Rear edge variable camber airfoil profile composed of bistable composite material structure and design and control method of rear edge variable camber airfoil profile

Abstract

The invention provides a trailing edge variable camber airfoil composed of a bistable composite material structure and a design and control method of the trailing edge variable camber airfoil, and relates to the technical field of aircraft design and aerodynamics. Wherein the trailing edge variable camber airfoil sequentially comprises a leading edge section, an airfoil middle section and a trailing edge section in the chordwise direction, a front beam is arranged between the leading edge section and the airfoil middle section, the trailing edge section is fixed to an airfoil rear beam, and the trailing edge section is a core function section and is formed by arranging multiple layers of glass / carbon hybrid bistable composite material structures from top to bottom in the wingspan direction; each glass / carbon hybrid bistable composite material structure has a first stable state and a second stable state which correspond to a first flight state and a second flight state of the airfoil profile respectively, one end of each glass / carbon hybrid bistable composite material structure is fixedly connected to the airfoil profile rear beam, and the other end of each glass / carbon hybrid bistable composite material structure is fixed into a whole through glue joint. According to the invention, the steady-state characteristic of the bistable composite material is utilized, so that non-driven conversion between two stable bending states is realized.

Description

Technical Field

[0001] This application relates to the fields of aircraft design and aerodynamics, and in particular to a trailing edge variable camber airfoil composed of a bistable composite material structure, its design, and control methods. Background Technology

[0002] Variable camber airfoils can adapt to different flight conditions by changing their airfoil shape, thereby improving the overall aerodynamic efficiency of an aircraft. There are two main traditional methods for achieving variable camber airfoils by altering the trailing edge shape and position: (1) Articulated control surfaces: flaps, ailerons, etc. with independent actuators (such as servos, hydraulic actuators) are set on the trailing edge of the airfoil. This scheme relies on mechanical joints, bearings and connecting rods to achieve rotation. It has a complex structure, heavy weight, and gaps, and cannot adapt to continuous deformation according to real-time airflow conditions.

[0003] (2) Flexible skin / flexible airfoil: Flexible materials or structures are used to achieve smooth and continuous deformation of the airfoil surface. Although this solution has the advantages of lightweight and adaptability, it usually has limited load-bearing capacity, complex actuator integration, and fatigue problems after long-term use.

[0004] In recent years, achieving airfoil deformation using bistable composite material structures has become an emerging technology. Bistable composite laminates, under specific layup conditions, possess two stable geometric configurations and can switch between these two stable states without requiring external energy; they only need to overcome a small energy barrier.

[0005] For example, patent application number 202311775012.7 discloses a trailing edge structure and design method of a variable camber wing based on multi-knot rotation. This solution adopts the design of an articulated control surface, which requires components such as mechanical joints, bearings, and movable parts. These components conflict with the requirements of lightweight design and cannot adapt to deformation.

[0006] Patent application number 201410677233.5 discloses an airfoil with an adaptive variable camber trailing edge. Although this design has the advantages of being lightweight and adaptive, its load-bearing capacity is still relatively insufficient.

[0007] Patent application number 202210410689.X discloses a trailing edge variable camber wing driven by a multistable composite material structure. This design uses a bistable composite material structure as the driving mechanism to drive trailing edge deformation, but it requires a buffer notch, resulting in a complex structure and difficulty in ensuring sealing. Furthermore, using adhesive connectors to fill the buffer notch causes bulges to form on the wing skin, creating localized turbulence. As can be seen from the above, existing technical solutions typically arrange a series of bistable composite material structures with the same material and ply density side-by-side on the trailing edge of the airfoil. In these solutions, the curvature change of each bistable structure is consistent in both steady states. However, the curvature change of each bistable composite material structure is inconsistent under different airfoil states. Therefore, a free space needs to be reserved at one end of the trailing edge structure to ensure the steady-state transition of each bistable composite material structure, which leads to complex manufacturing processes and makes it impossible to seal the airfoil. Summary of the Invention

[0008] This application provides a trailing-edge variable camber airfoil composed of a bistable composite material structure, along with its design and control methods. The aim is to provide an airfoil with a trailing-edge variable camber composed of a bistable composite material structure that requires no additional drive, has high load-bearing capacity, is lightweight, has a smooth aerodynamic shape, and whose trailing-edge ends can be fixed. Simultaneously, a design method is provided to customize the parameters of each bistable composite material structure at the trailing edge of this airfoil, ensuring stable operation and efficient transitions between two flight states.

