Supersonic velocity gas film distributed cooling method for near space waverider aircraft

Abstract

The invention provides a supersonic velocity gas film distributed cooling method for a near space waverider aircraft, and the method comprises the steps: arranging a first cooling device on the upper surface of a waverider close to the head of the waverider, and enabling the first cooling device to be used for generating a tangential supersonic velocity gas film; a second cooling device is arranged at the front edge of a sweepback wing of the waverider, and the second cooling device is used for generating a normal supersonic velocity gas film; the flowing direction of the tangential supersonic air film is consistent with the incoming air flow, and the flowing direction of the normal supersonic air film is perpendicular to the incoming air flow. The surface of the wave rider is covered with cooling gas through the supersonic cooling gas film, the coverage area is large, the cooling effect is stable, and the wave rider is suitable for optical windows and has no requirement for aircraft structural materials. And meanwhile, similar cooling devices can be arranged at different positions of the wave rider for combined cooling, and distributed large-area cooling is achieved.

Description

Technical Field

[0001] The invention mainly relates to the technical field of high-speed waverider aircraft, in particular to a supersonic air film distributed heat reduction method for a near-space waverider. Background Art

[0002] At present, waverider aircraft have excellent performance and are difficult to intercept because of their high flying speed, high speed and "water-skipping" flight trajectory. Military powers have launched fierce competition in the field of high-speed waverider aircraft, and have put forward higher requirements for the thermal protection of the aircraft surface.

[0003] High-speed flight brings serious aerodynamic heat problems to the waverider. Excessive heat flux density causes ablation and damage to the surface structure of the aircraft. Taking the Mach 6 waverider as an example, the heat flux density on the surface of the aircraft is about 7.8×10 4 W / m 2 ,like Figure 2 As shown. At present, ablation is usually used for thermal protection, which belongs to passive thermal protection and is suitable for ultra-high altitude (low temperature) and low Mach number (low speed) aircraft. For high-speed aircraft, especially during ascent, reentry and acceleration, severe heat flow poses a huge challenge to ablation and heat reduction. For example, the most serious accident in aviation history, the Columbia Space Shuttle, was caused by cracks in the carbon-carbon heat shield at the leading edge of the aircraft's wing, which led to ablation and disintegration of the aircraft when it re-entered the atmosphere.

[0004] At present, the thermal protection method of high-speed aircraft mainly relies on passive ablation, and convection cooling, sweat cooling and air film cooling are relatively less used. For high-speed waverider aircraft, the defects of the existing technology are mainly:

[0005] (1) Relying solely on ablation to reduce heat for passive thermal protection is not enough. The leading edge of the swept wing of a waverider aircraft has a significant impact on the aerodynamic performance of the entire aircraft. At the same time, this part has limited bluntness due to the need to wrap the shock wave, resulting in almost the highest heat flux density and the most serious ablation. For high-speed waveriders, other cooling methods must be combined to cool the leading edge of the swept wing.

[0006] (2) Convection internal cooling uses the cooling fluid to conduct heat under the aircraft skin and removes heat through internal circulation. It is suitable for cooling the aircraft's stationary points, head, and swept wings. However, the required cooling fluid flow rate is too large and the efficiency is low, so it cannot be used on a large area of ​​the aircraft.

[0007] (3) Transpiration cooling involves seepage through small pores on the surface of the vehicle and absorbing heat through phase change. However, it requires the use of special materials on the surface of the vehicle, which has a great impact on the structural strength. More importantly, it cannot be applied to optical windows.

[0008] (4) The optical window used in imaging guidance cannot rely on passive ablation and sweating cooling for thermal protection. Due to its good imaging and anti-interference capabilities, imaging guidance is one of the current air flow guidance methods for high-speed aircraft to achieve precision strikes. However, due to the particularity of the optical window, thermal protection methods such as passive ablation and sweating cooling cannot be used. At the same time, the tolerance of the optical window to high temperature is much lower than that of other parts of the aircraft. High-density heat flow can cause the optical signal to be submerged at best, and at worst cause the window to overheat and be damaged, causing the waverider aircraft to lose its guidance capability.

[0009] For the above reasons, distributed cooling is needed for high-speed waverider vehicles. Summary of the invention

[0010] In view of the problem in the prior art that the surface heat flux of existing near-space waverider vehicles is too high and it is difficult to perform global cooling, the present invention proposes a supersonic air film distributed heat reduction method for near-space waverider bodies.

