Design method and system for pneumatic stealth integrated air inlet channel

Abstract

According to the design method and system for the pneumatic stealth integrated air inlet channel, a designed bulge plays a role in decelerating and pressurizing airflow, and displacement of a low-energy boundary layer is achieved through the transverse pressure gradient generated by the wall face of the bulge; in addition, the bulge also plays a role in shielding the air inlet channel and weakening radar wave reflection of the air inlet channel cavity, so that the forward radar scattering cross section of the air inlet channel is effectively reduced; on one hand, the designed sawtooth lip is matched with a bump molded surface, so that shock waves generated when free incoming flow passes through the bump wall surface can be better intersected at the front edge of the lip, shock wave sealing is realized, and overflow is reduced; on the other hand, reflection echoes deviate from a forward key angular domain through the sawtooth design of the front edge of the lip, so that the forward radar scattering area is reduced, and the design requirement of the hypersonic speed air inlet channel can be well met through the integrated design of the bump and the sawtooth lip.

Description

Technical Field

[0001] This application relates to the field of aerospace technology, and particularly to a design method and system for an integrated aerodynamic and stealth inlet. Background Art

[0002] In the hypersonic flow field of an aircraft, the shock wave / boundary layer interference phenomenon is extremely likely to occur, resulting in an increase in the instability of the flow field, and then phenomena such as boundary layer separation, strong local heating, transonic flutter, and inlet surge occur. For example, when the inlet captures the airflow, due to the presence of the forebody fuselage, the oncoming flow will form a boundary layer near the fuselage wall, and as the length of the fuselage increases, the boundary layer continuously thickens. If the low-energy airflow in the boundary layer enters the engine, it will reduce the working efficiency of the engine. In addition, as the boundary layer thickens, the boundary layer is also prone to separation, affecting the flow stability, and even causing the aerodynamic performance to deviate from the designed state. Therefore, controlling the boundary layer flow is particularly important in the design of air-breathing hypersonic aircraft. Currently, the control of the boundary layer mainly includes active control technology and passive control technology. The boundary layer passive control technology has received extensive attention due to its simple structure, high method feasibility, and remarkable effect, and mainly includes boundary layer suction, bypass inlet, bump inlet, etc. Among them, the bump inlet has the characteristics of simple structure and excellent performance. The compression surface of the bump inlet has normal and transverse pressure gradients, and the combined action of the two is equivalent to a passive boundary layer blowing device, which can blow most of the boundary layer airflow on the fuselage outside the inlet. Therefore, good total pressure recovery coefficients can be obtained without taking boundary layer bypass or blowing / suction measures on the bump inlet. In addition, the hypersonic inlet faces a severe starting problem: when the oncoming flow Mach number is too low or the inlet back pressure is too high, the inlet will be in a non-starting state. At this time, large-scale separated flow will form in the contraction section of the inlet, the total pressure recovery of the airflow will drop significantly, and at the same time, the inlet capture flow will also drop significantly. At this time, the engine cannot generate enough thrust, and in severe cases, it will cause the engine to flame out. An important factor in the inlet starting problem is the interference between the lip shock wave and the boundary layer, resulting in flow separation and finally forming a large-scale separation zone. It can be seen that the boundary layer plays a key role in this process. Therefore, a supersonic bump is introduced into the hypersonic inlet, and the bump is used to displace the boundary layer in order to achieve the purpose of improving the starting performance.

[0003] There are mainly three major scattering sources in the radar stealth design of aircraft, namely the aircraft cockpit, the radome, and the inlet cavity. According to the different characteristics and importance of aircraft radar scattering, the radar scattering of aircraft is usually divided into the forward sector, the lateral sector, and the rearward sector according to the azimuth angle for analysis. Since the interception trajectories during penetration are mostly head-on attack methods for various aircraft, the forward sector is the key angular domain that needs to be considered in the stealth design for all types of aircraft. A large amount of data shows that the inlet is the scattering source with the widest influence and the strongest scattering in the forward sector of combat aircraft. Generally, the inlet of a single-engine aircraft can account for about 40% of the radar scattering area of the whole aircraft in the forward sector, while for a twin-engine aircraft, this ratio can reach 60% or even higher. Whether the radar scattering of the inlet can be effectively controlled is directly related to the RCS level of the whole aircraft in the forward sector. At the same time, due to the complexity of the radar scattering of the inlet and its close connection with the aircraft engine (which is related to the flight safety or flight performance of the aircraft), it makes the RCS control very difficult. Moreover, the layout forms of combat aircraft are diverse, and the inlet forms are also various, which further increases the difficulty of RCS control of the inlet system.

