Heat reduction stealth method for near space waverider

Abstract

In order to solve the problem of aerodynamic heat of a high-speed waverider aircraft, the invention provides a heat-reducing stealth method for a near-space waverider, which comprises the following steps: arranging a supersonic gas film generating device on the periphery of the surface of the waverider close to the head position of the waverider; the supersonic air film generated by the supersonic air film generating device can cover the wall surface of the waverider on the downstream portion of the supersonic air film, the high-temperature air incoming flow and the wall surface of the waverider are separated by the supersonic air film, and through heat conduction between the supersonic air film and the wall surface of the waverider, the supersonic air film cools heat of the wall surface of the waverider, so that the waverider is prevented from being damaged. The unit area radiation power of the waverider surface cooled by the supersonic air film is reduced in order of magnitude, and the stealth performance of the waverider is enhanced.

Description

Technical Field

[0001] The invention mainly relates to the technical field of high-speed waverider aircraft, in particular to a heat reduction and stealth 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 high flight speed and low detectability.

[0003] However, high-speed flight brings severe aerodynamic heat to the waverider, causing ablation and destruction of the aircraft surface structure. Taking the Mach 6 waverider aircraft as an example, the atmospheric temperature at an altitude of 20 km is about 216K, the turbulent Prandtl number is 0.9, and the adiabatic temperature of the aircraft surface is nearly 1720K; the atmospheric temperature at an altitude of 100 km is about 195K, and the laminar Prandtl number is 0.7, so the adiabatic temperature of the aircraft surface is about 1340K. Figure 2 As shown. At present, ablation is usually used for thermal protection, which is passive thermal protection and is suitable for ultra-high altitude (low temperature) and low Mach number (low speed) aircraft. For high-speed aircraft, passive thermal protection has limited effect. For example, the most serious accident in aviation history, the Columbia Space Shuttle, was caused by cracks in the carbon-carbon insulation board at the leading edge of the aircraft wing, which led to ablation and disintegration of the aircraft when it re-entered the atmosphere. At the same time, ablation thermal protection cannot reduce the thermal radiation on the surface of the aircraft.

[0004] Aerodynamic heat causes severe thermal radiation on the surface of high-speed waverider aircraft. The Stefan-Boltzmann law shows that the radiation power is proportional to the fourth power of the aircraft surface temperature (Equation 1.1). For example, the radiation coefficient of the aircraft surface is close to that of a black body, and the radiation energy flux density at 100K is 5.67W / m 2 , at 1000K it is 5.67×10 4 W / m 2 However, the surface temperature of a high-speed waverider aircraft is far more than 1000K, which makes the aircraft easily detected by the enemy and greatly reduces its penetration capability.

[0005] j * =εσT 4

[0006] Where j * is the radiance, ε is the material emissivity, T is the temperature, and σ is the Stefan-Boltzmann constant.

[0007] Due to its good imaging and anti-interference capabilities, imaging guidance is currently one of the mainstream guidance methods for high-speed aircraft to achieve precision strikes. However, the tolerance of optical windows to high temperatures is much lower than that of other components of waverider aircraft. Aerodynamic heat can cause the optical signal to be submerged, and the window to overheat and be damaged. However, due to the particularity of the optical window, passive ablation, sweating cooling and other thermal protection methods cannot be used. It is necessary to arrange a supersonic cooling air film at the front of the window to avoid direct heating of the optical window by the high-speed mainstream.

[0008] For the above reasons, thermal protection will become one of the bottlenecks restricting the development of high-speed waverider aircraft, and it is particularly important to carry out thermal protection design for high-speed waverider aircraft. Summary of the invention

[0009] In view of the above-mentioned defects existing in the prior art, the present invention proposes a heat reduction and stealth method for a near-space waverider.

