Aircraft active and passive combined thermal protection system

Abstract

This invention discloses a combined active and passive thermal protection system for aircraft, comprising: an aerodynamic heat transfer component, an aircraft shell, a local active cooling component, a main cooling source, a heat exchange and transport network, a secondary cooling source, a bypass transport network, a local cooling transport network, and a measurement and control system. The aerodynamic heat transfer component is located at the front end of the aircraft shell; the main cooling source and the measurement and control system are located inside the aircraft shell; the main cooling source, the heat exchanger of the aerodynamic heat transfer component, and the secondary cooling source are connected through the heat exchange and transport network; a local active cooling component is located in the middle of the aircraft shell; the main cooling source and the local active cooling component are connected through the bypass transport network and the local cooling transport network; the secondary cooling source and the local active cooling component are connected through the local cooling transport network. The system of this invention employs a flexible combination of active and passive thermal protection in space and active and passive thermal protection in time, achieving optimal overall thermal protection performance for long-duration, high-speed near-space aircraft.

Description

Technical Field

[0001] This invention belongs to the field of aircraft thermal protection technology, and particularly relates to an active-passive combined thermal protection system for aircraft. Background Technology

[0002] When an aircraft moves at high speed in the dense atmosphere, the air is subjected to intense compression and friction. Most of the aircraft's kinetic energy is converted into heat, causing a rapid increase in the temperature of the surrounding air. A significant temperature difference is created between the hot air and the aircraft's surface, and some of the heat is rapidly transferred to the surface. This heating phenomenon caused by an object flying at high speed in the atmosphere is called aerodynamic heating, and its direct consequence is an increase in the temperature of the aircraft's surface and interior. When the temperature is too high, it can damage the aircraft's instruments, cause control malfunctions, and in severe cases, directly lead to structural burnout or even explosion. The key to solving this problem is to design a reasonable thermal protection system to ensure that the aircraft is not burned out.

[0003] Traditional thermal protection methods often employ a passive approach, relying on the material's own heat absorption, ablation, and insulation properties to achieve thermal protection of the aircraft's walls. For example, Chinese patent CN117049787A discloses a thermally protective metallic enamel composite coating with low infrared emissivity, and patent CN117024905A discloses a fiber preform thermal protection material. With the rise of a new wave of high-speed aircraft development, aerodynamic shapes, flight speeds, flight environments, and flight times have all undergone profound changes compared to traditional aircraft. Aircraft need to withstand prolonged aerodynamic thermal environments with high enthalpy and medium-low heat flux, while maintaining a high lift-to-drag ratio and a sharp leading-edge shape. This places extremely stringent demands on the temperature resistance, durability, structural efficiency, and reliability of thermal protection materials. Faced with the constraints of ensuring survival under extremely high heat flux density and ultra-high temperature conditions, while also meeting requirements for structural stability, long-term operation, lightweight design, and reusability, existing passive thermal protection technologies are struggling to meet these performance requirements. Active thermal protection has received increasing attention in recent years because it can handle high heat flux density for extended periods. For example, patent CN107914862B proposes using sweating cooling to provide active thermal protection for high heat flux areas such as the tip, wing leading edge, and rudder leading edge.

[0004] In summary, existing patents primarily propose single thermal protection methods, either passive or active. Passive thermal protection has poor heat flux handling capabilities but is lightweight and reliable. Active thermal protection, while possessing strong heat flux handling capabilities, suffers from drawbacks such as large weight and system complexity, limiting its application to high-heat-flux areas within the aircraft. Both active and passive thermal protection have their advantages and disadvantages. A combination of spatially integrated active and passive solutions, along with temporal switching between them, can achieve optimal overall performance. Currently, there are few publicly reported integrated solutions combining active and passive thermal protection. Summary of the Invention

[0005] The technical problem solved by this invention is to overcome the shortcomings of the prior art and provide a combined active and passive thermal protection system for aircraft. This system adopts a flexible combination of active and passive thermal protection in space and active and passive thermal protection in time, achieving the best overall thermal protection performance for long-duration, high-speed near-space aircraft.

