A pneumatic and stealth integrated inlet design method and system

Abstract

This application proposes a design method and system for an integrated pneumatic and stealth inlet. The designed bump plays a role in decelerating and pressurizing the airflow, and through the lateral pressure gradient generated by the bump wall surface, the displacement of the low-energy boundary layer is achieved. In addition, the bump also plays a role in shielding the inlet, weakening the radar wave reflection of the inlet cavity, thereby effectively reducing the forward radar cross section of the inlet. The designed serrated lip, on the one hand, cooperates with the bump profile, enabling the shock wave generated by the free stream passing through the bump wall surface to better intersect at the leading edge of the lip, achieving shock wave sealing and thus reducing overflows. On the other hand, through the serration design at the leading edge of the lip, the reflected echo is deviated from the forward key angular domain, thereby reducing the forward radar scattering area. Thus, this integrated design of the bump and serrated lip can excellently meet the design requirements of hypersonic inlets.

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 flow field instability, and then phenomena such as boundary layer separation, strong local heating, transonic flutter, and inlet surge occur. For example, when the inlet captures air flow, 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 continues to thicken. If the low-energy air flow 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 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 air flow 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 air flow will drop significantly, and 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] In the radar stealth design of aircraft, there are mainly three major scattering sources, 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, the proportion 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 is very difficult to control its RCS. Moreover, the layout forms of combat aircraft are diverse, and the inlet forms are various, which further increases the difficulty of controlling the RCS of the inlet system.

[0004] The radar stealth design of inlets mainly includes methods such as S-bend inlet design, inlet grille design, beveled inlet design, and inlet conformal design. However, the inlet grille design is not applicable to the design of hypersonic inlets. The beveled inlet causes adverse effects such as flow separation at the upper part of the cut-off 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 method and system for integrated aerodynamic and stealth inlet design, 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:

[0007] In the first aspect, a method for integrated aerodynamic and stealth inlet design 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

[0008] The serrated lip and bump compression surface design part includes:

[0009] 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 the design height;

[0010] Determine the parameters of the first shock wave, including the first shock wave angle and the incoming flow Mach number;

[0011] Based on the parameters of the first shock wave, the Oswatisch optimal wave system theory, and the oblique shock wave relationship, iteratively solve for the parameters of the Mth shock wave, where M is a positive integer greater than 1;

[0012] Determine the two-dimensional position of the inlet profile using the shock wave parameters and the capture height; and

[0013] Based on the given sawtooth angle θ, the half-width of the inlet 、the number of points N for discretizing the half-width of the inlet 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, where the half-width of the inlet is half of the inlet width W,

[0014] The design part of the sawtooth lip and bump integrated inlet includes:

[0015] Based on the position coordinates, inner surface parameters, and outer surface parameters, call 3D modeling software to connect and generate the integrated inlet profile of the bump and sawtooth lip.

[0016] Based on the above technical solutions, the designed bump plays a role in decelerating and pressurizing the airflow, and through the lateral pressure gradient generated by the bump wall surface, the displacement of the low-energy boundary layer is realized; 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 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, through the sawtooth design at the leading edge of the lip, the reflected echo is deviated from the forward key angular domain, thereby reducing the forward radar scattering area. Therefore, this integrated design of the bump and sawtooth lip can excellently meet the design requirements of the hypersonic inlet.

[0017] In a possible design method of the first aspect, according to the design input parameters of the inlet, determine the capture height and the inlet width, specifically including:

[0018] According to the design Mach number and design height of the aircraft, obtain the incoming flow Mach number, incoming flow density, and local sound speed of the inlet;

[0019] According to the incoming flow Mach number and the local sound speed, determine the incoming flow velocity;

[0020] According to the thrust / drag matching requirement of the aircraft, determine the capture flow rate of the inlet;

[0021] Determine the capture area of the inlet through the capture flow rate, the incoming flow velocity, and the incoming flow density; and

[0022] Determine the capture height and the inlet duct width according to the capture area of the inlet duct.

