A method for designing a double-back-swept ogive flying craft with controllable back-sweeping angle

Abstract

The application provides a double-back-sweep ogive full-riding wave aircraft design method with controllable back-sweep angle, comprising the following steps: designing a reference flow field central body generatrix, solving the position coordinates and flow parameters of each point in the flow field area and shock wave profile by applying the characteristic line theory; designing the horizontal plane projection profile of the aircraft leading edge and air inlet lip, defining the size of the back-sweep angle and the aircraft size; designing the aircraft shock wave outlet profile, determining the angle between the shock wave shape on each ogive surface and the vertical plane and the flow field parameters; solving the shape and position coordinates of the three-dimensional leading edge profile and air inlet lip profile of the aircraft, and obtaining the position coordinates of the intersection points of each ogive surface and the leading edge profile and air inlet lip profile of the aircraft; solving the reference flow field in each ogive surface, and obtaining the flow lines on each component profile of the aircraft in the ogive surface by using flow line tracking; and smoothly connecting the flow lines to obtain the basic configuration of the aircraft. The application guarantees the performance of the aircraft on the basis of improving the design freedom.

Description

Technical Field

[0001] This invention relates to the field of hypersonic vehicle technology, and in particular to a design method for a double-sweep kiss cone full waverider vehicle with controllable sweep angle. Background Technology

[0002] Air-breathing hypersonic vehicles, as launch vehicles capable of efficient flight in near-space, have attracted considerable attention since their concept was first proposed. Currently, these vehicles are being developed towards higher speeds, greater maneuverability, and wider speed ranges.

[0003] The key to the efficient, stable, and wide-range flight of air-breathing hypersonic vehicles lies in their superior aerodynamic layout and the continuous thrust provided by their engines. Waveriders, with their advantages of high lift, high lift-to-drag ratio, and uniform flow on the lower surface, are highly valuable in hypersonic vehicle design and are currently a hot research topic in the field of hypersonic aerodynamic layout both domestically and internationally. Furthermore, waveriders are theoretical configurations based on characteristic line theory and streamline tracing methods, which are also applicable to guiding the design of three-dimensional internally contracting inlet ducts. Therefore, when designing the aerodynamic shape of an aircraft, the waverider and the three-dimensional internally contracting inlet duct are usually designed as an integrated unit.

[0004] Currently, the integrated design of waveriders and three-dimensional internal inlet ducts can be divided into two categories. One category uses the waverider solely as the pre-compression surface of the inlet's forebody, leveraging its strong compressibility to enhance the aircraft's capture performance; this is known as "waverider forebody / inlet integrated design." The second category not only uses a waverider forebody but also designs the entire aircraft configuration based on waverider theory, improving both the aircraft's capture compression performance and overall performance; this is known as "full waverider integrated design." The full waverider integrated design method first constructs a reference flow field based on the method of characteristics, encompassing both the external flow of the aircraft and the internal flow of the inlet. Then, streamline tracing is performed on this reference flow field to generate components such as the forebody, fuselage, wings, and inlet. This method achieves integrated design while ensuring that the entire frontal surface of the aircraft exhibits waverider characteristics, ultimately improving the overall lift-to-drag ratio of the aircraft. This is currently the commonly used design method in integrated design.

[0005] However, for the integrated design of air-breathing hypersonic vehicles, it is not only necessary to focus on the overall aerodynamic performance of the vehicle and the compressibility of the inlet, but also to conduct in-depth research on the complex flow characteristics caused by the coupling of the external and internal flow fields, such as the uniformity of the airflow at the inlet, the flow field structure, and the flow capture capability. Furthermore, to better perform wide-range flight missions, the vehicle needs to simultaneously consider high- and low-speed aerodynamic performance. Currently, a common approach is to use a double-sweep method on the leading edge of the vehicle, thereby utilizing the vortex structure on the upper surface at low speeds to improve the vehicle's aerodynamic performance. However, the lip shock profile in the general conical-guided full waverider design method is an arc, resulting in significant lateral flow phenomena. Moreover, the horizontal projection profile of the waverider's leading edge is an indirectly controllable curve, making it impossible to achieve controllable sweep angle design, which to some extent reduces the design freedom of the vehicle. Summary of the Invention

[0006] The purpose of this invention is to address the shortcomings of the aforementioned background technology by providing an improved design method. This method reduces the variation amplitude of the pressure flow field between different cut surfaces by changing the curvature distribution of the shock wave profile at the bottom of the aircraft, improves the lateral flow phenomenon of airflow at the inlet of the air intake, and enables controllable leading edge sweep angle, thereby increasing design freedom.