[0006] To address the aforementioned technical problems, this invention discloses a combined active and passive thermal protection system for aircraft, comprising: an aerodynamic heat conduction component, an aircraft shell, a local active cooling component, a main cooling source, a heat exchange and transport network, a secondary cooling source, a bypass transport network, a local cooling transport network, and a measurement and control system.

[0007] The aerodynamic heat transfer component is located at the front end of the aircraft shell; the aerodynamic heat transfer component includes: a high-temperature metal heat pipe and a heat exchanger; the high-temperature metal heat pipe is divided into an evaporation zone and a condensation zone located at the end of the evaporation zone; the heat exchanger is located in the condensation zone of the high-temperature metal heat pipe, is integrally formed with the condensation zone, and shares the same interface;

[0008] Both the main cooling source and the secondary cooling source are located inside the aircraft hull; the main cooling source is located on the side closest to the aerodynamic heat transfer components; the main cooling source, heat exchanger, and secondary cooling source are connected through a heat exchange and transport pipeline network.

[0009] A local active cooling system is installed in the middle of the aircraft hull; the main cooling source and the local active cooling system are connected through a bypass transport network and a local cooling transport network; the secondary cooling source and the local active cooling system are connected through a local cooling transport network.

[0010] The measurement and control system is located inside the aircraft's hull.

[0011] In the aforementioned active-passive combined thermal protection system for aircraft, the local active cooling component has a porous structure, including: a porous region consisting of several holes, and a liquid collection cavity on the inner wall side of the porous region; wherein, the porous region is connected to the liquid collection cavity, and the end of the local cooling transport network is connected to the liquid collection cavity.

[0012] In the aforementioned combined active and passive thermal protection system for aircraft, the main cooling source includes: high-pressure gas cylinder A, electronic valve A, pipeline A, and storage tank A;

[0013] Storage tank A includes: gas expansion tank A, cooling working fluid tank A, and movable sealing plate A; wherein, gas expansion tank A and cooling working fluid tank A are installed in a docking manner; and movable sealing plate A is provided between gas expansion tank A and cooling working fluid tank A.

[0014] High-pressure gas cylinder A is connected to the input end of gas expansion tank A via pipeline A; wherein, electronic valve A is installed on pipeline A;

[0015] The output end of the cooling working fluid tank A is connected to the heat exchange and transportation pipeline network.

[0016] In the aforementioned active-passive combined thermal protection system for aircraft, the heat exchange and transport network includes: electronic valve B, pipeline B, electronic valve C, and pipeline C;

[0017] The output end of the cooling working fluid tank A is connected to the heat exchanger through pipes B and C, and then connected to the cooling working fluid tank B of the secondary cooling source after passing through the heat exchanger; wherein, electronic valve B and electronic valve C are respectively installed on pipes B and C.

[0018] In the aforementioned combined active and passive thermal protection system for aircraft, the secondary cooling source includes: high-pressure gas cylinder B, electronic valve D, storage tank B, electronic valve E, pipeline D, pipeline E, and pipeline F.

[0019] Storage tank B includes: gas expansion tank B, movable sealing plate B, and cooling working fluid tank B; wherein, gas expansion tank B and cooling working fluid tank B are installed in a docking manner; and movable sealing plate B is provided between gas expansion tank B and cooling working fluid tank B.

[0020] High-pressure gas cylinder B is connected to the input end of gas expansion tank B via pipeline E; an electronic valve D is installed on pipeline E.

[0021] Pipeline F is connected to gas expansion tank B; an electronic valve E is installed on pipeline F.

[0022] The output end of the cooling medium tank B is connected to the local cooling transport network through pipe D.

[0023] In the aforementioned combined active and passive thermal protection system for aircraft, the bypass transport network includes: pipeline G, electronic valve F, electronic valve G, and pipeline H;

[0024] One end of pipe G is connected to the cooling medium tank A, and the other end is connected to the local cooling transport network; an electronic valve F is installed on pipe G.

[0025] One end of pipe H is connected to the cooling medium tank A, and the other end is connected to the local cooling transport network; an electronic valve G is installed on pipe H.

