant and pumping equipment. For servo systems, the tail section has limited space, and the volume for carrying coolant is limited. Maximizing the utilization of designable space to maximize coolant efficiency is a key issue in the thermal protection design of control wings / rudders during high-Mach flight, and it also plays a decisive role in the overall design of high-speed aircraft. Summary of the Invention

[0005] Therefore, it is necessary to provide a design method for a high-speed aircraft thermal management system based on aerodynamic heat utilization technology to address the above-mentioned technical problems. The high-speed aircraft thermal management system designed by the above method can achieve semi-active thermal control of gaps, connectors and other parts within a limited space.

[0006] A design method for a high-speed aircraft thermal management system based on aerodynamic heat utilization technology is disclosed. The system includes: a thermal control box disposed within a control surface; a cooling channel disposed at the bottom of the control surface; and a cooling assembly disposed within the fuselage. A gap exists between the bottom of the control surface and the fuselage, and an exhaust port is provided in the cooling channel facing this gap. One end of the cooling assembly is connected to the thermal control box via a pipe equipped with a pressure control element, and the other end is connected to the cooling channel. The thermal control box contains a cooling medium that absorbs heat from the control surface and undergoes a phase change.

[0007] On the other hand, a jet port is provided on the fuselage located upstream of the control surface. The jet port is connected to the cooling assembly through a pipe, and a pressure control component is provided on the pipe.

[0008] Furthermore, the cooling assembly is fixed to the bottom end of the control surface by a connector; the connector has a concave cavity structure, and the cooling assembly is placed in the concave cavity.

[0009] Furthermore, a temperature-sensitive device is installed inside the body, and the connector is connected to the temperature-sensitive device.

[0010] Furthermore, the cooling assembly includes: a sweating cooling end connected to the cooling channel, and a porous medium region, a cooling working fluid region, and a flexible diaphragm disposed below the sweating cooling end;

[0011] The flexible diaphragm and the connector form a sealed cavity, which is connected to the thermal control box via a pipe;

[0012] Along the path of the aforementioned pipe, another pipe extends to the jet outlet of the fuselage.

[0013] Furthermore, the exhaust port is a series of spaced holes or a groove.

[0014] Furthermore, the jet outlet is a single hole, or a plurality of holes arranged at intervals, or a groove.

[0015] Furthermore, the sweating cooling end has a plate-like structure and is made of sintered porous metal material.

[0016] Furthermore, the cooling channel is a series serpentine coil structure.

[0017] Furthermore, the porous medium region is provided with a lightweight porous thermal insulation felt.

[0018] Compared with existing technologies, the design method for a high-speed aircraft thermal management system based on aerodynamic heat utilization technology provided in this application has the following advantages:

[0019] 1. By absorbing heat through the thermal control box and cooling components and discharging the phase-change cooling medium from the exhaust port, semi-active thermal control of gaps, connecting parts and other parts can be achieved, effectively ensuring the safe and stable operating temperature of the servo mechanism.

[0020] 2. It effectively utilizes the limited design space within the servo system, without carrying additional large equipment, thus optimizing space utilization. Attached Figure Description

[0021] Figure 1 This is a schematic diagram of the thermal management system structure of a high-speed aircraft in one embodiment;

[0022] Explanation of reference numerals in the attached figures:

[0023] 1. Control surface, 2. Body, 3. Connector, 4. Temperature-sensitive device, 5. Thermal control box, 6a, 6b. Pressure control components, 7. Flexible diaphragm, 8. Cooling medium area, 81. Cooling medium, 9. Porous medium, 91. Sweating cooling end, 10. Cooling channel, 11. Exhaust port, 111. Jet port, 12. High-speed inflow, 13. Detailed Implementation

[0024] To make the objectives, technical solutions, and advantages of this application clearer, the following detailed description is provided in conjunction with the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are merely illustrative and not intended to limit the scope of this application.

[0025] It should be noted that all directional indications (such as up, down, left, right, front, back, etc.) in the embodiments of the present invention are only used to explain the relative positional relationship and movement of each component in a certain specific posture (as shown in the figure). If the specific posture changes, the directional indication will also change accordingly.

[0026] Furthermore, in this invention, descriptions involving "first," "second," etc., are for descriptive purposes only and should not be construed as indicating or implying their relative importance or implicitly specifying the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include at least one of that feature. In the description of this invention, "a plurality of" means at least two, such as two, three, etc., unless otherwise explicitly and specifically defined.

