Supersonic air film nozzle and design method, device, and near-space waverider aircraft

Abstract

The present invention proposes a supersonic air film nozzle and a design method, device, and near-space waverider aircraft, including obtaining the supersonic air film nozzle outlet pressure of the high-speed waverider aircraft under pressure matching conditions, determining the supersonic air film mass flow rate, the supersonic air film Mach number, the supersonic air film nozzle throat height, and the nozzle outlet height of the supersonic air film nozzle, obtaining the transonic solution of the supersonic air film nozzle, and setting the Mach number distribution of the supersonic air film nozzle axis. The supersonic air film nozzle transonic solution, the nozzle outlet, and the Mach number distribution of the supersonic air film nozzle axis are used as boundary conditions to determine the nozzle inviscid profile, solve the boundary layer displacement thickness, obtain the final nozzle profile, and then obtain a complete supersonic air film nozzle. The wave is completely eliminated in the nozzle, shock waves are fundamentally eliminated, and the distribution of flow field parameters such as pressure, temperature, and velocity is controllable.

Description

Technical Field

[0001] The present invention mainly relates to the technical field of near-space waverider aircraft, in particular to a supersonic air film nozzle and a design method and device, and a near-space waverider aircraft. Background Art

[0002] High-speed waverider aircraft are currently a hot topic of research worldwide due to their high altitude, high speed, and "water-skimming" trajectory, which makes them difficult to intercept. However, high-speed flight causes severe aerodynamic heating in the waverider, leading to ablation and damage to the vehicle's surface structures and optical windows. For example, a Mach 6 waverider aircraft has an atmospheric temperature of approximately 195K at 100km altitude and a laminar Prandtl number of 0.7, resulting in an adiabatic surface temperature of approximately 1340K.

[0003] Covering a high-speed waverider aircraft with a cooling air film is an effective way to improve infrared stealth performance and the only way to prevent optical window burnout. Currently, supersonic air films on high-speed aircraft are typically placed in front of the optical window. Supersonic air films used for infrared stealth, on the other hand, are placed within the aircraft's fuselage, resulting in wider nozzles and a greater effective cooling area. Both utilize wall-mounted nozzles.

[0004] Current supersonic air film nozzles are generally designed based on the basic theory of one-dimensional nozzle flow, such as Figure 1 As shown, Figure 1 A * is the throat area, A e is the outlet area, P e is the outlet pressure, M e The main design parameters for the outlet Mach number are the nozzle Mach number, air film height, total temperature and pressure, and mass flow rate. The implementation method is as follows:

[0005] First, relying on prior experiments or empirical curves, the supersonic air film Mach number and the air film mass flow rate required for the cooling length of the aircraft surface are determined.

[0006] Under the total temperature and total pressure conditions of the air source carried by the aircraft, the air film height and throat height at the nozzle outlet are determined by the air film mass flow rate, and the static temperature and static pressure at the nozzle outlet are determined by the isentropic relationship.

[0007] Given the nozzle length and setting the Mach number distribution along the nozzle axis, the nozzle height corresponding to the Mach number is determined according to the formula of Mach number and area ratio.

[0008] A supersonic nozzle is obtained by smoothly connecting a series of nozzle height profile points, which is arranged on the surface of the aircraft to generate a supersonic air film.

[0009] However, current supersonic film nozzles have several shortcomings. First, their design theory is flawed. Nozzle flow is inherently not one-dimensional, but at least two-dimensional. This is especially true for curved nozzles, where the flow is three-dimensional, or a two-dimensional flow theory combined with transverse curvature correction is employed. Existing supersonic film nozzles are designed using one-dimensional flow theory, making it impossible for the nozzle wall to perfectly eliminate waves. In other words, the internal flow channel inevitably contains concentrated compression waves. This results in low nozzle exit uniformity, exacerbating mixing with the mainstream after ejection and reducing thermal performance. Furthermore, the presence of weak shock waves leads to isentropic flow, resulting in inconsistent static temperature and pressure at the nozzle exit compared to the theoretical results of one-dimensional flow. This makes it difficult to effectively match the film pressure with the vehicle's mainstream pressure. This pressure mismatch leads to wave structure, which in turn enhances mixing and increases optical distortion. Summary of the Invention

[0010] In response to the problems existing in the prior art, the present invention proposes a supersonic air film nozzle and a design method, a device, and a near-space waverider aircraft.

