Supersonic velocity gas film nozzle, design method and device and near space waverider aircraft

Abstract

The invention provides a supersonic gas film nozzle, a design method and device and a near space waverider aircraft. The design method comprises the steps that the outlet pressure of the supersonic gas film nozzle of the high-speed waverider aircraft under the pressure matching condition is obtained; determining a supersonic velocity gas film mass flow rate, a supersonic velocity gas film Mach number, a supersonic velocity gas film nozzle throat height and a nozzle outlet height of the supersonic velocity gas film nozzle, solving a supersonic velocity solution of the supersonic velocity gas film nozzle, and setting Mach number distribution of an axis of the supersonic velocity gas film nozzle; the supersonic velocity gas film nozzle transonic solution, the nozzle outlet and the mach number distribution of the axis of the supersonic velocity gas film nozzle are used as boundary conditions, the nozzle non-stick molded line is determined, the displacement thickness of a boundary layer is solved, the final nozzle molded line is obtained, then the complete supersonic velocity gas film nozzle is obtained, wave absorption in the nozzle is complete, shock waves are eliminated fundamentally, and the service life of the nozzle is prolonged. And distribution of flow field parameters such as pressure, temperature and speed 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] At present, high-speed waverider aircraft have become a research hotspot in countries around the world due to their advantages such as high flying height, high speed, and "water-drifting" flight trajectory that is difficult to intercept. However, high-speed flight brings serious aerodynamic heat to the waverider, causing ablation and damage to the surface structure and optical windows of the aircraft. Taking the Mach 6 waverider aircraft as an example, the atmospheric temperature at an altitude of 100km is about 195K, and the laminar Prandtl number is 0.7, so the adiabatic temperature of the aircraft surface is about 1340K.

[0003] Covering a layer of cooling air film on the surface of a high-speed waverider aircraft is an effective way to improve infrared stealth performance and the only way to avoid burning of the optical window. At present, the supersonic air film of a high-speed aircraft is generally arranged in front of the optical window. The supersonic air film used for infrared stealth is arranged on the fuselage of the aircraft, and its nozzle width is larger and the effective cooling area is longer. Both are achieved by arranging nozzles on the wall.

[0004] The current supersonic air film nozzle is 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 of 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] Firstly, 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 Mach number and area ratio formula.

[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, the current supersonic air film nozzle has several shortcomings. First, there are defects in its design theory. The nozzle flow is not one-dimensional in nature, at least two-dimensional flow, especially for curved nozzles, the flow is three-dimensional, or two-dimensional flow theory combined with transverse curvature correction; the existing supersonic air film nozzle is designed using one-dimensional flow theory, and its nozzle wall cannot perfectly eliminate waves. In other words, there must be concentrated compression waves in the inner flow channel. On the one hand, the uniformity of the nozzle outlet is very low, and the mixing with the mainstream after injection is intensified, and the heat protection performance is reduced. On the other hand, the existence of weak shock waves leads to unequal entropy in the flow, and the static temperature and static pressure at the nozzle outlet are inconsistent with the results of one-dimensional flow theory. It is difficult to effectively control the air film pressure to match the mainstream pressure of the aircraft. The pressure mismatch leads to wave structure, which in turn enhances mixing and increases optical distortion. Summary of the invention

[0010] In view of 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 purpose, 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 air 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 air film nozzle is determined according to the supersonic air film mass flow rate and the supersonic air film nozzle outlet pressure;

[0016] 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;

[0017] 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.

[0018] 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.

[0019] 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.

[0020] According to the von Karman momentum integral relation, the displacement thickness of the boundary layer 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 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.

[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 the supersonic air film Mach number according to the cooling length requirements of the high-speed waverider aircraft;

[0026] The third module is used to determine the throat height of the supersonic air film nozzle 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 air film nozzle by using the series expansion method to obtain the transonic solution of the supersonic air film nozzle;

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

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

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

[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 contraction section profile of the preset 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, which is 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 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 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.

[0041] Figure 1 It 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 is a schematic diagram of a nozzle characteristic line network and a nozzle non-viscous profile line in one embodiment;

[0046] Figure 6 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 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.

[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 flow hyperbolic equation to design the supersonic air film nozzle. The designed supersonic air film nozzle has complete internal wave elimination, which fundamentally eliminates shock waves. 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 supersonic air film generator is suitable for use in surface heat reduction and stealth of high-speed waverider aircraft.

[0050] Specifically, refer to Figure 2 , a supersonic air film nozzle design method provided in one embodiment includes:

[0051] Obtain the supersonic air 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 air film nozzle is determined according to the supersonic air film mass flow rate and the supersonic air film nozzle outlet pressure;

[0054] 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;

[0055] 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.

[0056] 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.

[0057] 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.

[0058] According to the von Karman momentum integral relation, the displacement thickness of the boundary layer 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 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.

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

[0065] 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:

[0066]

[0067] 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.

[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 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.

[0071] For the transonic flow at the throat of the 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, as follows: Figure 3 shown. Figure 3 is the throat of the supersonic air film nozzle, y* is the throat height, and the dotted line is the speed of sound line, which is obtained by solving the transonic solution of the supersonic air film nozzle. The Mach number of the speed of sound line is 1. The upstream of the speed of sound line (left side) is the subsonic region, and the Mach number is less than 1. The downstream of the speed of sound line (right side) is the supersonic region, and the Mach number is greater than 1.