[0027] Understandably, in order to effectively utilize the limited design space within the servo system, this application designs a high-speed aircraft thermal management system. By introducing aerodynamic heat as a power source into the system, the system uses the cooling medium in the thermal control box to absorb the heat accumulated on the control surfaces, and the cooling components absorb the heat from the connectors, which is then discharged through the exhaust port, thus achieving thermal control at the connectors and gaps. On the other hand, by setting a jet port connected to the cooling components, the phase-change cooling medium is ejected through the jet port, thereby disrupting the shock wave at the leading edge of the control surface, and thus reducing drag and preventing heat at the leading edge.

[0028] The embodiments of the present invention will now be described in detail with reference to the accompanying drawings.

[0029] See Figure 1 The high-speed aircraft thermal management system provided in this embodiment includes: a thermal control box 5 disposed in the control surface 1, a cooling channel 11 disposed at the bottom of the control surface 1, and a cooling assembly disposed in the fuselage 2; there is a gap between the bottom of the control surface 1 and the fuselage 2, and the cooling channel 11 has an exhaust port 111 opening towards the gap; one end of the cooling assembly is connected to the thermal control box 5 through a pipe, and a pressure control component 6a is disposed on the pipe, and the other end is connected to the cooling channel 11; the thermal control box 5 contains a cooling medium, and the cooling medium absorbs heat from the control surface 1 to undergo a phase change.

[0030] It is understood that the cooling medium 81 in the thermal control box 5 absorbs the heat from the control surface 1 to produce a phase change. After the phase change, the cooling medium 81 generates pressure. When the pressure reaches a certain level, it breaks through the pressure control element 6a and flows into the cooling assembly through the pipeline, generating pressure on the cooling assembly. The cooling medium 81 in the cooling assembly absorbs the heat from the connector 3 to produce a phase change. Then, under the action of pressure, the phase-changed cooling medium 81 enters the cooling channel 11 and is discharged through the exhaust port 111, thereby providing thermal protection for the connector 3 and the gap.

[0031] On the other hand, a jet nozzle 12 is provided on the fuselage 2 located upstream of the control surface 1. The jet nozzle 12 is connected to the cooling assembly via a pipe, and a pressure control element 6b is installed on the pipe. When the pressure inside the cooling assembly reaches a certain level and exceeds the threshold of the pressure control element 6b, the pressure control element 6b opens, the pipe is connected, and the phase-change cooling medium 81 is ejected from the jet nozzle 12 through the pipe. The ejected jet interacts with the incoming flow, disrupting / pushing away the shock wave at the leading edge of the control surface, thereby reducing drag and preventing heat at the leading edge. It is worth noting that the jet nozzle 12 can be a single hole or a plurality of holes arranged at intervals. The shape of the holes can be polygonal, circular, or, depending on the situation, a rectangular groove.

[0032] Specifically, the thermal control box 5 is a sealed structure, connected to the cooling components only through pipes, and stores a cooling medium 81 inside. The number of pipes is set according to the situation; in this example, one pipe is used.

[0033] The cooling channel 11 can be a series serpentine coil structure or a rod-shaped hollow structure, preferably a series serpentine coil structure. This structure occupies less space, achieves more uniform fluid distribution, avoids localized temperature differences, and provides greater heat exchange efficiency. The cooling channel 11 is arranged parallel to the bottom end of the control surface 1, with one section connected to the cooling assembly and the other section serving as an exhaust port 111. The exhaust port 111 can be a series of spaced through holes or a long, narrow groove structure. The phase-change cooling medium 81 flows from the exhaust port 111 to the high-heat area of ​​the gap, thereby cooling the gap area.

[0034] The cooling assembly is connected to the bottom of the control surface 1 via connector 3. Connector 3 is a concave cavity structure, open at the top and sealed around the sides and bottom. The cooling assembly is placed within this cavity. The upper end of this cavity structure passes through the body 2 and connects to the control surface 1. Inside the body 2, a temperature-sensitive device 4 is installed. Connector 3 is connected to the temperature-sensitive device 4. Specifically, the outer surface near the concave end of connector 3 is in contact with the surface of the temperature-sensitive device 4. The temperature-sensitive device 4 is mainly used to drive the control surface 1 to rotate. Through the action of the temperature-sensitive device 4, torque is transmitted to the control surface 1 via connector 3, causing the control surface 1 to rotate. The heat of the control surface 1 is transferred to the temperature-sensitive device 4 via connector 3. Therefore, connector 3 plays the role of transmitting force and thermal load, and is one of the objects to be cooled. It is worth noting that the connection method between connector 3 and control surface 1 can be welding, threaded connection, or snap-fit ​​connection, depending on the requirements.