[0007] To achieve the above objectives, the present invention provides a design method for a double-sweep, kiss-shaped, full-waverider aircraft with controllable sweep angle, comprising the following steps:

[0008] S1, design the generatrix of the central body of the reference flow field, and use the characteristic line theory to solve the position coordinates and flow parameters of each point in the flow field region and shock wave profile.

[0009] S2, design the horizontal projection profiles of the aircraft's leading edge and air intake lip, and define the sweep angle and aircraft dimensions;

[0010] S3, design the shock wave exit profile of the aircraft, determine the shock wave shape and flow field parameters on each cross-section, as well as the angle between the shock wave and the vertical plane;

[0011] S4, solve for the shape and position coordinates of the three-dimensional leading edge profile of the aircraft and the lip profile of the air intake, and obtain the position coordinates of the intersection points of each cross-section with the leading edge profile of the aircraft and the lip profile of the air intake;

[0012] S5. The reference flow field in each tangent surface is solved by the method of rotational characteristic lines, and then the streamlines on the corresponding aircraft component surfaces in the tangent surface are obtained by using streamline tracing.

[0013] S6, obtain the corresponding streamlines within each cut surface, smoothly connect the streamlines to obtain the basic configuration of the aircraft.

[0014] Furthermore, a central body generatrix is ​​set in S1. Under supersonic incoming flow conditions, a shock wave profile is generated based on the central body to form a backflow reference field. Given the supersonic incoming flow conditions, the position coordinates and flow parameters at any location in the reference flow field are solved using the theory of swirling characteristic lines. The position coordinates are the axial and radial coordinates in the cylindrical coordinate system. The flow parameters include local static pressure, local density, local velocity, and local flow direction angle.

[0015] Furthermore, the projected profile in S2 consists of an arc and two straight line segments. The arc serves as the nose of the aircraft. The angle between the first straight line segment and the front and rear axes defines the magnitude of the first-stage sweep angle λ1. The angle between the second straight line segment and the front and rear axes defines the magnitude of the second-stage sweep angle λ2. At the same time, it determines the distance from the nose of the aircraft to the air intake lip in the front and rear axis directions, the length of the aircraft fuselage, and the wingspan of the aircraft.

[0016] Furthermore, in S3, the shock wave exit profile consists of a quartic curve and a third straight line segment. The third straight line segment is on the left and right axes. The length of the third straight line segment is set, the equation of the quartic curve is determined, and the curvature radius and curvature center coordinates corresponding to each discrete point on the shock wave exit profile are obtained. The shock wave shape and flow field parameters in the tangent plane corresponding to each discrete point are obtained by scaling the reference flow field in S1. The reference flow fields on each tangent plane are smoothly connected along the shock wave exit profile to obtain the three-dimensional tangent cone shock wave flow field parameters required for aircraft design.

[0017] Furthermore, in S4, the three-dimensional leading edge profile of the aircraft, the inlet lip profile, and the projected profile correspond. Based on the projected profile, the coordinates of each point on the three-dimensional leading edge profile of the aircraft and the inlet lip profile on the front-rear axis and left-right axis are obtained. Based on the flow field parameters and position coordinates in the cut surface obtained in S3, as well as the angle between the cut surface and the vertical plane, the three-dimensional coordinates of the intersection point of the cut surface with the leading edge profile of the aircraft and the inlet lip profile are obtained. The three-dimensional coordinates of the points on the leading edge profile of the aircraft and the inlet lip profile corresponding to each discrete point on the horizontal projected profile of the aircraft are determined.

