A method for supersonic film cooling of a spherical optical dome of a hypersonic vehicle

Abstract

The present invention provides a supersonic film cooling method for the spherical optical dome of a hypersonic vehicle, belonging to the technical field of thermal protection of hypersonic vehicles. The supersonic cooling film proposed by the present invention does not need to be in convective contact with the high-temperature, high-pressure and high-speed oncoming flow, which can reduce the film jet pressure, save the consumption of cooling gas in the high-pressure gas storage tank, and thus reduce the load of the vehicle. At the same time, after the interaction between the oncoming flow gas and the cooling gas behind the shock wave, the cooling gas is close to the spherical optical window body, effectively covering the window surface, effectively isolating the high-temperature oncoming flow gas, and realizing the heat reduction effect on the spherical optical dome through film / mainstream heat transfer. The cooling performance of the full-coverage optical dome window can effectively reduce the aerodynamic heat radiation effect, make the density distribution fluctuation of the surrounding flow field smaller, and reduce the optical distortion caused; at the same time, it enables the high-speed vehicle to be applicable to the flight environment with large angle-of-attack attitude changes, and ensures its high-speed and precise strike working performance.

Description

Technical Field

[0001] The present invention relates to the technical field of thermal protection for hypersonic vehicles, and particularly to a supersonic film cooling method for a spherical optical dome of a hypersonic vehicle. Background Art

[0002] For a hypersonic vehicle that flies at hypersonic speeds in the atmosphere and is equipped with infrared optical imaging detection and guidance, a bow shock wave is formed at its head due to the strongly compressed air, and the temperature of the gas passing through the shock wave rises significantly. When the gas passing through the shock wave contacts the vehicle wall, due to the wall viscous shear stress, the fluid velocity in the boundary layer is greatly reduced, and a large amount of kinetic energy is converted into heat energy, causing a strong aerodynamic heating effect on the vehicle. At the same time, since the optical dome window is directly in contact with the high-temperature gas and its surface is affected by the aerodynamic heating effect, the temperature rise increases rapidly, triggering a huge aerodynamic thermal-optical effect. Under the environment of thermo-optical coupled aerodynamic effects, the refractive index field distribution of the light on the window wall of the spherical optical dome of the hypersonic vehicle is no longer uniform, and serious aberrations occur in the imaging system inside the vehicle, greatly reducing the imaging quality of the infrared detection system and seriously threatening the high-precision strike performance of the hypersonic vehicle against ground-to-air and other targets.

[0003] In addition, compared with the traditional planar side window design, the spherical optical dome window of the hypersonic vehicle can not only avoid significantly damaging the original aerodynamic contour design of the hypersonic vehicle, but also has a nearly 360° wide circumferential imaging detection and guidance field of view, and is easy to introduce aberration correction. At the same time, the spherical optical window can effectively avoid the aerodynamic problems caused by the aerodynamic contour generated by the planar side window. Therefore, its "spherical conformal" characteristic plays a huge potential in significantly improving the working performance of the hypersonic vehicle.

[0004] In recent years, many domestic and foreign scholars have mainly focused on active protection methods for the thermal protection of the optical dome window of hypersonic vehicles, including two methods: thermal protection by gas injection and thermal protection using mechanical devices. For example, mechanical thermal protection is to equip sharp mechanical fairing cones (such as spikes), drag reduction rods, aerodynamic discs, etc. at the head of the vehicle, which will significantly change the flow field structure of the oncoming flow near the head of the vehicle, thereby protecting the spherical optical dome window of the vehicle; in addition, gas injection thermal protection mainly includes reverse jet (single-hole or micro-hole array jet at the head), tangential or vertical injection jet, transpiration cooling, combined jet, etc. Its main purpose is to eject cooling gas to cover the surface of the optical window to isolate the direct contact between the high-temperature mainstream gas and the optical window wall, thereby achieving the purpose of protecting the optical window of the vehicle. Among them, tangential or vertical injection jet gas injection thermal protection is mainly applied to the aerodynamic heating of the side optical window of hypersonic vehicles.

[0005] In pursuit of good aerodynamic performance and good imaging quality, the surface shapes of traditional hypersonic vehicle optical domes are mostly flat side window and conformal curved side window optical domes. Among them, the traditional flat side window optical dome seriously damages the aircraft's shape design and greatly changes the flow field characteristics of the aircraft's head; at the same time, both of them greatly limit the imaging range, and due to extreme aerodynamic heat, they are prone to serious deformation, which in turn affects the imaging quality of the internal optical system.