[0004] The main radar stealth design methods for inlets include S-bend inlet design, inlet grille design, beveled inlet design, inlet conformal design, etc. However, the inlet grille design is not applicable to the design of hypersonic inlets. The beveled inlet causes adverse effects such as flow separation in the upper part of the cut angle during ground takeoff, obvious single vortex flow on the outlet section, large flow distortion, and low total pressure recovery. These designs are difficult to match the design requirements of hypersonic inlets. Summary of the Invention

[0005] This application proposes a pneumatic and stealth integrated inlet design method and system, which can solve one of the problems existing in the background technology.

[0006] To achieve the above object, this application adopts the following technical solutions: In the first aspect, a pneumatic and stealth integrated inlet design method is provided. The method includes: a serrated lip and bump compression surface design part, and a serrated lip and bump integrated inlet design part, where, The serrated lip and bump compression surface design part includes: According to the design input parameters of the inlet, determine the incoming flow Mach number , capture height h, and inlet width W. The design input parameters include the design Mach number and design height; Determine the parameters of the first shock wave, including the first shock wave angle and the incoming flow Mach number; Based on the parameters of the first shock wave, Oswatisch's optimal wave system theory, and the oblique shock wave relationship, iteratively solve the parameters of the Mth shock wave, where M is a positive integer greater than 1; Determine the two-dimensional position of the inlet profile using the shock wave parameters and the capture height; and Based on the given sawtooth angle θ, the half-mode width of the inlet , the number of points N at which the half-mode of the inlet is discretized in the width direction, and the variation law of the first shock wave angle relative to the sawtooth angle θ, obtain the position coordinates of N inlet cross-sections, where the half-mode width of the inlet is half of the inlet width W, The sawtooth lip and bump integrated inlet design part includes: Based on the position coordinates, the inner surface parameters and the outer surface parameters, call a 3D modeling software to connect and generate the integrated inlet profile of the bump and the sawtooth lip.

[0007] Based on the above technical solutions, the designed bump plays a role in decelerating and pressurizing the air flow, and realizes the displacement of the low-energy boundary layer through the lateral pressure gradient generated by the bump wall surface; in addition, the bump also plays a role in shielding the inlet, weakening the radar wave reflection of the inlet cavity, and thus effectively reducing the forward radar cross-section of the inlet; the designed sawtooth lip, on the one hand, cooperates with the bump profile, so that the shock wave generated by the free oncoming flow passing through the bump wall surface can better intersect at the leading edge of the lip, realizing shock wave sealing, thereby reducing overflows; on the other hand, the sawtooth design at the leading edge of the lip makes the reflected echo deviate from the forward key angular domain, thereby reducing the forward radar scattering area. Therefore, this integrated design of the bump and the sawtooth lip can excellently meet the design requirements of the hypersonic inlet.

[0008] In a possible design manner of the first aspect, according to the design input parameters of the inlet, determine the capture height and the inlet width, specifically including: According to the design Mach number and the design height of the aircraft, obtain the oncoming flow Mach number, the oncoming flow density and the local sound speed of the inlet; According to the oncoming flow Mach number and the local sound speed, determine the oncoming flow velocity; According to the thrust / drag matching requirement of the aircraft, determine the capture flow rate of the inlet; Determine the capture area of the inlet through the capture flow rate, the oncoming flow velocity and the oncoming flow density; and According to the capture area of the inlet, determine the capture height and the inlet width.

[0009] In a possible design of the first aspect, a first relationship is established among the shock wave angle, the wavefront Mach number, and the airflow deflection angle. A second relationship is established among the normal Mach number ahead of the shock wave, the wavefront Mach number, and the shock wave angle. A third relationship is established between the normal Mach number ahead of the M-th shock wave and the normal Mach number behind the M-th shock wave. A fourth relationship is established among the shock wave angle of the M-th shock wave, the airflow deflection angle of the M-th shock wave, the Mach number behind the M-th shock wave, and the normal Mach number behind the M-th shock wave. A fifth relationship is established among the shock wave angle of the M-th shock wave, the wavefront Mach number, the Mach number behind the wave, and the shock wave angle of the (M + 1)-th shock wave. Based on the first relationship, the second relationship, the third relationship, the fourth relationship, and the fifth relationship, the parameters of the M-th shock wave are iteratively solved.