[0023] In a possible design manner 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 before the shock wave, the wavefront Mach number, and the shock wave angle, a third relationship is established between the normal Mach number before the M-th shock wave and the normal Mach number after 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 after the M-th shock wave, and the normal Mach number after the M-th shock wave, and a fifth relationship is established among the shock wave angle of the M-th shock wave, the wavefront Mach number, the Mach number after the wave, and the shock wave angle of the (M + 1)-th shock wave. Iteratively solve the parameters of the M-th shock wave based on the first relationship, the second relationship, the third relationship, the fourth relationship, and the fifth relationship.

[0024] In a possible design manner of the first aspect, use the shock wave parameters and the capture height to determine the two-dimensional position of the inlet duct profile, specifically including:

[0025] Use 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 duct profile; and

[0026] Based on the starting position, the ending position, the shock wave angle, and the airflow deflection angle, determine the positions of the points between the starting point and the ending point of the inlet duct profile.

[0027] In a possible design manner of the first aspect, the half model of the inlet duct 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 of the (N - 1)-th section, the distance variation 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.

[0028] In a possible design manner of the first aspect, the projected distance is jointly determined by the width of the half model of the inlet duct, the number of discrete points, and the sawtooth angle.

[0029] In a possible design manner of the first aspect, based on the position coordinates, the inner surface parameters, and the outer surface parameters, call a three-dimensional modeling software to connect and generate an integrated inlet duct profile of a bump and a sawtooth lip, specifically:

[0030] 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, and on the basis of the line segment extension and connection processing, perform lofting filling to obtain the inner surface of the integrated inlet duct profile; and

[0031] On the basis of the inner surface construction, based on the position coordinates, the preset outer surface curve thickening parameters and the surface angle parameters, construct the line segments, surfaces and planes of the outer surface, and perform lofting filling to obtain the outer surface of the integrated inlet duct profile.

[0032] In a second aspect, a pneumatic stealth integrated inlet duct design system is provided, and the system includes:

[0033] A serrated lip and bump compression surface design module, configured to determine the incoming flow Mach number , capture height and inlet duct width W according to the design input parameters of the inlet duct, where 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; based on the parameters of the first shock wave, the Oswatisch optimal wave system theory and the oblique shock wave relation formula, 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 duct profile; based on the given serration angle θ, the half-mode width of the inlet duct , the number of points N for discretizing the inlet duct width, and the variation law of the first shock wave angle relative to the serration angle θ, obtain the position coordinates of N inlet duct cross-sections, and the half-mode width of the inlet duct is half of the inlet duct width W; and

[0034] A serrated lip and bump integrated inlet duct design module, configured to call a three-dimensional modeling software to connect and generate a serrated lip and bump integrated inlet duct profile based on the position coordinates, inner surface parameters and outer surface parameters. Description of the Drawings

[0035] In order to more clearly illustrate the technical solutions in the embodiments of the present application, the following will briefly introduce the drawings required for use in the embodiments or related technical descriptions. Obviously, the following described drawings are only some embodiments of the present application. For those of ordinary skill in the art, without creative efforts, other drawings can also be obtained based on these drawings.

[0036] Figure 1 is the design flow chart of the serrated lip / bump integrated inlet duct provided by the embodiment of the present application;

[0037] Figure 2 is a schematic diagram of the parameters before and after the shock wave in the oblique shock wave theory adopted by the present application, where the red solid line is the shock wave surface;

[0038] Figure 3 is a schematic diagram of the flow parameters before and after the shock wave of the inlet duct provided by the embodiment of the present application, where the red solid line is the shock wave surface;

[0039] Figure 4It is a schematic diagram of the inlet duct profile position provided by an embodiment of the present application, where the solid dots are key points and the red solid line is the shock wave surface;