[0006] Regarding the existing publicly available thermal protection methods and technologies for hypersonic vehicle spherical optical domes, they are mainly mechanical (shock pin type, etc.) thermal protection. However, it bears the most intense aerodynamic heating effect itself, is extremely prone to damage and ablation, has very low usage efficiency, and seriously changes the contour structure of the high-speed aircraft. Summary of the Invention

[0007] In view of the above technical problems, the present invention proposes a supersonic film cooling solution for the spherical optical dome of a hypersonic vehicle.

[0008] The first aspect of the present invention discloses a supersonic film cooling method for the spherical optical dome of a hypersonic vehicle. The method realizes the heat reduction effect on the spherical optical dome of the hypersonic vehicle based on a supersonic film cooling system. The supersonic film cooling system includes a quick-connect plug 1, an air duct 2, a spherical optical dome window 3, a connection and fixing member 4, a film component 5, and a high-pressure gas source component 6. In the method:

[0009] The spherical optical dome window 3 and the connection and fixing member 4 are tightly connected by screws, and by adjusting the depth of the screws, the inner and outer contour lines of the two are made to coincide; the film component 5 and the connection and fixing member 4 are tightly connected and fixed through the internal structure of the air duct 2, and the jet outlet height h of the supersonic film component is limited through the internal structure. e ; At the same time, the film component 5 and the connection and fixing member 4 form a Laval nozzle profile for generating a supersonic cooling film. According to the flow parameters of the required supersonic film, the film component 5 is replaced on the premise of determining the jet outlet height h. e to form a stable supersonic cooling film.

[0010] The high-pressure gas source component 6 includes a stop valve 6-1, a pressure reducing valve 6-2, an electromagnetic controller 6-3, a mass flow meter 6-4, and a high-pressure gas source 6-5; among them: the stagnation pressure of its spherical optical dome is obtained based on the true flight speed and altitude of the hypersonic vehicle; according to the ratio between the static pressure at the jet outlet of the supersonic film and the stagnation pressure of the spherical optical dome, the total pressure of the supersonic film chamber is obtained; the total pressure supplied by the high-pressure gas source 6-6 to the supersonic film chamber through the high-pressure resistant air duct 6-5 is controlled by adjusting the stop valve 6-1, the pressure reducing valve 6-2, and the electromagnetic controller 6-3 in sequence.

[0011] According to the preferred embodiment, in the method, the total pressure of the supersonic film chamber is calculated as follows:

[0012] According to the true flight speed and altitude of the hypersonic vehicle, the oncoming flow state parameters are obtained by querying the standard atmosphere, including the total pressure p0, static pressure p1, static temperature T1, and total temperature T0;

[0013] According to the Newton-Lees Newton pressure formula, the stagnation total pressure p of the spherical optical dome of the hypersonic vehicle is obtained 0,2 ;

[0014]

[0015]

[0016] where C pmax is the correction coefficient, representing the stagnation point pressure coefficient after the normal shock; p 0,2 is the stagnation total pressure after the normal shock; p1 is the oncoming flow static pressure; p2 is the static pressure after the normal shock; Ma1 is the flight Mach number of the hypersonic vehicle, Ma2 is the Mach number after the bow shock of the hypersonic spherical optical dome, γ is the specific heat ratio, and θ is the local surface inclination angle;

[0017] According to the set jet pressure ratio NPR = p c,02 / p 0,2 , the static pressure p c,02 at the outlet of the supersonic film jet is obtained;

[0018] According to the isentropic relation and the designed Mach number Ma c at the outlet of the supersonic film jet, the total pressure p c,01 of the supersonic film chamber is determined:

[0019]

[0020] According to the preferred embodiment, in the method, the cooling effect of the supersonic cooling film on the hypersonic spherical optical dome is characterized by the cooling efficiency, and there are two calculation methods for the cooling efficiency:

[0021] Or

[0022] where T ∞ is the oncoming flow temperature, T w is the wall temperature of the spherical optical dome, T c is the coolant temperature, Qc is the heat flux on the wall of the spherical optical dome after film cooling, and Q0 is the heat flux on the wall of the spherical optical dome without film cooling; when η≥1, it represents the effective film insulation cooling area, otherwise it is the effective film heat reduction cooling area.