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

[0012] In one aspect, the present invention provides a method for designing a supersonic air film nozzle, comprising:

[0013] Obtain the supersonic film nozzle outlet pressure of a high-speed waverider vehicle under pressure matching conditions;

[0014] 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;

[0015] 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.

[0016] Determine the nozzle exit height of the supersonic film nozzle based on the supersonic film nozzle throat height and the supersonic film Mach number;

[0017] 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.

[0018] 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.

[0019] 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.

[0020] According to the von Karman momentum integral relationship, the boundary layer displacement thickness is solved;

[0021] The viscosity correction of the nozzle inviscid profile is performed based on the boundary layer displacement thickness to obtain the final nozzle profile;

[0022] 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.

[0023] In another aspect, the present invention provides a supersonic air film nozzle design device, comprising:

[0024] The first module is used to obtain the supersonic air film nozzle outlet pressure of the high-speed waverider aircraft under pressure matching conditions;

[0025] The second module is used to determine the supersonic air film mass flow rate and supersonic air film Mach number based on the cooling length requirements of the high-speed waverider aircraft;

[0026] The third module is used to determine the supersonic air film nozzle throat height according to the supersonic air film mass flow rate and the supersonic air film nozzle outlet pressure;

[0027] The fourth module is used to determine the nozzle outlet height of the supersonic air film nozzle according to the supersonic air film nozzle throat height and the supersonic air film Mach number;

[0028] The fifth module is used to solve the parabolic potential function equation for the transonic flow at the throat of the supersonic film nozzle using the series expansion method to obtain the transonic solution of the supersonic film nozzle;

[0029] The sixth module is used to set the Mach number distribution of the supersonic film nozzle axis with the supersonic film nozzle transonic solution and the nozzle outlet as the starting point and end point respectively;

[0030] The seventh module is used to construct a supersonic film nozzle characteristic line grid 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, solve the supersonic flow field inside the supersonic film nozzle, and then determine the nozzle inviscid profile based on the streamline control equation;

[0031] The eighth module is used to solve the boundary layer displacement thickness based on the von Karman momentum integral relationship;

[0032] The ninth module is used to perform viscosity correction on the nozzle inviscid profile based on the boundary layer displacement thickness to obtain the final nozzle profile;

[0033] The tenth module is used to connect the preset contraction section profile of the supersonic air film nozzle with the nozzle profile at the throat to obtain a complete supersonic air film nozzle.

[0034] On the other hand, the present invention provides a supersonic air film nozzle designed using the above-mentioned supersonic air film nozzle design method.

[0035] On the other hand, the present invention provides a near-space waverider aircraft, on the surface of which a supersonic air film nozzle is arranged. The supersonic air film nozzle is designed using the above-mentioned supersonic air film nozzle design method, and the supersonic air film nozzle is used to generate a supersonic air film.

[0036] Compared with the prior art, the technical effects of the present invention are at least reflected in the following aspects:

[0037] (1) The supersonic air film nozzle designed by the method of the present invention has perfect internal flow wave elimination and no concentrated compression wave or shock wave exists.

[0038] (2) The nozzle has been viscosity corrected, and the uniformity of the internal flow and nozzle outlet is very high.

[0039] (3) The nozzle outlet pressure is consistent with the theoretical value, which can achieve pressure matching between the supersonic air film and the high-speed waverider aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0040] 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.

[0041] Figure 1 This is a schematic diagram of a supersonic air film nozzle designed based on the basic theory of one-dimensional nozzle flow in the prior art;

[0042] Figure 2 is a flow chart of an embodiment of the present invention;

[0043] Figure 3 is a schematic diagram of a transonic solution of a supersonic nozzle in one embodiment of the present invention;

[0044] Figure 4 is a nozzle axis Mach number distribution diagram in one embodiment;

[0045] Figure 5 1. It is a schematic diagram of a nozzle characteristic line network and a nozzle inviscid profile line in one embodiment;

[0046] Figure 6 This is a nozzle Mach number cloud diagram in one embodiment;

[0047] Figure 7 1 is a nozzle outlet Mach number distribution diagram in one embodiment. DETAILED DESCRIPTION

[0048] 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.

