A flight test platform parameterization modeling method based on CST method

Abstract

The application discloses a flight test platform parameterization modeling method based on a CST method, relates to the field of aerospace vehicle modeling design, and comprises the following steps: S1, function division is performed according to the shape structure of a flight test platform vehicle, and each part after the division is divided into multiple different cross-sectional shapes according to the shape structure; S2, corresponding control points are arranged on each cross-sectional shape; S3, adjacent circumferential control points on each cross-sectional shape are connected one by one to generate a transverse cross-sectional curve through a five-parameter moded line design method, then each curve control point is connected one by one to generate a longitudinal curve based on adjacent transverse cross sections, and finally a curved surface is obtained, and parameterization modeling of each part is completed; and S4, each part independently completing parameterization modeling is assembled, and final parameterization modeling of the vehicle is completed. The flight test platform parameterization modeling method based on the CST method provided by the application significantly improves comprehensive performance, reliability and design iteration efficiency.

Description

Technical Field

[0001] The present invention relates to the field of aerospace vehicle modeling and design, and more particularly to a parametric modeling method for a flight test platform based on a CST method. Background Art

[0002] With the increasing demand for efficient flight test platforms in the aerospace sector, next-generation flight test platforms, as crucial flight test vehicles, have become a key research and development focus in various countries. These aircraft must meet two core requirements when performing flight test missions: a large underside surface area to accommodate various flight test payloads; and an optimized backside streamlined transition design to enhance lift performance. However, these two requirements present a conflicting design requirement: a large underside surface affects aerodynamic characteristics, while an optimized backside streamlined design places higher demands on structural layout and aerodynamic performance. Traditional design methods rely on experience and experimentation, making it difficult to fully and accurately consider the relationships between parameters. This leads to long design cycles, high costs, and difficulty in achieving the optimal solution. To address these issues, parametric modeling technology has proven to be invaluable. Based on mathematical models and computer technology, this technology defines parameters to describe geometric shape, structural characteristics, and performance parameters, and establishes mathematical relationships and constraints between them. This allows for rapid generation of multiple design options, facilitating comprehensive analysis of the impact of each parameter on performance. During the conceptual design phase, multiple options can be quickly generated and evaluated to shorten the cycle; during the detailed design phase, complex geometric shapes can be accurately described and combined with CFD analysis to optimize aerodynamic performance.

[0003] However, the main problem with existing technologies is that current parametric modeling almost only involves the design of some components and fails to expand to the design of the overall aircraft structure. The main reason is the limitations of component modeling. It only models the head configuration or other components, and fails to establish the global parameter transfer of the aircraft. As a result, the aircraft lacks cross-component parameter correlation and global coordination, resulting in system performance losses such as uncontrolled aerodynamic interference, weak structural connection areas, and failure of multi-objective optimization in the subsequent calculation process, making it impossible to achieve the overall optimal design of the aircraft. Summary of the Invention

[0004] One objective of the present invention is to solve the problem of overall parametric modeling of aircraft by driving global parameters and automatically coordinating across components to achieve multi-dimensional systematic optimal design of aerodynamics, structure, and mission, thereby significantly improving performance, reliability, and iteration efficiency. To achieve these objectives, the present invention provides a parametric modeling method for a flight test platform based on the CST method, comprising:

[0005] S1. Divide the functions of the flight test platform aircraft according to its external structure, and further divide the divided parts into multiple different cross-sectional shapes according to the external structure;

[0006] S2. Set corresponding control points on each cross-sectional shape;

[0007] S3. Using the five-parameter model line design method, first connect the adjacent circumferential control points on each cross-sectional shape one by one to generate a transverse curve. Then, based on the adjacent transverse cross-sections, connect the control points of each curve one by one to generate a longitudinal curve. Finally, the curved surface is obtained, completing the parametric modeling of each part.

[0008] S4, assembling the independently completed parametric modeling parts to complete the final parametric modeling of the aircraft;

[0009] In S3, the five-parameter mold line design method includes: first defining the boundary contour of each cross-sectional shape by using the top view power function and the front view parameters, and then using the CST method to finely control the cross-sectional shape;

[0010] Wherein, the refined control of the shape of each cross section on the aircraft is achieved through dimensionless curves.