[0023] According to a preferred embodiment, in the method, the position of the gas film assembly 5 and the connecting and fixing member 4 is determined by the height h* of the throat of the Laval nozzle in the gas film assembly 5, and its calculation method is as follows:

[0024]

[0025] Wherein, R A is the area ratio, Ma c is the Mach number at the outlet of the gas film jet, A e is the area of the annular gas film jet outlet, A * is the cross-sectional area of the annular Laval nozzle throat, h1 is the height of the gas film jet outlet from the central axis, h2 is the height of the Laval nozzle throat from the central axis, and h e is the height of the gas film jet outlet.

[0026] According to a preferred embodiment, in the method, when the aircraft is in a hypersonic flight state, the connecting and fixing member 4 and the gas film assembly 5 form a Laval nozzle profile, thereby generating a supersonic gas film. The supersonic gas film is squeezed and attached to the wall surface of the spherical optical dome through the bow shock wave in front of the spherical optical dome, so as to resist the aerodynamic heat received by the hypersonic spherical optical dome. By adjusting the flow parameters of the supersonic gas film, it is adapted to different flight environments of hypersonic aircraft.

[0027] The second aspect of the present invention discloses a supersonic gas film cooling system for a spherical optical dome of a hypersonic aircraft. The system includes: a quick-connect plug 1, a gas guide pipe 2, a spherical optical dome window 3, a connecting and fixing member 4, a gas film assembly 5, and a high-pressure gas source assembly 6; wherein:

[0028] The spherical optical dome window 3 and the connecting and fixing member 4 are tightly connected by screws, and the inner and outer contour lines of the two are made to coincide by adjusting the depth of the screws; the gas film assembly 5 and the connecting and fixing member 4 are tightly connected and fixed through the internal structure of the gas guide pipe 2, and the height h of the jet outlet of the supersonic gas film assembly is limited through the internal structure e ; at the same time, the gas film assembly 5 and the connecting and fixing member 4 form a Laval nozzle profile for generating a supersonic cooling gas film. According to the required flow parameters of the supersonic gas film, the gas film assembly 5 is replaced on the premise of determining the jet outlet height h e to form a stable supersonic cooling gas film.

[0029] The high-pressure gas source assembly 6 includes a stop valve 6-1, a pressure reducing valve 6-2, an electromagnetic controller 6-3, a mass flow meter 6-4, and a high-pressure gas source 6-5; among them: the stagnation pressure of the spherical optical dome of the hypersonic vehicle is obtained based on its true flight speed and altitude; according to the ratio between the static pressure at the outlet of the supersonic gas film jet and the stagnation pressure of the spherical optical dome, the total pressure of the supersonic gas film chamber is obtained; the total pressure supplied by the high-pressure gas source 6-6 to the supersonic gas film chamber through the high-pressure resistant gas pipe 6-5 is controlled by adjusting the stop valve 6-1, the pressure reducing valve 6-2, and the electromagnetic controller 6-3 in sequence.

[0030] According to the preferred embodiment, the calculation method of the total pressure of the supersonic gas film chamber is as follows:

[0031] According to the true flight speed and altitude of the hypersonic vehicle, the oncoming flow state parameters are obtained by querying the standard atmosphere, including the total pressure p0, the static pressure p1, the static temperature T1, and the total temperature T0;

[0032] According to the Newton-Lees Newton pressure formula, the stagnation total pressure p of the spherical optical dome of the hypersonic vehicle is obtained 0,2 ;

[0033]

[0034] Among them, C pmax is the correction coefficient, representing the stagnation point pressure coefficient after the normal shock; p 0,2 is the stagnation total pressure after the normal shock; p1 is the oncoming flow static pressure; p2 is the static pressure after the normal shock; Ma1 is the flight Mach number of the hypersonic vehicle, Ma2 is the Mach number after the bow shock of the hypersonic spherical optical dome, γ is the specific heat ratio, and θ is the local surface inclination angle;

[0035] According to the set jet pressure ratio NPR = p c,02 / p 0,2 , the static pressure p c,02 at the outlet of the supersonic gas film jet is obtained;

[0036] According to the isentropic relation and the designed Mach number Ma c at the outlet of the supersonic gas film jet, the total pressure p c,01 of the supersonic gas film chamber is determined:

[0037]

[0038] According to the preferred embodiment, the cooling effect of the supersonic cooling gas film on the hypersonic spherical optical dome is characterized by the cooling efficiency, and there are two calculation methods for the cooling efficiency:

[0039] Or

[0040] Among them, T∞ is the oncoming flow temperature, T w is the wall temperature of the spherical optical dome, T c is the coolant temperature, Qc is the heat flux on the wall of the spherical optical dome after film cooling, and Q0 is the heat flux on the wall of the spherical optical dome without film cooling; when η ≥ 1, it represents the effective film insulation cooling zone, otherwise it is the effective film heat reduction cooling zone.

