Heat reduction and stealth method for near-space waverider

Abstract

To address the aerodynamic heat problem of high-speed waverider aircraft, the present invention proposes a heat reduction and stealth method for near-space waveriders. 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. Through heat conduction between the supersonic air film and the waverider wall, the supersonic air film reduces the heat of the waverider wall. After being cooled by the supersonic air film, the radiation power per unit area of ​​the waverider surface is reduced by an order of magnitude, thereby enhancing the stealth performance of the waverider.

Description

Technical Field

[0001] The present 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] 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 high flight speed and low observability.

[0003] However, high-speed flight brings severe aerodynamic heat to the waverider, causing ablation and damage to the surface structure of the aircraft. Taking a Mach 6 waverider aircraft as an example, the atmospheric temperature at an altitude of 20km 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 100km is about 195K, the laminar Prandtl number is 0.7, and the adiabatic temperature of the aircraft surface is about 1340K. Figure 2 As shown. 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). Passive thermal protection is limited in effectiveness for high-speed aircraft. 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. Furthermore, ablation cannot reduce thermal radiation from the aircraft's surface.

[0004] Aerodynamic heating causes severe thermal radiation from the surface of high-speed waverider aircraft. The Stefan-Boltzmann law states that the radiated power is proportional to the fourth power of the aircraft surface temperature (Equation 1.1). For example, the emissivity of the aircraft surface is close to that of a blackbody, with a radiation energy flux density of 5.67 W / m at 100 K. 2 , at 1000K it is 5.67×10 4 W / m 2 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 emissivity, ε is the material emissivity, T is the temperature, and σ is the Stefan-Boltzmann constant.

[0007] Due to its excellent 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 optical window's tolerance to high temperatures is far lower than that of other waverider components. Aerodynamic heat can drown out the optical signal and damage the window due to overheating. However, due to the special characteristics of the optical window, passive ablation, transpiration cooling, and other thermal protection methods cannot be used. Instead, a supersonic cooling film must be placed at the leading edge of the window to prevent direct heating of the optical window by the high-speed mainstream airflow.

[0008] For the reasons mentioned above, 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 in the prior art, the present invention proposes a heat reduction and stealth method for near-space waveriders.

[0010] To achieve the above object, 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 achieving stealth for a near-space waverider. A supersonic air film generating device is disposed 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 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. After being cooled by the supersonic air film, the radiation power per unit area of ​​the waverider surface 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 injection 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 through 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.

[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 are passive ablation, convection 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 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 rate 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 aircraft surface and absorbs heat through phase change, but requires the use of special materials on the aircraft surface, which has a great impact on the structural strength. More importantly, it cannot be applied to optical windows. In response to the above-mentioned defects in the prior art, the present invention proposes a thermal reduction and stealth method for 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) The 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 achieve an order of magnitude reduction in the surface radiation power of the waverider aircraft, 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] Determine the supersonic air film mass flow rate and supersonic air film Mach number according to the cooling length requirements;

[0023] 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.

[0024] Determine the nozzle exit height of the supersonic film nozzle based on the supersonic film nozzle throat height and the supersonic film Mach number;

[0025] 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.

[0026] 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.

[0027] 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.

[0028] According to the von Karman momentum integral relationship, the boundary layer displacement thickness 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. The viscosity correction method is as follows: the viscosity correction is achieved by increasing the distance of the boundary layer displacement thickness in the normal direction of the nozzle inviscid profile.

[0030] 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.

[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, with 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 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.

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

[0037] Figure 2 This is a temperature distribution diagram of the aircraft surface in the background art at a Mach 6 waverider and a cruising altitude of 100 km;

[0038] Figure 3This is a schematic diagram of a supersonic nozzle generating a supersonic cooling film in one embodiment;

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

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

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

[0042] Figure 7 This 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 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, the present invention provides a method for reducing heat and achieving stealth for a near-space waverider. A supersonic air film generating device is disposed 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. Through heat conduction between the supersonic air film and the waverider wall, the supersonic air film reduces the heat of the waverider wall. After being cooled by the supersonic air film, the radiation power per unit area of ​​the waverider surface 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. An annular mounting area for the supersonic air film nozzles is provided on the surface of the waverider near the head of the waverider. The series of supersonic air film nozzles are arranged side by side within the mounting area, conforming to the surface of the waverider. The nozzles of each supersonic air film nozzle spray 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 via the air supply pipeline to provide injection gas to each supersonic air film nozzle. The air source and air supply pipeline are disposed within the waverider.

[0047] Through the above scheme, the supersonic air film isolates the high-temperature inflow of the waverider aircraft, preventing the high-temperature inflow from directly heating the waverider and the optical window. At the same time, the heat conduction between the supersonic air film and the wall of the waverider reduces the heat of the waverider wall, thereby achieving large-area cooling of the waverider surface. The four-bit unit area radiation power of the waverider surface after cooling by the supersonic air film 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 a supersonic air film in the supersonic air film generating device on the waverider 1. Generally, they are located downstream of the cooling area inside 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 restriction 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 of ​​the waverider where heat reduction is required. The supersonic air film nozzles are connected to the internal air source through pipelines, 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] This invention addresses the aerodynamic heating issues of high-speed waverider aircraft by employing a supersonic air film for active thermal protection, reducing surface thermal radiation and achieving both heat reduction and stealth. The invention deploys a Mach 3 supersonic air film on the surface of a Mach 6 cone-guided waverider aircraft, generated by a supersonic nozzle. Figure 3 A supersonic nozzle that generates a film of air was demonstrated. A high-pressure cooling air source is connected upstream of the nozzle, regulated by a pressure reducing valve to provide 0.2 MPa, 300K cooling air to the nozzle inlet. The cooling air passes through the perfectly wave-canceling nozzle profile, accelerating the sub- to trans-supersonic flow and further reducing the gas temperature, producing a uniform 100K supersonic cooling air film at the outlet, far cooler than the high-temperature incoming air.