[0011] To achieve the above purpose, the technical solution adopted by the present invention is as follows:

[0012] On the one hand, the present invention provides a distributed heat reduction method for supersonic air film of a near-space waverider, wherein a first heat reduction device is arranged on the upper surface of the waverider near the head position of the waverider, and the first heat reduction device is used to generate a tangential supersonic air film; a second heat reduction device is arranged on the leading edge of the swept wing of the waverider, and the second heat reduction device is used to generate a normal supersonic air film; the flow direction of the tangential supersonic air film is consistent with the incoming air flow, which slows down the development of the supersonic air film-incoming flow mixing layer, extends the cooling range of the supersonic air film, and realizes large-area coverage cooling of the waverider surface; the normal supersonic air film pushes the hot incoming air flow away from the leading edge of the swept wing, and the flow direction of the normal supersonic air film is perpendicular to the incoming air flow. The exchange of momentum causes the normal supersonic air film to flow downstream with the incoming air flow, thereby realizing coverage of the downstream of the swept wing of the waverider, and the exchange of energy reduces the temperature of the hot incoming air flow, thereby realizing cooling of the downstream of the swept wing of the waverider.

[0013] Furthermore, the first heat reduction device is arranged flush with the upper surface of the waverider body, the flow direction of the tangential supersonic air film generated by the first heat reduction device is parallel to the incoming air flow, the pressure of the tangential supersonic air film generated by the first heat reduction device is equal to the incoming air flow, and the temperature of the tangential supersonic air film generated by the first heat reduction device is much lower than the stagnation temperature of the incoming air flow.

[0014] Furthermore, the first heat reduction device and the second heat reduction device each include a series of supersonic air film nozzles, each supersonic air film nozzle is connected to a gas source and a gas supply pipeline for providing gas, and the gas source and the gas supply pipeline are arranged inside the waverider.

[0015] The above heat reduction method has the following advantages over the prior art:

[0016] (1) Distributed active heat dissipation of tangential supersonic air film and normal supersonic air film is adopted to effectively suppress the heat flux density of the fuselage and swept wings of the high-speed waverider aircraft.

[0017] (2) The tangential supersonic air film has no special requirements for the surface material of the high-speed waverider aircraft; the tangential supersonic air film can effectively suppress the cooling and aero-optical effects of the imaging guidance window, which is impossible to achieve with passive ablation, convection cooling and sweating cooling.

[0018] (3) The normal supersonic air film can effectively suppress the heat flux density on the leading edge and downstream surface of the swept wing of a high-speed waverider aircraft.

[0019] (4) Supersonic air film is an active heat-reduction method, which can avoid the structural damage of waverider aircraft caused by passive heat-reduction due to ablation.

[0020] (5) Furthermore, the first heat reduction device and the second heat reduction device each include a series of supersonic air film nozzles. The first heat reduction device, which is arranged flush with the upper surface of the waverider, generates a tangential supersonic air film through a series of supersonic air film nozzles, and covers the cooling gas on the surface of the waverider. The coverage area is large, the cooling effect is stable, and it is suitable for optical windows, and there is no requirement for the structural material of the aircraft. At the same time, based on this idea, similar heat reduction devices can be arranged at different positions of the waverider for combined cooling to achieve distributed large-area heat reduction.

[0021] Furthermore, the supersonic air film nozzle is designed by the following steps, including:

[0022] Obtain the supersonic air film nozzle outlet pressure under pressure matching conditions;

[0023] According to the cooling length requirements, the supersonic air film mass flow rate and supersonic air film Mach number are determined;

[0024] The throat height of the supersonic air film nozzle is determined according to the supersonic air film mass flow rate and the supersonic air film nozzle outlet pressure;

[0025] According to the supersonic air film nozzle throat height and the supersonic air film Mach number, the nozzle outlet height of the supersonic air film nozzle is determined;

[0026] For the transonic flow in the throat of supersonic film nozzle, the parabolic potential function equation is solved by series expansion method, and the transonic solution of supersonic film nozzle is obtained.

[0027] The Mach number distribution of the supersonic film nozzle axis is set with the supersonic film nozzle transonic solution and the nozzle outlet as the starting point and the end point respectively.