[0010] To achieve the above purpose, the technical solution adopted by the present invention is as follows:

[0011] On the one hand, the present invention provides a method for reducing heat and making a waverider in near-space invisible. A supersonic air film generating device is arranged on the periphery of the waverider surface near the head of the waverider. The supersonic air film generated by the supersonic air film generating device can cover the downstream wall of the waverider. The high-temperature air flow and the wall of the waverider are separated by the supersonic air film. The supersonic air film reduces the heat of the wall of the waverider through heat conduction between the supersonic air film and the wall of the waverider. The radiation power per unit area of ​​the waverider surface after being cooled by the supersonic air film is reduced by an order of magnitude, thereby enhancing the stealth performance of the waverider.

[0012] Furthermore, the supersonic air film generating device includes a series of supersonic air film nozzles. A ring-shaped installation area for installing the supersonic air film nozzles is provided on the surface of the waverider near the head of the waverider. A series of supersonic air film nozzles are arranged side by side in the installation area in contact with the surface of the waverider, and the spray direction of the nozzle of each supersonic air film nozzle is in the same direction as the incoming air flow.

[0013] Furthermore, the supersonic air film generating device includes an air source and an air supply pipeline, the air source is connected to each supersonic air film nozzle via the air supply pipeline to provide injection gas for each supersonic air film nozzle, and the air source and the air supply pipeline are arranged inside the waverider.

[0014] Furthermore, the flow direction of the supersonic air film generated by the supersonic air film generating device is parallel to the incoming air flow and the pressure of the supersonic air film is equal to the incoming air flow.

[0015] At present, the main thermal protection methods for high-speed aircraft include passive ablation, convection internal cooling, sweat cooling and film cooling. Among them: (1) Passive ablation is difficult to independently meet all the thermal protection requirements of the surface of high-speed waverider aircraft, cannot be applied to the optical window of the seeker, and cannot reduce the thermal radiation of the waverider surface. (2) Convection internal cooling uses the cooling medium to conduct heat under the aircraft skin and removes heat through internal circulation. It is suitable for cooling the aircraft's stationary point, head and swept wings, but the required cooling medium flow is too large and the efficiency is low, so it cannot be used on a large area of ​​the aircraft. (3) Sweating cooling seeps through small holes on the surface of the aircraft and absorbs heat through phase change, but 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. In view of the above-mentioned defects in the prior art, the present invention proposes a method for reducing heat and making it invisible in near-space waveriders. Compared with the prior art, it has the following advantages:

[0016] (1) Supersonic air film is suitable for large-area heat reduction in high-speed waverider aircraft;

[0017] (2) Supersonic air film has a large heat reduction range;

[0018] (3) The supersonic air film generated by the supersonic air film generating device can cover the downstream waverider wall. The high-temperature air flow and the waverider wall are separated by the supersonic air film. In this way, the supersonic air film is used to isolate the high-temperature mainstream, reduce heat, and reduce the radiation power of the waverider surface by an order of magnitude, thereby enhancing the stealth performance of the waverider.

[0019] (4) Supersonic air film has no requirements on aircraft materials and is suitable for optical windows.

[0020] Furthermore, the supersonic air film nozzle is designed by the following steps, including:

[0021] Obtain the supersonic air film nozzle outlet pressure under pressure matching conditions;

[0022] According to the cooling length requirements, the supersonic air film mass flow rate and supersonic air film Mach number are determined;

[0023] 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;

[0024] 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;

[0025] 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.

[0026] 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.

[0027] 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.

[0028] According to the von Karman momentum integral relation, the displacement thickness of the boundary layer is solved;

[0029] The viscosity correction is performed on the nozzle inviscid profile based on the boundary layer displacement thickness to obtain the final nozzle profile, wherein the viscosity correction method is: the nozzle inviscid profile is increased in the normal direction by the distance of the boundary layer displacement thickness to achieve viscosity correction;

[0030] 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.

[0031] The supersonic air film nozzle designed by the above method has the following advantages:

[0032] (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.

[0033] (2) The supersonic air film nozzle has been viscosity corrected, and the uniformity of the internal flow and nozzle outlet is very high.

[0034] (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

[0035] 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.