[0041] According to the preferred embodiment, the position of the film assembly 5 and the connecting and fixing member 4 is determined by the throat height h* of the Laval nozzle in the film assembly 5, and its calculation method is:

[0042]

[0043] wherein, R A is the area ratio, Ma c is the Mach number at the film jet outlet, A e is the area of the annular film jet outlet, A * is the cross-sectional area of the annular Laval nozzle throat, h1 is the height of the film jet outlet from the central axis, h2 is the height of the Laval nozzle throat from the central axis, h e is the height of the film jet outlet.

[0044] According to the preferred embodiment, when the aircraft is in the hypersonic flight state, the connecting and fixing member 4 and the film assembly 5 form a Laval nozzle profile, thereby generating a supersonic film. The supersonic film is squeezed and attached to the wall of the spherical optical dome by the bow shock wave in front of the spherical optical dome, so as to resist the aerodynamic heat received by the hypersonic spherical optical dome. By adjusting the flow parameters of the supersonic film, it can meet the different flight environments of hypersonic aircraft.

[0045] In summary, the supersonic cooling film proposed by the present invention does not need to be in convective contact with the high-temperature, high-pressure and high-speed oncoming flow, can reduce the jet pressure, save the cooling gas consumption of the high-pressure gas storage tank, and thus reduce the aircraft load; at the same time, after the oncoming flow gas and the cooling gas interact behind the shock wave, the cooling gas is close to the "shape" of the spherical optical window, effectively covering the window surface, effectively isolating the high-temperature oncoming flow gas, and achieving the heat reduction effect on the spherical optical dome through film / mainstream heat transfer. In addition, due to its cooling performance that fully covers the optical dome window, it can effectively reduce the aerodynamic heat radiation effect, make the density distribution fluctuation of the surrounding flow field smaller, and reduce the optical distortion caused; at the same time, it enables the high-speed aircraft to be applicable to the flight environment with large angle-of-attack attitude changes, and thus ensures its working performance of higher speed and more accurate strike. Brief Description of the Drawings

[0046] To more clearly illustrate the specific embodiments of the present invention or the technical solutions in the prior art, the following will briefly introduce the drawings required for the description of the specific embodiments or the prior art. Obviously, the drawings in the following description are some embodiments of the present invention. For those of ordinary skill in the art, without creative efforts, other drawings can also be obtained based on these drawings.

[0047] Figure 1 It is a three-dimensional structural schematic diagram of the film cooling system of the spherical optical dome of a hypersonic vehicle according to an embodiment of the present invention.

[0048] Figure 2 It is a two-dimensional schematic diagram of the film cooling structure and high-pressure gas source assembly of the spherical optical dome of a hypersonic vehicle according to an embodiment of the present invention.

[0049] Figure 3a It is a cloud map of the wall temperature change of the hypersonic spherical optical dome under different jet pressure ratios in the underpressure and pressure matching states according to an embodiment of the present invention.

[0050] Figure 3b It is a cloud map of the wall temperature change of the hypersonic spherical optical dome under different jet pressure ratios in the overpressure state according to an embodiment of the present invention.

[0051] Figure 4 It is a schematic diagram of the wall temperature change of the hypersonic spherical optical dome under different jet pressure ratios according to an embodiment of the present invention.

[0052] Figure 5 It is a schematic diagram of the film cooling efficiency of the hypersonic spherical optical dome under different jet pressure ratios according to an embodiment of the present invention.

[0053] Figure 6 It is a cloud map of the wall temperature of the hypersonic spherical optical dome under different angles of attack according to an embodiment of the present invention.

[0054] Figure 7 It is a schematic diagram of the wall temperature change of the hypersonic spherical optical dome under different angles of attack according to an embodiment of the present invention.

[0055] Figure 8 It is a schematic diagram of the film cooling efficiency of the hypersonic spherical optical dome under different angles of attack according to an embodiment of the present invention. Specific Embodiments

[0056] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the following will clearly and completely describe the technical solutions in the embodiments of the present invention with reference to the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are only a part of the embodiments of the present invention, rather than all of the embodiments. All other embodiments obtained by those of ordinary skill in the art based on the embodiments of the present invention without creative efforts shall fall within the scope of protection of the present invention.