[0028] Taking the transonic solution of the supersonic film nozzle, the Mach number distribution of the nozzle outlet and the axis of the supersonic film nozzle as boundary conditions, the characteristic line grid of the supersonic film nozzle is constructed, and the supersonic flow field inside the supersonic film nozzle is solved. Then, the inviscid profile of the nozzle is determined according to the streamline control equation.

[0029] According to the von Karman momentum integral relation, the displacement thickness of the boundary layer is solved;

[0030] The viscosity correction of the nozzle inviscid profile is performed based on the boundary layer displacement thickness to obtain the final nozzle profile;

[0031] The contraction section profile of the preset supersonic air film nozzle is connected with the nozzle profile at the throat to obtain a complete supersonic air film nozzle.

[0032] The supersonic air film nozzle designed by the above method has the following advantages:

[0033] (1) The internal flow of the designed supersonic air film nozzle is perfectly wave-free, and there are no concentrated compression waves or shock waves.

[0034] (2) The supersonic air film nozzle has been viscosity corrected, and the uniformity of the internal flow and nozzle outlet is very high.

[0035] (3) The nozzle outlet pressure is consistent with the theoretical value, which can achieve the matching of supersonic air film pressure and air flow pressure. BRIEF DESCRIPTION OF THE DRAWINGS

[0036] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings required for use in the embodiments or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on the structures shown in these drawings without paying creative work.

[0037] Figure 1 This is a schematic diagram of the arrangement position of the heat reduction device on the waverider in one embodiment;

[0038] Figure 2 is a heat flux distribution diagram on the surface of a Mach 6 waverider aircraft in one embodiment;

[0039] Figure 3 This is a heat flux distribution diagram on the surface of a Mach 6 waverider aircraft using a supersonic tangential cooling air film in one embodiment;

[0040] Figure 4 is a heat flux distribution diagram on the surface of a Mach 6 waverider aircraft using a supersonic normal cooling air film on the leading edge of the swept wing in one embodiment;

[0041] Figure 5This is a heat flux distribution diagram on the surface of a Mach 6 waverider aircraft after adopting supersonic air film distributed heat reduction in one embodiment;

[0042] Figure 6 is a nozzle Mach number cloud diagram in one embodiment;

[0043] Figure 7 1 is a nozzle outlet Mach number distribution diagram in one embodiment. DETAILED DESCRIPTION

[0044] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0045] In one embodiment, a distributed heat reduction method for supersonic air film near a space waverider is provided, wherein a first heat reduction device is arranged on the upper surface of the waverider near the head of the waverider, and the first heat reduction device is used to generate a tangential supersonic air film; a second heat reduction device is arranged on the leading edge of the swept wing of the waverider, and the second heat reduction device is used to generate a normal supersonic air film. In one embodiment, the arrangement positions of the first heat reduction device and the second heat reduction device on the waverider are as follows: Figure 1 As shown, it includes a first heat reduction device arranged at position 1 on the waverider body and a second heat reduction device arranged at position 2 on the waverider body.

[0046] It can be understood that the second heat reduction device is arranged on the leading edges of the swept wings on both sides of the waverider body.

[0047] The flow direction of the tangential supersonic air film is consistent with the incoming air flow, which slows down the development of the supersonic air film-incoming air mixing layer, extends the cooling range of the supersonic air film, and realizes large-area coverage cooling of the waverider surface; the normal supersonic air film pushes the hot incoming air flow away from the leading edge of the swept wing, and the flow direction of the normal supersonic air film is perpendicular to the incoming air flow. The exchange of momentum makes the normal supersonic air film flow downstream with the incoming air flow, thereby realizing coverage of the downstream of the swept wing of the waverider. The exchange of energy reduces the temperature of the hot incoming air flow, thereby realizing cooling of the downstream of the swept wing of the waverider.

[0048] The first heat reduction device and the second heat reduction device each include a series of supersonic air film nozzles, each supersonic air film nozzle is connected to a gas source and a gas supply pipeline for providing gas, and the gas source and the gas supply pipeline are arranged inside the waverider.

[0049] In one embodiment, a scheme for installing the supersonic air film nozzle in the first heat reduction device is provided: an installation area for installing each supersonic air film nozzle in the first heat reduction device is provided on the upper surface of the waverider near the head of the waverider, and each supersonic air film nozzle is arranged side by side in the installation area in contact with the upper surface of the waverider.