[0009] In a first aspect, this application provides a trailing edge variable camber airfoil composed of a bistable composite material structure, which includes a leading edge section, a mid-wing section and a trailing edge section in sequence along the chord direction. A front spars are provided between the leading edge section and the mid-wing section, and the trailing edge section is fixed to the airfoil's rear spars. The trailing edge section is the core functional section and is composed of a multilayer glass / carbon hybrid bistable composite material structure arranged from top to bottom along the wingspan direction. Each of the glass / carbon hybrid bistable composite material structures has a first stable state and a second stable state, which correspond to the first flight state and the second flight state of the airfoil, respectively. One end of each glass / carbon hybrid bistable composite material structure is fixedly connected to the rear spars of the airfoil, and the other end is fixed as a whole by adhesive bonding. The first flight state, such as hovering, corresponds to a trailing edge deflection angle of 0°; the second flight state, such as forward flight, corresponds to a trailing edge deflection angle of 0°. θ °.

[0010] In one possible design, the centerline of the first stable section of the glass / carbon hybrid bistable composite material structure is an arc, the centerline of the second stable section is a straight line, and the curvature variation of the glass / carbon hybrid bistable composite material structure at different positions is customized according to its installation position.

[0011] In one possible design, the curvature change of the trailing edge segment is triggered by a trailing edge distributed aerodynamic load. When the aerodynamic load exceeds the bistable jump critical load, all glass / carbon hybrid bistable composite material structures synchronously and automatically jump to the corresponding steady state, and can stably bear the maximum aerodynamic load in the corresponding steady state in both steady states.

[0012] Secondly, this application provides a control method for a trailing-edge variable camber airfoil composed of a bistable composite material structure as described in the first aspect and any possible design therein, the control method comprising: Automatic transition between two steady states is achieved through variations in trailing-edge distributed aerodynamic loads, where: In the first flight state, all glass / carbon hybrid bistable composite material structures are in the first stable state, and the airfoil exhibits a camber of 0° corresponding to the trailing edge deflection angle. When changes in flight conditions lead to changes in the aerodynamic load distribution acting on the trailing edge, and the aerodynamic load exceeds the critical load for bistable transition, all glass / carbon hybrid bistable composite material structures synchronously and automatically transition to the second stable state, and the airfoil exhibits the corresponding trailing edge deflection angle. θ A curvature of °, and maintain a stable state; When the flight state changes in the opposite direction, and the aerodynamic load direction or magnitude changes to another critical load, all glass / carbon hybrid bistable composite material structures synchronously and automatically jump back to the first stable state, completing the reverse conversion.

[0013] In one possible design, the critical load for the bistable transition is preset by customizing the interlaminar hybrid parameters and curvature variation of each glass / carbon hybrid bistable composite material structure.

[0014] Thirdly, this application provides a design method for a trailing-edge variable camber airfoil composed of a bistable composite material structure as described in the first aspect and any possible design thereof, characterized in that the design method includes: Calculate the aerodynamic loads on the trailing edge structure; Based on the established overall trailing edge stiffness calculation model, the equivalent bending stiffness is calculated, and the trailing edge deflection angle is calculated based on the equivalent bending stiffness. The radius of curvature of each layer of the bistable composite material structure is calculated based on the assumption that the centerline of the second steady-state section of each layer is a straight line, the centerline of the first steady-state section is a circular arc with constant curvature, and the displacement and deflection angle of the free ends of each layer are the same. The radius of curvature of the first steady state is then calculated by inverse equations. Calculate the interlaminar hybrid parameters of glass / carbon fiber hybrid bistable composite structures.

[0015] In one possible design, the aerodynamic load on the trailing edge structure is calculated using the following formula: In the formula, F a For the aerodynamic loads on the surface of the trailing edge section structure. p ( x , y ) is a surface structure point ( x , y The pressure component of ) τ( x , y ) is a surface structure point ( x , y The viscous shear stress tensor of ) x , y These are the coordinates of surface structure points in the trailing edge segment structure. n For the surface normal vector, S Let be the surface area of ​​the trailing edge segment.