[0010] In a possible design of the first aspect, using the shock wave parameters and the capture height, the two-dimensional position of the inlet profile is determined, specifically including: Using the capture height and the shock wave angle of the first shock wave to determine the starting position and the ending position of the inlet profile; and Based on the starting position, the ending position, the shock wave angle, and the airflow deflection angle, determine the positions of points between the starting point and the ending point of the inlet profile.

[0011] In a possible design of the first aspect, the half model of the inlet is discretized into N points along the width direction, and the variation law is determined by the following parameters: the shock wave angle of the first shock wave at the (N - 1)-th section, the distance change value of the ending position relative to the starting position, and the projected distance along the axial direction of the distance between the ending positions of adjacent sections.

[0012] In a possible design of the first aspect, the projected distance is jointly determined by the width of the inlet half model, the number of discrete points, and the sawtooth angle.

[0013] In a possible design of the first aspect, based on the position coordinates, the inner surface parameters, and the outer surface parameters, a three-dimensional modeling software is called to connect and generate an integrated inlet profile of the bump and the sawtooth lip, specifically: Based on the position coordinates and the preset length parameter of the extended line segment of the inner surface, perform line segment extension and connection processing on the points constituting the inner surface. On the basis of the line segment extension and connection processing, perform lofting filling to obtain the inner surface of the integrated inlet profile; and On the basis of constructing the inner surface, based on the position coordinates and the preset surface thickening parameter and surface included angle parameter of the outer surface, construct line segments, surfaces, and planes of the outer surface, and perform lofting filling to obtain the outer surface of the integrated inlet profile.

[0014] In the second aspect, a pneumatic and stealth integrated inlet design system is provided, and the system includes: The serrated lip and bump compression surface design module is used to determine the incoming flow Mach number, capture height, and inlet width W according to the design input parameters of the inlet. The design input parameters include the design Mach number and design height; determine the parameters of the first shock wave, including the first shock wave angle and the incoming flow Mach number; based on the parameters of the first shock wave, Oswatisch's optimal wave system theory, and the oblique shock relation, iteratively solve the parameters of the Mth shock wave, where M is a positive integer greater than 1; use the shock wave parameters and the capture height to determine the two-dimensional position of the inlet profile; based on the given serration angle θ, half-mode width of the inlet, number of points N for discretizing the inlet width, and the variation law of the first shock wave angle relative to the serration angle θ, obtain the position coordinates of N inlet cross-sections. The half-mode width of the inlet is half of the inlet width W; and The serrated lip and bump integrated inlet design module is used to call a 3D modeling software to connect and generate the integrated inlet profile of the bump and serrated lip based on the position coordinates, inner surface parameters, and outer surface parameters.

[0015]

[0016] Figure 1 Figure 2 Figure 3 Figure 4 Figure 5 Figure 6 Figure 7 ​​​​​​​​​​​Schematic diagram of the integrated sawtooth lip and bump inlet provided by an embodiment of the present application; Figure 8 Schematic diagram of the profile parameters of the integrated sawtooth lip and bump inlet provided by an embodiment of the present application, where (a) is the sectional view and (b) is the overall view; Figure 9 RCS distribution diagram of the 30° angular range in the forward direction of the horizontal plane of the inlet provided by an embodiment of the present application, where (a) is the vertical polarization and (b) is the horizontal polarization; Figure 10 RCS distribution diagram of the angular range of 10° depression angle - 30° elevation angle in the vertical plane of the inlet provided by an embodiment of the present application, where (a) is the vertical polarization and (b) is the horizontal polarization; Figure 11 Mach number contour map of the symmetry plane of the inlet provided by an embodiment of the present application, where (a) is the integrated bump inlet and (b) is the conventional inlet; Figure 12 Total pressure recovery coefficient contour map of the symmetry plane of the inlet provided by an embodiment of the present application, where (a) is the integrated bump inlet and (b) is the conventional inlet; Figure 13 Total pressure recovery coefficient contour map of the outlet section at x = 3800mm of the inlet provided by an embodiment of the present application, where (a) is the integrated bump inlet and (b) is the conventional inlet. Detailed implementation manners

[0017] In order to make the objectives, technical solutions and advantages of the present application clearer, the present application will be further described in detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present application and are not used to limit the present application.