[0018] Furthermore, in S5, the position coordinates of discrete points on the corresponding leading edge profile, inlet lip profile, and shock wave profile within the cut surface are determined in the cylindrical coordinate system of the flow field. Based on the position coordinates, the shock wave radius of the axisymmetric flow field on the cut surface is determined. The reference flow field on the cut surface is obtained by scaling the reference flow field in S1 proportionally. Using the discrete points of the leading edge profile and inlet lip profile as starting points, streamline tracing is used to solve for the streamlines passing through the points. The region enclosed by the streamlines of the discrete points of the leading edge profile is the flow field in the aircraft design process. The streamlines of the discrete points of the inlet lip profile divide the flow field into two parts: one part is the external flow field region used for solving the streamlines on the lower surface of the aircraft; the other part is the internal flow field region used for solving the streamlines on the forebody wall, upper inlet wall, and lower inlet wall.

[0019] The above-described solution of the present invention has the following beneficial effects:

[0020] The design method for a double-swept leading edge full-wave-riding aircraft with controllable sweep angle provided by this invention can directly and precisely control the sweep angle distribution of the double-swept leading edge of the aircraft according to requirements, thereby improving the design freedom. At the same time, the designed shock wave exit profile is no longer limited to a circular arc but is an arbitrary second-derivative continuous curve. By changing the curvature distribution on the shock wave exit profile, the variation amplitude of the pressure flow field between different slit surfaces can be reduced, thereby improving the lateral flow phenomenon of the airflow at the three-dimensional inlet lip, providing a stable high-energy gas source for the combustion chamber, and ultimately improving the performance of the designed aircraft.

[0021] Other beneficial effects of the present invention will be described in detail in the following detailed description section. Attached Figure Description

[0022] Figure 1 This is a flowchart of the steps of the present invention;

[0023] Figure 2 This is a schematic diagram of the coordinate system definition rules in an embodiment of the present invention;

[0024] Figure 3 This is a schematic diagram of the reference flow field model in an embodiment of the present invention;

[0025] Figure 4 This is a schematic diagram of the horizontal projection profile of the aircraft in an embodiment of the present invention;

[0026] Figure 5 This is a schematic diagram of the shock wave exit profile in an embodiment of the present invention;

[0027] Figure 6 This is a schematic diagram showing the positional correspondence between the leading edge profile and the horizontal projection profile of the aircraft in an embodiment of the present invention;

[0028] Figure 7This is a schematic diagram illustrating the method for determining the intersection point of a certain cross-section and the leading edge of an aircraft in an embodiment of the present invention;

[0029] Figure 8 This is a schematic diagram showing the correspondence between discrete points of the leading edge profile and the horizontal projection profile of the aircraft in an embodiment of the present invention;

[0030] Figure 9 This is a schematic diagram of the internal and external flow field division on the cut surface in an embodiment of the present invention;

[0031] Figure 10 This is a schematic diagram illustrating the solution of the position of the lip-reflected shock wave and the flow parameters in an embodiment of the present invention;

[0032] Figure 11 This is a schematic diagram of the flow field solution in the shock wave dependent region after the reflected shock wave in an embodiment of the present invention;

[0033] Figure 12 This is a schematic diagram of the flow field solution in the stable region inside the air intake in an embodiment of the present invention;

[0034] Figure 13 This is a schematic diagram of the streamlines of the various components that make up the aircraft in an embodiment of the present invention;

[0035] Figure 14 This is an isometric side view of the aircraft in an embodiment of the present invention. Detailed Implementation

[0036] The following specific examples illustrate the implementation of this disclosure. Those skilled in the art can easily understand other advantages and effects of this disclosure from the content disclosed in this specification. Obviously, the described embodiments are only a part of the embodiments of this disclosure, and not all of them. This disclosure can also be implemented or applied through other different specific embodiments, and the details in this specification can also be modified or changed based on different viewpoints and applications without departing from the spirit of this disclosure. It should be noted that, in the absence of conflict, the following embodiments and features in the embodiments can be combined with each other. Based on the embodiments in this disclosure, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of this disclosure.