[0040] Figure 5 It is a schematic diagram of the inlet duct profile provided by an embodiment of the present application;

[0041] Figure 6 It is a schematic diagram of the serrated lip provided by an embodiment of the present application, where the blue dots are the discrete points of the half mold along the width direction, the black dots are the axial projection points of the end position, and the red dots are the actual points of the end position;

[0042] Figure 7 It is a schematic diagram of the serrated lip and bump integrated inlet duct provided by an embodiment of the present application;

[0043] Figure 8 It is a schematic diagram of the profile parameters of the serrated lip and bump integrated inlet duct provided by an embodiment of the present application, where (a) is the sectional view and (b) is the overall view;

[0044] Figure 9 It is a RCS distribution diagram of the 30° angular domain in front of the horizontal plane of the inlet duct provided by an embodiment of the present application, where (a) is the vertical polarization and (b) is the horizontal polarization;

[0045] Figure 10 It is a RCS distribution diagram of the 10° - 30° angular domain of the depression and elevation angles of the vertical plane of the inlet duct provided by an embodiment of the present application, where (a) is the vertical polarization and (b) is the horizontal polarization;

[0046] Figure 11 It is a Mach number contour map of the symmetry plane of the inlet duct provided by an embodiment of the present application, where (a) is the bump integrated inlet duct and (b) is the conventional inlet duct;

[0047] Figure 12 It is a total pressure recovery coefficient contour map of the symmetry plane of the inlet duct provided by an embodiment of the present application, where (a) is the bump integrated inlet duct and (b) is the conventional inlet duct;

[0048] Figure 13 It is a total pressure recovery coefficient contour map of the outlet section at x = 3800mm of the inlet duct provided by an embodiment of the present application, where (a) is the bump integrated inlet duct and (b) is the conventional inlet duct. Detailed implementation manners

[0049] In order to make the purpose, 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.

[0050] It should be noted that although the functional modules are divided in the schematic diagram and the logical sequence is shown in the flowchart, in some cases, the steps shown or described can be executed in a different module division from that in the program or a different sequence from that in the flowchart. Terms such as "first" and "second" in the specification, claims and the above-mentioned drawings are used to distinguish similar objects and do not necessarily need to describe a specific order or sequence.

[0051] 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.

[0052] An integrated pneumatic and stealth inlet design method is proposed in the embodiments of this application, which is divided into two parts: the design of the serrated lip / bump pre-compression surface and the design of the integrated serrated lip / bump 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 design of the serrations at the leading edge of the lip deviates the reflected echo from the forward key angular domain, thereby reducing the forward radar scattering area.

[0053] I. Design Method

[0054] The integrated pneumatic and stealth inlet design method is divided into the design of the serrated lip / bump pre-compression surface and the design of the integrated serrated lip / bump inlet profile. As Figure 1 shown, the main steps are as follows: 1. Given the design parameters according to the mission requirements; 2. Obtain the airflow parameters of the wavefront; 3. Calculate the airflow parameters behind the wave through the oblique shock relation; 4. Calculate the next shock angle through the Oswatisch optimal wave system theory; 5. Calculate the coordinate positions of each wedge angle; 6. Calculate the shock angle of the incoming flow of the next section according to the serration control law; 7. Replace the shock angle in the design parameters with the shock angle of the incoming flow of the new section, and obtain the coordinate positions of each wedge angle on 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 integrated serrated lip / bump inlet profile through the inner and outer surface parameters in the 3D modeling software.

[0055] 1. Design of the serrated lip / bump pre-compression surface

[0056] First, determine the design input parameters of the inlet, including the inlet flow parameters and the inlet capture area. According to the flight Mach number and flight altitude of the aircraft, the incoming flow Mach number of the inlet can be obtained , the incoming flow static pressure , the incoming flow density , the incoming flow static temperature , the local speed of sound . According to the thrust / drag matching requirements of the aircraft, the inlet capture flow rate can be determined . Through the capture flow rate, the incoming flow velocity and the incoming flow density , the inlet capture area A 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. ; .

