A hypersonic vehicle head thermal protection structure

Abstract

The application discloses a hypersonic aircraft head thermal protection structure, and belongs to the technical field of hot end component cooling. The hypersonic aircraft head thermal protection structure comprises a cooling system, a preheating system, a CBC power generation system and a secondary cooler. The cooling system comprises a coolant storage tank, an electromagnetic adjustable pump and a cooling channel arranged in the wall surface of the aircraft head. The cooling channel is provided with a cooling medium which flows reversely to the high-temperature gas outside. The preheating system comprises a fuel tank, a fuel adjustable pump and a fuel channel. The secondary cooler is used for transferring heat in the cooling channel to the fuel channel. The CBC power generation system comprises a heater, a primary cooler and a circulating working medium. The heater is used for transferring heat in the cooling channel to the circulating working medium. The primary cooler is used for transferring heat in the circulating working medium to the fuel channel. The application can realize efficient thermal protection of the head of the hypersonic aircraft, reduce the risk of structural failure caused by high temperature, and improve the safety of the aircraft.

Description

Technical Field

[0001] The present application relates to the technical field of hot end component cooling, and in particular to a supersonic aircraft head thermal protection structure. Background Art

[0002] Hypersonic vehicles refer to winged or wingless aircraft, such as aircraft, missiles, and artillery shells, that fly at speeds exceeding Mach 5. These vehicles, when performing missions in near-space, possess both the advantages of aviation technology and the incomparable strengths of spacecraft. They can cruise at hypersonic speeds within the atmosphere and penetrate the atmosphere for re-entry.

[0003] Hypersonic aircraft will encounter extreme high temperature environments during flight, especially in the head area. Due to the aerodynamic heating effect, the temperature can rise to extremely high values ​​and may be subjected to high thermal loads, which can cause component degradation and failure. Summary of the Invention

[0004] In order to improve the above problems, the present application provides a supersonic aircraft head thermal protection structure.

[0005] The present application provides a supersonic aircraft head thermal protection structure adopting the following technical solutions:

[0006] A thermal protection structure for the head of a supersonic aircraft includes a cooling system, a preheating system, a CBC power generation system and a secondary cooler. The cooling system includes a coolant storage tank, an electromagnetic adjustable pump and a cooling channel arranged in the wall of the aircraft head. A cooling medium that flows in the opposite direction to the external high-temperature gas is provided in the cooling channel. The coolant storage tank is used to store the cooling medium, and the electromagnetic adjustable pump is used to control the flow rate of the cooling medium in the cooling channel; the preheating system includes a fuel tank, an adjustable fuel pump and a fuel channel, the adjustable fuel pump is used to control the amount of fuel in the fuel channel, and the secondary cooler is used to transfer heat in the cooling channel to the fuel channel; the CBC power generation system includes a heater, a primary cooler and a circulating working fluid. The heater is used to transfer heat in the cooling channel to the circulating working fluid, and the primary cooler is used to transfer heat in the circulating working fluid to the fuel channel.

[0007] Optionally, a temperature sensor is further included, which is arranged in the aircraft air inlet duct. The temperature sensor is used to collect a temperature detection signal in the aircraft air inlet duct. The electromagnetic adjustable pump controls the flow rate of the cooling medium in the cooling channel in response to the temperature detection signal, and the fuel adjustable pump controls the amount of fuel in the fuel channel in response to the temperature detection signal.

[0008] Optionally, the cooling medium is liquid metal.

[0009] Optionally, the material of the cooling channel is a high-temperature alloy, the inner wall of the cooling channel is coated with an anti-corrosion coating, and the cross-sectional shape of the cooling channel is circular or polygonal.

[0010] Optionally, the cross-sectional area of ​​the cooling channel close to the leading edge of the aircraft nose is smaller than the cross-sectional area of ​​the cooling channel away from the leading edge of the aircraft nose.