[0037] It should be noted that various aspects of embodiments within the scope of the appended claims are described below. It will be apparent that the aspects described herein can be embodied in a wide variety of forms, and any particular structure and / or function described herein is merely illustrative. Based on this disclosure, those skilled in the art will understand that one aspect described herein can be implemented independently of any other aspect, and two or more of these aspects can be combined in various ways. For example, any number of aspects set forth herein can be used to implement the device and / or practice the method. Additionally, this device and / or method can be implemented using structures and / or functionalities other than one or more of the aspects set forth herein.

[0038] It should also be noted that the illustrations provided in the following embodiments are merely schematic representations of the basic concept of this disclosure. The illustrations only show components relevant to this disclosure and are not drawn according to the actual number, shape, and size of components in implementation. In actual implementation, the type, quantity, and proportion of each component can be arbitrarily changed, and the component layout may be more complex. Furthermore, specific details are provided in the following description to facilitate a thorough understanding of the examples. However, those skilled in the art will understand that the described aspects can be practiced without these specific details.

[0039] like Figure 1 As shown, embodiments of the present invention provide a design method for a double-sweep, kiss-shaped, full-waverider aircraft with controllable sweep angle, the corresponding coordinate system being as follows: Figure 2 As shown, the Ox axis coincides with the longitudinal axis of the aircraft, and the direction pointing to the bottom of the aircraft is positive; the Oy axis is in the longitudinal plane of the aircraft and is perpendicular to the Ox axis; the Oz axis is perpendicular to xOy, and its direction is determined by the right-hand rule, with the bottom surface of the aircraft on the yOz plane.

[0040] The design methodology specifically includes:

[0041] S1, design the generatrix of the central body of the reference flow field, and use the characteristic line theory to solve for the position coordinates and flow parameters of each point in the flow field region and shock wave profile.

[0042] like Figure 3As shown, curve 1-2 is the generatrix of the central body, curve 1-3 is the shock wave profile generated by the central body under supersonic inflow conditions, and region 1-2-3 is the reference flow field after the wave. The reference flow field is located in a cylindrical coordinate system, with the x-axis coinciding with the axis of symmetry of the central body and pointing from the front of the aircraft to the rear. The r-axis is perpendicular to the x-axis and points downwards. The intersection of the cutoff line of the reference flow field and the x-axis is the origin O. Given the supersonic inflow conditions, the position coordinates and flow parameters at any location in the reference flow field can be solved using the theory of swirling characteristic lines (this theory is common knowledge in the field; see "Gas Dynamics," MJ Zöcklough, JD Hoffmann, National Defense Industry Press, 1984, pp. 138-195). The position coordinates are the axial and radial coordinates in the cylindrical coordinate system, and the flow parameters include local static pressure, local density, local velocity, and local flow direction angle.

[0043] S2 designs the horizontal projection profiles of the aircraft's leading edge and air intake lip, and defines the sweep angle and aircraft dimensions.

[0044] like Figure 4 As shown, line 4-5-6-7-8-O is the projected profile of the aircraft on the horizontal plane (xOz), and line segment 6-9 is the projected profile of the air intake lip on the horizontal plane. Due to the high surface symmetry of the hypersonic aircraft's geometry, only the design parameters of the half-model need to be defined to obtain the aircraft's geometric shape. The projected profile consists of an arc and two straight line segments. Arc 4-5 serves as the blunted nose of the aircraft, preventing severe aerodynamic heating of the sharp nose by shock waves under hypersonic conditions and protecting the aircraft's fuselage structure. The angle between straight line segment 5-6-7 and the z-axis defines the magnitude of the first-order sweep angle λ1; the angle between straight line segment 7-8 and the z-axis defines the magnitude of the second-order sweep angle λ2. The distance L1 between line segments 4-9 is the distance from the aircraft's nose to the air intake lip in the x-axis direction; the distance L between line segments 4-O is the aircraft's fuselage length; and the distance between line segments O-8 is W / 2, which is half the aircraft's wingspan. Ultimately, the (x,z) coordinates of the three-dimensional leading edge of the aircraft and the inlet lip profile in this coordinate system can be completely determined.

[0045] S3. Design the shock wave exit profile of the aircraft, determine the shock wave shape and flow field parameters on each cross-section, and the angle between the shock wave and the plane xOy.