[0057] Solve the shock wave angle and the airflow deflection angle through the Oswatisch optimal wave system theory and the oblique shock wave relation, as shown in Figure 3 . The Oswatisch optimal wave system theory is: Given a series of n consecutive shock waves, these shock waves 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 incoming flow total pressure, 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 incoming flow Mach number, is the Mach number before the (i + 1)-th shock wave, is the shock wave angle of the (i + 1)-th shock wave.

[0058] Determine the first shock wave angle , and obtain the airflow deflection angle from the oblique shock wave relation.

[0059] The normal Mach number before the shock wave is .

[0060] The relationship between the flow parameters before and after the wave: ; ; ; ; ; ; 。

[0061] Among them, γ is the specific heat ratio of the fluid. is the difference between the shock wave angle and the airflow deflection angle.

[0062] First, substitute the incoming flow Mach number into the equation to obtain the normal Mach number behind the shock wave. Then substitute the normal Mach number , static pressure , density , velocity , and temperature into the shock wave relationship respectively to obtain the normal Mach number , static pressure , density , velocity , and temperature behind the wave. Then, use the equation to find the second shock wave angle . Substitute into the equation to obtain the airflow deflection angle . Finally, use the flow parameters after the first shock wave as the pre-wave flow parameters before the second shock wave and repeat the above calculation steps to derive the flow parameters , static pressure , density , velocity , and temperature after the second shock wave.

[0063] From the equation: calculate the internal compression angle . Then substitute , into the equation to obtain the internal compression shock wave angle .

[0064] After calculating the shock wave angle and the airflow deflection angle, the wall position can be obtained according to the following formula, as shown in Figure 4 . ;

[0065] ; ; 。

[0066] 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 relation. For a traditional two-dimensional inlet, the width of the inlet is calculated based on the captured flow rate, 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 bump-shaped. The following presents the design process.

[0067] First, given the serration angle θ and the half-model width of the inlet , where the half-model width of the inlet is half of the inlet width. The half-model 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 them using 3D modeling software. The formula for the variation law of the first shock wave angle is as follows, where the subscript k represents the k-th cross-section. Connect , , …, , … to obtain the curve. Connect , , …, , … to obtain the curve. Connect , , …, , … to obtain the curve. Connect , , …, , … to obtain the curve. Then connect the curves , , , , , as shown in Figure 5 .

[0068] Formulas (18)-(20) are as follows: ; ; .

[0069] 2. Design of serrated lip / bump integrated inlet

[0070] Subsequently, the surface construction is completed in 3D modeling software, and an inlet with aerodynamic and stealth integration can be obtained, asFigure 7 as shown. The specific steps are as Figure 8 shown. In the first step, the inner surface is generated. Specifically, on the basis of the serrated lip / pre-compression bulge surface, starting from point and extending backward to obtain line segment ; starting from point and extending horizontally by W / 2 to the symmetry plane position to obtain line segment ; respectively starting from , and extending backward to the position of point section to obtain line segments , ; connecting line segments , ; starting from and extending forward to obtain line segment , connecting points , to obtain line segment ; taking a point on line segment such that the length of line segment is equal to , connecting points , to obtain line segment ; using the lofting and filling tools in the modeling software, the smooth inner surface of the intake duct , , , can be obtained. Then the outer surface is generated. Specifically, a straight line segment is created at the position of point E. 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 . On line segment , points B and C with the same x coordinate as points , , are intercepted to obtain line segment and line segment ; the planes , are thickened into a curved surface , and then beveled at the lip , 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, through lofting and filling, planes , , and the curved surface Thus, the design of the sawtooth lip / bump integrated inlet is completed.