[0011] Optionally, the cooling channel includes a mainstream channel and several diversion channels, the mainstream channel is arranged along the busbar direction of the aircraft head, the diversion channel is arranged along the circumferential side of the aircraft head, the inlet end of the diversion channel and the outlet end of the diversion channel are both connected to the mainstream channel, the inlet end of the diversion channel is located on the side of the outlet end of the diversion channel away from the leading edge of the aircraft head, and the outlet end of the diversion channel and the inlet end of the diversion channel are both provided with a one-way valve.

[0012] Optionally, the cooling channel is arranged in a spiral shape along the circumference of the aircraft head, the inlet of the cooling channel is located at the trailing edge of the aircraft head, and the outlet of the cooling channel is located at the leading edge of the aircraft head.

[0013] Optionally, the cooling channel includes a plurality of annular channels, which are arranged along the circumferential side of the aircraft head, and the annular channel includes two semi-annular channels, the inlet end of the semi-annular channel is located on the side of the outlet end of the semi-annular channel away from the leading edge of the aircraft head, the inlet ends of the two semi-annular channels are connected, the outlet ends of the two semi-annular channels are connected, and the outlet end of the semi-annular channel away from the leading edge of the aircraft head is connected to the inlet end of the semi-annular channel close to the leading edge of the aircraft head.

[0014] Optionally, the cooling channel includes a feed channel, a discharge channel and several circular channels, the feed channel and the discharge channel are both arranged along the busbar direction of the aircraft head, the circular channel is arranged along the circumferential side of the aircraft head, the circular channel includes two semicircular channels, the inlet end of the semicircular channel is located on the side of the outlet end of the semicircular channel away from the leading edge of the aircraft head, the inlet end of the semicircular channel is connected to the feed channel, the outlet end of the semicircular channel is connected to the feed channel, and the inlet end of the semicircular channel and the outlet end of the semicircular channel are both provided with a one-way valve 2.

[0015] Optionally, it also includes a material delivery channel, which is arranged along the busbar direction of the aircraft head, the outlet end is connected to the inlet end of the semicircular channel located at the leading edge of the aircraft head, the outlet end of the material delivery channel is connected to the inlet end of the feed channel, and the feed channel is closed at the end located at the trailing edge of the aircraft.

[0016] In summary, the present application includes at least one of the following beneficial technical effects:

[0017] 1. During the flight of a hypersonic aircraft, the electromagnetic adjustable pump is started, and the liquid metal is pumped into the cooling channel at the head of the aircraft, achieving efficient thermal protection of the head of the hypersonic aircraft, reducing the risk of structural failure caused by high temperature, and improving the safety of the aircraft; the liquid metal can transfer the heat absorbed by the liquid metal to the fuel in the fuel channel through the secondary cooler, and the fuel directly injected into the combustion chamber after completing heat absorption can improve the temperature of the fuel, thereby improving the combustion efficiency and thrust, and further increasing the flight Mach number of the aircraft;

[0018] 2. The cooling channel is set as a main flow channel and a sub-flow channel, realizing the setting of single inlet and multiple outlets of the cooling medium, which can shorten the flow distance of the cooling medium, reduce the flow resistance of the cooling medium and the energy required for pumping, and especially reduce the problem of poor heat carrying caused by low specific heat of the cooling medium, and the multiple outlet design can increase the contact area between the cooling medium and the cooling surface, and improve the heat exchange efficiency;

[0019] 3. According to the temperature detection signal detected by the temperature sensor of the aircraft inlet, the flow of the cooling medium is accurately adjusted by the adjustable electromagnetic pump, which can adjust the cooling intensity according to the actual thermal load demand of the aircraft at different flight altitudes, save energy and improve system efficiency, and the fuel adjustable pump adjusts the amount of fuel, which works with the cooling system to enhance the aircraft's management ability of thermal energy and improve the utilization rate of thermal energy. BRIEF DESCRIPTION OF DRAWINGS

[0020] Figure 1 is a schematic diagram of the overall structure of embodiment 1 of the present application.

[0021] Figure 2 is a schematic diagram of the structure of the cooling pipe in embodiment 1 of the present application.

[0022] Figure 3 is a schematic diagram of the structure of the cooling pipe in embodiment 2 of the present application.