[0018] It should be noted that although functional module division is performed in the schematic diagram and the logical sequence is shown in the flowchart, in some cases, the steps shown or described may be executed in a different order from the module division in the program or the sequence in the flowchart. Terms such as "first" and "second" in the description, claims and the above-mentioned drawings are used to distinguish similar objects and do not necessarily need to describe a specific order or sequence.

[0019] Unless otherwise defined, all technical and scientific terms used herein have the same meaning as commonly understood by those skilled in the technical field to which this application belongs. The terms used herein are only for the purpose of describing the embodiments of this application and are not intended to limit this application.

[0020] An embodiment of the present application proposes a design method for an aerodynamic and stealth integrated inlet, which is divided into two parts: the design of the serrated lip / bump pre-compression surface and the design of the serrated lip / bump integrated inlet profile. The bump designed in this embodiment plays a role in decelerating and pressurizing the airflow, and realizes the displacement of the low-energy boundary layer through the lateral pressure gradient generated by the bump wall surface; in addition, the bump also plays a role in shielding the inlet, weakening the radar wave reflection of the inlet cavity, and thus effectively reducing the forward radar cross-section of the inlet. On the one hand, the serrated lip designed in this embodiment cooperates with the bump profile, so that the shock wave generated by the free incoming flow passing through the bump wall surface can better intersect at the leading edge of the lip, realizing shock wave sealing, thereby reducing overflows; on the other hand, the serrated design at the leading edge of the lip makes the reflected echo deviate from the forward key angular domain, thereby reducing the forward radar scattering area.

[0021] I. Design Method The design method for the aerodynamic and stealth integrated inlet is divided into the design of the serrated lip / bump pre-compression surface and the design of the serrated lip / bump integrated inlet profile. As Figure 1 shown, the main steps are: 1. Given the design parameters according to the mission requirements; 2. Obtain the wavefront airflow parameters; 3. Calculate the post-wave airflow parameters through the oblique shock wave relation; 4. Calculate the next shock wave angle through the Oswatisch optimal wave system theory; 5. Calculate the coordinate positions of each wedge angle; 6. Calculate the shock wave angle of the incoming flow at the next section according to the serration control law; 7. Replace the shock wave angle in the design parameters with the shock wave angle of the incoming flow at the new section, and obtain the coordinate positions of each wedge angle at a series of sections through steps 2, 3, 4, 5, and 6; 8. Obtain the serrated lip and the bump pre-compression surface through lofting; 9. Generate the serrated lip / bump integrated inlet profile through the inner and outer surface parameters in the 3D modeling software.

[0022] 1. Design of the serrated lip / bump pre-compression surface First, determine the design input parameters of the inlet, including the incoming flow parameters of the inlet and the capture area of the inlet. According to the flight Mach number and flight altitude of the aircraft, the incoming flow Mach number , incoming flow static pressure , incoming flow density , incoming flow static temperature , local sound speed of the inlet can be obtained. According to the thrust / drag matching requirements of the aircraft, the capture flow rate of the inlet can be determined. Through the capture flow rate, incoming flow velocity and incoming flow density , the capture area A of the inlet can be determined, where the incoming flow velocity is obtained through the formula. After obtaining the capture area, the capture height h and the inlet width W can be determined. ; 。

[0023] Solve the shock wave angle and the airflow deflection angle through the Oswatisch optimal wave system theory and the oblique shock wave relation, as Figure 3 shown. The Oswatisch optimal wave system theory is: Given a series of n consecutive shock waves, which include (n - 1) oblique shock waves and a terminal normal shock wave, when the total pressure recovery coefficients of each oblique shock wave are equal, the total pressure recovery coefficient of the entire shock wave system is the highest. The mathematical representation is as follows: ; where is the total pressure of the oncoming flow, is the total pressure of the airflow before passing through the (i + 1)-th shock wave. The formula can also be changed to the following form: ; where is the Mach number of the oncoming flow, is the Mach number before the (i + 1)-th shock wave, is the shock wave angle of the (i + 1)-th shock wave.

[0024] Determine the first shock wave angle , and obtain the airflow deflection angle from the oblique shock wave relation.

[0025] The normal Mach number before the shock wave is .

[0026] Relationships between flow parameters before and after the wave: ; ; ; ; ; ; .

[0027] Among them, γ is the specific heat ratio of the fluid. is the difference between the shock wave angle and the airflow deflection angle.