[0050] The flow direction of the tangential supersonic air film generated by a series of supersonic air film nozzles in the first heat reduction device is parallel to the incoming air flow, the pressure of the tangential supersonic air film generated by the first heat reduction device is equal to the incoming air flow, and the temperature of the tangential supersonic air film generated by the first heat reduction device is much lower than the stagnation temperature of the incoming air flow. In this way, the pressure of the tangential supersonic air film generated by the first heat reduction device matches the pressure of the incoming air flow and the flow direction is consistent, which fundamentally avoids the condition of the tangential supersonic air film generated by the first heat reduction device and the supersonic air flow generating shock waves, so that the tangential supersonic air film generated by the first heat reduction device and the incoming air flow will not be strongly mixed due to the shock wave, and has a better and longer distance heat reduction effect.

[0051] In one embodiment, an installation scheme for the supersonic air film nozzle in the second heat reduction device is provided: a series of supersonic air film nozzles in the second heat reduction device are arranged side by side on the leading edge of the swept wing of the waverider. Without loss of generality, the series of supersonic air film nozzles in the second heat reduction device can be built into the leading edge of the swept wing of the waverider, and the nozzle of each supersonic air film nozzle faces outward and the injection direction of each supersonic air film nozzle is perpendicular to the incoming air flow.

[0052] Figure 3 The numerical simulation results of heat flux on the surface of a high-speed waverider with a supersonic tangential cooling film are presented. Figure 2 By comparison, it can be seen that the heat flux density downstream of the tangential air film increases from 7.8×10 4 W / m 2 Down to 6.4×10 3 W / m 2 The tangential air film covers the downstream along the wall of the waverider aircraft, and the effective coverage length can reach the tail of the aircraft. In this process, although the air film-air flow structure is constantly mixed and the air film is continuously heated by the air flow, the cooling effect is at the tail of the aircraft, and the heat flux density can still be controlled at 2.56×10 4 W / m 2 About 5×10 3 W / m 2 The results show that the supersonic tangential air film can effectively cool the downstream wall of the head of the high-speed waverider.

[0053] Figure 4 The numerical simulation results of heat flux on the surface of a high-speed waverider with a supersonic normal cooling film are presented. Figure 2 In comparison, the normal air film successfully cooled the heat flux density of the swept wing leading edge and its downstream wall to 6.4×10 3 W / m 2 ~3.2×10 4 W / m 2 , which is 4.8×10 lower than that of the high-speed waverider without normal air film. 3 W / m 2 The results show that the supersonic normal air film successfully cools the leading edge of the high-speed waverider swept wing and its downstream wall.

[0054] Figure 5 The numerical simulation results of distributed heat reduction of high-speed waverider aircraft using supersonic air film distributed heat reduction are shown. It can be seen that the distributed cooling based on supersonic air film successfully reduces the surface heat flux density of Mach 6 waverider aircraft to 3.2×10 4 W / m 2 The distributed air film has a better cooling effect than the tangential or normal air film alone, and the heat flux density on the surface of the waverider is lower. The heat flux density of the fuselage and swept wings of the high-speed waverider can be effectively suppressed by the distributed heat reduction of the supersonic air film.

[0055] The present invention has been simulated by numerical simulation, such as Figure 2-5 The present invention has undergone one round of high enthalpy wind tunnel tests and has achieved the expected design, and the technology is feasible.

[0056] In one embodiment, a supersonic air film nozzle design method is provided, comprising:

[0057] Obtain the supersonic air film nozzle outlet pressure of a high-speed waverider vehicle under pressure matching conditions;

[0058] According to the cooling length requirements of high-speed waverider aircraft, the supersonic air film mass flow rate and supersonic air film Mach number are determined;

[0059] The throat height of the supersonic air film nozzle is determined according to the supersonic air film mass flow rate and the supersonic air film nozzle outlet pressure;

[0060] According to the supersonic air film nozzle throat height and the supersonic air film Mach number, the nozzle outlet height of the supersonic air film nozzle is determined;

[0061] For the transonic flow in the throat of supersonic film nozzle, the parabolic potential function equation is solved by series expansion method, and the transonic solution of supersonic film nozzle is obtained.

[0062] The Mach number distribution of the supersonic film nozzle axis is set with the supersonic film nozzle transonic solution and the nozzle outlet as the starting point and the end point respectively.

