A hypersonic vehicle aerodynamic configuration with a bendable nose and a method of designing the same

Abstract

The application discloses a supersonic aircraft aerodynamic configuration with a bendable head and a design method thereof. The aerodynamic configuration comprises a waverider head, a lifting body fuselage and a bendable transition region connecting the two. The waverider head is used to generate compression lift, the lifting body fuselage provides main lift, and the bendable transition region is used to realize continuous deflection of the head relative to the fuselage in the symmetry plane and keep the configuration smooth. By adjusting the deflection angle of the head, the forebody pressure and aerodynamic moment distribution of the aircraft can be changed, and dynamic regulation of lift-drag characteristics and stability can be realized. The application organically integrates the waverider aerodynamic layout and the head deflection function, and constructs the transition region through a parameterized Bezier surface, thereby overcoming the technical obstacles of disconnection of the aerodynamic layout and flow separation in the transition region in the traditional head deflection scheme, and significantly improving the aerodynamic efficiency, maneuvering ability and flight stability of the aircraft under multiple working conditions.

Description

Technical Field

[0001] This invention belongs to the field of aerodynamic layout design technology for aircraft, specifically relating to the aerodynamic shape of a supersonic aircraft with a bendable nose and its design method. Background Technology

[0002] With the continuous development of supersonic aircraft technology, higher requirements are being placed on the aerodynamic performance, maneuverability, and long-range flight capability of aircraft under wide speed range conditions. Traditional supersonic aircraft typically adopt a conical nose and an integrated lifting body or slender body configuration, and their overall aerodynamic performance mainly depends on the combined performance of the fixed shape at different flight stages.

[0003] However, under various flight conditions, a fixed aerodynamic shape cannot simultaneously meet the demands of cruise efficiency and maneuverability. For example, during high-speed cruise flight, a smaller frontal area and optimized shock wave structure are typically required to reduce drag; while during maneuvering or attitude adjustments, a shape characteristic capable of providing a large aerodynamic torque is needed. Because traditional aircraft use a rigid integrated design between the nose and fuselage, the forebody flow field and aerodynamic torque distribution are difficult to actively adjust, which limits the aircraft's aerodynamic efficiency, control torque distribution, and flight stability, thus restricting further improvements in its overall performance.

[0004] To address this issue, variable aerodynamic shape technology has emerged. Compared to traditional rigid aircraft, aircraft with a flexible nose shape can dynamically adjust their aerodynamic profile, thereby achieving superior aerodynamic characteristics under different flight conditions. By actively altering the aircraft's aerodynamic shape, lift, drag, and torque characteristics can be improved, enhancing the aircraft's maneuverability and aerodynamic efficiency. Simultaneously, the synergistic effect of the deformable nose and control system helps improve attitude adjustment capabilities and system stability margin, thereby enhancing controllability and stability during flight. Several solutions regarding aircraft nose deformation have already been proposed in the existing technology.

[0005] For example, prior art document CN113232828A discloses a deflection control mechanism for a deformable nose structure of a supersonic aircraft. This mechanism uses a double ball joint structure to achieve deflection of the nose at any angle in the heading, thereby generating a bias torque to change the flight direction. However, this technical solution mainly focuses on the deflection implementation at the mechanism level and does not delve into and solve the following key issues:

[0006] 1. Coordinated design of aerodynamic shape: The scheme does not specify what kind of efficient aerodynamic shape (such as a wave rider) should be used for the head, nor does it involve how to design the fuselage aerodynamic shape that matches the head deflection, which makes it impossible for the overall aerodynamic layout to fully utilize the performance gain brought by the deflection.

[0007] 2. Issues with the continuity of the deflection transition: This design does not provide an aerodynamic shape scheme for achieving a smooth transition between the nose and fuselage after deflection. In actual supersonic flight, if a simple rigid deflection is used, a step or gap will inevitably be generated at the connection between the nose and fuselage, leading to severe shock wave / boundary layer interference, flow separation, and localized aerodynamic heating problems, which will greatly impair the aerodynamic performance and safety of the aircraft.

[0008] 3. Quantitative control of aerodynamic characteristics: The scheme does not establish a quantitative response relationship between the nose deflection angle and key aerodynamic parameters such as the overall lift, drag, and pitch moment of the aircraft, which makes it difficult to accurately control the attitude and stability of the aircraft.

[0009] In addition, there are other variant head cone solutions in the prior art, such as multi-stage driven linkage variant head cones or reconfigurable head cones, but they also focus on the deformation mechanism itself, or have fundamental differences in the deformation method from the overall head deflection solution of the present invention. None of them provide a complete technical solution that organically integrates the efficient aerodynamic layout of the waverider with the head deflection function and solves the problem of aerodynamic continuity in the transition region.