[0049] In one embodiment, a method for designing a supersonic air film nozzle for a near-space waverider aircraft is provided. The characteristic line method is used to solve the two-dimensional hyperbolic flow equation to design the supersonic air film nozzle. The designed supersonic air film nozzle completely eliminates internal waves, fundamentally eliminating shock waves. The distribution of flow field parameters such as pressure, temperature, and velocity are controllable, and viscosity correction can be performed based on the von Karman momentum equation. The nozzle is suitable for use as a supersonic air film generator for surface heat reduction and stealth of high-speed waverider aircraft.

[0050] Specifically, refer to Figure 2 In one embodiment, a method for designing a supersonic air film nozzle is provided, comprising:

[0051] Obtain the supersonic film nozzle outlet pressure of a high-speed waverider vehicle under pressure matching conditions;

[0052] 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;

[0053] 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.

[0054] Determine the nozzle exit height of the supersonic film nozzle based on the supersonic film nozzle throat height and the supersonic film Mach number;

[0055] 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.

[0056] 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.

[0057] 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.

[0058] According to the von Karman momentum integral relationship, the boundary layer displacement thickness is solved;

[0059] The viscosity correction of the nozzle inviscid profile is performed based on the boundary layer displacement thickness to obtain the final nozzle profile;

[0060] 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.

[0061] It can be understood that the pressure matching condition in the present invention means that the mainstream pressure is equal to the air film pressure.

[0062] 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:

[0063]

[0064] 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.

[0065] 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:

[0066]

[0067] 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.

[0068] In a supersonic air film nozzle design method provided in one embodiment, the parabolic potential function equation used is:

[0069]

[0070] 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.

[0071] For the transonic flow in the throat of the 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, as shown in the following example: Figure 3 shown. Figure 3 is the throat of the supersonic film nozzle, y* is the throat height, and the dashed line is the sonic line, obtained by solving the transonic solution for the supersonic film nozzle. The Mach number of the sonic line is 1. The area upstream (left) of the sonic line is the subsonic region, with a Mach number less than 1. The area downstream (right) of the sonic line is the supersonic region, with a Mach number greater than 1.

[0072] In a supersonic air film nozzle design method provided in one 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, such as Figure 4 As shown in the figure, according to the boundary conditions of the transonic solution and the exit boundary conditions of the supersonic air film nozzle, that is, the transonic solution and the nozzle exit of the supersonic air film nozzle are taken 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.

[0073] In a supersonic film nozzle design method provided in one embodiment, the transonic solution of the supersonic film nozzle and the Mach number distribution at the nozzle outlet and the axis of the supersonic film nozzle are used as boundary conditions. A characteristic line grid of the supersonic film nozzle is constructed based on the characteristic line method. The supersonic flow field inside the supersonic film nozzle is solved, and then the inviscid profile of the nozzle is determined based on the streamline control equation, such as Figure 5 The thick black solid line shown represents the nozzle's inviscid profile. The characteristic line method is a commonly used technique in this field. It solves the two-dimensional hyperbolic equations for designing supersonic film nozzles. This two-dimensional supersonic film nozzle achieves complete internal wave cancellation, fundamentally eliminating shock waves and making the distribution of flow field parameters such as pressure, temperature, and velocity controllable. Two characteristic line equations are given below:

[0074] Characteristic line method equation 1:

[0075]

[0076] Characteristic line method equation 2:

[0077]

[0078] 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.

[0079] The streamline control equation is: dy / dx=v / u

[0080] Where u is the axial velocity and v is the radial velocity.

[0081] 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: * :

[0082]

[0083]

[0084]

[0085] 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.

[0086] 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.

[0087] 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.

[0088] 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:

[0089]

[0090] 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.