[0063] Taking the transonic solution of the supersonic film nozzle, the Mach number distribution of the nozzle outlet and the axis of the supersonic film nozzle as boundary conditions, the characteristic line grid of the supersonic film nozzle is constructed, and the supersonic flow field inside the supersonic film nozzle is solved. Then, the inviscid profile of the nozzle is determined according to the streamline control equation.

[0064] According to the von Karman momentum integral relation, the displacement thickness of the boundary layer is solved;

[0065] The viscosity correction of the nozzle inviscid profile is performed based on the boundary layer displacement thickness to obtain the final nozzle profile;

[0066] The contraction section profile of the preset supersonic air film nozzle is connected with the nozzle profile at the throat to obtain a complete supersonic air film nozzle.

[0067] It can be understood that the pressure matching condition in the present invention means that the incoming air pressure is equal to the supersonic air film pressure.

[0068] In one embodiment, a method for calculating the throat height of a supersonic air film nozzle is proposed. Specifically, the throat height y of the supersonic air film nozzle is * Determined by the following formula:

[0069]

[0070] Where γ is the specific heat ratio of the gas, R is the gas constant, T is the static temperature, P is the supersonic air film nozzle outlet pressure, and Me is the supersonic air film Mach number.

[0071] In one embodiment, a method for calculating the nozzle outlet height of a supersonic air film nozzle is provided. Specifically, the nozzle outlet height y of the supersonic air film nozzle is e Determined by the following formula:

[0072]

[0073] Where y* is the throat height of the supersonic air film nozzle, γ is the specific heat ratio of the gas, and Me is the supersonic air film Mach number.

[0074] In a supersonic air film nozzle design method provided in one embodiment, the parabolic potential function equation used is:

[0075]

[0076] Where a is the critical sound speed of the sound speed line at the throat of the supersonic air film nozzle, u is the axial speed of the sound speed line at the throat of the supersonic air film nozzle, v is the radial speed of the sound speed line at the throat of the supersonic air film nozzle, and x and y represent the x-axis coordinate and y-axis coordinate of the sound speed line at the throat of the supersonic air film nozzle, respectively.

[0077] For the transonic flow in the throat of a supersonic air film nozzle, the series expansion method is used to solve the parabolic potential function equation and the transonic solution of the supersonic air film nozzle is obtained.

[0078] In a supersonic air film nozzle design method provided in an embodiment, the supersonic air film nozzle transonic solution and the nozzle outlet are used as the starting point and the end point respectively, and the Mach number distribution of the supersonic air film nozzle axis is set. According to the supersonic air film nozzle transonic solution boundary condition and the outlet boundary condition, that is, the supersonic air film nozzle transonic solution and the nozzle outlet are used as the starting point and the end point respectively, the position coordinates, Mach number, and the first derivative of the Mach number of the starting point and the end point are substituted into the B-spline curve expression to obtain the Mach number distribution of the supersonic air film nozzle axis.

[0079] In a supersonic air film nozzle design method provided in one embodiment, the transonic solution of the supersonic air film nozzle, the Mach number distribution of the nozzle outlet and the axis of the supersonic air film nozzle are used as boundary conditions, and the characteristic line grid of the supersonic air film nozzle is constructed based on the characteristic line method to solve the supersonic flow field inside the supersonic air film nozzle, and then determine the nozzle inviscid line according to the streamline control equation. The characteristic line method is a conventional technical means in this field. The characteristic line method is used to solve the two-dimensional flow hyperbolic equation to design a supersonic air film nozzle. The wave is completely eliminated in the two-dimensional structure of the supersonic air film nozzle, which fundamentally eliminates the shock wave, and the distribution of flow field parameters such as pressure, temperature, and velocity is controllable. The two characteristic line equations are given below:

[0080] Characteristic line method equation 1:

[0081]

[0082] Characteristic line method equation 2:

[0083]

[0084] Where θ is the momentum loss thickness, μ is the Mach angle, Ma is the local Mach number, γ is the specific heat ratio of the gas, and δ is the flow characteristic factor.

[0085] The streamline control equation is: dy / dx=v / u

[0086] Where u is the axial velocity and v is the radial velocity.

[0087] In a supersonic air film nozzle design method provided in an embodiment, the boundary layer displacement thickness δ is solved by combining the following formulas: * :

[0088]

[0089]

[0090]

[0091] Where θ is the momentum loss thickness, δ * is the boundary layer displacement thickness, is the wall angle, H is the boundary layer shape factor, C f is the compressible friction coefficient, Ma is the local Mach number, and γ is the specific heat ratio of the gas.