[0046] like Figure 5As shown, profile 10-11-12 represents the shock wave exit profile of the aircraft. The hollow dots in the diagram represent a series of discrete points on the shock wave exit profile, the dashed lines represent the radii of curvature corresponding to these discrete points, and the solid origin represents the centers of curvature corresponding to these discrete points. In this embodiment, the shock wave exit profile consists of a quartic curve (10-11) and a straight line segment (11-12). The straight line segment lies on the z-axis. Given the length of the straight line segment 11-12, the coordinates of point 11 can be determined. Therefore, the equation of the quartic curve 10-11 can also be uniquely determined (the general equation of a quartic curve is y = a*(zb)). 4 The straight line segment lies on the z-axis. To ensure the continuity of the flow field, the second derivative of the shock wave exit profile must be continuous, meaning the derivative of the quartic curve at point 11 must be zero. Therefore, given the length of the straight line segment 11-12, the coordinates of point 11 are determined, and the equation of the quartic curve 10-11 is uniquely determined. Based on the equation of the quartic curve 10-11, the magnitude of the radius of curvature and the coordinates of the center of curvature at each discrete point can be calculated.

[0047] Therefore, the shock wave shape and flow field parameters within the cut surface corresponding to each discrete point can be obtained by scaling the reference flow field in S1. By smoothly connecting the reference flow fields on each cut surface along the shock wave exit profile, the three-dimensional cut-cone shock wave flow field parameters required for aircraft design can be obtained. It should be noted that to ensure the continuity of the three-dimensional cut-cone shock wave flow field, the second derivative of the shock wave exit profile 10-11-12 must be continuous. The straight segment 11-12 can be considered as a circular arc with an infinite radius of curvature, and its corresponding shock wave flow field parameters can be solved using the two-dimensional oblique shock wave relation.

[0048] S4, solve for the shape and position coordinates of the three-dimensional leading edge profile and the inlet lip profile of the aircraft.

[0049] Figure 6 This demonstrates the positional correspondence between the three-dimensional configuration of the aircraft and its projected profiles on the horizontal plane. Curves 13-14-15-16-10 represent the three-dimensional leading edge profile of the aircraft, corresponding to projected profiles 4-5-6-7-8; curve 15-17 represents the inlet lip profile, corresponding to projected profiles 6-9. Based on the projected profile equations designed in S2, the coordinates of each point on the three-dimensional leading edge and inlet lip profiles in the x and z directions can be obtained. By solving for the coordinates of each point in the y direction, the design input curves of the aircraft can be uniquely determined.

[0050] like Figure 7 As shown, in this embodiment, taking the tangent plane 18-22-20 corresponding to a discrete point 22 on the shock wave exit profile as an example, the flow parameters and position coordinates of the flow field 18-22-21 within the tangent plane, as well as the angle between the tangent plane and the plane xOy, are obtained through S3. The intersection point of the cut surface and the leading edge of the aircraft is 24, which corresponds to point 23 on the projected profile (the intersection of the projection curve of shock wave profile 18-22 on the horizontal plane and the projected profile of the aircraft's leading edge). By combining the equations of the two intersection lines on the horizontal plane, the coordinates of point 23 in the x and z directions can be obtained, which are also the coordinates of point 24 in the x and z directions. Since the coordinates of all points in the flow field within the cut surface are known, the coordinates of point 24 in the flow field coordinate system can be solved based on its x-coordinate. 24 r 24 The coordinates of point 24 along the y-axis can be represented as follows: That is, obtain the three-dimensional coordinates (x, y) of point 24. 24 y 24 , z 24 ).

[0051] Repeat this step to obtain the position coordinates of the intersection points of each discretized cut surface in S3 with the leading edge profile of the aircraft and the inlet lip profile. For example... Figure 8 As shown, the three-dimensional coordinates of the points (solid dots) on the leading edge profile and the inlet lip profile corresponding to each discrete point (hollow dot) on the horizontal projection profile of the aircraft are completely determined.