[0010] Therefore, there is an urgent need to propose a new type of aerodynamic shape for supersonic aircraft, which can not only achieve nose deflection to provide maneuver control torque, but also systematically optimize the shape of the nose, fuselage and transition section from the aerodynamic layout level, ensuring the continuous smoothness of the aerodynamic shape of the whole aircraft under different deflection states, thereby realizing fine control of the forebody flow field and quantitative optimization of aerodynamic characteristics. Summary of the Invention

[0011] This invention aims to address the problems in existing supersonic aircraft aerodynamic shapes, such as the rigid connection between the waverider head and the lifting body fuselage, which leads to difficulties in adjusting the forebody flow field, fixed aerodynamic layout, insufficient maneuverability, and limited flight stability margin. It also addresses the technical problems in existing technologies, such as the disconnect between the head deflection scheme and efficient aerodynamic layout design, aerodynamic discontinuities in the deflection transition region, and a lack of quantitative control basis for aerodynamic characteristics. The invention proposes a supersonic aircraft aerodynamic shape with a flexible head and its design method. A flexible transition section connects the waverider head to the mid-to-rear section of the lifting body fuselage, allowing the waverider head to deflect relative to the lifting body fuselage within a certain range. This enables the aircraft to adaptively adjust its aerodynamic shape according to different flight conditions during flight, improving overall aerodynamic performance and control capabilities.

[0012] The technical solution of this invention is as follows:

[0013] This invention provides an aerodynamic shape for a supersonic aircraft with a bendable nose, comprising:

[0014] The waverider head is used to control the shock wave structure of the forebody and generate compressive lift under supersonic flight conditions.

[0015] The lifting body fuselage provides the main volume and main lift of the aircraft;

[0016] A flexible transition region connects the waverider head and the lifting body fuselage, enabling continuous deflection of the waverider head relative to the lifting body fuselage within the aircraft's plane of symmetry, and maintaining aerodynamic continuity and surface smoothness between the waverider head and the lifting body fuselage under different deflection states.

[0017] The deflection angle of the waverider head is configured to be adjustable to change the pressure distribution and aerodynamic torque distribution in the forebody region of the aircraft.

[0018] In a further preferred embodiment, the geometry of the bendable transition region is constructed by a parametric Bezier surface, the position of the control points of which is dynamically adjusted as the deflection angle changes to satisfy the boundary continuity condition.

[0019] In a further preferred embodiment, the boundary continuity condition includes: at least first-order geometric continuity or second-order geometric continuity is satisfied at the connection between the waverider head and the flexible transition region, and at the connection between the flexible transition region and the lifting body fuselage.

[0020] In a further preferred embodiment, the waverider head is designed based on the theory of cone-guided waveriders, with a design Mach number range of Ma = 4 to Ma = 7, and the leading edge of the waverider head is passivated.

[0021] In a further preferred embodiment, the cross-sectional profile of the lifting body fuselage at any axial position is composed of an upper curved segment, a middle transition curved segment, and a lower curved segment, and each cross-section gradually widens along the axial direction.

[0022] In a further preferred embodiment, the adjustment of the deflection angle of the waverider head enables the aircraft to achieve a pitch moment coefficient change of not less than 0.02 when the deflection angle changes from 0° to 15° under the conditions of Ma=6 and angle of attack 4°, and the peak pressure coefficient of the upper surface of the aircraft is reduced by not less than 15%.

[0023] Furthermore, the present invention also provides a method for designing the aerodynamic shape of the aforementioned supersonic aircraft, the method comprising the following steps:

[0024] Step 1: Overall configuration parameterization definition: Establish the spatial coordinate system of the aircraft's aerodynamic shape, divide the aerodynamic shape into the waverider head, the lifting body fuselage, and the flexible transition region connecting the two, and define the range of values ​​for the head deflection angle.

[0025] Step 2: Constructing the aerodynamic shape of the lifting body fuselage: Using a multi-segment curve fusion section parameterization construction method, the three-dimensional aerodynamic shape of the lifting body fuselage is constructed by defining the geometric parameters of multiple control sections and performing surface interpolation along the axial direction.

[0026] Step 3: Construct the aerodynamic shape of the waverider head: Based on the theory of conical waveriders, the shock surface is established under the given design Mach number. The surface of the waverider head is generated by defining the upper edge line and tracing the streamline along the incoming flow direction, and the leading edge is passivated.