[0026] In the aforementioned combined active and passive thermal protection system for aircraft, the local cooling transport network includes: pipe I, electronic valve H, electronic valve I, and pipe J;

[0027] The end of pipe I is connected to the liquid collection chamber of the local active cooling component; an electronic valve H is installed on pipe I.

[0028] The end of pipe J is connected to the liquid collection chamber of the local active cooling component; wherein, an electronic valve I is installed on pipe J.

[0029] In the aforementioned combined active and passive thermal protection system for aircraft, the measurement and control system includes: a data acquisition unit, a controller, a temperature sensor for the end area, a temperature sensor for the local cooling area, and a liquid level sensor.

[0030] The end-area temperature sensor is embedded within the evaporation zone;

[0031] The temperature sensor for the localized cooling zone is embedded in a porous area.

[0032] The liquid level sensor is installed inside the cooling medium tank B;

[0033] The data acquisition unit is connected to the end area temperature sensor, the local cooling area temperature sensor, and the liquid level sensor via cables.

[0034] The data acquisition unit is connected to the controller via RS-485 communication.

[0035] The controller is connected to each electronic valve via cable.

[0036] In the aforementioned combined active and passive thermal protection system for aircraft

[0037] The data acquisition unit is used to collect the temperature measurement value T from the temperature sensor in the end area. a Temperature measurement value T of the local cooling area temperature sensor b The liquid level measurement value H from the liquid level sensor; the temperature measurement value T. a Temperature measurement value T b The liquid level measurement value H is transmitted to the controller;

[0038] The controller is used to determine the temperature measurement value T based on the received temperature measurement value. a Temperature measurement value T b The liquid level measurement value H controls the opening and closing of each electronic valve:

[0039] Step 1, Initial state: All electronic valves are in the closed state;

[0040] Step 2, determine T a Is it greater than the set upper limit temperature T at the end? a—max If T a Greater than T a—max If yes, proceed to step 3; otherwise, proceed to step 5.

[0041] Step 3, determine T b Is it greater than the set local cooling temperature limit T? b—max If T b Greater than T b—max If yes, proceed to step 4; otherwise, proceed to step 8.

[0042] Step 4: Open electronic valves A, B, C, H, and I, then return to step 2;

[0043] Step 5, determine T b Is it greater than the set local cooling temperature limit T? b—max If T b Greater than T b—max If yes, proceed to step 6; otherwise, return to step 2.

[0044] Step 6: Determine if H is greater than the set upper limit of the liquid level H. max If H is greater than H max If yes, proceed to step 7; otherwise, proceed to step 9.

[0045] Step 7: Open electronic valves D, H, and I, then return to step 2;

[0046] Step 8: Open electronic valves A, B, C, and E, then return to step 1;

[0047] Step 9: Open electronic valves A, F, G, H, and I, then return to step 2.

[0048] In the aforementioned combined active and passive thermal protection system for aircraft

[0049] When an aircraft flies at high speed, due to the intense compression of air and the fierce friction between air and the wall, the aerodynamic heat generated at the end of the aircraft is transported from the evaporation zone to the condensation zone through high-temperature metal heat pipes; the heat in the condensation zone is transferred to the external environment through radiation and / or to the cooling medium through a heat exchanger.

[0050] The high-pressure gas cylinder A of the main cold source transfers gas to the gas expansion tank A; after the gas expands, it pushes the moving sealing plate A to move. Under the push of the moving sealing plate A, the cooling medium in the cooling medium tank A is transported to the heat exchange and transport network. After exchanging heat with the condensation zone of the high-temperature metal heat pipe through the heat exchanger, it is transported to the cooling medium tank B of the secondary cold source; and / or, it is transported to the local cooling transport network through the bypass transport network; wherein, when the heat exchange and transport network transports cooling medium to the cooling medium tank B of the secondary cold source, the gas expansion tank B opens the electronic valve E to exhaust gas;

[0051] The high-pressure gas cylinder B, the secondary cooling source, transfers gas to the gas expansion chamber B. After the gas expands, it pushes the moving sealing plate B to move. Under the push of the moving sealing plate B, the cooling medium enters the local cooling transport network, then enters the liquid collection chamber of the local active cooling component, and then enters the porous area. After sufficient heat exchange with the porous area, it flows out of the aircraft wall and spreads into a gas film in the downstream direction, continuing to block the transfer of aerodynamic heat to the aircraft wall.