[0036] Figure 1 It is a schematic diagram of the arrangement position of a supersonic air film nozzle in a supersonic air film generating device on a waverider in one embodiment;

[0037] Figure 2 It is the surface temperature distribution diagram of the aircraft in the background technology under the Mach 6 waverider-100km cruising state;

[0038] Figure 3Schematic diagram of a supersonic nozzle generating a supersonic cooling air film in one embodiment;

[0039] Figure 4 This is a temperature distribution diagram of the aircraft surface of a Mach 6 waverider with a supersonic air film in a cruising state at an altitude of 100 km in one embodiment;

[0040] Figure 5 It is a schematic diagram of the radiation power per unit area of ​​an aircraft in a Mach 6 waverider-100km cruising state without supersonic film cooling in one embodiment;

[0041] Figure 6 It is a schematic diagram of the radiation power per unit area of ​​an aircraft in a Mach 6 waverider-100km cruising state with supersonic film cooling in one embodiment;

[0042] Figure 7 is a nozzle Mach number cloud diagram in one embodiment;

[0043] Figure 8 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, the present invention provides a method for reducing heat and making a waverider in near space invisible. A supersonic air film generating device is arranged on the periphery of the waverider surface near the head of the waverider. The supersonic air film generated by the supersonic air film generating device can cover the downstream wall of the waverider. The high-temperature air flow and the wall of the waverider are separated by the supersonic air film. The supersonic air film reduces the heat of the wall of the waverider through heat conduction between the supersonic air film and the wall of the waverider. The radiation power per unit area of ​​the waverider surface after being cooled by the supersonic air film is reduced by an order of magnitude, thereby enhancing the stealth performance of the waverider.

[0046] In one embodiment, a supersonic air film generating device is provided: the supersonic air film generating device includes a series of supersonic air film nozzles, a ring-shaped installation area for installing the supersonic air film nozzles is provided on the surface of the waverider near the head of the waverider, a series of supersonic air film nozzles are arranged side by side in the installation area in close contact with the surface of the waverider, and the spray direction of the nozzle of each supersonic air film nozzle is in the same direction as the incoming air flow. The supersonic air film generating device also includes an air source and an air supply pipeline, the air source is connected to each supersonic air film nozzle through the air supply pipeline, and provides spray gas for each supersonic air film nozzle, and the air source and the air supply pipeline are arranged inside the waverider.

[0047] Through the above scheme, the supersonic air film isolates the high-temperature inflow of the waverider aircraft, avoiding direct heating of the waverider and the optical window by the high-temperature inflow. At the same time, the heat of the waverider wall is reduced through heat conduction between the supersonic air film and the wall of the waverider, thereby achieving large-area cooling of the waverider surface. The four-bit unit area radiation power of the waverider surface after supersonic air film cooling is reduced by an order of magnitude, thereby enhancing the stealth performance of the waverider.

[0048] There is no strict limitation on the installation position of a series of supersonic air film nozzles for spraying and forming supersonic air film in the supersonic air film generating device on the waverider 1. Generally, they are located downstream of the inner cooling area of ​​the waverider head 2, such as Figure 1 As shown, the installation area 3 for installing the supersonic air film nozzle is located at a distance from the waverider head 2 that is 1 / 5 of the length of the waverider.

[0049] The flow direction of the supersonic air film generated by the supersonic air film generating device is parallel to the incoming air flow, and the pressure of the supersonic air film is equal to the incoming air flow.

[0050] It can be understood that the number of supersonic air film nozzles is not limited. In addition, there is no limitation on whether the supersonic air film nozzles are arranged in a full ring around the surface of the waverider or only on the upper surface of the waverider. Those skilled in the art can set it according to the specific shape of the waverider and the heat reduction requirements. A series of supersonic air film nozzles for generating supersonic air film can be arranged upstream of the area where the waverider needs to cool down. The supersonic air film nozzles are connected to the internal air source through a pipeline, and can be turned on to generate supersonic air film when heat reduction is required to achieve the purpose of heat reduction.

[0051] A supersonic cooling air film is generated through a supersonic air film nozzle, and the cooling gas covers the surface of the waverider. It has a large coverage area and a stable cooling effect. It is suitable for optical windows and has no requirements for the structural materials of the aircraft. Combined cooling can be performed to achieve distributed large-area heat reduction.