[0027] Step 4: Construct the aerodynamic shape of the flexible transition region: Based on the constructed waverider head and lifting body fuselage, the head deflection angle is introduced as a control parameter. The head deflection is achieved through rigid body geometric transformation, and a continuous and smooth transition surface is constructed between the two boundaries using parametric Bezier curves or surfaces.

[0028] Step 5: Aerodynamic characteristic parameter response and verification: Based on the constructed aerodynamic shape, the aerodynamic force and aerodynamic torque characteristics of the aircraft are analyzed under different incoming Mach numbers Ma, angle of attack α and nose deflection angle θ, and the effect of nose deflection on the forebody flow field structure and aerodynamic characteristics is verified.

[0029] In a further preferred embodiment, in step 2, at least three control sections are provided along the axial direction of the aircraft. The upper curve characteristic radius and the lower curve characteristic radius of each control section increase monotonically along the axial direction to form an aerodynamic shape that gradually widens along the axial direction.

[0030] In a further preferred embodiment, in step 4, the control point position of the Bezier surface is dynamically calculated based on the deflection angle using an interpolation function to ensure that the geometry of the bendable transition region continuously evolves with the deflection angle and always satisfies the boundary matching conditions with the waverider head and the lifting body fuselage.

[0031] Beneficial effects

[0032] Compared with the prior art, the present invention has the following beneficial effects:

[0033] 1. This invention achieves the ability to adjust the aerodynamic shape of the head relative to the fuselage: by constructing a flexible transition section between the waverider head section and the lifting body fuselage section, and introducing the head deflection angle parameter θ, the aerodynamic shape of the head relative to the fuselage can be continuously and adjustable in space; under different θ values, the aerodynamic shape of the aircraft forebody changes, thereby achieving the adjustment of the forebody flow field structure and aerodynamic force distribution, and improving the adaptability of the aerodynamic shape to different flight states.

[0034] 2. This invention achieves overall collaborative design: This invention parametrically and collaboratively designs the waverider head, lifting body fuselage, and flexible transition region as an organic whole, solving the problem of the disconnect between the deformation mechanism and aerodynamic layout design in existing technologies. The waverider head ensures a high lift-to-drag ratio during high-speed cruise, the lifting body fuselage provides efficient internal loading space and lift contribution, while the flexible transition section ensures the continuity and smoothness of the overall aerodynamic shape when the head deflects.

[0035] 3. This invention achieves refined flow field control capabilities: by changing the nose deflection angle θ, the pressure distribution C on the aircraft surface can be induced. p Changes in these parameters affect aerodynamic forces and aerodynamic torque characteristics, thereby enabling the adjustment of aerodynamic torque and improving the aircraft's attitude control under different operating conditions. Compared to prior art document CN113232828A, which only provides the offset torque, this invention establishes the relationship between the deflection angle θ and aerodynamic parameters (such as the pitching moment coefficient C). m Pressure coefficient C p The quantitative response relationship of the shock wave distribution (pressure distribution) enables precise and predictable adjustment of the forebody shock wave structure, pressure distribution, and aerodynamic center. Experimental data show that, under the conditions of Ma=6 and angle of attack of 4°, when θ changes from 0° to 15°, the pitching moment coefficient C m It can produce a change of 0.03 and reduce the peak pressure on the upper surface of the aircraft by more than 20%, significantly improving the aircraft's trim capability and maneuverability.

[0036] 3. This invention overcomes the technical obstacles of flow separation in the transition region: This invention designs a bendable transition region constructed from a Bezier surface and applies strict geometric continuity constraints, solving the problems of steps and gaps caused by rigid deflection in existing technologies, and the resulting flow separation and local heat flux peaks. Simulation comparisons show that the smooth transition scheme of this invention can shorten the boundary layer separation zone length in the transition region by more than 50% and reduce the local heat flux peak by more than 30%.

[0037] 4. This invention improves the aerodynamic center position and longitudinal stability characteristics: by deflecting the nose, the pressure distribution of the forebody is redistributed along the axial direction, thereby changing the aerodynamic center position; and by adjusting θ, the relative positional relationship between the aerodynamic center and the center of mass can be changed within a certain range, thereby affecting the longitudinal stability characteristics of the aircraft and enabling it to maintain a relatively stable aerodynamic response under different angles of attack and Mach numbers.

[0038] 5. Significantly enhanced multi-condition adaptability: By actively adjusting the nose deflection angle θ, the aircraft of this invention can flexibly switch between efficient cruise mode (small θ angle, drag reduction) and high maneuverability mode (large θ angle, providing large control torque), so that the aerodynamic shape has different geometric shapes in different flight states. Under different θ conditions, the aerodynamic shape of the aircraft can be adjusted between lift characteristics and drag characteristics, which greatly expands the airworthiness envelope and mission flexibility of the aircraft and improves its adaptability in various flight conditions.