[0035] The cooling assembly includes: a sweating cooling end 10 connected to the cooling channel 11, and a porous medium region 9, a cooling working fluid region 8 and a flexible diaphragm 7 disposed below the sweating cooling end 10; the flexible diaphragm 7 is connected to the thermal control box 5 through a pipe, and another pipe extends along the path of the pipe to the jet port 12 of the body 2.

[0036] Specifically, within the recess of the connector 3, from top to bottom, are a sweating cooling end 10, a porous medium region 9, a cooling working fluid region 8, and a flexible diaphragm 7. The sweating cooling end 10 has a plate-like structure and is generally made of a porous liquid-absorbing material, preferably a sintered porous metal material.

[0037] A porous medium is disposed within the porous medium region 9. This porous medium is typically a lightweight porous insulating felt, which serves as the carrier for storing the cooling medium 81. Preferably, the lightweight porous insulating felt has a porosity of 85-90% and a thermal conductivity of 0.02-0.06 kW / m².2 .

[0038] Cooling medium 81 is stored in cooling medium region 8. Cooling medium 81 is generally a liquid material capable of phase change, such as ammonia, hydrofluorocarbons, liquid water, etc., with liquid water being preferred. After absorbing heat, liquid water changes from a liquid phase to a vapor phase, thereby carrying away heat.

[0039] The flexible diaphragm 7 and the bottom surface of the concave cavity of the connector 3 form a sealed cavity. This sealed cavity is connected to the heat control box 5 only through a pipe and is completely isolated from the cooling medium region 8. When the cooling medium 81 in the heat control box 5 vaporizes, the resulting pressure exceeds the threshold of the pressure control device 6a. The pressure control device 6a opens, and the phase-change cooling medium 81 enters the sealed cavity formed by the flexible diaphragm 7 and the connector 3 through the pipe, causing the flexible diaphragm 7 to expand. This forces the cooling medium 81 in the cooling medium region 8 to penetrate into the porous medium in the porous medium region 9, and then enter the cooling channel 11 from the evaporating cooling end 10.

[0040] On the pipe connecting the flexible diaphragm 7 and the thermal control box 5, another pipe extends to the jet outlet 12. A pressure control element 6b is installed on this pipe. As the cooling medium 81 in the thermal control box 5 continues to vaporize and flow into the flexible diaphragm 7, the pressure inside the flexible diaphragm 7 gradually increases. When the pressure increases to a certain level, it exceeds the threshold of the pressure control element 6b, and the pressure control element 6b opens. The phase-change cooling medium 81 is then ejected from the jet outlet 12, thereby reducing drag and preventing heat at the leading edge.

[0041] During operation, the high-speed incoming flow 13 acts on the surface of the control surface 1, and is subjected to friction and obstruction. The aerodynamic heat is generated and accumulates on the control surface 1, causing the temperature of the control surface 1 to rise sharply. As the temperature of the control surface 1 rises, the cooling medium 81 in the thermal control box 5 absorbs heat and undergoes a phase change. The pressure difference caused by the vaporization expansion drives the flow of the medium. When the pressure reaches a certain level, the pressure control component 6a is opened, and the phase-changed cooling medium 81 flows through the pipe into the sealed cavity formed by the flexible diaphragm 7 and the connector 3, and continuously accumulates in the sealed cavity. Meanwhile, as a high heat flux zone, control surface 1 will transfer some heat to connector 3. In porous medium region 9, the cooling medium 81 stored in porous medium 91 absorbs heat and vaporizes. Under the action of vaporization pressure difference, steam and liquid cooling medium exchange heat in the porous medium 91 through convection, and the root of control surface 1 is cooled by sweating through sweating cooling end 10, which reduces the heat transferred from control surface 1 to connector 3 to a certain extent. The steam near sweating cooling end 10 flows to exhaust port 111 through cooling channel 11 and is discharged through exhaust port 111, thereby improving the high heat environment of the gap.

[0042] Furthermore, the volume of the sealed cavity formed by the flexible diaphragm 7 and the connector 3 continuously increases due to the entry of the cooling medium 81 after phase change, causing the flexible diaphragm 7 to expand. This, in turn, forces the cooling medium 81 in the cooling medium region 8 to permeate into the porous medium 91 above it. The porous medium 91 receives effective replenishment and further absorbs heat and vaporizes. The two-phase mixed cooling medium 81 continues to move towards the sweating cooling end 10 and flows from the exhaust port 111 to the high-heat gap region, continuously and effectively cooling the gap region.