[0057] Refer to the attached Figure 1-2 , a supersonic film cooling system for the spherical optical dome of a hypersonic vehicle, which consists of six parts: a quick-connect plug 1, a gas duct 2, a spherical optical dome window 3, a connecting and fixing member 4 (spherical optical dome window - gas duct), a film component 5, and a high-pressure gas source component 6. The spherical optical dome window 3 and the connecting and fixing member 4 are tightly connected by screws, and the depth of the screws is adjusted to ensure that the inner and outer contour lines of the two coincide; the film component 5 and the connecting and fixing member 4 are tightly connected and fixed through the internal structure of the gas duct 2, and the height h of the jet outlet of the supersonic film component is limited through the internal structure e . At the same time, the film component 5 and the connecting and fixing member 4 can form a nozzle profile for generating a supersonic cooling film, that is, the film component 5 can be replaced based on the required flow parameters of the supersonic film on the premise of determining the jet outlet height h e to form a stable supersonic film.

[0058] Refer to the attached Figure 3a-3b , the high-pressure gas source component 6 provided by the present invention includes a stop valve 6-1, a pressure reducing valve 6-2, an electromagnetic controller 6-3, a mass flow meter 6-4, and a high-pressure gas source 6-5. Based on the actual flight speed and altitude of the hypersonic vehicle, the stagnation pressure of its spherical optical dome is obtained, and then according to the ratio between the static pressure at the jet outlet of the supersonic film and the stagnation pressure of the spherical optical dome, that is, the jet pressure ratio (Nozzle Pressure Ratio, NPR = p c,02 / p 0,2 ), the total pressure of the supersonic film chamber is further obtained. Finally, the stop valve 6-1, the pressure reducing valve 6-2, and the electromagnetic controller 6-3 are adjusted in sequence to control the total pressure supplied by the high-pressure gas source 6-6 to the supersonic film chamber through the high-pressure-resistant gas duct 6-5.

[0059] The solution steps for the total pressure p c,0 of the supersonic film chamber of the present invention are as follows:

[0060] (1) According to the actual flight speed and altitude of the hypersonic vehicle, by querying the standard atmosphere, the oncoming flow state parameters can be obtained, such as the total pressure p0, static pressure p1, static temperature T1, and total temperature T0, etc.

[0061] (2) Obtained the total stagnation pressure p of the spherical optical dome of the hypersonic vehicle by correcting according to the Newton-Lees pressure formula 0,2 , and the specific derivation steps are as follows:

[0062]

[0063]

[0064] In the formula, C pmax is the correction coefficient, which is the stagnation point pressure coefficient after the normal shock wave (referred to as the maximum pressure coefficient); p 0,2 is the total stagnation pressure after the normal shock wave; p1 is the static pressure of the oncoming flow; p2 is the static pressure after the normal shock wave; Ma1 is the flight Mach number of the hypersonic vehicle, Ma2 is the Mach number after the bow shock wave of the hypersonic spherical optical dome, γ is the specific heat ratio, and θ is the local surface inclination angle.

[0065] (3) According to the set jet pressure ratio NPR = p c,02 / p 0,2 , the static pressure p c,02 at the outlet of the supersonic film jet can be obtained.

[0066] (4) According to the isentropic relation and the designed Mach number Ma c at the outlet of the supersonic film jet, the total pressure p c,01 of the supersonic film chamber can be obtained, and the specific formula is as follows:

[0067]

[0068] Refer to Appendix Figure 5 and 8 , the present invention characterizes the advantages and disadvantages of the designed supersonic film for the hypersonic spherical optical dome by the cooling efficiency (adiabatic wall, numerical calculation) or (isothermal wall, flight condition). In the formula, T∞ is the oncoming flow temperature, Tw is the wall temperature of the spherical optical dome, Tc is the coolant temperature, Qc is the heat flux of the wall of the spherical optical dome after film cooling, and Q0 is the heat flux of the wall of the spherical optical dome without film cooling. When η ≥ 1, it represents an effective film insulation cooling area, and when η < 1, it represents an effective film heat exchange cooling area.

[0069] Refer to Appendix Figure 2 , the position of the film assembly 5 and the connecting and fixing member 4 is mainly determined by the throat height h* of the Laval nozzle in the film assembly 5, and the solution expression of this dimension is as follows:

[0070]

[0071] In the formula, RA is the area ratio, Mac is the Mach number at the outlet of the film jet, A e is the outlet area of the annular film jet, A * is the cross-sectional area of the throat of the annular Laval nozzle. h1 is the height of the film jet outlet from the central axis, h2 is the height of the Laval nozzle throat from the central axis, h e is the height of the film jet outlet.