[0052] The present invention has the following advantages:

[0053] (1) This invention discloses a combined active and passive thermal protection system for aircraft, which combines passive thermal protection with active thermal protection and can be flexibly switched according to actual usage requirements. It can not only meet a series of constraints for future aerospace vehicles such as the stability of the thermal protection structure prototype, long-term operation, light weight and reusability, but also achieve the best overall performance.

[0054] (2) The present invention discloses an active and passive combined thermal protection system for aircraft. The local active cooling components can be arranged in multiple areas according to actual usage requirements. The system has good scalability and flexible arrangement. Attached Figure Description

[0055] Figure 1 This is a schematic diagram of the composition of an active-passive combined thermal protection system for an aircraft according to an embodiment of the present invention;

[0056] Figure 2 This is a flowchart of a control strategy in an embodiment of the present invention. Detailed Implementation

[0057] To make the objectives, technical solutions, and advantages of the present invention clearer, the embodiments disclosed in the present invention will be described in further detail below with reference to the accompanying drawings.

[0058] like Figure 1 In this embodiment, the aircraft's active-passive combined thermal protection system includes: an aerodynamic heat transfer component 1, an aircraft shell 2, a local active cooling component 3, a main cooling source 4, a heat exchange and transport network 5, a secondary cooling source 6, a bypass transport network 7, a local cooling transport network 8, and a measurement and control system 9. The specific connections are as follows:

[0059] A pneumatic heat transfer component 1 is located at the front end of the aircraft shell 2. The pneumatic heat transfer component 1 includes a high-temperature metal heat pipe 11 and a heat exchanger 12. The high-temperature metal heat pipe 11 is divided into an evaporation zone 111 and a condensation zone 112 located at the end of the evaporation zone 111. The heat exchanger 12 is located in the condensation zone 112 of the high-temperature metal heat pipe 11, is integrally formed with the condensation zone 112, and shares a common interface.

[0060] Both the main cooling source 4 and the secondary cooling source 6 are located inside the aircraft shell 2; the main cooling source 4 is located on the side closer to the aerodynamic heat transfer component 1; the main cooling source 4, the heat exchanger 12 and the secondary cooling source 6 are connected through the heat exchange and transport pipeline network 5.

[0061] A local active cooling component 3 is provided in the middle of the aircraft hull 2; the main cooling source 4 is connected to the local active cooling component 3 through the bypass transport network 7 and the local cooling transport network 8; the secondary cooling source 6 is connected to the local active cooling component 3 through the local cooling transport network 8.

[0062] The measurement and control system 9 is located inside the aircraft hull 2.

[0063] In this embodiment, the local active cooling component 3 has a porous structure, including: a porous region 31 composed of a plurality of holes, and a liquid collection cavity 32 on the inner wall side of the porous region 31; wherein, the porous region 31 is connected to the liquid collection cavity 32, and the end of the local cooling transport pipeline 8 is connected to the liquid collection cavity 32.

[0064] In this embodiment, the main cooling source 4 may specifically include: a high-pressure gas cylinder A41, an electronic valve A42, a pipeline A43, and a storage tank A44. Further, the storage tank A44 may specifically include: a gas expansion tank A441, a cooling medium tank A442, and a movable sealing plate A443. The gas expansion tank A441 and the cooling medium tank A442 are connected in a mating manner; a movable sealing plate A443 is provided between the gas expansion tank A441 and the cooling medium tank A442. The high-pressure gas cylinder A41 is connected to the input end of the gas expansion tank A441 via the pipeline A43; the pipeline A43 is equipped with the electronic valve A42. The output end of the cooling medium tank A442 is connected to the heat exchange and transport pipeline network 5.