[0028] First, substitute the oncoming flow Mach number into the formula to obtain the normal Mach number after the shock wave. Then substitute the normal Mach number , static pressure , density , velocity , and temperature into the shock wave relation respectively to obtain the normal Mach number after the wave , static pressure , density , velocity , temperature , and then the second shock wave angle is obtained from the formula . Substitute into the formula to obtain the flow deflection angle . Finally, take the flow parameters after the first shock wave as the wavefront flow parameters before the second shock wave and repeat the above calculation steps to derive the flow parameters after the second shock wave , static pressure , density , velocity , temperature .

[0029] From the formula: Calculate the internal compression angle , and then substitute , into the formula to obtain the internal compression shock wave angle .

[0030] After calculating the shock wave angle and the flow deflection angle, the wall position can be obtained according to the following formula, as shown in Figure 4 . ; ; ; .

[0031] Given a shock wave angle, the two-dimensional profile of the inlet can be obtained according to the Oswatisch optimal wave system theory and the oblique shock wave intensity relationship. For a traditional two-dimensional inlet, the inlet width is calculated by the capture flow and then the two-dimensional profile is stretched to obtain a three-dimensional two-dimensional inlet. For a serrated lip / bump integrated inlet, the first shock wave angle is controlled to make the inlet lip serrated and the compression surface bulged. The following gives the design process

[0032] First, given the serration angle θ and the half-mode width of the inlet , the half-mode width of the inlet is half of the inlet width. The half-mode width of the inlet is discretized into N points, and then by controlling the first shock wave angle , the position coordinates of N transverse inlet cross-sections are obtained (as shown in Figure 6 ). Finally, the serrated lip / bump integrated inlet profile is generated by connecting with 3D modeling software. The formula is the variation law of the first shock wave angle , where the subscript k represents the kth cross-section. Connect , , …, , … Get curve, connect , , …, , … Get curve, connect , , …, , … Get curve, connect , , …, , … Get curve, then connect the curves , , , , , as Figure 5 shown.

[0033] Formulas (18)-(20) are as follows: ; ; .

[0034] 2. Integrated design of serrated lip / dome inlet Subsequently, complete the surface construction in 3D modeling software to obtain an inlet with aerodynamic stealth integration, as Figure 7 shown. The specific steps are as Figure 8 shown. The first step is to generate the inner surface. The specific method is to extend backward from the serrated lip / pre-compression dome surface with point as the starting point for to obtain line segment ; extend horizontally from point by W / 2 to the symmetry plane position to obtain line segment ; respectively extend backward from , as the starting points to the point cross-section position to obtain line segments , ; connect line segments , ; extend forward from as the starting point for to obtain line segment , connect points , to obtain line segment ; take a point on line segment ​ , make the length of line segment equal to , connect points and to obtain line segment ; The smooth inner surface of the inlet can be obtained by using the lofting and filling tools in the modeling software , , , . Then generate the outer surface. The specific method is to create a straight line segment at the position of point . The x - coordinate of point E is equal to the x - coordinate of point . The included angle between line segment and the vertical plane is , and the included angle with the horizontal plane is . Intercept points B and C with the same x - coordinate as points on line segment and to obtain line segments and line segment ; Thicken the surfaces of planes and by , and then perform an oblique cut at the lip and positions at an angle of to obtain the outer surfaces of the lower plate and the side plates; Connect point F on the outer surface of the side plate and point E on the leading edge profile of the fuselage to obtain line segment EF; Finally, obtain planes and , and surface through lofting and filling. Thus, the design of the serrated lip / bump integrated inlet is completed.

[0035] This embodiment realizes the displacement of the boundary layer by the pre - compression surface, and effectively controls the boundary layer thickness of the inlet; The forward radar cross - section of the inlet is effectively reduced by the shielding of the bump to the inlet cavity.