[0072] Refer to Appendix Figure 3a 、 3b 、4-8. A hypersonic spherical optical dome supersonic film cooling working method disclosed by the present invention is as follows: When the aircraft is flying at hypersonic speed, it is placed in the connecting and fixing member 4 and the film assembly 5 to form a Laval nozzle profile, thereby generating a supersonic film. The supersonic film is squeezed and "attached" to the wall surface of the spherical optical dome through the bow shock wave in front of the spherical optical dome, thereby resisting the severe aerodynamic heat received by the hypersonic spherical optical dome, improving the flow structure around the spherical optical dome, and greatly enhancing its optical imaging characteristics. At the same time, the flow parameters of the supersonic film are adjusted according to formulas (1)-(6) to meet different flight environments of hypersonic aircraft.

[0073] In addition, according to the flight Mach number of the hypersonic vehicle being 7.1, the total pressure being 5 MPa, and the total incoming flow temperature being 670 K, and the supersonic film Mach number being 3 and the total temperature being 300 K, numerical simulation verification of different jet pressure ratios NPR (0-2) and large angles of attack (0°-30°) was carried out on the present invention, and the schematic diagrams of the heat reduction effect and cooling efficiency shown in Figures 3-8 of the accompanying drawings were obtained. As can be seen from the figures, based on the flight condition of an angle of attack of 0°, the hypersonic spherical optical dome supersonic film cooling method in the present invention can cause a significant decrease in the wall temperature of the optical dome. When NPR ≤ 1, the wall temperature of the spherical optical dome drops from 670 K to about 450 K - 325 K, and the heat reduction amplitude reaches 32.83% - 51.49%; when 1 < NPR ≤ 2, the wall temperature of the spherical optical dome drops to about 320 K - 250 K, and the heat reduction amplitude reaches 52.23% - 62.68%; and when NPR ≥ 1.6, the wall temperature of the spherical optical dome drops below 300 K. Based on the flight condition of NPR = 1, when the flight angle of attack is 25°, the wall temperature of the windward side spherical optical dome still drops to 400 - 450 K, and the supersonic film cooling efficiency is as high as 50%. Therefore, the hypersonic spherical optical dome supersonic film cooling method in the present invention has excellent cooling performance.

[0074] It can be seen that the supersonic cooling gas film of the present invention does not need to be in convective contact with the high-temperature, high-pressure and high-speed oncoming flow, greatly reducing the jet pressure, saving the consumption of cooling gas in the high-pressure gas storage tank, and thus effectively reducing the load of the aircraft. At the same time, after the interaction between the oncoming flow gas and the cooling gas behind the shock wave, the cooling gas adheres to the "shape" of the spherical optical window, effectively covering the window surface, isolating the high-temperature oncoming flow gas, and instantaneously enhancing heat dissipation and transporting heat flux through convective heat transfer. In addition, due to its cooling performance that fully covers the optical hood window, the density distribution fluctuation of the surrounding flow field becomes smaller, and the optical distortion caused is reduced. At the same time, the high-speed aircraft can be applicable to the flight environment with large angle-of-attack attitude changes, thereby ensuring its working performance of higher speed and more accurate strike.

[0075] Please note that the technical features of the above embodiments can be combined arbitrarily. For the sake of brevity of description, 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, it should be considered as the scope described in this specification. The above-described embodiments only represent several implementation manners of the present application. Their descriptions are relatively specific and detailed, but they should not be construed as limiting the scope of the invention patent. It should be pointed out that for those of ordinary skill in the art, without departing from the concept of the present application, several deformations and improvements can still be made, and these all belong to the protection scope of the present application. Therefore, the protection scope of the patent of the present application should be subject to the appended claims.