[0072] 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, such as Figure 4 As shown, 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 respectively used as the starting point and the end point, 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 air film nozzle design method provided in an embodiment, the supersonic air film nozzle transonic solution, the Mach number distribution of the nozzle outlet and the supersonic air film nozzle axis are used as boundary conditions, and a supersonic air film nozzle characteristic line grid is constructed based on the characteristic line method to solve the supersonic flow field inside the supersonic air film nozzle, and then the nozzle inviscid line is determined according to the streamline control equation, such as Figure 5 The thick black solid line shown represents the inviscid line of the nozzle. The characteristic line method is a common technical means in this field. The characteristic line method is used to solve the two-dimensional flow hyperbolic equation to design the supersonic air film nozzle. The two-dimensional structure of the supersonic air film nozzle has complete internal wave elimination, which fundamentally eliminates the shock wave, and the distribution of flow field parameters such as pressure, temperature, and velocity is controllable. The following are two characteristic line equations:

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

[0087] 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.

[0088] 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:

[0089]

[0090] 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.

[0091] In order 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 air film nozzle design method provided by the present invention is used to design a Mach 3 nozzle, so that the tangential jet passing through the Mach 3 nozzle avoids direct friction of the hypersonic mainstream on the surface of the waverider, thereby reducing the surface temperature of the waverider and achieving the effect of reducing heat and drag. The input conditions of the Mach 3 nozzle are controlled at a total temperature of 300K, a total pressure of 100000Pa, a nozzle outlet temperature of 108K, and a pressure of 2820Pa. The design results are as follows: Figure 6 , Figure 7 As shown, Figure 6 It is the nozzle Mach number cloud diagram; Figure 7 This is the distribution diagram of the Mach number at the nozzle outlet. It can be seen that the flow is steadily accelerated from subsonic to supersonic flow, the wave is completely eliminated inside the nozzle, there is no concentrated shock wave, and the outlet Mach number is particularly uniform, which reflects the feasibility of the present invention. The present invention has been tested in the wind tunnel of the unit, and has achieved the expected 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 the supersonic air film Mach number according to the cooling length requirements of the high-speed waverider aircraft;

[0096] The third module is used to determine the throat height of the supersonic air film nozzle 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 air film nozzle by using the series expansion method to obtain the transonic solution of the supersonic air film nozzle;

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

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

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

[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 contraction section profile of the preset 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, including a memory and a processor, the memory storing a computer program, and the processor implementing the steps of the supersonic air film nozzle design method provided in any of the above embodiments when executing the computer program. The computer device may be a server. The computer device includes a processor, a memory, a network interface and a database connected via a system bus. Among them, the processor of the computer device is used to provide computing and control capabilities. The memory of the computer device includes 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 can understand that all or part of the processes in the above-mentioned embodiment methods can be completed by instructing the relevant hardware through a computer program, and 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. As an illustration and not limitation, RAM is available in many 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, and 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 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.

[0112] 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.

[0113] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. For those skilled in the art, the present invention may have various modifications and variations. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention shall be included in the protection scope of the present invention.

Claims

1. A supersonic air film nozzle design method, characterized in that: include: Obtain the supersonic air 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 air film nozzle is determined according to the supersonic air film mass flow rate and the supersonic air film nozzle outlet pressure; According to the supersonic air film nozzle throat height and the supersonic air film Mach number, the nozzle outlet height of the supersonic air film nozzle is determined; For the transonic flow in the throat of supersonic film nozzle, the parabolic potential function equation is solved by series expansion method, and the transonic solution of supersonic film nozzle is obtained. The Mach number distribution of the supersonic film nozzle axis is set with the supersonic film nozzle transonic solution and the nozzle outlet as the starting point and the end point respectively. Taking the transonic solution of the supersonic film nozzle, the Mach number distribution of the nozzle outlet and the axis of the supersonic film nozzle as boundary conditions, the characteristic line grid of the supersonic film nozzle is constructed, and the supersonic flow field inside the supersonic film nozzle is solved. Then, the inviscid profile of the nozzle is determined according to the streamline control equation. According to the von Karman momentum integral relation, the displacement thickness of the boundary layer is solved; The viscosity correction of the nozzle inviscid profile is performed based on the boundary layer displacement thickness to obtain the final nozzle profile; The contraction section profile of the preset supersonic air film nozzle is connected with the nozzle profile at the throat to obtain a complete supersonic air film nozzle.

2. The supersonic air film nozzle design method according to claim 1, 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.

3. The supersonic air film nozzle design method according to claim 1, 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.

4. The method for designing a supersonic air film nozzle according to claim 1, 2 or 3, 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.

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

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 air 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 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 the supersonic air film Mach number according to the cooling length requirements of the high-speed waverider aircraft; The third module is used to determine the throat height of the supersonic air film nozzle 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 air film nozzle by using the series expansion method to obtain the transonic solution of the supersonic air film nozzle; The sixth module is used to set the Mach number distribution of the supersonic air film nozzle axis with the supersonic air film nozzle transonic solution and the nozzle outlet as the starting point and the end point respectively; The seventh module is used to construct the characteristic line grid of the supersonic film nozzle with 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, solve the supersonic flow field inside the supersonic film nozzle, and then determine the nozzle inviscid line according to the streamline control equation; The eighth module is used to solve the boundary layer displacement thickness based on the von Karman momentum integral relation; 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 contraction section profile of the preset 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 by adopting the supersonic air film nozzle design method as claimed in claim 1.

10. A near-space waverider aircraft, the surface of which is provided with a supersonic air film nozzle, characterized in that: The supersonic air film nozzle is designed by 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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