Supersonic air film distributed heat reduction method for near-space waverider aircraft

Abstract

The present invention proposes a distributed supersonic air film cooling method for near-space waverider aircraft. A first cooling device is positioned on the upper surface of the waverider near the head of the waverider. This device generates a tangential supersonic air film. A second cooling device is positioned on the leading edge of the waverider's swept wings. This device generates a normal supersonic air film. The tangential supersonic air film flows in the same direction as the incoming air flow, while the normal supersonic air film flows perpendicular to the incoming air flow. This supersonic cooling air film coats the waverider surface with cooling gas, resulting in a large coverage area, stable cooling effect, and applicability to optical windows, without requiring high-quality aircraft structural materials. Similar cooling devices can also be positioned at different locations on the waverider for combined cooling, achieving distributed, large-area heat reduction.

Description

Technical Field

[0001] The present 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] Currently, waverider aircraft, due to their high altitude, high speed, and "water-skimming" flight trajectory, have excellent performance and are difficult to intercept. Military powers are engaged in fierce competition in the field of high-speed waverider aircraft, placing higher demands on the thermal protection of aircraft surfaces.

[0003] High-speed flight brings serious aerodynamic heat problems to waveriders. Excessive heat flux density causes ablation and damage to the surface structure of the aircraft. Taking a Mach 6 waverider aircraft as an example, the heat flux density on the aircraft surface is about 7.8×10 4 W / m 2 ,like Figure 2 Currently, ablation is commonly used for thermal protection, a passive method suitable for aircraft operating at very high altitudes (low temperatures) and low Mach numbers (low speeds). For high-speed aircraft, severe heat fluxes, especially during ascent, reentry, and acceleration, pose a significant challenge to ablation-based heat reduction. For example, the Columbia Space Shuttle, the worst accident in aviation history, was caused by cracks in the carbon-carbon heat shielding on the leading edge of the aircraft's wings, leading to ablation and disintegration during re-entry.

[0004] Currently, the thermal protection method for high-speed aircraft mainly relies on passive ablation, and the use of convection cooling, perspiration cooling and film cooling is relatively rare. For high-speed waverider aircraft, the main defects of the existing technology are:

[0005] (1) Relying solely on ablation to reduce heat for passive thermal protection is insufficient. The swept wing leading edge 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 severe ablation. For high-speed waveriders, other cooling methods must be combined to cool the swept wing leading edge.

[0006] (2) Convection internal cooling uses the cooling medium to conduct heat under the aircraft skin and removes the heat through internal circulation. It is suitable for cooling the aircraft's stationary points, head, and swept wings. However, the required cooling medium 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 aircraft and absorbing heat through phase change. However, this requires the use of special materials on the surface of the aircraft, 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 or sweating cooling for thermal protection. Due to its excellent 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 optical window's tolerance to high temperatures is much lower than that of other aircraft components. High-density heat flow can at best cause the optical signal to be submerged, 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 necessary for high-speed waverider aircraft. Summary of the Invention

[0010] Aiming at the problem in the prior art that the surface heat flux of existing near-space waverider aircraft is too high and global cooling is difficult to achieve, the present invention proposes a supersonic air film distributed heat reduction method for near-space waverider.

[0011] To achieve the above object, 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 a 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 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-oncoming flow mixing layer, extends the cooling range of the supersonic air film, and achieves 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 achieving 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 achieving 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 gas supply pipeline are arranged inside the waverider.

[0015] The above-mentioned fever reduction method has the following advantages over the prior art:

[0016] (1) Distributed active heat dissipation of tangential supersonic air films and normal supersonic air films is used to effectively suppress the heat flux density of the fuselage and swept wings of 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 by 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 of the swept wing and its downstream surface of the high-speed waverider aircraft.