[0011] Preferably, in S1, the flight test platform aircraft is divided into a body and wings according to its external structure;

[0012] The projectile body is divided into four cross-sectional shapes: head cross-sectional, first cross-sectional, second cross-sectional, and third cross-sectional;

[0013] The wing includes a horizontal wing and / or a vertical wing. The horizontal wing is divided into two cross-sectional shapes: a wing root section and a wing end section. The vertical wing is divided into two cross-sectional shapes: a lower section and an upper section. The CST curve corresponding to the lower section of the vertical wing is consistent with the design value of the curve corresponding to the wing end section of the horizontal wing.

[0014] Preferably, in S2, the number of circumferential control points of the head section is 41, and the number of axial control points of the corresponding curved surface of the head section is 20;

[0015] The first section, the second section, and the third section are each divided into an upper section and a lower section, and the upper section and the lower section each have 41 circumferential control nodes. The first section, the second section, and the third section are adjacent to each other to form an inter-section curved surface I, and the inter-section curved surface I has 30 axial control points.

[0016] The number of circumferential control nodes of the wing root section and the wing end section is set to 60, and the number of axial control points of the curved surfaces corresponding to the wing root section and the wing end section is set to 30;

[0017] The number of circumferential control nodes of the lower and upper sections is set to 50, the number of axial control points of the curved surface corresponding to the upper section is 10, and an inter-section curved surface II is constructed between the upper and lower sections, and the number of axial control points of the inter-section curved surface II is 25;

[0018] The dimensionless curve parameters include: upper section type parameter Nu, height parameter Hu, lower section type parameter Nl, height parameter Hl, width parameter Yu, and axial position Xu.

[0019] Preferably, in S3, the design parameters of the head cross-section curve are: width parameter top_Yu=width_y, where width_y is the design width of the aircraft, height parameter top_Hu=3, type function top_Nu=1, shape function top_Sn=2, and Z-axis position Zu=0;

[0020] The design parameters of the cross-sectional curve of the upper half of the first section are: shape function sec1_Sn=2, type function sec1_Nu=1, curve height sec1_Hu=0.8, curve width sec1_Yu=1.7, and x-axis position sec1_Xu=4;

[0021] The design parameters of the lower half of the cross-sectional curve in the first section are: shape function sec1_Sn1=2, type function sec1_Nu1=1, curve height sec1_Hu1=0.03, and x-axis position sec1_Xu1=5. The design parameters of the lower half of the cross-sectional curve in the second and third sections are consistent with those of the lower half of the cross-sectional curve in the first section.

[0022] The design parameters of the upper section curve in the second section are: shape function sec2_Sn=2, type function sec2_Nu=1, curve height sec2_Hu=0.8, curve width sec2_Yu=1.7, and x-axis position sec2_Xu=6;

[0023] The design parameters of the upper section of the third section are: shape function sec3_Sn=2, type function sec3_Nu=1, curve height sec2_Hu=1, curve width sec2_Yu=1.7, and x-axis position sec3_Xu=10;

[0024] The design parameters of the wing end section curve are: shape parameter wings1_Sn=2, type function wings1_Nu=1, curve height wings1_Hu=0.05, curve length wings1_length=0.8167, horizontal wingspan wings1Y_H=0.56;

[0025] The design parameters of the wing root cross-section curve are: wing length wings1_L=4.9, wing width wings1_width=0.03;

[0026] The design parameters of the upper section curve are: shape parameter wings2_Sn=2, type parameter wings2_Nu=1, height parameter wings2_Hu=0.05, curve length wings2_length=0.245, and the vertical height of the upper and lower sections wings2Y_Hu=1.1.

[0027] Preferably, in S3, the curved surface I between each cross section is generated based on the corresponding transverse curve;

[0028] The curved surface between the head section and the first section adopts the upper half curved surface and the lower half curved surface to obtain the corresponding head curved surface;

[0029] The upper half curved surface is generated based on the transverse curve of the head section and the upper half of the first section; the lower half curved surface is generated based on the transverse curve of the head section and the lower half of the first section;

[0030] The head surface and the surface I between each section are combined according to the X-axis position to complete the generation of the aircraft body surface.

[0031] Preferably, in S3, the curved surface between the wing tip section and the wing root section is modeled based on the corresponding transverse curves, and the control points of the wing root section are sequentially connected from the wing tip section along the Y axis using straight lines to complete the modeling of the curved surface between the wing tip and the wing root;

[0032] The sealing surfaces of the upper and lower sections are generated by diffusing outward from the center points of the corresponding transverse curves, and the inter-section surface II is generated by connecting the circumferential control points corresponding to the upper and lower sections in a straight line manner.