[0065] In this embodiment, the heat exchange and transport pipeline network 5 may specifically include: electronic valve B51, pipeline B52, electronic valve C53, and pipeline C54. The output end of the cooling medium tank A442 is connected to the heat exchanger 12 via pipelines B52 and C54, and then connected to the cooling medium tank B633 of the secondary cooling source 6 via the heat exchanger 12. Electronic valves B51 and C53 are respectively installed on pipelines B52 and C54.

[0066] In this embodiment, the secondary cooling source 6 may specifically include: a high-pressure gas cylinder B61, an electronic valve D62, a storage tank B63, an electronic valve E64, a pipeline D65, a pipeline E66, and a pipeline F67. Further, the storage tank B63 may specifically include: a gas expansion tank B631, a movable sealing plate B632, and a cooling medium tank B633. The gas expansion tank B631 and the cooling medium tank B633 are connected in a mating manner; a movable sealing plate B632 is provided between the gas expansion tank B631 and the cooling medium tank B633. The high-pressure gas cylinder B61 is connected to the input end of the gas expansion tank B631 via pipeline E66; an electronic valve D62 is installed on pipeline E66. Pipeline F67 is connected to the gas expansion tank B631; an electronic valve E64 is installed on pipeline F67. The output end of the cooling medium tank B633 is connected to the local cooling transport network 8 via pipeline D65.

[0067] In this embodiment, the bypass transport network 7 may specifically include: pipeline G71, electronic valve F72, electronic valve G73, and pipeline H74. One end of pipeline G71 is connected to the cooling medium tank A442, and the other end is connected to the local cooling transport network 8; electronic valve F72 is installed on pipeline G71. One end of pipeline H74 is connected to the cooling medium tank A442, and the other end is connected to the local cooling transport network 8; electronic valve G73 is installed on pipeline H74.

[0068] In this embodiment, the local cooling transport network 8 may specifically include: pipe I81, electronic valve H82, electronic valve I83, and pipe J84. Pipe I81 is connected at its end to the liquid collection chamber 32 of the local active cooling component 3; electronic valve H82 is installed on pipe I81. Pipe J84 is also connected at its end to the liquid collection chamber 32 of the local active cooling component 3; electronic valve I83 is installed on pipe J84.

[0069] In this embodiment, the measurement and control system 9 may specifically include: a data acquisition unit 91, a controller 92, an end-area temperature sensor 93, a local cooling area temperature sensor 94, and a liquid level sensor 95. The end-area temperature sensor 93 is embedded within the evaporation zone 111; the local cooling area temperature sensor 94 is embedded within the porous area 31; and the liquid level sensor 95 is disposed within the cooling medium tank B633. The data acquisition unit 91 is connected to the end-area temperature sensor 93, the local cooling area temperature sensor 94, and the liquid level sensor 95 via cables. The data acquisition unit 91 is connected to the controller 92 via RS-485 communication. The controller 92 is connected to each electronic valve via cables.

[0070] Preferably, the data acquisition unit 91 is used to acquire the temperature measurement value T of the end area temperature sensor 93. a Temperature measurement value T of the local cooling area temperature sensor 94 b The liquid level sensor 95 measures the liquid level H; the temperature measurement value T is used. a Temperature measurement value T b The liquid level measurement value H is transmitted to the controller 92. The controller 92 is used to transmit the received temperature measurement value T according to the liquid level measurement value H. a Temperature measurement value T b The liquid level measurement value H controls the opening and closing of each electronic valve. For example... Figure 2 As shown, the specific control flow is as follows:

[0071] Step 1, initial state: all electronic valves are closed.

[0072] Step 2, determine T a Is it greater than the set upper limit temperature T at the end? a—max If Ta Greater than T a—max If yes, proceed to step 3; otherwise, proceed to step 5.

[0073] Step 3, determine T b Is it greater than the set local cooling temperature limit T? b—max If T b Greater than T b—max If yes, proceed to step 4; otherwise, proceed to step 8.

[0074] Step 4: Open electronic valves A42, B51, C53, H82, and I83, then return to step 2.

[0075] Step 5, determine T b Is it greater than the set local cooling temperature limit T? b—max If T b Greater than T b—max If yes, proceed to step 6; otherwise, return to step 2.