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

[0020] (5) Furthermore, the first and second heat reduction devices each include a series of supersonic air film nozzles. The first heat reduction device, which is flush with the upper surface of the waverider, generates a tangential supersonic air film through a series of supersonic air film nozzles, covering the surface of the waverider with cooling gas. This provides a large coverage area, a stable cooling effect, and is suitable for optical windows, without requiring any requirements for aircraft structural materials. Based on this concept, similar heat reduction devices can be installed at different locations on the waverider for combined cooling, achieving 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] Determine the supersonic air film mass flow rate and supersonic air film Mach number according to the cooling length requirements;

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

[0025] Determine the nozzle exit height of the supersonic film nozzle based on the supersonic film nozzle throat height and the supersonic film Mach number;

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

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

[0028] Using the transonic solution of the supersonic film nozzle and the Mach number distribution at the nozzle exit and the axis of the supersonic film nozzle as boundary conditions, a characteristic line grid of the supersonic film nozzle is constructed to solve the supersonic flow field inside the supersonic film nozzle. Then, the inviscid profile of the nozzle is determined based on the streamline governing equations.

[0029] According to the von Karman momentum integral relationship, the boundary layer displacement thickness 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 preset contraction section profile of the 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, with 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 following briefly introduces the drawings required for use in the embodiments or the description of the prior art. 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 any creative work.

[0037] Figure 1 Schematic diagram of the arrangement position of the heat reduction device on the waverider in one embodiment;

[0038] Figure 2 This is a diagram of heat flux distribution 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 This 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 using a supersonic air film distributed heat reduction in one embodiment;

[0042] Figure 6 This 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 clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. All other embodiments obtained by ordinary technicians in this field based on the embodiments of the present invention without making any creative efforts shall fall within the scope of protection of the present invention.

[0045] In one embodiment, a method for distributing heat reduction of supersonic air film near a space waverider is provided. A first heat reduction device is arranged on the upper surface of the waverider near the head of the waverider. 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. 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, the first heat reduction device is arranged at position 1 on the waverider and the second heat reduction device is arranged at position 2 on the waverider.

[0046] It can be understood that the second heat reduction device is arranged on the leading edge 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. 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 waverider swept wing. The exchange of energy reduces the temperature of the hot incoming air flow, thereby realizing cooling of the downstream of the waverider swept wing.

[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 gas supply pipeline are arranged inside the waverider.

[0049] In one embodiment, a solution 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 tangential supersonic air film generated by a series of supersonic air film nozzles in the first heat reduction device flows parallel to the incoming air flow. The pressure of the tangential supersonic air film generated by the first heat reduction device is equal to that of the incoming air flow, and the temperature of the tangential supersonic air film generated by the first heat reduction device is far below the stagnation temperature of the incoming air flow. Thus, 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. This fundamentally avoids the conditions for the tangential supersonic air film generated by the first heat reduction device and the supersonic air flow to generate shock waves. This prevents strong mixing between the tangential supersonic air film generated by the first heat reduction device and the incoming air flow due to shock waves, resulting in a better and longer-range heat reduction effect.

[0051] In one embodiment, a solution for installing the supersonic air film nozzles 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, with the nozzle of each supersonic air film nozzle facing outward and the injection direction of each supersonic air film nozzle being 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 In 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 wall of the waverider aircraft, with an effective coverage length reaching the tail of the aircraft. During this process, although the air film and air flow structure are 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 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 aircraft without normal air film. 3 W / m 2 The results show that the supersonic normal air film successfully cooled 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 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 Furthermore, distributed air films offer greater cooling effectiveness than single tangential or normal air films, lowering the surface heat flux of waverider aircraft. Distributed supersonic air film cooling can effectively suppress the heat flux of the fuselage and swept wings of high-speed waverider aircraft.

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

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

[0057] Obtain the supersonic 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 film nozzle is determined according to the supersonic film mass flow rate and the supersonic film nozzle outlet pressure.