[0052] S5 takes the three-dimensional coordinates of each solid circle obtained in S4 as input, solves the reference flow field in each tangent surface according to the method of rotational characteristic lines, and then uses streamline tracing (streamline tracing refers to the process of solving the streamline passing through a certain point in the flow field, see "Fundamentals of Aerodynamics", JD Anderson, Aviation Industry Press, 2006) to obtain the streamlines on the surface of each component of the aircraft in the tangent surface.

[0053] This embodiment takes the streamline solution process on a certain cutting plane as an example, such as... Figure 9 As shown. First, determine the position coordinates of discrete points on the corresponding leading edge profile, inlet lip profile, and shock wave profile within the cut surface in the cylindrical coordinate system of the flow field (points 26, 28, and 30, respectively). Based on the position coordinate of point 30, the shock wave radius R of the axisymmetric flow field on the cut surface can be determined. n Then the reference flow field O-25-30 on the cut surface can be expressed by the proportional relationship R. n / R0 scales the baseline flow field in S1 proportionally. Then, using points 26 and 28 as starting points, the streamline tracing method is used to solve for streamlines 26-27 passing through point 26 and streamlines 28-29 passing through point 28. At this point, the region enclosed by 26-27-30 is the flow field for the subsequent aircraft design process. Streamline 28-29 divides it into two parts: 28-29-30 is the external flow field region, used for solving the streamlines on the lower surface of the aircraft; 26-27-28-29 is the internal flow field region, used for solving the streamlines on the forebody wall, the upper wall of the air intake, and the lower wall of the air intake.

[0054] like Figure 10 As shown, in the flow field of region 26-27-28-29, the internal flow field region is reconstructed using the method of characteristics. First, the leading edge appendage shock wave O-30 intersects with the inlet lip point 28, generating a reflected shock wave 28-31. The reflected shock wave intersects with the characteristic line grid, discretizing a series of intersection points. Since the front flow parameters of the reflected shock wave 28-31 are known, only the back flow angle as shown in equation (1) needs to be given, then the local shock wave angle can be uniquely determined by the oblique shock wave relation (2).

[0055] θ 28-31,2 =θ 28-31,2 (x), x∈[x 28 ,x 31 (1)

[0056]

[0057] Where, θ 28-31,1 θ 28-31,2 M represents the flow direction angles before and after the reflected shock wave, respectively. 28-31,1 Let β be the local Mach number of the reflected shock wave front, and β be the local shock wave angle. Taking point 28 as the starting point of the reflected shock wave at the lip, repeat the above steps to solve for the position and shape of the reflected shock wave 28-31 until it intersects with streamline 26-27 at shoulder point 31. Finally, the region enclosed by the leading edge attached shock wave 26-28, streamline 26-31, and reflected shock wave 28-31 is defined as the isentropic compression region of the forebody shock wave in the internal flow field.

[0058] Solve for the flow field inside the air intake. This region is divided into two parts: the region affected by the reflected shock wave and the lower wall of the air intake, and the region affected by the upper and lower walls of the air intake. For example... Figure 11 As shown, after passing through the upstream point n of the reflected shock wave i,j The streamline and the point n downstream i+1,j The rightward-moving Mach lines intersect at a point n. i+1,j+1 The position coordinates and flow parameters of this point can be solved using the theory of swirling characteristic lines. Therefore, using the position and flow parameters of the unit point on the reflected shock wave 28-31 as initial conditions, the position and shape of streamline 28-31 can be determined using the method of swirling characteristic lines. The region enclosed by the reflected shock wave 28-31, the streamline 28-32 passing through point 28, and the right-handed Mach line 31-32 passing through point 31 is considered as the lip-reflected shock wave dependent region in the internal flow field.

[0059] Figure 12The process of solving the stable flow field region inside the inlet is demonstrated. Given the position and flow parameters of the rightward Mach line 31-32 passing through point 31, if the central body wall curve 31-33 passing through point 31 and the flow parameters on this curve are given, the parameters of the stable flow field region can be solved according to the method of swirling characteristic lines. According to the compatibility equation (3) along the streamline, for the flow parameters v, p, and ρ at each point on the same streamline, the values ​​of the other two parameters can be obtained by giving any one parameter. In this embodiment, the tilt angle and Mach number distribution on the central body wall 31-33 are given, as shown in equations (4) and (5). Among them, the tilt angle distribution determines the position and shape of the curve 31-33, and the Mach number distribution determines the flow parameters on the curve 31-33.