[0076] Step 6: Determine if H is greater than the set upper limit of the liquid level H. max If H is greater than H max If yes, proceed to step 7; otherwise, proceed to step 9.

[0077] Step 7: Open electronic valves D62, H82, and I83, then return to step 2.

[0078] Step 8: Open electronic valves A42, B51, C53, and E64, then return to step 1.

[0079] Step 9: Open electronic valves A42, F72, G73, H82, and I83, then return to step 2.

[0080] In this embodiment, the working principle of the aircraft's active-passive combined thermal protection system is as follows:

[0081] When the aircraft is flying at high speed, due to the intense compression of the air and the intense friction between the air and the wall, the aerodynamic heat generated at the end of the aircraft is transported from the evaporation zone 111 to the condensation zone 112 through the high-temperature metal heat pipe 11; wherein, the heat of the condensation zone 112 is transferred to the external environment through radiation and / or transferred to the cooling medium through the heat exchanger 12.

[0082] The high-pressure gas cylinder A41 of the main cooling source 4 transfers gas to the gas expansion tank A441. After the gas expands, it pushes the moving sealing plate A443 to move. Under the push of the moving sealing plate A443, the cooling medium in the cooling medium tank A442 is transported to the heat exchange and transport network 5. After exchanging heat with the condensation zone 112 of the high-temperature metal heat pipe 11 through the heat exchanger 12, it is transported to the cooling medium tank B633 of the secondary cooling source 6. And / or, it is transported to the local cooling transport network 8 through the bypass transport network 7. When the heat exchange and transport network 5 transfers the cooling medium to the cooling medium tank B633 of the secondary cooling source 6, the gas expansion tank B631 opens the electronic valve E64 to exhaust gas.

[0083] The high-pressure gas cylinder B61 of the secondary cooling source 6 transmits gas to the gas expansion chamber B631. After the gas expands, it pushes the moving sealing plate B632 to move. Under the push of the moving sealing plate B632, the cooling working fluid enters the local cooling transport network 8, then enters the liquid collection chamber 32 of the local active cooling component 3, and then enters the porous region 31. After sufficient heat exchange with the porous region 31, it flows out of the aircraft wall and spreads into a layer of gas film in the downstream direction, continuing to block the transfer of aerodynamic heat to the aircraft wall.

[0084] In this embodiment, the gas in the gas expansion chamber can be helium. The cooling medium in the cooling medium chamber can be water. The working medium of the high-temperature metal heat pipe can be sodium. The porous medium material in the porous region can be titanium alloy TC4, with a porosity of 50μm to 100μm. The local active cooling components are suitable for placement in high heat flux regions such as the wing leading edge and rudder leading edge.

[0085] Although the present invention has been disclosed above with reference to preferred embodiments, it is not intended to limit the present invention. Any person skilled in the art can make possible changes and modifications to the technical solutions of the present invention by utilizing the methods and techniques disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and alterations made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solutions of the present invention shall fall within the protection scope of the technical solutions of the present invention.

[0086] The contents not described in detail in this specification are common knowledge to those skilled in the art.