[0060] Determine the nozzle exit height of the supersonic film nozzle based on the supersonic film nozzle throat height and the supersonic film Mach number;

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

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

[0063] Using the transonic solution of the supersonic film nozzle and the Mach number distribution at the nozzle exit and the axis of the supersonic film nozzle as boundary conditions, a characteristic line grid of the supersonic film nozzle is constructed to solve the supersonic flow field inside the supersonic film nozzle. Then, the inviscid profile of the nozzle is determined based on the streamline governing equations.

[0064] According to the von Karman momentum integral relationship, the boundary layer displacement thickness 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 preset contraction section profile of the 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 film nozzle outlet pressure, and Me is the supersonic film Mach number.

[0071] In one embodiment, a method for calculating the nozzle outlet height of a supersonic air film nozzle is proposed. 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 film nozzle, γ is the specific heat ratio of the gas, and Me is the supersonic 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 velocity of the sound line at the throat of the supersonic air film nozzle, u is the axial velocity of the sound line at the throat of the supersonic air film nozzle, v is the radial velocity of the sound 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 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 above parabolic potential function equation and the transonic solution of the supersonic air film nozzle is obtained.

[0078] In a supersonic film nozzle design method provided in one embodiment, a Mach number distribution along the axis of the supersonic film nozzle is set, using the transonic solution of the supersonic film nozzle and the nozzle outlet as the starting point and end point, respectively. Based on the boundary conditions of the transonic solution and the outlet of the supersonic film nozzle, i.e., using the transonic solution of the supersonic film nozzle and the nozzle outlet as the starting point and end point, respectively, the position coordinates, Mach number, and first derivative of the Mach number of the starting point and end point are substituted into a B-spline curve expression to obtain the Mach number distribution along the axis of the supersonic film nozzle.

[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 supersonic air film nozzle axis are used as boundary conditions. A 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. Then, based on the streamline control equation, the inviscid profile of the nozzle is determined. 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 supersonic air film nozzle, fundamentally eliminating 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 one 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] Furthermore, the boundary layer displacement thickness δ is obtained according to the above method. * Then, the nozzle non-viscous profile is increased by δ in the normal direction. * The distance is used to achieve viscosity correction.

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

[0094] In a preferred embodiment, a 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 a quintic curve as follows:

[0095]

[0096] where y i is the inlet height of the supersonic film nozzle, y * is the supersonic air 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 film nozzle is provided. The coordinates (x, y) of each point on the contraction profile of the supersonic film nozzle are determined by the Witoshinsky curve as follows:

[0098]

[0099] where y i is the inlet height of the supersonic film nozzle, y * is the supersonic air 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] Using the supersonic film nozzle design method provided in the above embodiment, a Mach 3 nozzle was designed. The tangential jet through the Mach 3 nozzle avoids direct friction between the hypersonic main flow and the waverider surface, 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 were controlled at a total temperature of 300K, a total pressure of 100,000Pa, a nozzle outlet temperature of 108K, and a pressure of 2820Pa. The design results are shown in Figure 1. 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. It can be seen that the flow accelerates smoothly from subsonic to supersonic, with complete wave dissipation inside the nozzle and no concentrated shock waves. The exit Mach number is very uniform, demonstrating the feasibility of this method.

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

[0102] The technical features of the above embodiments can 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-described embodiments merely represent several implementation methods of the present application. While the descriptions are relatively specific and detailed, they should not be construed as limiting the scope of the present invention. It should be noted that a person skilled in the art could make various modifications and improvements without departing from the spirit of the present application, all of which fall within the scope of protection of the present application. Therefore, the scope of protection of the present patent application shall be determined by the appended claims.

[0104] The foregoing description is merely a preferred embodiment of the present invention and is not intended to limit the present invention. Those skilled in the art will readily appreciate that the present invention is susceptible to various modifications and variations. Any modifications, equivalent substitutions, or improvements made within the spirit and principles of the present invention shall be included within the scope of protection of the present invention.