[0052] The present invention aims at the aerodynamic heat problem of high-speed waverider aircraft, adopts supersonic air film for active thermal protection, reduces the thermal radiation on the aircraft surface, and achieves the purpose of heat reduction and stealth. The present invention arranges Mach 3 supersonic air film on the surface of Mach 6 cone-guided waverider aircraft, and the air film is generated by a supersonic nozzle. Figure 3 The supersonic nozzle that produces air film was demonstrated. The upstream of the nozzle is connected to a high-pressure cooling gas source, which is regulated by a pressure reducing valve to provide 0.2MPa, 300K cooling gas to the nozzle inlet. The cooling gas passes through the nozzle profile with perfect wave elimination to achieve sub-trans-supersonic flow acceleration, further reducing the gas temperature and generating a uniform 100K supersonic cooling air film at the outlet, with a temperature much lower than the high-temperature incoming flow.

[0053] On the one hand, the supersonic air film isolates the high-temperature incoming flow of the waverider aircraft, preventing the high-temperature incoming flow from directly heating the waverider body and the optical window. On the other hand, the heat conduction between the supersonic air film and the aircraft wall cools the wall surface, thereby achieving large-area cooling of the waverider surface. Figure 4 As shown in Figure 2, it can be seen that the surface temperature of the waverider downstream of the nozzle changes from Figure 2 The temperature of the tail of the aircraft is reduced from 1300K to 400K, with a significant heat reduction effect. The supersonic 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-mainstream flow structure is constantly mixed and the air film is continuously heated by the mainstream, the cooling effect can still control the temperature below 800K at the tail of the aircraft, which is about 500K lower than that of the waverider aircraft without an air film.

[0054] The radiation power per unit area on the surface of a high-speed waverider is calculated according to the Stefan-Boltzmann law. The results are as follows: Figure 5 and Figure 6 As shown. Figure 5 In the case of no supersonic air film cooling, due to the high surface temperature of the waverider without air film heat protection, its radiation power per unit area is generally around 1.7×10 5 W / m 2 , the penetration concealment is poor. Figure 6 The high-speed waverider with supersonic film cooling has a large-scale and large-scale cooling on the surface, and its radiation power per unit area has dropped by an order of magnitude. Downstream of the nozzle, the radiation power can be reduced to 1000W / m 2 ; At the tail of the waverider, the radiation power can be suppressed to 1.7×10 4 W / m 2 This greatly enhances the stealth performance of the waverider.

[0055] Convective internal cooling can suppress the temperature and radiation power of the head and swept wings of the waverider aircraft. Combining supersonic air film with convection internal cooling to achieve combined cooling can ultimately effectively achieve overall heat reduction and stealth of high-speed waverider aircraft.

[0056] The present invention has been simulated by numerical simulation, such as Figure 2 , 4 , 5 and 6. The present invention has been subjected to two rounds of high enthalpy wind tunnel tests, achieving the expected design and being technically feasible.

[0057] In one embodiment, a supersonic air film nozzle design method is provided, in which the characteristic line method is used to solve the two-dimensional flow hyperbolic equation to design the nozzle. The wave is completely eliminated inside the nozzle, and the shock wave is fundamentally eliminated. The distribution of flow field parameters such as pressure, temperature, and velocity is controllable, and viscosity correction can be performed based on the von Karman momentum equation. The method is suitable for a supersonic air film generator used for heat reduction and stealth of a high-speed waverider aircraft surface.

[0058] Specifically, it includes:

[0059] Obtain the supersonic air film nozzle outlet pressure under pressure matching conditions;

[0060] According to the cooling length requirements, the supersonic air film mass flow rate and supersonic air film Mach number are determined;

[0061] 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;

[0062] 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;

[0063] 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.

[0064] 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.

[0065] 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.

[0066] According to the von Karman momentum integral relation, the displacement thickness of the boundary layer is solved;

[0067] The viscosity correction is performed on the nozzle inviscid profile based on the boundary layer displacement thickness to obtain the final nozzle profile, wherein the viscosity correction method is: the nozzle inviscid profile is increased in the normal direction by the distance of the boundary layer displacement thickness to achieve viscosity correction;

[0068] 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.