[0033] The present invention includes at least the following beneficial effects: the overall parametric modeling of the aircraft is driven by the collaborative operation of global parameters to achieve efficient optimization in multiple dimensions such as aerodynamics, structure, and mission, significantly improving the overall performance, reliability, and design iteration efficiency.

[0034] Other advantages, objectives and features of the present invention will be reflected in part from the following description and will be understood by those skilled in the art through study and practice of the present invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0035] Figure 1 This is a schematic diagram of the structure of an aircraft in an embodiment of the present invention;

[0036] Figure 2 This is a dimensionless curve diagram of CST in an embodiment of the present invention;

[0037] Figure 3 This is a schematic diagram of a control cross section of an aircraft body in an embodiment of the present invention;

[0038] Figure 4 This is a schematic diagram of a head cross-section curve in an embodiment of the present invention;

[0039] Figure 5 A schematic diagram of a plane rotation of a head cross-section curve in an embodiment of the present invention;

[0040] Figure 6 This is a schematic diagram of a first cross-sectional curve in an embodiment of the present invention;

[0041] Figure 7 This is a schematic diagram of a curve corresponding to the generation of a head curved surface in an embodiment of the present invention;

[0042] Figure 8 A schematic diagram of a head curved surface in an embodiment of the present invention;

[0043] Figure 9 Schematic diagram of a curve corresponding to the generation of the inter-section curved surface I between the first section, the second section, and the third section in an embodiment of the present invention;

[0044] Figure 10 Schematic diagram of the inter-sectional curved surface I between the first section, the second section, and the third section in an embodiment of the present invention;

[0045] Figure 11 Schematic diagram of parameterized modeling of the bullet body in an embodiment of the present invention;

[0046] Figure 12 Schematic diagram of the control curve position of the horizontal wing in an embodiment of the present invention;

[0047] Figure 13 Schematic diagram of the position of the vertical wing control curve in an embodiment of the present invention;

[0048] Figure 14 Schematic diagram of the offset angle position of the upper section in an embodiment of the present invention;

[0049] Figure 15 This is a schematic diagram of a single horizontal wing curved surface in an embodiment of the present invention;

[0050] Figure 16 This is a schematic diagram of two horizontal wing curved surfaces distributed on both sides of the missile body in an embodiment of the present invention;

[0051] Figure 17 Schematic diagram of the sealing curved surface of the upper section of the vertical wing in an embodiment of the present invention;

[0052] Figure 18 This is a schematic diagram of two vertical wing curved surfaces distributed on both sides of the missile body in an embodiment of the present invention;

[0053] Figure 19 This is a schematic diagram of the parametric shape design of an aircraft in an embodiment of the present invention. DETAILED DESCRIPTION

[0054] The present invention will be described in further detail below in conjunction with the accompanying drawings so that those skilled in the art can implement the invention with reference to the description.

[0055] The present invention establishes different cross-sectional shapes based on the basic shape of a flight test platform aircraft. For example, control points are set for the XY plane cross-sectional shape of the head and the YZ plane cross-sectional shape of the body. Five-parameter and mold line design methods are then used to connect the corresponding control points for each cross-sectional shape. This five-parameter and mold line design method is then used to generate a curved surface, ultimately completing the parametric modeling of the aircraft. Building on the successful experience of the CST method in existing patents, its application scope is expanded from single components to the overall design of entire flight test platforms or other types of aircraft. This requires adjustments and optimization of the CST method to accommodate more complex geometric shapes.

[0056] In terms of parameter setting, the CST method has increased flexibility and adjustability. It can produce different shapes based on the design values ​​of different cross-sections and the design values ​​of the aircraft's top-view curve shape. This includes the shape function, type function, curve height, width, and number of curve control points of the head cross-section curve. This allows the generated shape to be adjusted to specific mission requirements. This can be achieved by introducing dynamic parameters and multi-mode switching, thereby enhancing the adaptability of the flight test platform.

[0057] Different control section shapes can be designed independently, as can the surface generation between sections. Overall exterior design often requires dealing with multiple distinct surface areas, such as the fuselage, wingtips, and empennage. By combining the five-parameter method with a mold line design approach, the various surfaces generated by the CST method are organically combined to form a coherent and efficient overall structure, effectively improving the performance of the flight test platform. This provides new insights into the development of parametric modeling for flight test platforms.