Claims

1. A supersonic air film distributed heat reduction method for a near-space waverider, characterized in that: 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; 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 an air supply pipeline for providing gas, and the gas source and the air supply pipeline are arranged inside the waverider; 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. 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 achieving coverage of the downstream of the waverider swept wing. The exchange of energy reduces the temperature of the hot incoming air flow, thereby achieving cooling of the downstream of the waverider swept wing. The supersonic air film nozzle is designed by the following steps, including: Obtain the supersonic air film nozzle outlet pressure under pressure matching conditions; Determine the supersonic air film mass flow rate and supersonic air film Mach number according to the cooling length requirements; The throat height of the supersonic film nozzle is determined according to the supersonic film mass flow rate and the supersonic film nozzle outlet pressure. Determine the nozzle exit height of the supersonic film nozzle based on the supersonic film nozzle throat height and the supersonic film Mach number; For the transonic flow in the throat of a supersonic film nozzle, the parabolic potential function equation is solved by using the series expansion method, and the transonic solution of the supersonic film nozzle is obtained. The Mach number distribution of the supersonic film nozzle axis is set with the transonic solution of the supersonic film nozzle and the nozzle outlet as the starting point and the end point respectively. Using the transonic solution of the supersonic film nozzle and the Mach number distribution at the nozzle exit and the axis of the supersonic film nozzle as boundary conditions, a characteristic line grid of the supersonic film nozzle is constructed to solve the supersonic flow field inside the supersonic film nozzle. Then, the inviscid profile of the nozzle is determined based on the streamline governing equations. According to the von Karman momentum integral relationship, the boundary layer displacement thickness 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 preset contraction section profile of the supersonic air film nozzle is connected with the nozzle profile at the throat to obtain a complete supersonic air film nozzle.

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 air flow, the pressure of the tangential supersonic air film generated by the first heat reduction device is equal to the 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 air flow.

3. The supersonic air film distributed heat reduction method for near-space waverider according to claim 1 or 2, characterized in that: Supersonic film nozzle throat height Determined by the following formula: In the formula is the specific heat ratio of the gas, R is the gas constant, T It is a static temperature. P is the supersonic film nozzle outlet pressure, is the supersonic air film Mach number.

4. The supersonic air film distributed heat reduction method for near-space waverider according to claim 3, characterized in that: Nozzle exit height of supersonic air film nozzle Determined by the following formula: In the formula is the supersonic film nozzle throat height, is the specific heat ratio of the gas, is the supersonic air film Mach number.

5. The supersonic air film distributed heat reduction method for a near-space waverider according to claim 1, 2 or 4, characterized in that: The parabolic potential function equation is: In the formula is the critical speed of sound of the throat of the supersonic air film nozzle. is the axial velocity of the sonic line at the throat of the supersonic film nozzle, is the radial velocity of the sound 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 line at the throat of the supersonic air film nozzle, respectively.

6. The supersonic air film distributed heat reduction method for near-space waverider according to claim 5, characterized in that: Boundary layer displacement thickness By combining the following formulas to solve: In the formula is the momentum loss thickness, is the boundary layer displacement thickness, is the wall angle, is the boundary layer shape factor, is the compressible friction coefficient, is the local Mach number, is the specific heat ratio of the gas.

7. The supersonic air film distributed heat reduction method for near-space waverider according to claim 6, characterized in that: The method of making viscosity correction for the nozzle non-viscous profile is: the nozzle non-viscous profile increases in the normal direction The distance is used to achieve viscosity correction.

8. The supersonic air film distributed heat reduction method for a near-space waverider according to claim 1, 2, 4, 6, or 7, characterized in that: Coordinates of each point on the contraction profile of the supersonic air film nozzle ( x , y ) is determined by the Witoshinsky curve as follows: in is the inlet height of the supersonic film nozzle, is the supersonic film nozzle throat height, is the given contraction length.

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