[0039] Additional aspects and advantages of the invention will be set forth in part in the description which follows, and in part will be obvious from the description, or may be learned by practice of the invention. Attached Figure Description

[0040] The above and / or additional aspects and advantages of the present invention will become apparent and readily understood from the description of the embodiments taken in conjunction with the following drawings, in which:

[0041] Figure 1 This is a design flowchart of the aircraft of this invention.

[0042] Figure 2 These are three-view drawings of the supersonic aircraft with a variable-bend nose designed according to the present invention.

[0043] Figure 3 This is a schematic diagram of the cross-sectional parameters of the lifting body fuselage of the present invention.

[0044] Figure 4 This is a schematic diagram of the design of the cone-guided waverider head of the present invention.

[0045] Figure 5 This is a schematic diagram of the leading edge passivation of the cone-guided waverider head of the present invention.

[0046] Figure 6 The pitch moment coefficient C of the aircraft of this invention at Mach 6 m Relationship with angle of attack α.

[0047] Figure 7 This is a pressure cloud diagram of the upper surface and the plane of symmetry of the aircraft at an angle of attack of 6 Ma and 4°.

[0048] Figure 8 This is a pressure cloud diagram of the lower surface and the plane of symmetry of the aircraft at an angle of attack of 6 Ma and 4°. Detailed Implementation

[0049] The present invention will now be described in further detail with reference to the accompanying drawings. The description is intended to explain the present invention and should not be construed as limiting the present invention.

[0050] Design process description:

[0051] like Figure 1 As shown, the aerodynamic shape of the supersonic aircraft with a bendable nose in this embodiment is constructed according to the parametric design process, which mainly includes the design of the lifting body fuselage, the design of the waverider nose, and the construction of the bendable transition section.

[0052] like Figure 2 As shown, the overall aerodynamic shape of the aircraft includes a waverider head section, a lifting body fuselage section, and a flexible transition section connecting the two; wherein, the waverider head section is located at the front of the aircraft, the lifting body fuselage section is located at the middle and rear of the aircraft, and the flexible transition section is located between the head section and the fuselage section.

[0053] Step 1: Establishing the overall configuration and defining the coordinate system:

[0054] like Figure 1 As shown in the design flow, a spatial rectangular coordinate system O-XYZ is first established. The X-axis points along the aircraft's axial direction towards the flight direction, the Y-axis points along the wing span towards the wingtip, and the Z-axis is perpendicular to both the X and Y axes. The aircraft's aerodynamic shape is symmetrical about the XOZ plane.

[0055] The aerodynamic shape is segmented along the X-axis of the aircraft, such as... Figure 2 As shown, aircraft are divided into:

[0056] Waverider head;

[0057] Flexible transition area;

[0058] Lifting body fuselage;

[0059] Where: L is the total length of the aircraft, L h L is the length of the head segment. t This is the endpoint of the transition zone. In this embodiment, L = 4.5m. h =1.1m, L t =1.65m.

[0060] like Figure 2 As shown, the waverider nose section can deflect relative to the lifting body fuselage section within the aircraft's plane of symmetry; the nose deflection angle is defined as θ, and its range is:

[0061]

[0062] in The maximum deflection angle is set; in this embodiment, the deflection is a rotation about the Y-axis, and its geometric transformation relationship is as follows:

[0063]

[0064] in: The spatial parameter equations for the head of the waverider in the undeflected state are given. The spatial parameter equations for the head of the deflected waverider are as follows. For rotation matrix:

[0065]

[0066] In addition, to ensure the continuity and smoothness of the aircraft's aerodynamic shape, the following conditions must be met at the joints between each segment:

[0067] At the connection point between the head segment and the transition segment, x=L h Location:

[0068]

[0069] At the connection point between the transition section and the body section, x=L t Location:

[0070]

[0071] in The spatial parameter equations for the bendable transition section are as follows: The equations for the spatial parameters of a lifting body fuselage are given.

[0072] Step 2: Constructing the aerodynamic shape of the lifting body fuselage:

[0073] (1) Three control sections are selected along the aircraft axis in the lifting body fuselage region, namely the front section, the middle section and the rear section, and their positions are defined as:

[0074] , ,

[0075] In this embodiment, we take L1 = 1.2m, then x2 = 2.85m.

[0076] The three sections are parallel to each other, their axes of symmetry coincide with the X-axis, and the top of each section is located on the same horizontal reference plane.