[0058] The research contents of the adopted parameterization method are as follows: CST method: CST (Class Function and Shapefunction transformation) method is an exponential curve parameterization method based on class function and shape function. Its core is to generate the basic geometric shape through the class function, and then use the shape function to modify it, and finally obtain the desired geometric parameterized shape. CST parameterization has the advantages of wide design space, easy adjustment, few variables, and the ability to generate continuous and robust geometric shapes. Therefore, it is widely used in aircraft shape modeling. In aircraft design, the CST method can be divided into cross-section control type (such as fuselage and lifting body) and airfoil control type (such as wings, vertical tail, etc.). By selecting the appropriate type function to determine the basic shape, and then using the shape function to accurately adjust it, the efficient design of the shape is finally achieved. For the airfoil control type shape, the mathematical expression of the CST method is as follows:

[0059]

[0060] In the above formula, ψ=x / c, ξ=z / c, ξT=Δz / c, c is the chord length, Δz is the thickness of the trailing edge, is used to describe the rounded leading edge, 1-ψ is used to describe the sharp trailing edge, and ψ∙ξT is used to describe the thickness of the trailing edge. is the mathematical expression of the CST method, ψ represents the normalized unit X-axis length, N represents the order of the Bernstein polynomial, represents the normalized X-axis length coordinate, represents the Bernstein polynomial expression, represents the airfoil thickness, z represents the Z-axis coordinate, x represents the X-axis coordinate, and ξ represents the CST curve coordinate.

[0061] definition is a type function, and ,definition is a shape function, and , N1, N2 represent geometric shape category parameters, Represents a type function expression 1, Represents a function expression of type 2.

[0062] Then the mathematical expression of the CST method can be expressed as:

[0063]

[0064] when At this time The cross-sectional curve is generated as a mathematical expression for the CST cross-sectional shape design. For a simple fuselage cross-section, the corresponding shape type can be obtained by determining the index Nu of the upper half and the index Nl of the lower half.

[0065] This invention proposes a five-parameter mold line design method. In practice, the five parameters primarily refer to combining contour parameters with CST technology, aiming to simplify three-dimensional shape modeling and reduce the number of parameters. The core approach is: first, the boundary contour of the aircraft is defined using a top-view power function (exponent N) and front-view parameters, and then the CST method is used to fine-tune the bottom cross-section. This fine-tuning is primarily achieved through dimensionless curves. For the missile body, the dimensionless curves for the nose section and the upper half of other sections primarily include six key parameters: shape parameter Sn, height parameter Hu, type parameter Nu, width parameter Yu, axial position, and top-view curve exponent parameter N. The dimensionless curves for the lower half of other sections primarily include five key parameters: shape parameter Sn, height parameter Hl, type parameter Nl, axial position, and top-view curve exponent parameter N. Together, these five parameters enable independent adjustment of the curvature and thickness of the upper and lower surfaces.

[0066] The mold line design in the five-parameter mold line design method mainly describes the shape characteristics through adjustable parameters, achieving efficient modeling while ensuring longitudinal smoothness. The specific process is: divide the fuselage longitudinally into 5 to 10 control stations, arrange the key control points of the cross section at each station and connect them into a closed curve; then connect the corresponding control points of each station with a continuous longitudinal curve to form a complete parametric model. This method is based on a small number of core control stations and generates the remaining cross-sectional data through longitudinal curve interpolation. It can not only accurately control the shape characteristics of key areas, but also significantly reduce the complexity of modeling. It is a key technology for missile body design that takes into account both engineering efficiency and aerodynamic performance.

[0067] Example:

[0068] The process of parametric modeling based on the image shape is as follows:

[0069] This study determined the basic shape of the flight test platform aircraft as follows Figure 1 As shown, it has a unique head shape. During the actual operation, the aircraft is divided into two parts: the body and the wings. Each part completes the parametric modeling design independently, and finally completes the parametric modeling design of the aircraft.

[0070] At the beginning of the parametric design of the aircraft, the CST method draws the curve size as a dimensionless size. The specific expression is:

[0071]

[0072] The CST curve drawn using the above formula is as follows Figure 2 As shown, Figure 1 N1=N2=0.5, , and when N1=N2, the curve is symmetrical on both sides.