Claims

1. A method for supersonic film cooling of a spherical optical head cover of a hypersonic aircraft, characterized in that: The method realizes the heat reduction effect on the spherical optical head cover of a hypersonic aircraft based on a supersonic air film cooling system. The supersonic air film cooling system comprises a quick-connect plug (1), an air guide tube (2), a spherical optical head cover window (3), a connecting fixture (4), an air film assembly (5), and a high-pressure air source assembly (6). In the method: The spherical optical head cover window (3) is fastened to the connecting fixture (4) by means of screws, and the inner and outer contours of the two are matched by adjusting the depth of the screws; the air film assembly (5) and the connecting fixture (4) are pressed and connected and fixed by means of the internal structure of the air guide tube (2), and the jet outlet height h of the supersonic air film assembly is further limited by means of the internal structure. e At the same time, the air film assembly (5) and the connecting fixture (4) form a Laval nozzle profile that generates a supersonic cooling air film. According to the required supersonic air film flow parameters, the jet outlet height h is determined based on e Replace the air film assembly (5) under the premise of forming a stable supersonic cooling air film; The high-pressure gas source assembly 6 comprises a stop valve (6-1), a pressure reducing valve (6-2), an electromagnetic controller (6-3), a mass flow meter (6-4), and a high-pressure gas source (6-5); wherein: based on the actual flight speed and altitude of the hypersonic aircraft, the stagnation point pressure of the spherical optical head cover is obtained; according to the ratio between the static pressure of the supersonic air film jet outlet and the stagnation point pressure of the spherical optical head cover, the total pressure of the supersonic air film chamber is obtained; and the total pressure supplied by the high-pressure gas source (6-5) to the supersonic air film chamber through the high-pressure resistant air guide pipe is controlled by adjusting the stop valve (6-1), the pressure reducing valve (6-2), and the electromagnetic controller (6-3) in sequence; In the method, the total pressure of the supersonic air film chamber is calculated as: According to the actual flight speed and altitude of the hypersonic aircraft, the incoming flow state parameters are obtained by querying the standard atmosphere, including total pressure p0, static pressure p1, static temperature T1 and total temperature T0; According to the Newton-Lees pressure formula, the total stagnation pressure p of the spherical optical head cover of the hypersonic vehicle is obtained. 0,2 ; Among them, C pmax is the correction coefficient, which represents the stagnation point pressure coefficient after the positive shock wave; p 0,2 is the total pressure at the stagnation point after the normal shock wave; p1 is the static pressure of the incoming flow; p2 is the static pressure after the normal shock wave; Ma1 is the flight Mach number of the hypersonic vehicle, Ma2 is the Mach number after the bow shock wave of the hypersonic spherical optical head cover, γ is the specific heat ratio, and θ is the inclination angle of the local object surface; According to the set jet pressure ratio NPR = p c,02 / p 0,2 , and obtain the static pressure p of the supersonic air film jet outlet c,02 ; According to the isentropic relationship and the designed supersonic air film jet outlet Mach number Ma c , determine the total pressure p of the supersonic air film chamber c,01 :

2. The method for supersonic film cooling of a spherical optical head cover of a hypersonic aircraft according to claim 1, characterized in that: In the method, the cooling efficiency is used to characterize the cooling effect of the supersonic cooling film on the hypersonic spherical optical head cover. There are two ways to calculate the cooling efficiency: or Among them, T ∞ is the incoming flow temperature, T w is the wall temperature of the spherical optical head cover, T c is the coolant temperature, Q c is the heat flux on the wall of the spherical optical head cover after air film cooling, Q0 is the heat flux on the wall of the spherical optical head cover without air film cooling; when η≥1, it indicates an effective air film insulation cooling zone, otherwise it is an effective air film heat reduction cooling zone.

3. The method for supersonic film cooling of a spherical optical head cover of a hypersonic aircraft according to claim 2, characterized in that: In the method, the position of the air film assembly (5) and the connecting fixture (4) is determined by the Laval nozzle throat height h* in the air film assembly (5), which is calculated as follows: Among them, R A is the area ratio, Ma c is the Mach number of the air film jet outlet, A e is the annular air film jet outlet area, A * is the cross-sectional area of ​​the annular Laval nozzle throat, h1 is the height of the air film jet outlet from the central axis, h2 is the height of the Laval nozzle throat from the central axis, and h e is the height of the air film jet outlet.

4. A method for supersonic film cooling of a spherical optical head cover of a hypersonic aircraft according to claim 3, characterized in that: In the method, when the aircraft is in a hypersonic flight state, the fixing member (4) is connected to the air film assembly (5) to form a Laval nozzle profile, thereby generating a supersonic air film, which is squeezed and attached to the wall surface of the spherical optical head cover through the front bow shock wave of the spherical optical head cover, thereby resisting the aerodynamic heat received by the hypersonic spherical optical head cover, and adjusting the flow parameters of the supersonic air film to meet the needs of adapting to different flight environments of the hypersonic aircraft.