[0077] (2) such as Figure 3 As shown, each control section consists of three curve segments: the upper curve segment (arc apex), the middle transition curve segment, and the lower curve segment (circular arc surface). Each segment is controlled by the following parameters:

[0078] Upper curve segment: radius R1, corresponding central angle θ r1 The middle transition curve segment has a radius of R2; the lower curve segment has a radius of R3, corresponding to a central angle θ. r3Passivation transition: radius R4; the same parameter of each section increases monotonically along the axial direction to form a gradually widening shape.

[0079] In this example, the parameters of each section are as follows:

[0080] Front section (x=1.65m):

[0081] R1=0.25m, θ r1 =100°, R2=0.2m, R3=0.15m, θ r3 =120°, R4=0.05m

[0082] Mid-section (x=2.85m):

[0083] R1=0.4m, θ r1 =110°, R2=0.32m, R3=0.25m, θ r3 =130°, R4=0.08m

[0084] Front section (x=45m):

[0085] R1=0.6m, θ r1 =120°, R2=0.5m, R3=0.4m, θ r3 =140°, R4=0.1m

[0086] (3) For each cross section, the curve construction method is as follows:

[0087] The upper curve segment is represented by an arc, and its expression is:

[0088]

[0089] Its angle range is:

[0090]

[0091] The lower curve segment is represented by an arc:

[0092]

[0093] The corresponding angle range is:

[0094]

[0095] Where y c1 z c1 Let x and y be the x and y coordinates of the center of the upper arc of the cross section, respectively. c3 z c3 These are the x and y coordinates of the center of the lower arc of the cross section, respectively.

[0096] The transition curve in the middle section uses either a circular arc or a Bezier curve to connect the upper and lower curve segments. In this example, a circular arc is used for the connection.

[0097] (4) To ensure the smoothness of the shape, curve connections and passivation treatment are performed:

[0098] The upper curve is tangent to the middle curve:

[0099]

[0100] The middle curve is tangent to the lower curve, and a blunt arc is set at the connection between the middle and lower curves:

[0101]

[0102] To avoid abrupt changes in local curvature, where y c4 z c4 These are the x-coordinate and y-coordinate of the center of the blunted circular arc of the cross section, respectively.

[0103] (5) Generation of three-dimensional body shape

[0104] By performing surface interpolation on the control sections, the three-dimensional aerodynamic shape of the aircraft is constructed:

[0105]

[0106] Where г is the surface interpolation function, and its corresponding spatial parameters are expressed as:

[0107]

[0108] The final aerodynamic shape of the lifting body fuselage section is obtained, such as Figure 2 As shown

[0109] Step 3: Construct the aerodynamic shape of the waverider head

[0110] like Figure 4 and Figure 5 As shown, in this embodiment, the head section of the waverider is designed based on the theory of conical waveriders. Under a given design Mach number, the aerodynamic shape of the head is generated by constructing a shock wave-constrained flow field and combining it with the streamline tracing method.

[0111] (1) Solving the conical flow field:

[0112] Based on the incoming Mach number Ma and the cone half-cone angle δ, the inviscid conical flow field is solved using the fourth-order Runge-Kutta numerical method, thereby determining the corresponding shock wave angle β.

[0113] In the specific solution process, we first assume an initial value for the shock angle β, and under the given shock angle condition, we reverse the calculation to find the corresponding cone half-cone angle δ. Through iterative calculation, we make the solved δ consistent with the target value, thereby determining the shock angle β that satisfies the condition.

[0114] The conical flow field is established in a polar coordinate system and numerically integrated along the polar angle direction. The integration step size is set to Δψ, and the initial integration starting point is taken at the shock wave surface. The initial flow parameters behind the wave are determined according to the shock wave relationship, including the components of velocity in the radial and tangential directions.

[0115] Subsequently, the fourth-order Runge-Kutta method was used to integrate the governing equations step by step to obtain the velocity distribution of the flow field at each polar angle, thus establishing the complete solution of the inviscid conical flow field. The specific solution process is as follows:

[0116]

[0117] Where v is the velocity component, and V is the velocity. ρ is the critical speed of sound, γ is the air density, γ is the specific heat ratio of air, and the subscript 1 indicates the wavefront parameter and 2 indicates the waveback parameter.

[0118] (2) Construction of the upper surface of the waverider:

[0119] After determining the conical flow field, the bottom profile of the upper surface of the waverider is defined. A set of flow surfaces parallel to the direction of the free flow is constructed through this bottom profile, and the intersection of this surface with the conical shock surface is the leading edge curve of the waverider.

[0120] The flow surface is intercepted from the leading edge curve to obtain the shape of the upper surface of the waverider.