[0091] To verify the effectiveness of the supersonic air film nozzle design method provided by the present invention, a simulation example based on the supersonic air film nozzle design method provided by the present invention is provided below:

[0092] The supersonic film nozzle design method provided by the present invention is used to design a Mach 3 nozzle. The tangential jet through the Mach 3 nozzle avoids direct friction between the hypersonic mainstream 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 are 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 as follows: Figure 6 、 Figure 7 As shown, Figure 6 It is the nozzle Mach number cloud diagram; Figure 7 This is a diagram of the nozzle exit Mach number distribution. It shows that the flow accelerates smoothly from subsonic to supersonic speeds, with complete wave dissipation within the nozzle and the absence of any concentrated shock waves. The Mach number at the exit is remarkably uniform, demonstrating the feasibility of this invention. This invention has been wind tunnel tested by our unit and has achieved the intended design and is technically feasible.

[0093] In one embodiment, a supersonic air film nozzle design device is provided, comprising:

[0094] The first module is used to obtain the supersonic air film nozzle outlet pressure of the high-speed waverider aircraft under pressure matching conditions;

[0095] The second module is used to determine the supersonic air film mass flow rate and supersonic air film Mach number based on the cooling length requirements of the high-speed waverider aircraft;

[0096] The third module is used to determine the supersonic air film nozzle throat height according to the supersonic air film mass flow rate and the supersonic air film nozzle outlet pressure;

[0097] The fourth module is used to determine the nozzle outlet height of the supersonic air film nozzle according to the supersonic air film nozzle throat height and the supersonic air film Mach number;

[0098] The fifth module is used to solve the parabolic potential function equation for the transonic flow at the throat of the supersonic film nozzle using the series expansion method to obtain the transonic solution of the supersonic film nozzle;

[0099] The sixth module is used to set the Mach number distribution of the supersonic film nozzle axis with the supersonic film nozzle transonic solution and the nozzle outlet as the starting point and end point respectively;

[0100] The seventh module is used to construct a supersonic film nozzle characteristic line grid 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, solve the supersonic flow field inside the supersonic film nozzle, and then determine the nozzle inviscid profile based on the streamline control equation;

[0101] The eighth module is used to solve the boundary layer displacement thickness based on the von Karman momentum integral relationship;

[0102] The ninth module is used to perform viscosity correction on the nozzle inviscid profile based on the boundary layer displacement thickness to obtain the final nozzle profile;

[0103] The tenth module is used to connect the preset contraction section profile of the supersonic air film nozzle with the nozzle profile at the throat to obtain a complete supersonic air film nozzle.

[0104] The implementation methods of the above modules and the construction of the model can adopt the methods described in any of the above embodiments, which will not be repeated here.

[0105] On the other hand, the present invention provides a computer device comprising a memory and a processor, wherein the memory stores a computer program, and when the processor executes the computer program, the steps of the supersonic air film nozzle design method provided in any of the above embodiments are implemented. The computer device may be a server. The computer device comprises a processor, a memory, a network interface, and a database connected via a system bus. The processor of the computer device is used to provide computing and control capabilities. The memory of the computer device comprises a non-volatile storage medium and an internal memory. The non-volatile storage medium stores an operating system, a computer program, and a database. The internal memory provides an environment for the operation of the operating system and the computer program in the non-volatile storage medium. The database of the computer device is used to store sample data. The network interface of the computer device is used to communicate with an external terminal via a network connection.

[0106] On the other hand, the present invention provides a computer-readable storage medium having a computer program stored thereon, which, when executed by a processor, implements the steps of the supersonic air film nozzle design method provided in any of the above embodiments.

[0107] Those skilled in the art will appreciate that all or part of the processes in the above-mentioned embodiments can be implemented by instructing the relevant hardware through a computer program. The computer program can be stored in a non-volatile computer-readable storage medium. When the computer program is executed, it can include the processes of the embodiments of the above-mentioned methods. Among them, any reference to memory, storage, database or other media used in the embodiments provided in this application can include non-volatile and / or volatile memory. Non-volatile memory can include read-only memory (ROM), programmable ROM (PROM), electrically programmable ROM (EPROM), electrically erasable programmable ROM (EEPROM) or flash memory. Volatile memory can include random access memory (RAM) or external cache memory. By way of illustration and not limitation, RAM is available in various forms, such as static RAM (SRAM), dynamic RAM (DRAM), synchronous DRAM (SDRAM), double data rate SDRAM (DDRSDRAM), enhanced SDRAM (ESDRAM), synchronous link (Synchlink) DRAM (SLDRAM), memory bus (Rambus) direct RAM (RDRAM), direct memory bus dynamic RAM (DRDRAM), and memory bus dynamic RAM (RDRAM).