[0036] When the design Mach number is 6, the design altitude is 25 km, the capture flow rate is 34.45 kg / s, the first shock wave angle is 12 degrees, the capture height of the inlet is 400 mm, the width of the inlet is 600 mm, the serration angle is 120 degrees, the two shock waves M = 2, and the number of discrete points N = 51; The outer surface design parameters = 1300 mm, = 500 mm, the wall thickness of the inlet = 10 mm, the lip cut angle = 10°, the included angle between the leading edge profile of the fuselage and the vertical plane = 2°, the included angle with the horizontal plane = 5°, as shown in Figure 9 andFigure 10 As shown in the figure, compared with the conventional inlet, the vertical polarization RCS logarithm mean value of the inlet designed by the design method of the present application is reduced by 2.1757dBsm in the 30° angle domain facing the horizontal plane, with a reduction of 14.01%; the horizontal polarization RCS logarithm mean value is reduced by 0.05 dBsm, with a reduction of 0.37%. In the vertical plane depression angle range of 10° to 30° elevation angle, compared with the conventional inlet, the vertical polarization RCS logarithm mean value is reduced by 2.50dBsm, with a reduction of 94.01%; the horizontal polarization RCS logarithm mean value is reduced by 2.44dBsm, with a reduction of 47.69%.

[0037] Compared with conventional air inlets, the air inlet designed by the present application can effectively realize the displacement of the low-energy boundary layer. By intercepting the symmetric planes z=0mm, z=100mm, z=200mm, and z=300mm compression surfaces 0~3Ma, and counting the average thickness of the third wedge angle, it is found that the low-energy boundary layer thickness of the air inlet designed by the present application is reduced by 0.8mm compared with the conventional air inlet. Figure 11 This is the Mach number cloud diagram at the symmetry plane.

[0038] like Figure 12 , Figure 13 As shown, at 6Ma, the average total pressure recovery coefficient of the conventional air inlet at the section x=3800mm is 64%, and the flow coefficient is 82%; the average total pressure recovery coefficient of the sawtooth lip bulge integrated air inlet at the section x=3800mm is 53%, and the flow coefficient is 76%. Although the aerodynamic performance of the air inlet designed by the method of this application is reduced, its stealth performance is significantly improved. In general, the stealth performance is significantly improved in exchange for sacrificing some aerodynamic performance, so as to achieve the effect of enhancing the stealth performance of the aircraft.

[0039] The present application also provides a pneumatic stealth integrated air inlet design system, the system comprising: The serrated lip and bulge compression surface design module is used to determine the incoming flow Mach number, capture height and inlet width according to the design input parameters of the inlet, wherein the design input parameters include the design Mach number and the design height; determine the parameters of the first shock wave, including the first shock wave angle, and the incoming flow Mach number; iteratively solve the parameters of the Mth shock wave based on the parameters of the first shock wave, Oswatisch optimal wave system theory and the oblique shock wave relationship, where M is a positive integer greater than 1; use the shock wave parameters and the capture height to determine the two-dimensional position of the inlet profile; based on the given sawtooth angle θ, the inlet half-mode width , the number of points N at which the width of the inlet is discretized, and the variation law of the first shock wave angle relative to the sawtooth angle θ, the position coordinates of N inlet cross sections are obtained, the inlet half-mode width is half of the width W of the air inlet duct; and The integrated inlet design module with serrated lip and bump is used to call 3D modeling software to connect and generate the integrated inlet profile of the bump and serrated lip based on the position coordinates, inner surface parameters and outer surface parameters.

[0040] The embodiment of the present application also provides a computer program product, including: a computer program or instruction, when the computer program or instruction runs on a computer, enabling the computer to execute the method of any one of the above possible implementation manners.

[0041] The above is the preferred implementation manner of the present application. It should be noted that for those of ordinary skill in the art, without departing from the principle of the present application, several improvements and refinements can be made, and these improvements and refinements are also regarded as the protection scope of the present application.

Claims

1. A method for designing an aerodynamic stealth integrated air inlet, characterized in that: The method comprises: a serrated lip and a bulge compression surface design part, and a serrated lip and a bulge integrated air intake duct design part, wherein: The serrated lip and bulge compression surface design parts include: Determine the incoming flow Mach number based on the design input parameters of the inlet duct , capture height h and inlet width W, wherein the design input parameters include a design Mach number and a design height; Determining parameters of the first shock wave, including the first shock wave angle and the incoming flow Mach number; Based on the parameters of the first shock wave, the Oswatisch optimal wave system theory and the oblique shock wave relationship, the parameters of the Mth shock wave are iteratively solved, where M is a positive integer greater than 1; Determining the two-dimensional position of the inlet profile using the shock wave parameters and the capture height; and Based on the given sawtooth angle θ, the inlet half-mold width , the number of points N of discretization of the inlet half-mold along the width direction, and the variation law of the first shock wave angle relative to the sawtooth angle θ, obtain the position coordinates of N inlet cross sections, the inlet half-mold width is half of the inlet duct width W, The serrated lip and bulge integrated air intake design part includes: Based on the position coordinates, inner surface parameters and outer surface parameters, three-dimensional modeling software is called to connect and generate an integrated air intake surface of the bulge and the serrated lip.