[0121] In this embodiment, the bottom profile of the upper surface is described using a quartic function:

[0122]

[0123] And introduce parameters:

[0124]

[0125] Where: k is the shape control parameter; ψ is the angle between the starting direction of the shape and the horizontal axis; θ0 is the angle between the tangent at the bottom end of the shape and the horizontal axis; by determining the parameters k, ψ, and θ, the bottom shape of the upper surface can be uniquely determined.

[0126] (3) Generation on the lower surface of the waverider

[0127] Using the leading edge curve as the starting boundary, streamline tracing is performed in the conical flow field to generate the lower surface of the waverider.

[0128] During streamline tracing, the time step Δt is first set, and the current velocity is used as an approximation of the flow velocity. Based on the flow control equations, the differential equations of particle motion are established:

[0129]

[0130] The fourth-order Runge-Kutta method was then used to numerically integrate the above equations, and the spatial coordinates of each discrete point of the streamline were solved step by step.

[0131] By performing streamline tracing on multiple starting points, a set of streamlined surfaces can be obtained; by smoothing these surfaces, the lower surface of the waverider can be formed.

[0132] (4) Passivation treatment of the leading edge of the waverider

[0133] The waverider generated using the above method has a sharp leading edge structure. To improve aerodynamic thermal adaptability, the leading edge is passivated.

[0134] In this embodiment, the entire upper surface is raised, and a radius of R is introduced in the leading edge region. n A 10 mm circular arc transition is used to form a blunt leading edge.

[0135] Finally, the constructed waverider head is geometrically connected to the flexible transition section to form the complete aerodynamic shape of the aircraft forebody, such as... Figure 5 As shown.

[0136] Step 4: Construct the aerodynamic shape of the flexible transition section:

[0137] like Figure 2 As shown, this embodiment achieves a smooth aerodynamic shape transition between the waverider nose section and the lifting body fuselage section by introducing a flexible transition section. The aerodynamic shape of this transition section is dynamically adjusted according to the nose deflection angle θ, thereby optimizing the aerodynamic performance of the aircraft.

[0138] (1) Spatial geometric relationship of the transition section

[0139] The transition segment begins at the end of the head segment, x=L. h =1.1 m, the termination position is the starting position of the lifting body fuselage section x=L t =1.65 m.

[0140] To realize the effect of head deflection on aerodynamic shape, the geometric transformation of the transition section is defined as being controlled by the head deflection angle θ of the head section. Specifically, the connection between the head section and the transition section is given by the following spatial transformation formula:

[0141]

[0142] The rotation matrix is:

[0143]

[0144] This formula shows that the shape of the transition section changes with the deflection angle θ, generating a smooth aerodynamic shape between the two sections.

[0145] (2) Constructing control points and curves for the transition section:

[0146] like Figure 2 As shown, the transition section is controlled by a Bezier curve. To ensure the continuity and smoothness of the transition section, control points are defined. It changes dynamically with the deflection angle θ.

[0147] Control Points The following interpolation curve is formed as the deflection angle θ changes:

[0148]

[0149] u∈[0,1] represents the curve parameters of the transition segment. Bernstein basis functions:

[0150]

[0151] Between each section, the control points are smoothly connected by the Bezier interpolation function to ensure that the shape of the transition section remains continuous and smooth as the deflection angle θ changes.

[0152] (3) Construct the transition surface:

[0153] Based on control points Construct the three-dimensional surface of the transition section:

[0154]

[0155] Q ij (θ) represents the control points of the surface. Here, is the Bernstein basis function, and u and v are the transition surface parameters.

[0156] By adjusting the deflection angle θ, different transition section shapes can be generated under different flight conditions, thereby optimizing the aerodynamic performance of the aircraft.

[0157] (4) Geometric constraints of the transition section

[0158] To ensure the smoothness and stability of the transition section's shape, the transition section satisfies the following constraints at its start and end positions, and the transition section satisfies boundary matching conditions at both ends:

[0159] Head connection

[0160] Body connection

[0161] These constraints ensure a smooth geometric transition between the head section, transition section, and body section, preventing aerodynamic discontinuities.

[0162] (5) Aerodynamic response of shape as deflection angle changes:

[0163] By adjusting the nose deflection angle θ, the aerodynamic shape of the aircraft changes, thereby affecting its lift, drag, and aerodynamic stability. At different deflection angles θ, the flow field in the aircraft's forebody changes, influencing the aerodynamic torque distribution.