[0092] Further, the boundary layer displacement thickness δ is obtained according to the above method. * After that, the nozzle non-viscous profile is increased by δ in the normal direction. * The distance is used to achieve viscosity correction.

[0093] It can be understood that those skilled in the art can predetermine the contraction section profile of the designed supersonic air film nozzle by using existing methods or directly given in a given manner based on experience. The present application does not limit the specific design method of the contraction section profile of the supersonic air film nozzle.

[0094] In a preferred embodiment, a method for designing a contraction profile of a supersonic air film nozzle is provided, wherein the coordinates (x, y) of each point on the contraction profile of the supersonic air film nozzle are determined by a quintic curve, as follows:

[0095]

[0096] where y i is the inlet height of the supersonic film nozzle, y * is the supersonic film nozzle throat height, L c is the given contraction length.

[0097] In a preferred embodiment, another method for designing the contraction profile of a supersonic air film nozzle is provided. The coordinates (x, y) of each point on the contraction profile of the supersonic air film nozzle are determined by the Witosinski curve, as follows:

[0098]

[0099] where y i is the inlet height of the supersonic film nozzle, y * is the supersonic film nozzle throat height, L c It is the given length of the contraction section, which is given by the designer. The contraction section does not affect the design results of the supersonic air film.

[0100] A Mach 3 nozzle is designed by using the supersonic air film nozzle design method provided in the above embodiment, so that the tangential jet passing through the Mach 3 nozzle avoids direct friction of the hypersonic mainstream on the surface of the waverider, thereby reducing the surface temperature of the waverider and achieving the effect of reducing heat and drag. The input conditions of the Mach 3 nozzle are controlled at a total temperature of 300K, a total pressure of 100000Pa, a nozzle outlet temperature of 108K, and a pressure of 2820Pa. The design results are as follows: Figure 6 , Figure 7 As shown, Figure 6 It is the nozzle Mach number cloud diagram; Figure 7 This is the nozzle exit Mach number distribution diagram. It can be seen that the flow is steadily accelerated from subsonic to supersonic flow, the nozzle internal wave is completely eliminated, there is no concentrated shock wave, and the exit Mach number is particularly uniform, which reflects the feasibility of this method.

[0101] Matters not covered by the present invention are known technologies.

[0102] The technical features of the above embodiments may be combined arbitrarily. To make the description concise, 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.

[0103] The above-mentioned embodiments only express several implementation methods of the present application, and the descriptions thereof are relatively specific and detailed, but they cannot be understood as limiting the scope of the invention patent. It should be pointed out that, for a person of ordinary skill in the art, several variations and improvements can be made without departing from the concept of the present application, and these all belong to the protection scope of the present application. Therefore, the protection scope of the patent of the present application shall be subject to the attached claims.

[0104] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. For those skilled in the art, the present invention may have various modifications and variations. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention shall be included in the protection scope of the present invention.

Claims

1. A supersonic air film distributed heat reduction method for near-space waveriders, characterized in that: A first heat reduction device is arranged on the upper surface of the waverider body near the head of the waverider body, and the first heat reduction device is used to generate a tangential supersonic air film; a second heat reduction device is arranged on the leading edge of the swept wing of the waverider body, and the second heat reduction device is used to generate a normal supersonic air film; the flow direction of the tangential supersonic air film is consistent with the incoming air flow, which slows down the development of the supersonic air film-incoming flow mixing layer, extends the cooling range of the supersonic air film, and realizes large-area coverage cooling of the waverider surface; the normal supersonic air film pushes the hot incoming air flow away from the leading edge of the swept wing, and the flow direction of the normal supersonic air film is perpendicular to the incoming air flow. The exchange of momentum makes the normal supersonic air film flow downstream with the incoming air flow, thereby realizing coverage of the downstream of the swept wing of the waverider body, and the exchange of energy reduces the temperature of the hot incoming air flow, thereby realizing cooling of the downstream of the swept wing of the waverider body.

2. The supersonic air film distributed heat reduction method for near-space waverider according to claim 1, characterized in that: The first heat reduction device is arranged flush with the upper surface of the waverider body, the flow direction of the tangential supersonic air film generated by the first heat reduction device is parallel to the incoming air flow, the pressure of the tangential supersonic air film generated by the first heat reduction device is equal to the incoming air flow, and the temperature of the tangential supersonic air film generated by the first heat reduction device is much lower than the stagnation temperature of the incoming air flow.