[0060]

[0061] θ 31-33 =θ 31-33 (x), x∈[x 31 ,x 33 (4)

[0062] M 31-33 =M 31-33 (x), x∈[x 31 ,x 33 (5)

[0063] It should be noted that the wall inclination angle at shoulder point 31 should coincide with the local flow direction angle behind the reflected shock wave. This prevents further reflection of the reflected shock wave after it reaches the shoulder point, thus achieving the purpose of wave dissipation inside the intake duct. Then, the distance x between the intake duct outlet section and shoulder point 31 in the x-axis direction is given. 31-33 This determines the length of the air intake along the x-axis.

[0064] Streamlines 26-31 are used as streamlines on the forebody wall, curves 31-33 on the central body wall are used as streamlines on the upper wall of the air intake, streamlines 28-32-34 are used as streamlines on the lower wall of the air intake, and streamlines 28-29 are used as streamlines on the lower surface of the aircraft. It should be noted that when the cut surface does not pass through the internal air intake of the aircraft (i.e., it does not intersect with the air intake lip line), the intersection of the cut surface and the leading edge of the wing is used as the starting point for streamline tracing in the external reference flow field to obtain the streamlines on the lower surface of the wing.

[0065] S6. Repeat the above steps in other matching cross-sections to obtain the corresponding streamlines in each plane, such as... Figure 13 As shown, these streamlines are smoothly connected to obtain the basic configuration of the aircraft.

[0066] Ultimately, the designed aircraft has a fuselage length of L, a wingspan of W, a double sweep at the leading edge, a first-stage sweep angle of λ1, and a second-stage sweep angle of λ2, which are the same as the design parameters given in S2, thus achieving controllable double sweep. Figure 14 The image displays a 3D view of the aircraft. In the figure, 35 represents the upper surface of the fuselage, 36 the forebody compression surface, 37 the upper surface of the air intake, 38 the upper surface of the wing, 39 the lower surface of the air intake, 40 the lower surface of the wing, 41 the bottom surface of the aircraft, 42 the lower surface of the fuselage, and 43 the kissing shock wave profile. The aircraft's external surface and internal flow channel surface are highly integrated and exhibit full-wave characteristics.

[0067] The technical features of the above embodiments can be combined in any way. For the sake of brevity, not all possible combinations of the technical features in the above embodiments are described. However, as long as there is no contradiction in the combination of these technical features, they should be considered to be within the scope of this specification.

[0068] The above embodiments are merely illustrative of several implementation methods of this application, and their descriptions are relatively specific and detailed, but they should not be construed as limiting the scope of the application. It should be noted that those skilled in the art can make various modifications and improvements without departing from the concept of this application, and these all fall within the protection scope of this application. Therefore, the protection scope of this application should be determined by the appended claims.