[0164] Numerical simulations and experimental verifications can be used to analyze the changes in the aerodynamic performance of an aircraft under different values ​​of θ. For example, when the deflection angle θ = 15°, the lift of the aircraft increases and the drag decreases at high Mach numbers; while when the deflection angle θ = -15°, the aerodynamic stability and handling performance of the aircraft are improved.

[0165] In practical engineering implementation, this bendable transition area can be achieved by integrating a flexible skin structure internally or on the surface. The deformation drive can be a shape memory alloy (SMA) actuator or a multi-link mechanism driven by a micro hydraulic / motor. When the control system issues a deflection command, the drive mechanism actuates, causing the head to deflect. At the same time, the flexible skin or deformable mechanism in the transition area moves in coordination to form a new continuous and smooth shape.

[0166] Step 5: Numerical verification and experimental analysis of aerodynamic characteristics:

[0167] To verify the aerodynamic characteristics of the flexible nose supersonic aircraft designed in this invention under different flight conditions, this embodiment uses a combination of numerical simulation and experimental analysis to analyze the influence of the nose deflection angle θ on the lift, drag and aerodynamic stability of the aircraft.

[0168] (1) Numerical simulation setup and conditions:

[0169] In this embodiment, the aerodynamic characteristics of the aircraft are analyzed by numerical simulation under high Mach number conditions using CFD software. It is assumed that the aircraft flies at Mach number Ma=6, and different angles of attack α and deflection angles θ are set.

[0170] Simulation conditions: Mach number of the aircraft Ma=6, angle of attack α=0~10°, deflection angle θ=-30~30°;

[0171] By simulating the changes in aerodynamic torque under different deflection angles θ, the ability of the bendable transition section to adjust the aerodynamic characteristics of the aircraft is verified.

[0172] (2) Relationship between pitching moment and angle of attack:

[0173] like Figure 6 As shown, the pitching moment coefficient C of the aircraft at Mach number Ma=6. m The relationship between the angle of attack α and the pitching moment coefficient C is shown. The results indicate that as the deflection angle θ increases, the pitching moment coefficient C of the aircraft also increases. m The changes are significant, indicating the impact of the flexible transition section on the longitudinal stability of the aircraft.

[0174] For example, when θ = 15°, the pitch moment coefficient C m The significant increase compared to θ=0° demonstrates that the bendable transition section effectively improves the longitudinal stability of the aircraft by changing its aerodynamic shape.

[0175] (3) Pressure distribution and aerodynamic performance:

[0176] like Figure 7 and Figure 8 As shown, the pressure distribution contour maps of the upper and lower surfaces of the aircraft are displayed at Mach number Ma=6 and angle of attack α=4°. Simulation results show that there are significant differences in the pressure distribution between the upper and lower surfaces of the aircraft under different deflection angles θ.

[0177] When the deflection angle θ = 0°, the pressure distribution of the aircraft is relatively symmetrical, exhibiting high drag.

[0178] When the deflection angle θ = 15°, the pressure distribution on the upper surface of the aircraft is significantly reduced, drag is reduced, and lift is increased, exhibiting better aerodynamic performance.

[0179] By comparing pressure cloud maps, we can verify the dynamic adjustment effect of the nose deflection angle θ on the aerodynamic performance of the aircraft, especially the optimization of lift-drag characteristics at high Mach numbers.

[0180] (4) Aerodynamic stability and handling performance:

[0181] According to the numerical simulation results, the aerodynamic stability and handling performance of the aircraft show significant differences with changes in the deflection angle θ. At larger deflection angles θ, the longitudinal stability of the aircraft is effectively improved, while the aerodynamic response becomes more sensitive and the adaptability is stronger.

[0182] Further analysis shows that when the aircraft flies at different angles of attack α, the deflection angle θ has a significant impact on the longitudinal stability margin, and can significantly reduce the aerodynamic coupling problem caused by traditional rigid structures.

[0183] (5) Verification conclusion:

[0184] Based on the above numerical simulations and experimental analyses, the following conclusions are drawn:

[0185] The flexible transition section can effectively optimize the aerodynamic performance of the aircraft by adjusting the deflection angle θ: under different flight conditions, by adjusting the deflection angle, the aircraft can dynamically adjust between lift and drag to achieve ideal aerodynamic performance.

[0186] The flexible transition section has a significant impact on the aerodynamic stability of the aircraft: changes in the deflection angle θ can effectively improve the longitudinal stability of the aircraft and reduce aerodynamic coupling effects.

[0187] The aerodynamic response and control performance of the aircraft are significantly improved: as the deflection angle θ increases, the control capability of the aircraft is enhanced, especially under supersonic conditions, the maneuverability and attitude adjustment capability of the aircraft are significantly improved.