[0023] Figure 4 is a schematic diagram of the structure of the cooling pipe in embodiment 3 of the present application.

[0024] Figure 5 is a schematic diagram of the structure of the cooling pipe in embodiment 4 of the present application.

[0025] Figure 6 is a schematic diagram of the structure of the cooling pipe in embodiment 5 of the present application.

[0026] Figure numerals: 1. Cooling system; 11. Coolant storage tank; 111. Separator; 112. Collector; 12. Electromagnetic adjustable pump; 13. Cooling channel; 2. Preheating system; 21. Fuel tank; 22. Fuel adjustable pump; 23. Fuel channel; 3. Secondary cooler; 4. Main channel; 41. Diversion channel; 42. One-way valve 1; 5. Annular channel; 51. Semi-annular channel; 6. Feed channel; 61. Discharge channel; 62. Circular channel; 621. Semi-circular channel; 622. One-way valve 2; 7. Feed channel; 8. CBC power generation system; 81. Heater; 82. Primary cooler. DETAILED DESCRIPTION

[0027] The following is combined with Figure 1-6 This application is described in further detail.

[0028] The embodiment of the present application discloses a supersonic aircraft head thermal protection structure.

[0029] Example 1

[0030] Reference Figure 1 A supersonic vehicle nose thermal protection structure includes a cooling system 1, a preheating system 2, a CBC power generation system 8, a secondary cooler 3, and a temperature sensor. Cooling system 1 comprises a coolant reservoir 11, an electromagnetic adjustable pump 12, and a cooling channel 13 disposed within the vehicle's nose wall. Cooling channel 13 contains a coolant medium that flows countercurrently to the external high-temperature gas, absorbing heat through forced convection, thereby reducing the temperature of the vehicle's nose wall.

[0031] Reference Figure 1 The cooling medium is liquid metal. Its high thermal conductivity and specific heat capacity can effectively improve aircraft performance. A coolant reservoir 11 stores the cooling medium, while an electromagnetic adjustable pump 12 controls the flow of the cooling medium within the cooling channel 13.

[0032] Reference Figure 1 Cooling channel 13 is made of a high-temperature alloy, such as a nickel-based or cobalt-based alloy. This high-temperature alloy improves the heat resistance and structural strength of cooling channel 13, making it suitable for operation in extremely high-temperature environments and extending the service life of the thermal protection structure. The inner wall of cooling channel 13 is coated with an anti-corrosion coating, such as an oxide coating or a silicon nitride coating. This coating prevents corrosion of the inner wall of cooling channel 13 by the cooling medium, enhancing the durability and reliability of cooling channel 13.

[0033] Preferably, the cross-sectional shape of cooling channel 13 is circular or polygonal, and the cross-sectional area of ​​cooling channel 13 near the leading edge of the aircraft nose is smaller than the cross-sectional area of ​​cooling channel 13 away from the leading edge of the aircraft nose. The leading edge of the aircraft nose bears a higher heat load. By changing the cross-sectional area of ​​cooling channel 13, the flow rate of the cooling medium can be increased, thereby improving cooling efficiency and ensuring effective thermal protection for the leading edge of the aircraft nose.

[0034] Reference Figure 1 The preheating system 2 includes a fuel tank 21, an adjustable fuel pump 22 and a fuel channel 23. The adjustable fuel pump 22 is used to control the amount of fuel in the fuel channel 23. The secondary cooler 3 is used to transfer the heat in the cooling channel 13 to the fuel channel 23.

[0035] Reference Figure 1 The CBC power generation system 8 includes a heater 81, a primary cooler 82, and a circulating working fluid. The circulating working fluid in the CBC power generation system 8 is supercritical CO2. The heater 81 is used to transfer heat from the cooling channel 13 to the circulating working fluid, while the primary cooler 82 is used to transfer heat from the circulating working fluid to the fuel channel 23. The coolant storage tank 11 includes a separator 111 and a collector 112. The liquid metal passes through the separator 111, separating the heat from the aircraft's nose into two parts. One portion of the heat is transferred to the fuel through the secondary cooler 3 and then injected into the combustion chamber for heat reuse. The other portion of the heat is transferred to the CBC power generation system 8 through the heater 81, achieving thermoelectric conversion.