Claims

1. A design method for a double-sweep, kiss-shaped, full-waverider aircraft with controllable sweep angle, characterized in that, Includes the following steps: S1, design the generatrix of the central body of the reference flow field, and use the characteristic line theory to solve the position coordinates and flow parameters of each point in the flow field region and shock wave profile. S2, design the horizontal projection profiles of the aircraft's leading edge and air intake lip, and define the sweep angle and aircraft dimensions; S3, design the shock wave exit profile of the aircraft, determine the shock wave shape and flow field parameters on each cross-section, as well as the angle between the shock wave and the vertical plane; S4, solve for the shape and position coordinates of the three-dimensional leading edge profile of the aircraft and the lip profile of the air intake, and obtain the position coordinates of the intersection points of each cross-section with the leading edge profile of the aircraft and the lip profile of the air intake; S5. The reference flow field in each tangent surface is solved by the method of rotational characteristic lines, and then the streamlines on the corresponding aircraft component surfaces in the tangent surface are obtained by using streamline tracing. In S5, the position coordinates of discrete points on the corresponding leading edge profile, inlet lip profile and shock wave profile in the cut surface are determined in the cylindrical coordinate system of the flow field. The shock wave radius of the axisymmetric flow field on the cut surface is determined based on the position coordinates. The reference flow field on the cut surface is obtained by scaling the reference flow field in S1 proportionally through the proportional relationship. Starting from the discrete points of the leading edge profile and the inlet lip profile, streamline tracing is used to solve for the streamlines passing through these points. The region enclosed by the streamlines of the discrete points of the leading edge profile is the flow field in the aircraft design process. The streamlines of the discrete points of the inlet lip profile divide the flow field into two parts: one part is the external flow field region used for solving the streamlines on the lower surface of the aircraft; the other part is the internal flow field region used for solving the streamlines on the forebody wall, the upper wall of the inlet, and the lower wall of the inlet. In the internal flow field region, the leading edge appendage shock wave intersects with the inlet lip point to generate a reflected shock wave. The reflected shock wave intersects with the streamline at the shoulder point. The region enclosed by the leading edge appendage shock wave, streamline, and reflected shock wave is the isentropic compression region of the leading edge shock wave in the internal flow field. The region enclosed by the reflected shock wave, the streamline passing through the lip point, and the rightward Mach line passing through the shoulder point is the lip reflected shock wave dependent region in the internal flow field. Given the central body wall curve and the flow parameters on the curve, the parameters of the stable flow field region are solved according to the swirling characteristic line method. S6, the corresponding streamline in each tangent plane is obtained, and the streamlines are smoothly connected to obtain the basic configuration of the aircraft.

2. The design method for a double-sweep, kiss-shaped, full-waverider aircraft with controllable sweep angle according to claim 1, characterized in that, In S1, a central body generatrix is ​​set. Under supersonic incoming flow conditions, a shock wave profile is generated based on the central body to form a backflow reference field. Given the supersonic incoming flow conditions, the position coordinates and flow parameters at any position in the reference flow field are solved using the theory of swirling characteristic lines. The position coordinates are the axial and radial coordinates in the cylindrical coordinate system. The flow parameters include local static pressure, local density, local velocity, and local flow direction angle.

3. The design method for a double-sweep, kiss-shaped, full-waverider aircraft with controllable sweep angle according to claim 2, characterized in that, In S2, the projected profile consists of an arc and two straight line segments. The arc represents the nose of the aircraft, and the angle between the first straight line segment and the front and rear axes defines the first-order sweep angle. The size; the angle between the second straight segment and the front and rear axes defines the second-order sweep angle. The size of the aircraft is determined, along with the distance from the nose of the aircraft to the air intake lip in the longitudinal axis direction, the length of the aircraft fuselage, and the wingspan of the aircraft.

4. The design method for a double-sweep, kiss-shaped, full-waverider aircraft with controllable sweep angle according to claim 3, characterized in that, In S3, the shock wave exit profile consists of a quartic curve and a third straight line segment. The third straight line segment is on the left and right axes. The length of the third straight line segment is set, the equation of the quartic curve is determined, and the curvature radius and curvature center coordinates corresponding to each discrete point on the shock wave exit profile are obtained. The shock wave shape and flow field parameters in the tangent plane corresponding to each discrete point are obtained by scaling the reference flow field in S1. The reference flow fields on each tangent plane are smoothly connected along the shock wave exit profile to obtain the three-dimensional tangent cone shock wave flow field parameters required for aircraft design.

5. The design method for a double-sweep, kiss-shaped, full-waverider aircraft with controllable sweep angle according to claim 4, characterized in that, In S4, the three-dimensional leading edge profile of the aircraft, the inlet lip profile, and the projected profile correspond. Based on the projected profile, the coordinates of each point on the three-dimensional leading edge profile of the aircraft and the inlet lip profile on the front-rear axis and left-right axis are obtained. Based on the flow field parameters and position coordinates in the cut surface obtained in S3, as well as the angle between the cut surface and the vertical plane, the three-dimensional coordinates of the intersection point of the cut surface with the leading edge profile of the aircraft and the inlet lip profile are obtained. The three-dimensional coordinates of the points on the leading edge profile of the aircraft and the inlet lip profile corresponding to each discrete point on the horizontal projected profile of the aircraft are determined.

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