[0043] Furthermore, when the volume of the sealed cavity formed by the flexible diaphragm 7 and the connector 3 increases to a certain extent, the pressure inside the cavity exceeds the threshold of the pressure control component 6b, the pressure control component 6b is opened, the pipeline is connected, and the cooling medium 81 in the flexible diaphragm area flows to the jet port 12 through the pipeline and is jetted at a position upstream of the control surface 1; the jet interacts with the incoming flow, destroying / pushing away the shock wave at the front edge of the control surface, thereby playing a role in reducing drag and preventing heat at the front edge.

[0044] Compared to existing technologies, the beneficial effects of this application include at least the following:

[0045] 1. Aerodynamic heat is introduced into the thermal management system as a power source, and the heat generated drives the phase change of the thermal control working fluid, thereby generating the pressure difference required for the working fluid flow and heat exchange.

[0046] 2. Lightweight porous insulation felt has strong hydrophilicity, which can adsorb and store the cooling working fluid, effectively avoiding the problem of excessive boiling caused by excessive heat flow at the gaps. At the same time, its lightweight and high thermal resistance characteristics effectively reduce heat conduction and reduce the overall weight.

[0047] 3. The thermal management system also features semi-active thermal control of connectors and gaps, as well as drag reduction and heat protection at the leading edge of the wing / rudder, effectively ensuring the safe and stable operating temperature of the servo mechanism.

[0048] 4. The thermal management system effectively utilizes the limited space within the servo system, achieving space optimization, and is a design method suitable for the thermal management system of high-speed aircraft.

[0049] The technical features of the above embodiments can be combined in any way. For the sake of brevity, not all possible combinations of the technical features in the above embodiments are described. However, as long as there is no contradiction in the combination of these technical features, they should be considered to be within the scope of this specification.

[0050] The embodiments described above are merely illustrative of several implementation methods of this application, and while the descriptions are specific and detailed, they should not be construed as limiting the scope of the invention. It should be noted that those skilled in the art can make various modifications and improvements without departing from the concept of this application, and these modifications and improvements all fall within the scope of protection of this application. Therefore, the scope of protection of this application should be determined by the appended claims.

Claims

1. A design method for a thermal management system of a high-speed aircraft based on aerodynamic heat utilization technology, characterized in that, The method described above is used to design a thermal management system for a high-speed aircraft. The system includes: a thermal control box disposed within the control surface, a cooling channel disposed at the bottom of the control surface, and a cooling assembly disposed within the fuselage. There is a gap between the bottom of the control surface and the body, and the cooling channel has an exhaust port facing the gap; One end of the cooling component is connected to the thermal control box via a pipe, and a pressure control device is installed on the pipe. The other end is connected to the cooling channel. The thermal control box contains a cooling medium, which absorbs heat from the control surface to undergo a phase change. A jet outlet is provided on the fuselage located upstream of the control surface. The jet outlet is connected to the cooling assembly through a pipe, and a pressure control component is provided on the pipe. The cooling assembly is fixed to the bottom of the control surface by a connector; the connector has a concave cavity structure, and the cooling assembly is placed in the concave cavity; The cooling assembly includes: a sweating cooling end connected to the cooling channel, and a porous medium region, a cooling working fluid region, and a flexible diaphragm disposed below the sweating cooling end; The flexible diaphragm and the connector form a sealed cavity, which is connected to the thermal control box via a pipe; Along the path of the aforementioned pipe, another pipe extends to the jet outlet of the fuselage.

2. The design method for a high-speed aircraft thermal management system based on aerodynamic heat utilization technology according to claim 1, characterized in that, A temperature-sensitive device is installed inside the body, and the connector is connected to the temperature-sensitive device.

3. A design method for a high-speed aircraft thermal management system based on aerodynamic heat utilization technology according to claim 1 or 2, characterized in that, The exhaust port is either a series of spaced holes or a groove.

4. A design method for a high-speed aircraft thermal management system based on aerodynamic heat utilization technology according to claim 1 or 2, characterized in that, The jet outlet is a single hole, or a plurality of holes arranged at intervals, or a groove.

5. A design method for a high-speed aircraft thermal management system based on aerodynamic heat utilization technology according to claim 1 or 2, characterized in that, The sweating cooling end has a plate-like structure and is made of sintered porous metal material.

6. A design method for a high-speed aircraft thermal management system based on aerodynamic heat utilization technology according to claim 1 or 2, characterized in that, The cooling channel is a series serpentine coil structure.

7. A design method for a high-speed aircraft thermal management system based on aerodynamic heat utilization technology according to claim 1 or 2, characterized in that, The porous medium region is provided with a lightweight porous thermal insulation felt.

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