Claims

1. A combined active and passive thermal protection system for aircraft, characterized in that, include: Aerodynamic heat transfer components (1), aircraft shell (2), local active cooling components (3), main cooling source (4), heat exchange and transport network (5), secondary cooling source (6), bypass transport network (7), local cooling transport network (8), and measurement and control system (9); A pneumatic heat transfer component (1) is installed at the front end of the aircraft shell (2); wherein, the pneumatic heat transfer component (1) includes: a high-temperature metal heat pipe (11) and a heat exchanger (12); the high-temperature metal heat pipe (11) is divided into an evaporation zone (111) and a condensation zone (112) located at the end of the evaporation zone (111); the heat exchanger (12) is located in the condensation zone (112) of the high-temperature metal heat pipe (11), and is integrally formed with the condensation zone (112), and the interface is shared; The main cooling source (4) and the secondary cooling source (6) are both located inside the aircraft shell (2); the main cooling source (4) is located on the side close to the aerodynamic heat transfer component (1); the main cooling source (4), the heat exchanger (12) and the secondary cooling source (6) are connected through the heat exchange transport network (5); A local active cooling component (3) is provided in the middle of the aircraft shell (2); the main cooling source (4) and the local active cooling component (3) are connected through the bypass transport network (7) and the local cooling transport network (8); the secondary cooling source (6) and the local active cooling component (3) are connected through the local cooling transport network (8); The measurement and control system (9) is installed inside the aircraft hull (2); The local active cooling component (3) has a porous structure, including: a porous region (31) consisting of several holes, and a liquid collection chamber (32) on the inner wall side of the porous region (31); wherein, the porous region (31) is connected to the liquid collection chamber (32), and the end of the local cooling transport pipeline (8) is connected to the liquid collection chamber (32); When the aircraft is flying at high speed, due to the strong compression of the air and the intense friction between the air and the wall, the aerodynamic heat generated at the end of the aircraft is transported from the evaporation zone (111) to the condensation zone (112) through the high-temperature metal heat pipe (11); wherein, the heat of the condensation zone (112) is transferred to the external environment through radiation and / or transferred to the cooling medium through the heat exchanger (12); The high-pressure gas cylinder A (41) of the main cold source (4) transmits gas to the gas expansion tank A (441); after the gas expands, it pushes the moving sealing plate A (443) to move. Under the push of the moving sealing plate A (443), the cooling medium in the cooling medium tank A (442) is transported to the heat exchange transport network (5), and after heat exchange with the condensation zone (112) of the high-temperature metal heat pipe (11) through the heat exchanger (12), it is transported to the cooling medium tank B (633) of the secondary cold source (6); and / or, it is transported to the local cooling transport network (8) through the bypass transport network (7); wherein, when the heat exchange transport network (5) transmits the cooling medium to the cooling medium tank B (633) of the secondary cold source (6), the gas expansion tank B (631) opens the electronic valve E (64) to exhaust gas; The high-pressure gas cylinder B (61) of the secondary cooling source (6) transmits gas to the gas expansion box B (631). After the gas expands, it pushes the moving sealing plate B (632) to move. Under the push of the moving sealing plate B (632), the cooling working medium enters the local cooling transport network (8), then enters the liquid collection chamber (32) of the local active cooling component (3), and then enters the porous area (31). After sufficient heat exchange with the porous area (31), it flows out of the aircraft wall and spreads into a layer of gas film in the downstream direction, continuing to block the transfer of aerodynamic heat to the aircraft wall.

2. The aircraft active-passive combined thermal protection system according to claim 1, characterized in that, The main cold source (4) includes: high-pressure gas cylinder A (41), electronic valve A (42), pipeline A (43) and storage tank A (44). Storage tank A (44) includes: gas expansion tank A (441), cooling working medium tank A (442) and movable sealing plate A (443); wherein, gas expansion tank A (441) and cooling working medium tank A (442) are installed in a docking manner; a movable sealing plate A (443) is provided between gas expansion tank A (441) and cooling working medium tank A (442). High-pressure gas cylinder A (41) is connected to the input end of gas expansion tank A (441) through pipeline A (43); wherein, an electronic valve A (42) is installed on pipeline A (43). The output end of the cooling working medium tank A (442) is connected to the heat exchange and transport pipeline (5).

3. The aircraft active-passive combined thermal protection system according to claim 2, characterized in that, The heat exchange and transport network (5) includes: electronic valve B (51), pipeline B (52), electronic valve C (53) and pipeline C (54); The output end of the cooling working medium tank A (442) is connected to the heat exchanger (12) through pipes B (52) and C (54), and then connected to the cooling working medium tank B (633) of the secondary cooling source (6) through the heat exchanger (12); wherein, electronic valve B (51) and electronic valve C (53) are respectively installed on pipes B (52) and C (54).