[0108] In one embodiment, a supersonic air film nozzle is provided, which is designed using the supersonic air film nozzle design method provided in any of the above embodiments.

[0109] In one embodiment, a near-space waverider aircraft is provided, wherein a supersonic air film nozzle is arranged on the surface of the aircraft. The supersonic air film nozzle is designed using the supersonic air film nozzle design method provided in any of the above embodiments, and the supersonic air film nozzle is used to generate a supersonic air film.

[0110] Matters not covered by the present invention are known technologies.

[0111] 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.

[0112] 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.

[0113] 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 nozzle design method, characterized in that: include: Obtain the supersonic film nozzle outlet pressure of a high-speed waverider vehicle under pressure matching conditions; 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; 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 nozzle design method according to claim 1, characterized in that: Supersonic 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 film nozzle outlet pressure, and Me is the supersonic film Mach number.

3. The supersonic air film nozzle design method according to claim 1, characterized in that: Nozzle exit height y of supersonic air film nozzle e Determined by the following formula: 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.

4. The method for designing a supersonic air film nozzle according to claim 1, 2 or 3, wherein: The parabolic potential function equation is: 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.

5. The supersonic air film nozzle design method according to claim 4, 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.

6. The supersonic air film nozzle design method according to claim 5, characterized in that: The method of making viscosity correction for the nozzle inviscid profile is: the nozzle inviscid profile increases δ in the normal direction. * The distance is used to achieve viscosity correction.

7. The method for designing a supersonic air film nozzle according to claim 1, 2, 3, 5 or 6, characterized in that: The coordinates (x, y) of each point on the contraction profile of the supersonic film nozzle are determined by the quintic curve as follows: 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.

8. A supersonic air film nozzle design device, characterized in that: include: The first module is used to obtain the supersonic air film nozzle outlet pressure of the high-speed waverider aircraft under pressure matching conditions; The second module is used to determine the supersonic air film mass flow rate and supersonic air film Mach number based on the cooling length requirements of the high-speed waverider aircraft; The third module is used to determine the supersonic air film nozzle throat height according to the supersonic air film mass flow rate and the supersonic air film nozzle outlet pressure; The fourth module is used to determine the nozzle outlet height of the supersonic air film nozzle according to the supersonic air film nozzle throat height and the supersonic air film Mach number; The fifth module is used to solve the parabolic potential function equation for the transonic flow at the throat of the supersonic film nozzle using the series expansion method to obtain the transonic solution of the supersonic film nozzle; The sixth module is used to set the Mach number distribution of the supersonic film nozzle axis with the supersonic film nozzle transonic solution and the nozzle outlet as the starting point and end point respectively; The seventh module is used to construct a supersonic film nozzle characteristic line grid 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, solve the supersonic flow field inside the supersonic film nozzle, and then determine the nozzle inviscid profile based on the streamline control equation; The eighth module is used to solve the boundary layer displacement thickness based on the von Karman momentum integral relationship; The ninth module is used to perform viscosity correction on the nozzle inviscid profile based on the boundary layer displacement thickness to obtain the final nozzle profile; The tenth module is used to connect the preset contraction section profile of the supersonic air film nozzle with the nozzle profile at the throat to obtain a complete supersonic air film nozzle.

9. A supersonic air film nozzle, characterized in that: It is designed using the supersonic air film nozzle design method as claimed in claim 1.

10. A near-space waverider aircraft, wherein a supersonic air film nozzle is arranged on the surface of the aircraft, characterized in that: The supersonic air film nozzle is designed using the supersonic air film nozzle design method according to claim 1, and the supersonic air film nozzle is used to generate a supersonic air film.

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