[0071] In this embodiment, the displacement of the boundary layer on the pre-compression surface is realized, and the thickness of the inlet boundary layer is effectively controlled; the forward radar cross-section area of the inlet is effectively reduced by the shielding of the bump on the inlet cavity.

[0072] 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 inlet capture height is 400 mm, the inlet width is 600 mm, the sawtooth angle is 120 degrees, the two shock waves M = 2, 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 vertical plane of the fuselage edge profile = 2°, the included angle in the horizontal plane = 5°, as Figure 9 , Figure 10 shown, compared with the conventional inlet, the logarithmic mean value of the vertical polarization RCS in the 30° angular domain in front of the horizontal plane of the inlet designed by the design method of the present application is reduced by 2.1757 dBsm, and the reduction amplitude is 14.01%; the logarithmic mean value of the horizontal polarization RCS is reduced by 0.05 dBsm, and the reduction amplitude is 0.37%. In the range of 10° to 30° of the vertical plane depression angle, compared with the conventional inlet, the logarithmic mean value of the vertical polarization RCS is reduced by 2.50 dBsm, and the reduction amplitude is 94.01%; the logarithmic mean value of the horizontal polarization RCS is reduced by 2.44 dBsm, and the reduction amplitude is 47.69%.

[0073] Compared with the conventional inlet, the inlet designed by the present application can effectively realize the displacement of the low-energy boundary layer. By intercepting the compression surface 0~3Ma at the symmetry planes z = 0 mm, z = 100 mm, z = 200 mm, and z = 300 mm and statistically analyzing the average thickness of the 3rd wedge angle, it is found that the thickness of the low-energy boundary layer of the inlet designed by the present application is reduced by 0.8 mm compared with the conventional inlet. Figure 11 It is the Mach number cloud map at the symmetry plane.

[0074] As Figure 12 , Figure 13As shown in the figure, at 6 Ma, the average total pressure recovery coefficient at the x = 3800 mm cross-section of the conventional inlet is 64%, and the flow coefficient is 82%; for the integrated sawtooth lip and bump inlet, the average total pressure recovery coefficient at the x = 3800 mm cross-section is 53%, and the flow coefficient is 76%. Although the aerodynamic performance of the inlet designed by the method of the present application has decreased, its stealth performance has been significantly improved. Generally speaking, it sacrifices some aerodynamic performance in exchange for a significant improvement in stealth performance, achieving the effect of enhancing the stealth performance of the aircraft.

[0075] The embodiment of the present application also provides a design system for an aerodynamic and stealth integrated inlet, and the system includes:

[0076] A sawtooth lip and bump compression surface design module, which is used to determine the incoming flow Mach number, capture height, and inlet width 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 wave relationship, 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 sawtooth angle θ, the half-mode width of the inlet 、the number of points N for discretizing the inlet width, 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, and the half-mode width of the inlet is half of the inlet width W; and

[0077] An integrated sawtooth lip and bump inlet design module, which is used to call a three-dimensional modeling software to connect and generate an integrated inlet profile of the bump and sawtooth lip based on the position coordinates, inner surface parameters, and outer surface parameters.

[0078] 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 of the above possible implementation manners.