4. The aircraft active-passive combined thermal protection system according to claim 3, characterized in that, The secondary cooling source (6) includes: high-pressure gas cylinder B (61), electronic valve D (62), storage tank B (63), electronic valve E (64), pipeline D (65), pipeline E (66) and pipeline F (67). Storage tank B (63) includes: gas expansion tank B (631), movable sealing plate B (632) and cooling working medium tank B (633); wherein, gas expansion tank B (631) and cooling working medium tank B (633) are installed in a docking manner; movable sealing plate B (632) is provided between gas expansion tank B (631) and cooling working medium tank B (633). High-pressure gas cylinder B (61) is connected to the input end of gas expansion tank B (631) through pipeline E (66); wherein, an electronic valve D (62) is installed on pipeline E (66). Pipeline F (67) is connected to gas expansion tank B (631); wherein, electronic valve E (64) is installed on pipeline F (67). The output end of the cooling working medium tank B (633) is connected to the local cooling transport network (8) through the pipeline D (65).

5. The aircraft active-passive combined thermal protection system according to claim 4, characterized in that, The bypass transport network (7) includes: pipeline G (71), electronic valve F (72), electronic valve G (73) and pipeline H (74); One end of the pipeline G (71) is connected to the cooling medium tank A (442), and the other end is connected to the local cooling transport network (8); wherein, an electronic valve F (72) is installed on the pipeline G (71). One end of the pipe H (74) is connected to the cooling working medium tank A (442), and the other end is connected to the local cooling transport network (8); wherein, an electronic valve G (73) is installed on the pipe H (74).

6. The aircraft active-passive combined thermal protection system according to claim 5, characterized in that, The local cooling transport network (8) includes: pipe I (81), electronic valve H (82), electronic valve I (83) and pipe J (84). The end of pipe I (81) is connected to the liquid collection chamber (32) of the local active cooling component (3); wherein, an electronic valve H (82) is provided on pipe I (81); The end of pipe J (84) is connected to the liquid collection chamber (32) of the local active cooling component (3); wherein, an electronic valve I (83) is provided on pipe J (84).

7. The aircraft active-passive combined thermal protection system according to claim 6, characterized in that, The measurement and control system (9) includes: a data acquisition unit (91), a controller (92), an end area temperature sensor (93), a local cooling area temperature sensor (94), and a liquid level sensor (95). The end area temperature sensor (93) is embedded in the evaporation zone (111); The local cooling area temperature sensor (94) is embedded in the porous area (31); The liquid level sensor (95) is installed inside the cooling medium tank B (633); The data acquisition unit (91) is connected to the end area temperature sensor (93), the local cooling area temperature sensor (94) and the liquid level sensor (95) via cables; The data acquisition unit (91) is connected to the controller (92) via RS-485 communication. The controller (92) is connected to each electronic valve via cable.

8. The aircraft active-passive combined thermal protection system according to claim 7, characterized in that, The data acquisition unit (91) is used to acquire the temperature measurement value of the end area temperature sensor (93). Temperature measurement value of local cooling area temperature sensor (94) Liquid level measurement value of liquid level sensor (95) H ; temperature measurement value Temperature measurement value and liquid level measurement value Transmitted to the controller (92); Controller (92) for using the received temperature measurement value Temperature measurement value and liquid level measurement value Controls the opening and closing of each electronic valve: Step 1, Initial state: All electronic valves are in the closed state; Step 2, Determine Is it greater than the set upper limit of the end temperature? ;like Greater than If yes, proceed to step 3; otherwise, proceed to step 5. Step 3, Determine Is it greater than the set local cooling temperature limit? ,like Greater than If yes, proceed to step 4; otherwise, proceed to step 8. Step 4: Open electronic valve A (42), electronic valve B (51), electronic valve C (53), electronic valve H (82), and electronic valve I (83), then return to step 2; Step 5, Determine Is it greater than the set local cooling temperature limit? ,like Greater than If yes, proceed to step 6; otherwise, return to step 2. Step 6, Determine Is it greater than the set upper limit of liquid level? ,like Greater than If yes, proceed to step 7; otherwise, proceed to step 9. Step 7: Open electronic valve D (62), electronic valve H (82), and electronic valve I (83), then return to step 2; Step 8: Open electronic valve A (42), electronic valve B (51), electronic valve C (53), and electronic valve E (64), then return to step 1; Step 9: Open electronic valve A (42), electronic valve F (72), electronic valve G (73), electronic valve H (82), and electronic valve I (83), then return to step 2.

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