[0036] In the entire circulation system, the fuel first serves as a cold source for the CBC power generation system 8, absorbing heat in the primary cooler 82. At this time, the fuel does not crack. When the fuel passes through the secondary cooler 3 and exchanges heat with the liquid metal to reach the cracking temperature, the hydrocarbon fuel is injected into the combustion chamber to burn and release heat.

[0037] A temperature sensor is installed in the aircraft's air inlet and is used to collect temperature detection signals from the aircraft's air inlet. The electromagnetic adjustable pump 12 controls the flow of the coolant in the cooling channel 13 in response to the temperature detection signal, and the fuel adjustable pump 22 controls the amount of fuel in the fuel channel 23 in response to the temperature detection signal. By precisely adjusting the flow of the coolant based on the temperature detection signal detected by the temperature sensor in the aircraft's air inlet, the adjustable electromagnetic pump can adjust the cooling intensity according to the aircraft's actual heat load requirements, saving energy and improving system efficiency. The fuel adjustable pump 22 then adjusts the fuel amount, working in conjunction with the cooling system 1, enhancing the aircraft's ability to manage thermal energy and improving thermal energy utilization.

[0038] Reference Figure 2The cooling channel 13 includes a main flow channel 4 and a plurality of branch flow channels 41. The main flow channel 4 is arranged along the generatrix of the aircraft nose, and the branch flow channels 41 are arranged along the circumference of the aircraft nose. The inlet and outlet of the branch flow channels 41 are both connected to the main flow channel 4. The inlet of the branch flow channel 41 is located on the side of the outlet of the branch flow channel 41 away from the leading edge of the aircraft nose. Both the outlet and inlet of the branch flow channel 41 are provided with a one-way valve 42.

[0039] The cooling channel 13 is set as a mainstream channel 4 and a branch channel 41 to realize a single inlet and multiple outlets for the cooling medium. This can shorten the flow distance of the cooling medium, reduce the flow resistance of the cooling medium and the energy required for pumping, and especially reduce the problem of poor heat carrying capacity caused by the low specific heat of the cooling medium. At the same time, the multi-outlet design can increase the contact area between the cooling medium and the cooling surface, thereby improving the heat exchange efficiency.

[0040] The implementation principle of Example 1 is as follows: during the flight of a hypersonic aircraft, the electromagnetic adjustable pump 12 is started, and liquid metal is pumped into the cooling channel 13 of the aircraft head, thereby achieving efficient thermal protection for the head of the hypersonic aircraft, reducing the risk of structural failure due to high temperature, and improving the safety of the aircraft; the liquid metal can transfer the heat absorbed by the liquid metal to the fuel in the fuel channel 23 through the secondary cooler 3, and the fuel is directly injected into the combustion chamber for combustion after completing heat absorption, which can increase the temperature of the fuel, thereby improving the combustion efficiency and thrust, and further increasing the flight Mach number of the aircraft.

[0041] Example 2

[0042] Reference Figure 3 The difference between this embodiment and embodiment 1 is that the cooling channel 13 is arranged in a spiral shape along the circumferential side of the aircraft head, the inlet of the cooling channel 13 is located at the trailing edge of the aircraft head, and the outlet of the cooling channel 13 is located at the leading edge of the aircraft head.

[0043] The implementation principle of Example 2 is: the cooling channel 13 is set to a spiral shape, and the cooling medium flows from the trailing edge of the aircraft head around the circumference of the aircraft head to the leading edge of the aircraft head, thereby increasing the flow path of the cooling medium, improving the heat exchange efficiency, and thus improving the thermal protection effect.

[0044] Example 3

[0045] Reference Figure 4This embodiment differs from Embodiment 1 in that the cooling channel 13 includes a plurality of annular channels 5, which are arranged along the circumference of the aircraft nose. The annular channels 5 include two semi-annular channels 51, the inlet ends of the semi-annular channels 51 being located on the side of the semi-annular channels 51 away from the leading edge of the aircraft nose. The inlet ends of the two semi-annular channels 51 are interconnected, and the outlet ends of the two semi-annular channels 51 are interconnected. The outlet end of the semi-annular channel 51 away from the leading edge of the aircraft nose is interconnected with the inlet end of the semi-annular channel 51 near the leading edge of the aircraft nose.