2. The method according to claim 1, characterized in that The capture height and inlet width are determined based on the design input parameters of the inlet, including: According to the design Mach number and design altitude of the aircraft, the incoming flow Mach number, incoming flow density and local sound speed of the inlet are obtained; Determining the incoming flow velocity according to the incoming flow Mach number and the local sound speed; Determine the inlet capture flow rate based on the thrust / drag matching requirements of the aircraft; Determining an intake duct capture area according to the capture flow, the incoming flow velocity and the incoming flow density; and The capture height and the width of the air intake duct are determined according to the capture area of ​​the air intake duct.

3. The method according to claim 1, characterized in that A first relationship is established between the shock wave angle, the wavefront Mach number and the airflow deflection angle, a second relationship is established between the shock wave front normal Mach number, the wavefront Mach number and the shock wave angle, a third relationship is established between the shock wave front normal Mach number of the Mth shock wave and the post-shock normal Mach number of the Mth shock wave, a fourth relationship is established between the shock wave angle of the Mth shock wave, the airflow deflection angle of the Mth shock wave, the post-shock Mach number of the Mth shock wave, and the post-shock normal Mach number of the Mth shock wave, a fifth relationship is established between the shock wave angle, the wavefront Mach number, the post-wave Mach number of the Mth shock wave and the shock wave angle of the M+1th shock wave, and the parameters of the Mth shock wave are iteratively solved based on the first relationship, the second relationship, the third relationship, the fourth relationship and the fifth relationship.

4. The method according to claim 1, characterized in that The two-dimensional position of the inlet profile is determined by using the shock wave parameters and the capture height, specifically including: Determining the starting position and the ending position of the inlet profile by using the capture height and the first shock wave angle; and Based on the starting point position, the ending point position, the shock wave angle and the airflow deflection angle, the position of each point between the starting point and the ending point of the air inlet profile is determined.

5. The method according to claim 4, characterized in that The inlet half mold is discretized into N points along the width direction, and the change rule is determined by the following parameters: the first shock wave angle of the N-1th section, the distance change value of the end position relative to the starting position, and the projection distance of the distance between the end positions of adjacent sections along the axial direction.

6. The method according to claim 5, characterized in that The projection distance is determined by the half-mold width of the air inlet, the number of discrete points and the sawtooth angle.

7. The method according to claim 6, characterized in that Based on the position coordinates, inner surface parameters and outer surface parameters, the three-dimensional modeling software is called to connect and generate the integrated intake duct profile of the bulge and the serrated lip, specifically: Based on the position coordinates and the preset inner surface extension line segment length parameters, line segment extension and connection processing are performed on the points constituting the inner surface, and on the basis of the line segment extension and connection processing, lofting and filling are performed to obtain the inner surface of the integrated air inlet profile; as well as On the basis of the inner surface construction, based on the position coordinates and the preset outer surface curved surface thickening parameters and surface angle parameters, the line segments, curved surfaces and planes of the outer surface are constructed, and lofting and filling are performed to obtain the outer surface of the integrated air inlet profile.

8. An aerodynamic stealth integrated air intake design system, characterized in that: The system comprises: The serrated lip and bulge compression surface design module is used to determine the incoming flow Mach number based on the design input parameters of the inlet duct. , capture height h and inlet width W, wherein the design input parameters include the design Mach number and the design height; determining the parameters of the first shock wave, including the first shock wave angle, and the incoming flow Mach number; iteratively solving the parameters of the Mth shock wave based on the parameters of the first shock wave, Oswatisch optimal wave system theory and the oblique shock wave relationship, where M is a positive integer greater than 1; determining the two-dimensional position of the inlet profile using the shock wave parameters and the capture height; based on the given sawtooth angle θ, the inlet half-mode width , the number of points N of discretization of the inlet half-mold along the width direction, and the variation law of the first shock wave angle relative to the sawtooth angle θ, obtain the position coordinates of N inlet cross sections, the inlet half-mold width is half of the width W of the air inlet duct; and The serrated lip and bulge integrated air intake design module is used to call the three-dimensional modeling software to connect and generate the bulge and serrated lip integrated air intake surface based on the position coordinates, inner surface parameters and outer surface parameters.

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