[0073] 1. Curved design of the bullet body

[0074] Based on the basic shape of the aircraft, refer to other related shape designs, such as Figure 3 The design of each control section of the aircraft body in this embodiment can be expressed as: head section ( Figure 3 The black curve in the middle), the first section ( Figure 3 The red curve in the middle), the second section ( Figure 3 The blue curve in the middle), the third section ( Figure 3 (As shown by the yellow curve in the middle), there are a total of four control sections and the corresponding head shape curve design curve, first section curve, second section curve, and third section curve.

[0075] Among them, the head cross-section curve is in the XY plane, the overall length of the curve is top_Xu=3, the width is top_Yu=1.7, the Z-axis position is 0, and the vertex of the curve is at the origin.

[0076] The remaining sections are all in the YZ plane, and the bottom of the curve is a flat design. The X-axis position of the first section is sec1_Xu=4, the overall height of the curve is sec1_Z=0.8, and the width is sec1_Y=top_Yu.

[0077] The X-axis position of the second section is sec2_Xu=6, the overall height of the curve is sec2_Z=0.8, and the width is sec2_Yu=top_Yu.

[0078] The X-axis position of the third section is sec3_Xu=10, the overall height of the curve is sec3_Z=1, and the width is sec3_Yu=top_Yu.

[0079] Among the various design parameters of the head cross-section curve, its y-axis width is equal to the design width of the aircraft, that is, top_Yu = width_y; the length of the x-axis is controlled by the height variable value of the CST curve, top_Hu = 3; the larger the design value, the slender and sharper the aircraft's head configuration; the curve shape is mainly controlled by the type function top_Nu = 1, and different type function values ​​correspond to different head configurations, which also enables the aircraft to achieve complex and diverse aerodynamic layouts; the shape function top_Sn = 2 is a fixed value; the z-axis control position is 0; the number of circumferential control points of the head curve is point = 41, which can also be used as a design parameter, and different design values ​​have different sparsity levels of the curve. At this time, the CST mathematical expression corresponding to the head configuration is: ;

[0080] Substituting the design parameters of the head cross-section curve into the CST mathematical expression I, it can be rewritten as , and get Figure 4 The head cross-section is shown with a curved shape.

[0081] Furthermore, the head cross-section curve is designed to be rotatable along the y-axis with a rotation angle of theta. Figure 5 As shown, the z-axis of the graph is placed at the origin, and the plane rotation is achieved by rotating the y-axis. The blue curve is the original position. In the xz plane, the gray curve is the position after rotating -10 degrees (the rotation angle is positive when it is upward and negative when it is downward).

[0082] like Figure 6 As shown, the first section curve is designed to be divided into two parts, the upper half of the first section curve (i.e. Figure 6 The middle green curve, serving as the leeward side of the aircraft, requires optimized streamlined transitions. The planar curve in the lower half of the first section requires a large flat surface area, designed as a flat surface with the same width as the first section. The design parameters for the curve in the upper half of the first section are shape function sec1_Sn=2, type function sec1_Nu=1, curve height sec1_Hu=0.8, curve width sec1_Yu=1.7, x-axis position sec1_Xu=4, and 41 curve design control points. According to CST Mathematical Expression I, the CST parameterization formula for the curve in the upper half of the first section can be expressed as:

[0083]

[0084] The plane curve of the lower half of the first section (i.e. Figure 6 The design parameters of the red curve in the middle are shape function sec1_Sn1=2, type function sec1_Nu1=1, curve height sec1_Hu1=0.03, x-axial position of the first section sec1_Xu1=5, 41 curve design control points, and the CST parameterization formula of the plane curve of the lower half of the first section can be expressed as: In order to unify the shape and height of the plane, the plane curves in the lower half of the second and third sections are designed to be the same as the plane curves in the lower half of the first section.

[0085] The design parameters of the curve of the upper half of the second section are shape function sec2_Sn=2, type function sec2_Nu=1, curve height sec2_Hu=0.8, curve width sec2_Yu=1.7, x-axis position sec2_Xu=6, and 41 curve design control points. According to CST mathematical expression I, the CST parameterization formula of the curve of the upper half of the second section can be expressed as: The design parameters and curve shape of the lower half of the second section are the same as those of the lower half of the first section.