[0046] The implementation principle of Example 3 is: dividing the annular channel 5 into two semi-annular channels 51 can shorten the distance that the liquid metal flows, effectively reduce the adverse effects caused by the specific heat capacity of the liquid metal, and enable the cooling medium to circulate around the entire aircraft head, thereby improving the uniformity and efficiency of thermal protection.

[0047] Example 4

[0048] Reference Figure 5 The difference between this embodiment and embodiment 1 is that the cooling channel 13 includes a feed channel 6, a discharge channel 61, and a plurality of circular channels 62. The feed channel 6 and the discharge channel 61 are both arranged along the generatrix direction of the aircraft head. The circular channel 62 is arranged along the circumference of the aircraft head. The circular channel 62 includes two semicircular channels 621. The inlet end of the semicircular channel 621 is located on the side of the semicircular channel 621 away from the leading edge of the aircraft head. The inlet end of the semicircular channel 621 is connected to the feed channel 6, and the outlet end of the semicircular channel 621 is connected to the feed channel 6. The inlet end and the outlet end of the semicircular channel 621 are both provided with a second one-way valve 622. The second one-way valve 622 can prevent the backflow of the cooling medium and ensure the stable operation of the cooling system 1.

[0049] The implementation principle of Example 4 is as follows: the cooling medium flows into the feed channel 6, the cooling medium in the feed channel 6 enters the discharge channel 61 through the semicircular channel 621, the cooling medium in several semicircular channels 621 converges in the discharge channel 61 and then flows out of the aircraft head, and several circular channels 62 are closely distributed on the inner wall of the flight head, which can improve the thermal protection effect of the cooling channel 13.

[0050] Example 5

[0051] Reference Figure 6 This embodiment differs from Embodiment 4 in that it further includes a feed channel 7, which is arranged along the generatrix of the aircraft head. The outlet of the feed channel 7 is connected to the inlet of the semicircular channel 621 located at the leading edge of the aircraft head. The outlet of the feed channel 7 is connected to the inlet of the feed channel 6, and the end of the feed channel 6 located at the trailing edge of the aircraft head is closed.

[0052] The implementation principle of Example 5 is as follows: the temperature of the cooling medium at the inlet end of the cooling channel 13 is lower than the temperature at the outlet end of the cooling channel 13. By adding a feed channel 7, the cooling medium first enters the semicircular channel 621 at the leading edge of the aircraft head through the feed channel 7, thereby achieving effective thermal protection for the leading edge of the aircraft head.

[0053] The above are all preferred embodiments of the present application, and are not intended to limit the scope of protection of the present application. Therefore, any equivalent changes made based on the structure, shape, and principle of the present application should be included in the scope of protection of the present application.

Claims

1. A supersonic aircraft head thermal protection structure, characterized by: The invention comprises a cooling system (1), a preheating system (2), a CBC power generation system (8) and a secondary cooler (3), wherein the cooling system (1) comprises a coolant storage tank (11), an electromagnetic adjustable pump (12) and a cooling channel (13) arranged in the wall surface of the aircraft head, wherein a cooling medium that flows in the opposite direction to the external high-temperature gas is arranged in the cooling channel (13), the coolant storage tank (11) is used to store the cooling medium, and the electromagnetic adjustable pump (12) is used to control the flow of the cooling medium in the cooling channel (13); the preheating system (2) comprises a fuel tank (21), a fuel adjustable pump (22) and a fuel channel (23), wherein the fuel adjustable pump (22) is used to control the amount of fuel in the fuel channel (23), and the secondary cooler (3) is used to transfer the heat in the cooling channel (13) to the fuel channel (23); the CBC power generation system ( 8) includes a heater (81), a primary cooler (82) and a circulating working fluid, wherein the heater (81) is used to transfer heat in the cooling channel (13) to the circulating working fluid, and the primary cooler (82) is used to transfer heat in the circulating working fluid to the fuel channel (23); the cooling channel (13) includes a main channel (4) and a plurality of branch channels (41), wherein the main channel (4) is arranged along the busbar direction of the aircraft head, and the branch channel (41) is arranged along the peripheral side of the aircraft head, and the inlet end of the branch channel (41) and the outlet end of the branch channel (41) are both connected to the main channel (4), and the inlet end of the branch channel (41) is located on the side of the outlet end of the branch channel (41) away from the leading edge of the aircraft head, and the outlet end of the branch channel (41) and the inlet end of the branch channel (41) are both provided with a one-way valve (42).