3. The supersonic air film distributed heat reduction method for near-space waverider according to claim 1 or 2, characterized in that: The first heat reduction device and the second heat reduction device each include a series of supersonic air film nozzles, each supersonic air film nozzle is connected to a gas source and a gas supply pipeline for providing gas, and the gas source and the gas supply pipeline are arranged inside the waverider.

4. The supersonic air film distributed heat reduction method for near-space waverider according to claim 3 is characterized in that: The supersonic air film nozzle is designed by the following steps, including: Obtain the supersonic air film nozzle outlet pressure under pressure matching conditions; According to the cooling length requirements, the supersonic air film mass flow rate and supersonic air film Mach number are determined; The throat height of the supersonic air film nozzle is determined according to the supersonic air film mass flow rate and the supersonic air film nozzle outlet pressure; According to the supersonic air film nozzle throat height and the supersonic air film Mach number, the nozzle outlet height of the supersonic air film nozzle is determined; For the transonic flow in the throat of supersonic film nozzle, the parabolic potential function equation is solved by series expansion method, and the transonic solution of supersonic film nozzle is obtained. The Mach number distribution of the supersonic film nozzle axis is set with the supersonic film nozzle transonic solution and the nozzle outlet as the starting point and the end point respectively. Taking the transonic solution of the supersonic film nozzle, the Mach number distribution of the nozzle outlet and the axis of the supersonic film nozzle as boundary conditions, the characteristic line grid of the supersonic film nozzle is constructed, and the supersonic flow field inside the supersonic film nozzle is solved. Then, the inviscid profile of the nozzle is determined according to the streamline control equation. According to the von Karman momentum integral relation, the displacement thickness of the boundary layer is solved; The viscosity correction of the nozzle inviscid profile is performed based on the boundary layer displacement thickness to obtain the final nozzle profile; The contraction section profile of the preset supersonic air film nozzle is connected with the nozzle profile at the throat to obtain a complete supersonic air film nozzle.

5. The supersonic air film distributed heat reduction method for near-space waverider according to claim 4, characterized in that: Supersonic air film nozzle throat height y * Determined by the following formula: Where γ is the specific heat ratio of the gas, R is the gas constant, T is the static temperature, P is the supersonic air film nozzle outlet pressure, and Me is the supersonic air film Mach number.

6. The supersonic air film distributed heat reduction method for near-space waverider according to claim 4, characterized in that: The nozzle exit height y of the supersonic air film nozzle e Determined by the following formula: Where y* is the throat height of the supersonic air film nozzle, γ is the specific heat ratio of the gas, and Me is the supersonic air film Mach number.

7. The supersonic air film distributed heat reduction method for near-space waverider according to claim 4, 5 or 6, characterized in that: The parabolic potential function equation is: Where a is the critical sound speed of the sound speed line at the throat of the supersonic air film nozzle, u is the axial speed of the sound speed line at the throat of the supersonic air film nozzle, v is the radial speed of the sound speed line at the throat of the supersonic air film nozzle, and x and y represent the x-axis coordinate and y-axis coordinate of the sound speed line at the throat of the supersonic air film nozzle, respectively.

8. The supersonic air film distributed heat reduction method for near-space waverider according to claim 7, characterized in that: Boundary layer displacement thickness δ * By combining the following formulas to solve: Where θ is the momentum loss thickness, δ * is the boundary layer displacement thickness, is the wall angle, H is the boundary layer shape factor, C f is the compressible friction coefficient, Ma is the local Mach number, and γ is the specific heat ratio of the gas.

9. The supersonic air film distributed heat reduction method for near-space waverider according to claim 7, characterized in that: The method of making viscosity correction to the nozzle inviscid profile is: the nozzle inviscid profile increases δ in the normal direction. * The distance is used to achieve viscosity correction.

10. The supersonic air film distributed heat reduction method for near-space waverider according to claim 4, 5, 6, 8 or 9, characterized in that: The coordinates (x, y) of each point on the contraction profile of the supersonic air film nozzle are determined by the Witoshinsky curve as follows: where y i is the inlet height of the supersonic film nozzle, y * is the supersonic film nozzle throat height, L c is the given contraction length.

Images