[0086] The design parameters of the curve of the upper half of the third section are shape function sec3_Sn=2, type function sec3_Nuu=1, curve height sec2_Hu=1, curve width sec2_Yu=1.7, x-axis position sec3_Xu=10, and 41 curve design control points. According to CST mathematical expression I, the CST parameterization formula of the curve of the upper half of the third section can be expressed as: The design parameters and curve shape of the lower half of the third section are the same as those of the lower half of the first section.

[0087] 2. Body surface generation

[0088] In this aircraft body design, based on the shape of each control section, the five-parameter design method and the mold line design method were used to generate the surface between two adjacent control sections, and the parametric modeling of the final aircraft body was completed. Figure 8 As shown, the head surface between the head curve and the first section is generated by the upper and lower parts; the upper half of the surface is generated by the head curve and the upper half of the first section; the lower half of the surface is generated by the head curve and the lower half of the first section; thus, the upper and lower parts are combined to generate the following Figure 8 The head surface shown in the gray grid has 20 axial station points and 82 circumferential station points. The corresponding curve design when the head section is generated is as follows: Figure 7 As shown ( Figure 7 The red lines in the middle are the head curve and the first section curve, respectively. The blue lines are the curve between the head section and the first section. The curve type is controlled by the top view curve index N).

[0089] like Figure 9- Figure 10 As shown ( Figure 9 、 Figure 10 The red lines are the head curve and each section curve respectively, and the blue lines are the curves between the first and second sections and the second and third sections. The curve type is controlled by the top view curve index N. Figure 10 The medium gray grid represents the corresponding inter-section surface I. Inter-section surface I is constructed between the first, second, and third sections. Similarly, the five-parameter + model line design method was used to generate the surface. Each inter-section surface I has 30 axial design points and 82 circumferential design points.

[0090] like Figure 11 As shown, the head surface and the surface I between each cross section are combined according to the X-axis position to obtain the surface parametric modeling of the aircraft body.

[0091] 3. Wing curve generation

[0092] Through the collection of relevant literature and information, combined with Figure 1The general shape of the aircraft wing determines the parameterization direction of this wing. The wing shape is divided into two parts: horizontal wing and vertical wing. Different parts are modeled separately. Based on the mathematical principles of the CST method, the basic airfoil of the aircraft is described, and the wing initialization shape expression is established.

[0093] like Figure 12 As shown, the horizontal wing control surface curve is divided into the wing root section curve and the wing end section curve ( Figure 12 The shorter red line D is the wing end section curve, the longer red line G is the wing root section curve, and the blue lines are the corresponding curves of the first section, the second section, and the third section). In the XZ plane, the wing root section curve is represented by a rectangular shape, with a wing length of wings1_L=4.9, a wing width of wings1_width=0.03, and 60 curve design control points. The wing end section curve is described using the CST method, with design variables including curve shape wings1_Sn=2, type function wings1_Nu=1, curve height wings1_Hu=0.05, and curve length wings1_length=0.8167. The horizontal span design parameter is wings1Y_H=0.56. The CST expression for the wing end curve is: , the curve design control points are 60.

[0094] like Figure 13 As shown ( Figure 13 The blue line X in the middle is the lower section curve, the blue line S is the upper section curve, the gray lines are the corresponding curves for the first, second, and third sections, and the red line is the horizontal wing control section curve. The CST method is used for the design curves of the upper and lower sections of the vertical wing. In the XY plane, the CST curve of the lower section has the same design value as the wing end section curve of the horizontal wing. The lower section closed curve is generated by symmetry of the curve endpoints, and the curve design control points are 50. The shape of the upper section is designed separately. The specific design parameters are curve shape wings2_Sn=2, type wings2_Nu=1, height wings2_Hu=0.05 and curve length wings2_length=0.245. The vertical height of the upper and lower sections of the vertical wing is wings2Y_Hu=1.1. The CST expression of the upper section curve is: , the curve design control points are 50. Figure 14 As shown ( Figure 14The red line in the middle represents the horizontal wing control surface curve, the blue line represents the vertical wing control surface curve, the large arc-shaped gray line represents the second and third section control surface curves, the gray line near theta angle represents the control surface position curve before the deflection of the upper half of the vertical wing, the gray circle is the position of the vertical wing before deflection, and the blue circle is the position of the vertical wing after deflection). In order to make the vertical wing coordinated, an upper section offset angle theta = 15° along the Z axis is added to the center point of the upper and lower sections, which can make the upper section translate a certain angle.