[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 optical windows. 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 aircraft surface. Figure 4 As shown in the figure, it can be seen that the surface temperature of the waverider downstream of the nozzle changes from Figure 2 The supersonic air film covers the waverider's wall downstream, extending its effective coverage to the tail of the vehicle. During this process, despite the continuous mixing of the air film and the mainstream flow structure, and the continuous heating of the air film by the mainstream, the cooling effect still keeps the temperature below 800K at the tail of the vehicle, approximately 500K lower than that of a waverider without an air film.

[0054] The radiation power per unit area of ​​the high-speed waverider surface is calculated according to the Stefan-Boltzmann law. The results are as follows: Figure 5 and Figure 6 As shown. You can see, Figure 5 In the case of no supersonic film cooling, the surface temperature of the waverider without air film heat protection is very high, and 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 substantial temperature reduction on its surface, and its radiation power per unit area has been reduced 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 waverider's nose and swept wings. Combining supersonic air film with convection internal cooling to achieve combined cooling can ultimately effectively reduce the overall heat and achieve stealth for high-speed waverider aircraft.

[0056] The present invention has been numerically simulated, such as Figure 2 、 4 , 5 and 6. The present invention has undergone two rounds of high enthalpy wind tunnel tests and has achieved the expected design and is technically feasible.

[0057] In one embodiment, a supersonic air film nozzle design method is provided. The characteristic line method is used to solve the two-dimensional flow hyperbolic equation to design the nozzle. The wave is completely eliminated in the nozzle, fundamentally eliminating the shock wave. 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 use as a supersonic air film generator for surface heat reduction and stealth of high-speed waverider aircraft.

[0058] Specifically, it includes:

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

[0060] Determine the supersonic air film mass flow rate and supersonic air film Mach number according to the cooling length requirements;

[0061] 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.

[0062] Determine the nozzle exit height of the supersonic film nozzle based on the supersonic film nozzle throat height and the supersonic film Mach number;

[0063] 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.

[0064] 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.

[0065] 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.

[0066] According to the von Karman momentum integral relationship, the boundary layer displacement thickness 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. The viscosity correction method is as follows: the viscosity correction is achieved by increasing the distance of the boundary layer displacement thickness in the normal direction of the nozzle inviscid profile.

[0068] 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.

[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 film nozzle outlet pressure, and Me is the supersonic film Mach number.

[0073] 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:

[0074]

[0075] 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.

[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 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.

[0079] 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.

[0080] 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.

[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 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:

[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 one 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 * Then, the nozzle non-viscous profile is increased by δ in the normal direction. * The distance is used to achieve viscosity correction.

[0095] 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.

[0096] 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:

[0097]

[0098] 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.

[0099] 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:

[0100]

[0101] 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.

[0102] 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 a bicubic curve as follows:

[0103]

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

[0105] 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 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. 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.

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

[0107] 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.

[0108] 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.

[0109] 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 method for reducing heat and making a waverider invisible in near-space, characterized in that: A supersonic air film generating device is provided on the periphery of the waverider surface near the head of the waverider. The supersonic air film generating device includes a series of supersonic air film nozzles. The jet direction of the nozzle of each supersonic air film nozzle is in the same direction as the incoming air flow. The supersonic air film generated by the supersonic air film generating device can cover the downstream waverider wall. The high-temperature incoming 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. 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 that of the incoming air flow. 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 is performed on the nozzle inviscid profile based on the boundary layer displacement thickness to obtain the final nozzle profile. The viscosity correction method is as follows: the viscosity correction is achieved by increasing the distance of the boundary layer displacement thickness in the normal direction of the nozzle inviscid 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 heat reduction stealth method for a near-space waverider according to claim 1, characterized in that: A ring-shaped installation area for installing 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 accordance with the surface of the waverider.

3. The heat reduction stealth method for a near-space waverider according to claim 2, characterized in that: 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 through 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 method for reducing heat and making a near-space waverider invisible according to any one of claims 1 to 3, 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.

5. The heat reduction stealth method for a near-space waverider according to claim 4, 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.

6. The heat reduction and stealth method for a near-space waverider according to claim 5, characterized in that: The parabolic potential function equation is: In the formula is the critical speed of sound of the sonic line at 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.

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

8. The heat reduction stealth method for a near-space waverider according to claim 1, 2, 3, 5, 6, or 7, characterized in that: The coordinates of each point on the contraction line of the supersonic air film nozzle ( x , y ) is determined by the bicubic 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, is the relative position of the bicubic curve connection point, take .

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