2. A supersonic vehicle head thermal protection structure according to claim 1, characterized in that: The invention also includes a temperature sensor, which is arranged in the aircraft air inlet duct and is used to collect a temperature detection signal in the aircraft air inlet duct. The electromagnetic adjustable pump (12) controls the flow rate of the cooling medium in the cooling channel (13) in response to the temperature detection signal, and the fuel adjustable pump (22) controls the amount of fuel in the fuel channel (23) in response to the temperature detection signal.

3. The supersonic vehicle head thermal protection structure according to claim 1, characterized in that: The cooling medium is liquid metal.

4. The supersonic vehicle head thermal protection structure according to claim 3, characterized in that: The material of the cooling channel (13) is a high-temperature alloy, the inner wall of the cooling channel (13) is coated with an anti-corrosion coating, and the cross-sectional shape of the cooling channel (13) is circular or polygonal.

5. The supersonic vehicle head thermal protection structure according to claim 4, characterized in that: The cross-sectional area of ​​the cooling channel (13) close to the leading edge of the aircraft head is smaller than the cross-sectional area of ​​the cooling channel (13) away from the leading edge of the aircraft head.

6. The supersonic vehicle head thermal protection structure according to claim 1, characterized in that: The cooling channel (13) is arranged in a spiral shape along the circumference of the aircraft head, the inlet of the cooling channel (13) is located at the rear edge of the aircraft head, and the outlet of the cooling channel (13) is located at the front edge of the aircraft head.

7. A supersonic vehicle head thermal protection structure according to claim 1, characterized in that: The cooling channel (13) includes a plurality of annular channels (5), the annular channels (5) are arranged along the circumference of the aircraft head, the annular channels (5) include two semi-annular channels (51), the inlet end of the semi-annular channel (51) is located on the side of the outlet end of the semi-annular channel (51) away from the leading edge of the aircraft head, the inlet ends of the two semi-annular channels (51) are connected, the outlet ends of the two semi-annular channels (51) are connected, and the outlet end of the semi-annular channel (51) away from the leading edge of the aircraft head is connected to the inlet end of the semi-annular channel (51) close to the leading edge of the aircraft head.

8. The supersonic vehicle head heat protection structure according to claim 1, characterized in that: The cooling channel (13) includes a feed channel (6), a discharge channel (61) and a plurality of circular channels (62). The feed channel (6) and the discharge channel (61) are both arranged along the busbar direction of the aircraft head. The circular channel (62) is arranged along the circumferential side of the aircraft head. The circular channel (62) includes two semicircular channels (621). The inlet end of the semicircular channel (621) is located on the side of the outlet end of the semicircular channel (621) away from the leading edge of the aircraft head. The inlet end of the semicircular channel (621) is connected to the feed channel (6), and the outlet end of the semicircular channel (621) is connected to the feed channel (6). Both the inlet end of the semicircular channel (621) and the outlet end of the semicircular channel (621) are provided with a second one-way valve (622).

9. A supersonic vehicle head heat protection structure according to claim 8, characterized in that: It also includes a material delivery channel (7), which is arranged along the busbar direction of the aircraft head, the outlet end of the material delivery channel (7) is connected to the inlet end of the semicircular channel (621) located at the front edge of the aircraft head, the outlet end of the material delivery channel (7) is connected to the inlet end of the feed channel (6), and the end of the feed channel (6) located at the rear edge of the aircraft is closed.

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