[0095] 4. Wing surface generation

[0096] like Figure 15- Figure 16 As shown ( Figure 15 YD is the surface corresponding to the wing end section, and YDG is the surface corresponding to the wing end section and the wing root section. Figure 16 The medium gray grid represents the horizontal wing surface. After completing the horizontal wing section curve design, the model line design method is used to determine the distribution of each point on the wing tip and root sections. Using straight lines along the Y-axis, the control points of the wing root section are connected from the wing tip section to the wing root section, completing the surface modeling between the wing tip and root. Similarly, the wing tip surface is modeled. At this point, the horizontal wing surface has 30 axial control points and 60 circumferential control points.

[0097] After completing the vertical wing curve design, the mold line design method was used to complete the sealing surface design of the upper and lower sections of the vertical wing. Figure 17 As shown ( Figure 17 The middle gray grid is the cross-section sealing surface on the horizontal wing surface. Taking the above cross section as an example, the upper cross-section sealing surface is obtained by diffusing outward from the center point of the curve. The surface has 10 points in the axial direction and 50 points in the circumferential direction. The specific surface generation is as follows: Figure 18 As shown ( Figure 18 The middle gray grid is the vertical wing surface, the gray lines are the curves of the three sections, and the red lines are the horizontal wing control surface curves. After determining the shape, position and curve control points of the upper and lower sections of the vertical wing, the model line design method is used to connect the corresponding control points of the two sections in a straight line to complete the generation of the inter-section surface II between the upper and lower sections. The surface has 25 points in the axial direction and 50 points in the circumferential direction. The generated surface diagram Figure 15 Finally, the lower section surface + inter-section surface II + upper section surface are combined to complete the parametric modeling of the vertical wing surface, with 45 points in the axial direction and 50 points in the circumferential direction.

[0098] 5. Aircraft parametric modeling generation

[0099] After completing the parametric modeling of the aircraft body and wing surfaces separately, the wing surfaces are placed at the appropriate position on the body surface to form the parametric shape of the aircraft, such as Figure 19This parametric generation method for aircraft, through the control of key aircraft parameters, enables high flexibility and variability, enabling the rapid generation of diverse appearance options based on design requirements. Geometrically, transitions from simple, flowing lines to complex and varied curved surfaces can all be precisely controlled through parameter changes. Furthermore, leveraging the associative nature of parametric design, the entire aircraft model automatically updates when a parameter is modified, effectively completing design optimization and significantly improving design efficiency and quality.

[0100] At this point, a parametric modeling design of the flight test platform based on the CST method was completed.

[0101] The above solution is only an illustration of a preferred embodiment, but is not limited thereto. When implementing the present invention, appropriate replacements and / or modifications can be made according to user needs.

[0102] Although the embodiments of the present invention have been disclosed above, they are not limited to the applications listed in the specification and exemplary embodiments. They can be applied to a variety of fields suitable for the present invention. Further modifications will be readily apparent to those skilled in the art. Therefore, the present invention is not limited to the specific details and illustrations shown and described herein without departing from the general concept defined by the claims and their equivalents.