5. A hypersonic film cooling system for a spherical optical head cover of a hypersonic aircraft, characterized in that: The system comprises: a quick-connect plug (1), an air guide tube (2), a spherical optical head cover window (3), a connecting fixture (4), an air film assembly (5), and a high-pressure air source assembly (6); wherein: The spherical optical head cover window (3) is fastened to the connecting fixture (4) by means of screws, and the inner and outer contours of the two are matched by adjusting the depth of the screws; the air film assembly (5) and the connecting fixture (4) are pressed and connected and fixed by means of the internal structure of the air guide tube (2), and the jet outlet height h of the supersonic air film assembly is further limited by means of the internal structure. e At the same time, the air film assembly (5) and the connecting fixture (4) form a Laval nozzle profile that generates a supersonic cooling air film. According to the required supersonic air film flow parameters, the jet outlet height h is determined based on e Replace the air film assembly (5) under the premise of forming a stable supersonic cooling air film; The high-pressure gas source assembly 6 comprises a stop valve (6-1), a pressure reducing valve (6-2), an electromagnetic controller (6-3), a mass flow meter (6-4), and a high-pressure gas source (6-5); wherein: based on the actual flight speed and altitude of the hypersonic aircraft, the stagnation point pressure of the spherical optical head cover is obtained; according to the ratio between the static pressure of the supersonic air film jet outlet and the stagnation point pressure of the spherical optical head cover, the total pressure of the supersonic air film chamber is obtained; and the total pressure supplied by the high-pressure gas source (6-5) to the supersonic air film chamber through the high-pressure resistant air guide pipe is controlled by adjusting the stop valve (6-1), the pressure reducing valve (6-2), and the electromagnetic controller (6-3) in sequence; Among them, the total pressure of the supersonic air film chamber is calculated as follows: According to the actual flight speed and altitude of the hypersonic aircraft, the incoming flow state parameters are obtained by querying the standard atmosphere, including total pressure p0, static pressure p1, static temperature T1 and total temperature T0; According to the Newton-Lees pressure formula, the total stagnation pressure p of the spherical optical head cover of the hypersonic vehicle is obtained. 0,2 ; Among them, C pmax is the correction coefficient, which represents the stagnation point pressure coefficient after the positive shock wave; p 0,2 is the total pressure at the stagnation point after the normal shock wave; p1 is the static pressure of the incoming flow; p2 is the static pressure after the normal shock wave; Ma1 is the flight Mach number of the hypersonic vehicle, Ma2 is the Mach number after the bow shock wave of the hypersonic spherical optical head cover, γ is the specific heat ratio, and θ is the inclination angle of the local object surface; According to the set jet pressure ratio NPR = p c,02 / p 0,2 , and obtain the static pressure p at the supersonic air film jet outlet c, 02; According to the isentropic relationship and the designed supersonic air film jet outlet Mach number Ma c , determine the total pressure p of the supersonic air film chamber c,01 :

6. A hypersonic aircraft spherical optical head cover supersonic air film cooling system according to claim 5, characterized in that: The cooling efficiency is used to characterize the cooling effect of the supersonic cooling film on the hypersonic spherical optical head cover. There are two ways to calculate the cooling efficiency: or Among them, T ∞ is the incoming flow temperature, T w is the wall temperature of the spherical optical head cover, T c is the coolant temperature, Qc is the heat flux on the wall of the spherical optical head cover after air film cooling, and Q0 is the heat flux on the wall of the spherical optical head cover without air film cooling; when η≥1, it indicates an effective air film insulation cooling zone, otherwise it is an effective air film heat reduction cooling zone.

7. A hypersonic film cooling system for a spherical optical head cover of a hypersonic aircraft according to claim 6, characterized in that: The position of the air film assembly (5) and the connecting fixture (4) is determined by the Laval nozzle throat height h* in the air film assembly (5), which is calculated as follows: Among them, R A is the area ratio, Ma c is the Mach number of the air film jet outlet, A e is the annular air film jet outlet area, A * is the cross-sectional area of ​​the annular Laval nozzle throat, h1 is the height of the air film jet outlet from the central axis, h2 is the height of the Laval nozzle throat from the central axis, and h e is the height of the air film jet outlet.

8. A hypersonic aircraft spherical optical head cover supersonic air film cooling system according to claim 7, characterized in that: When the aircraft is in a hypersonic flight state, the connecting fixture (4) and the air film assembly (5) form a Laval nozzle profile, thereby generating a supersonic air film, which is squeezed and attached to the wall surface of the spherical optical head cover through the front bow shock wave of the spherical optical head cover, thereby resisting the aerodynamic heat received by the hypersonic spherical optical head cover, and adjusting the flow parameters of the supersonic air film to meet the needs of adapting to different flight environments of the hypersonic aircraft.

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