[0079] The above is the preferred implementation manner of the present application. It should be noted that for those of ordinary skill in the art in this technical field, 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 pneumatic stealth integrated inlet design method, characterized in that, 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: Determine the incoming flow Mach number according to the designed Mach number and designed altitude of the aircraft , capture altitude h, and inlet width W; Determining 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, the Oswatitsch optimal wave system theory, and the oblique shock wave relation, iteratively solving the parameters of the Mth shock wave, where M is a positive integer greater than 1; Using the shock wave parameters and the capture height to determine the two-dimensional position of the inlet profile; and Based on the given sawtooth angle θ, the half-mode width of the inlet duct , the number of points N at which the half-mode of the inlet duct is discretized in the width direction, and the variation law of the first shock wave angle relative to the sawtooth angle θ, the position coordinates of N inlet duct cross-sections are obtained, and the half-mode width of the inlet duct is half of the width W of the inlet duct The serrated lip and bump integrated inlet design part includes: Based on the position coordinates, the inner surface parameters, and the outer surface parameters, calling a three-dimensional modeling software to connect and generate the integrated inlet profile of the bump and the serrated lip. Using the shock wave parameters and the capture height to determine the two-dimensional position of the inlet profile, specifically including: Using the capture height h and the first shock wave angle , determine the starting position and the ending position ; and Based on the starting position, the ending position, the shock wave angle ε, and the airflow deflection angle δ, determining the positions of each point between the starting point and the ending point of the inlet profile; The half model of the inlet is discretized into N points in the width direction, and the variation law is determined by the following parameters: the first shock wave angle of the N-1th section, the distance change value of the ending position relative to the starting position, and the projected distance of the spacing between the ending positions of adjacent sections along the axial direction; The projected distance is jointly determined by the width of the inlet half model, the number of discrete points, and the serration angle; Based on the position coordinates, the inner surface parameters, and the outer surface parameters, calling a three-dimensional modeling software to connect and generate the integrated inlet profile of the bump and the serrated lip, specifically: Based on the position coordinates and the preset length parameter of the extended line segment of the inner surface, performing line segment extension and connection processing on the points constituting the inner surface. On the basis of the line segment extension and connection processing, performing 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 outer surface curve thickening parameter and surface included angle parameter, constructing the line segments, surfaces, and planes of the outer surface, and performing lofting filling to obtain the outer surface of the integrated inlet profile.

2. The method according to claim 1, characterized in that, According to the design input parameters of the inlet, determining the capture height and the inlet width, specifically including: According to the design Mach number and the design height of the aircraft, obtaining the incoming flow Mach number, the incoming flow density, and the local sound speed of the inlet; According to the incoming flow Mach number and the local sound speed, determining the incoming flow velocity; According to the thrust / drag matching requirement of the aircraft, determining the capture flow rate of the inlet; Determining the capture area of the inlet through the capture flow rate, the incoming flow velocity, and the incoming flow density; and According to the capture area of the inlet, determining the capture height and the inlet width.

3. The method according to claim 1, wherein A first relationship is established among the shock wave angle, the wavefront Mach number, and the flow 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 flow 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.

4. A pneumatic and stealth integrated inlet design system, characterized in that, The system includes: A serrated lip and bump compression surface design module, which is used to determine the incoming flow Mach number according to the design Mach number and design altitude of the aircraft 、the capture altitude h and the inlet width W; 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, the Oswatitsch 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; use the shock wave parameters and the capture altitude to determine the two-dimensional position of the inlet profile; based on the given serration angle θ, the half-mode width of the inlet 、the number of points N at which the half-mode of the inlet is discretized along the width direction, 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, and the half-mode width of the inlet is half of the inlet width W; and A design module for an integrated inlet with a serrated lip and bump, which is used to call 3D modeling software to connect and generate the profile of the integrated inlet with a bump and a serrated lip based on the position coordinates, the inner surface parameters, and the outer surface parameters, and determine the two-dimensional position of the inlet profile by using the shock wave parameters and the capture height. Specifically, it includes: Using the capture height h and the first shock wave angle , 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 ε, and the flow deflection angle δ, determine the positions of points between the starting point and the ending point of the inlet profile. The half-model of the inlet is discretized into N points in the width direction. The variation law is determined by the following parameters: the first shock wave angle of 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. The projected distance is jointly determined by the width of the inlet half-model, the number of discrete points, and the serration angle. Based on the position coordinates, the inner surface parameters, and the outer surface parameters, call 3D modeling software to connect and generate the profile of the integrated inlet with a bump and a serrated 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 the line segments, surfaces, and planes of the outer surface, and perform lofting filling to obtain the outer surface of the integrated inlet profile.

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