Claims

1. A parametric modeling method for a flight test platform based on the CST method, characterized in that: include: S1. Divide the functions of the flight test platform aircraft according to its external structure, and divide the divided parts into multiple different cross-sectional shapes according to the external structure; S2. Set corresponding control points on each cross-sectional shape; S3. Using the five-parameter model line design method, first connect the adjacent circumferential control points on each cross-sectional shape one by one to generate a transverse curve. Then, based on the adjacent transverse cross-sections, connect the control points of each curve one by one to generate a longitudinal curve. Finally, the curved surface is obtained, completing the parametric modeling of each part. S4, assembling the independently completed parametric modeling parts to complete the final parametric modeling of the aircraft; In S3, the five-parameter mold line design method includes: first defining the boundary contour of each cross-sectional shape by using the top view power function and the front view parameters, and then using the CST method to finely control the cross-sectional shape; Mathematical expression of CST method for airfoil control shape for: In the above formula, is the type function, N1, N2 represent the geometric shape category parameters, is the shape function, represents the airfoil thickness, ψ represents the normalized unit X-axis length; Mathematical expression of CST method for cross-section control shape for: Wherein, the refined control of the shape of each cross section on the aircraft is achieved through dimensionless curves; In S1, the flight test platform aircraft is divided into the body and wings according to its external structure; The projectile body is divided into four cross-sectional shapes: head cross-sectional, first cross-sectional, second cross-sectional, and third cross-sectional; The wing includes a horizontal wing and / or a vertical wing, wherein the horizontal wing has two cross-sectional shapes: a wing root cross-sectional shape and a wing end cross-sectional shape, and the vertical wing has two cross-sectional shapes: a lower cross-sectional shape and an upper cross-sectional shape, and the CST curve corresponding to the lower cross-sectional shape of the vertical wing is consistent with the design value of the curve corresponding to the wing end cross-sectional shape of the horizontal wing; In S2, there are 41 circumferential control points on the head section and 20 axial control points on the corresponding curved surface of the head section; The first section, the second section, and the third section are each divided into an upper section and a lower section, and the upper section and the lower section each have 41 circumferential control nodes. The first section, the second section, and the third section are adjacent to each other to form an inter-section curved surface I, and the inter-section curved surface I has 30 axial control points. The number of circumferential control nodes of the wing root section and the wing end section is set to 60, and the number of axial control points of the curved surfaces corresponding to the wing root section and the wing end section is set to 30; The number of circumferential control nodes of the lower and upper sections is set to 50, the number of axial control points of the curved surface corresponding to the upper section is 10, and an inter-section curved surface II is constructed between the upper and lower sections, and the number of axial control points of the inter-section curved surface II is 25; The dimensionless curve parameters include: upper section type parameter Nu, height parameter Hu, lower section type parameter Nl, height parameter Hl, width parameter Yu, and axial position Xu; In S3, the design parameters of the head section curve are: width parameter top_Yu=width_y, where width_y is the design width of the aircraft, height parameter top_Hu=3, type function top_Nu=1, shape function top_Sn=2, and Z-axis position Zu=0; The design parameters of the cross-sectional curve of the upper half of the first section are: shape function sec1_Sn=2, type function sec1_Nu=1, curve height sec1_Hu=0.8, curve width sec1_Yu=1.7, and x-axis position sec1_Xu=4; The design parameters of the lower half of the cross-sectional curve in the first section are: shape function sec1_Sn1=2, type function sec1_Nu1=1, curve height sec1_Hu1=0.03, and x-axis position sec1_Xu1=5. The design parameters of the lower half of the cross-sectional curve in the second and third sections are consistent with those of the lower half of the cross-sectional curve in the first section. The design parameters of the cross-sectional curve of the upper half of the second section are: shape function sec2_Sn=2, type function sec2_Nu=1, curve height sec2_Hu=0.8, curve width sec2_Yu=1.7, and x-axis position sec2_Xu=6; The design parameters of the upper section of the third section are: shape function sec3_Sn=2, type function sec3_Nu=1, curve height sec2_Hu=1, curve width sec2_Yu=1.7, and x-axis position sec3_Xu=10; The design parameters of the wing end section curve are: shape parameter wings1_Sn=2, type function wings1_Nu=1, curve height wings1_Hu=0.05, curve length wings1_length=0.8167, horizontal wingspan wings1Y_H=0.56; The design parameters of the wing root cross-section curve are: wing length wings1_L=4.9, wing width wings1_width=0.03; The design parameters of the upper section curve are: shape parameter wings2_Sn=2, type parameter wings2_Nu=1, height parameter wings2_Hu=0.05, curve length wings2_length=0.245, and the vertical height of the upper and lower sections wings2Y_Hu=1.

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2. The parametric modeling method for a flight test platform based on the CST method according to claim 1, characterized in that: In S3, the surface I between each section is generated based on the corresponding transverse curve; The curved surface between the head section and the first section adopts the upper half curved surface and the lower half curved surface to obtain the corresponding head curved surface; The upper half curved surface is generated based on the transverse curve of the head section and the upper half of the first section; the lower half curved surface is generated based on the transverse curve of the head section and the lower half of the first section; The head surface and the surface I between each cross section are combined according to the X-axis position to complete the generation of the aircraft body surface.

3. The parametric modeling method for a flight test platform based on the CST method according to claim 2, characterized in that: In S3, the surface between the wing tip section and the wing root section is modeled based on the corresponding transverse curves. Straight lines are used to connect the control points of the wing root section along the Y axis from the wing tip section to complete the surface modeling. The sealing surfaces of the upper and lower sections are generated by diffusing outward from the center points of the corresponding transverse curves, and the inter-section surface II is generated by connecting the circumferential control points corresponding to the upper and lower sections in a straight line manner.

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