[0069] 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.

[0070] 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:

[0071]

[0072] 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.

[0073] 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:

[0074]

[0075] 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.

[0076] In a supersonic air film nozzle design method provided in one embodiment, the parabolic potential function equation used is:

[0077]

[0078] 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.

[0079] 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.

[0080] 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.

[0081] 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:

[0082] Characteristic line method equation 1:

[0083]

[0084] Characteristic line method equation 2:

[0085]

[0086] 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.

[0087] The streamline control equation is: dy / dx=v / u

[0088] Where u is the axial velocity and v is the radial velocity.

[0089] 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: * :

[0090]

[0091]

[0092]

[0093] 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.

[0094] According to the above method, the boundary layer displacement thickness δ is obtained * After that, the nozzle non-viscous profile is increased by δ in the normal direction. * The distance is used to achieve viscosity correction.

[0095] 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.

[0096] 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:

[0097]

[0098] 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.

[0099] 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:

[0100]

[0101] 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.

[0102] 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 a bicubic curve, as follows:

[0103]

[0104] where y i is the inlet height of the supersonic film nozzle, y * is the supersonic film nozzle throat height, L cis the given contraction length, x m is the relative position of the bicubic curve connection point, taking x m =0.5.

[0105] 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 7 , Figure 8 As shown, Figure 7 It is the nozzle Mach number cloud diagram; Figure 8 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.

[0106] Matters not covered by the present invention are known technologies.

[0107] 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.

[0108] 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.

[0109] 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 method for reducing heat and making a waverider invisible in near space, characterized in that: A supersonic air film generating device is arranged on the outer periphery of the waverider surface near the head of the waverider. The supersonic air film generated by the supersonic air film generating device can cover the downstream waverider wall. The high-temperature air flow and the waverider wall are separated by the supersonic air film. The supersonic air film reduces the heat of the waverider wall through heat conduction between the supersonic air film and the waverider wall. The radiation power per unit area of ​​the waverider surface after being cooled by the supersonic air film is reduced by an order of magnitude, thereby enhancing the stealth performance of the waverider.

2. The heat reduction stealth method of a near-space waverider according to claim 1, characterized in that: The supersonic air film generating device includes a series of supersonic air film nozzles. A ring-shaped installation area for installing the supersonic air film nozzles is provided on the surface of the waverider near the head of the waverider. A series of supersonic air film nozzles are arranged side by side in the installation area in contact with the surface of the waverider, and the spray direction of the nozzle of each supersonic air film nozzle is in the same direction as the incoming air flow.

3. The heat reduction stealth method of near-space waverider according to claim 2, characterized in that: The supersonic air film generating device comprises an air source and an air supply pipeline. The air source is connected to each supersonic air film nozzle via the air supply pipeline to provide injection gas for each supersonic air film nozzle. The air source and the air supply pipeline are arranged inside the waverider.

4. The heat reduction stealth method for near-space waverider according to claim 1, 2 or 3, characterized in that: The flow direction of the supersonic air film generated by the supersonic air film generating device is parallel to the incoming air flow, and the pressure of the supersonic air film is equal to the incoming air flow.

5. The heat reduction stealth method of near-space waverider according to claim 4, 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 is performed on the nozzle inviscid profile based on the boundary layer displacement thickness to obtain the final nozzle profile, wherein the viscosity correction method is: the nozzle inviscid profile is increased in the normal direction by the distance of the boundary layer displacement thickness to achieve viscosity correction; 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.

6. The heat reduction and stealth method of a near-space waverider according to claim 5, 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.

7. The heat reduction and stealth method of a near-space waverider according to claim 5, 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.

8. The heat reduction stealth method of a near-space waverider according to claim 5, 6 or 7, 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.

9. The heat reduction and stealth method of a near-space waverider according to claim 8, 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.

10. The supersonic air film distributed heat reduction method for near-space waverider according to claim 5, 6, 7 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 bicubic 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, x m is the relative position of the bicubic curve connection point, taking x m =0.5.

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