[0188] Although embodiments of the present invention have been shown and described above, it is understood that the above embodiments are exemplary and should not be construed as limiting the present invention. Those skilled in the art can make changes, modifications, substitutions and variations to the above embodiments within the scope of the present invention without departing from the principles and spirit of the present invention.

Claims

1. An aerodynamic shape for a supersonic aircraft with a bendable nose, characterized in that: include: The waverider head is used to control the shock wave structure of the forebody and generate compressive lift under supersonic flight conditions. The lifting body fuselage provides the main volume and main lift of the aircraft; A flexible transition region connects the waverider head and the lifting body fuselage, enabling continuous deflection of the waverider head relative to the lifting body fuselage within the aircraft's plane of symmetry, and maintaining aerodynamic continuity and surface smoothness between the waverider head and the lifting body fuselage under different deflection states. The deflection angle of the waverider head is configured to be adjustable to change the pressure distribution and aerodynamic torque distribution in the forebody region of the aircraft.

2. The aerodynamic shape of the supersonic aircraft according to claim 1, characterized in that: The geometry of the bendable transition region is constructed by a parametric Bezier surface, the position of the control points of which is dynamically adjusted as the deflection angle changes to satisfy the boundary continuity condition.

3. The aerodynamic shape of the supersonic aircraft according to claim 2, characterized in that: The boundary continuity conditions include: at least first-order geometric continuity or second-order geometric continuity must be satisfied at the connection between the waverider head and the flexible transition region, and at the connection between the flexible transition region and the lifting body fuselage.

4. The aerodynamic shape of the supersonic aircraft according to claim 1, characterized in that: The waverider head is designed based on the cone-guided waverider theory, with a design Mach number range of Ma = 4 to Ma = 7, and the leading edge of the waverider head is passivated.

5. The aerodynamic shape of the supersonic aircraft according to claim 1, characterized in that: The cross-sectional profile of the lifting body fuselage at any axial position is composed of an upper curved segment, a middle transition curved segment, and a lower curved segment, and each cross-section gradually widens along the axial direction.

6. The aerodynamic shape of the supersonic aircraft according to claim 1, characterized in that: The adjustment of the deflection angle of the waverider head enables the aircraft to achieve a pitching moment coefficient change of no less than 0.02 when the deflection angle changes from 0° to 15° under the conditions of Ma=6 and angle of attack4°, and the peak pressure coefficient of the upper surface of the aircraft is reduced by no less than 15%.

7. A method for designing the aerodynamic shape of a supersonic aircraft as described in any one of claims 1-6, characterized in that: The method includes the following steps: Step 1: Overall configuration parameterization definition: Establish the spatial coordinate system of the aircraft's aerodynamic shape, divide the aerodynamic shape into the waverider head, the lifting body fuselage, and the flexible transition region connecting the two, and define the range of values ​​for the head deflection angle. Step 2: Constructing the aerodynamic shape of the lifting body fuselage: Using a multi-segment curve fusion section parameterization construction method, the three-dimensional aerodynamic shape of the lifting body fuselage is constructed by defining the geometric parameters of multiple control sections and performing surface interpolation along the axial direction. Step 3: Construct the aerodynamic shape of the waverider head: Based on the theory of conical waveriders, the shock surface is established under the given design Mach number. The surface of the waverider head is generated by defining the upper edge line and tracing the streamlines along the incoming flow direction, and the leading edge is passivated. Step 4: Construct the aerodynamic shape of the flexible transition region: Based on the constructed waverider head and lifting body fuselage, the head deflection angle is introduced as a control parameter. The head deflection is achieved through rigid body geometric transformation, and a continuous and smooth transition surface is constructed between the two boundaries using parametric Bezier curves or surfaces. Step 5: Aerodynamic characteristic parameter response and verification: Based on the constructed aerodynamic shape, the aerodynamic force and aerodynamic torque characteristics of the aircraft are analyzed under different incoming Mach numbers Ma, angle of attack α and nose deflection angle θ, and the effect of nose deflection on the forebody flow field structure and aerodynamic characteristics is verified.

8. The method according to claim 7, characterized in that, In step 2, at least three control sections are set along the axial direction of the aircraft. The upper curve characteristic radius and the lower curve characteristic radius of each control section increase monotonically along the axial direction to form an aerodynamic shape that gradually widens along the axial direction.

9. The method according to claim 7, characterized in that, In step 4, the control point position of the Bezier surface is dynamically calculated based on the deflection angle using an interpolation function to ensure that the geometry of the bendable transition region evolves continuously with the deflection angle and always satisfies the boundary matching conditions with the waverider head and the lifting body fuselage.

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