[0016] In one possible design, based on the established trailing edge global stiffness calculation model, the equivalent bending stiffness is calculated using the following formula: In the formula, The equivalent bending stiffness of the trailing edge segment structure; Number of floors; For the first i The equivalent elastic modulus of the composite material along its length; For the first i The moment of inertia of a section about its own centroidal axis; For the first i Layer cross-sectional area; For the first i The distance from the centroid of the layer to the reference axis; The position of the neutral axis of the entire cross section; Based on the equivalent bending stiffness, the trailing edge deflection angle is calculated using the following formula. θ : In the formula, The arc length of the centerline of the rear edge section; This represents the bending moment distribution at each section of the trailing edge segment; For the aerodynamic load on the surface of the trailing edge section structure; EI ) eq This is the equivalent bending stiffness of the trailing edge segment.

[0017] In one possible design, the system of equations is expressed as: In the formula, The displacement of the free end of each layer of the bistable composite material structure in the trailing edge segment; This represents the displacement of the free end of the neutral layer in the trailing segment; θ i The free end deflection angle of each layer of the bistable composite material structure in the trailing edge segment; L i Let be the arc length of the centerline of the cross-section of each layer of the bistable composite material structure in the rear edge segment, which is a known term in the second steady state; R iThe radius of curvature of the centerline of the first steady-state section of each layer of the bistable composite material structure in the trailing edge segment; R 0 represents the radius of curvature of the centerline of the first steady-state section of the neutral layer in the trailing edge segment; θ The overall deflection angle of the trailing edge segment structure.

[0018] In one possible design, the interlaminar hybrid parameters of the glass / carbon fiber hybrid bistable composite structure are calculated as follows: Establish a steady-state configuration prediction model and solve for the interlayer hybridity parameters using the following formula: In the formula, The difference between the curing temperature and the operating temperature; w i The weighting is affected by the layer sequence type; For the thickness of carbon fiber and glass fiber layup; It represents the elastic modulus in the fiber direction of the unidirectional layup of carbon fiber and glass fiber; is the coefficient of thermal expansion in the fiber direction of the unidirectional layup of carbon fiber and glass fiber; k The curvature of the centerline of the first steady-state section of the bistable composite material structure; R The radius of curvature of the centerline of the first steady-state section of the bistable composite material structure; The interlayer hybridization parameter includes the thickness ratio of the carbon fiber layup to the glass fiber layup. And the layer sequence type.

[0019] This application provides a trailing-edge variable camber airfoil composed of a bistable composite material structure, its design, and a control method thereof, which have at least the following beneficial effects: 1. Driveless Dual-Switch: Utilizing the steady-state characteristics of bistable composite materials, it achieves a driveless transition between two stable camber states and can be passively locked in either state, capable of withstanding the maximum aerodynamic load perceived in that state, with a load-bearing capacity superior to traditional flexible airfoils.

[0020] 2. Significant lightweight advantages: By eliminating heavy mechanical components such as hinges, bearings, and actuators, and using lightweight composite materials, the weight of the airfoil trailing edge structure is greatly reduced.

[0021] 3. Excellent aerodynamic performance: The trailing edge curvature changes smoothly and continuously, avoiding airflow separation and interference caused by traditional articulated control surfaces, and can provide better lift-to-drag ratio and aerodynamic efficiency.

[0022] 4. Improved state transition performance: During the transition between two flight states, the trailing edge can quickly and synchronously adapt and jump, optimizing the aerodynamic performance during the transition phase.

[0023] 5. Solving sealing and process challenges: By customizing the curvature of each bistable structure, all structures can still work in coordination under the constraint of fixed ends, achieving complete sealing of the trailing edge, simplifying manufacturing and assembly processes, and improving the overall integrity and reliability of the structure. Attached Figure Description

[0024] The accompanying drawings, which are incorporated in and form part of this specification, illustrate embodiments consistent with this application and, together with the description, serve to explain the principles of this application.

[0025] Figure 1 This is a schematic diagram of a trailing-edge variable camber airfoil composed of a bistable composite material structure, provided in an embodiment of this application. Figure 2 A schematic diagram of a trailing-edge variable camber airfoil composed of a bistable composite material structure provided in this application when the trailing-edge deflection angle is 0°. Figure 3 An embodiment of this application provides a trailing-edge variable camber airfoil composed of a bistable composite material structure with a trailing-edge deflection angle of... θ Schematic diagram of the structure at °; Figure 4 A flowchart illustrating a design method for a trailing edge variable camber airfoil composed of a bistable composite material structure, as provided in this application embodiment.

[0026] Explanation of reference numerals in the attached figures: 1. Leading edge section; 2. Mid-wing section; 3. Trailing edge section; 4. Front spars; 5. Glass / carbon hybrid bistable composite material structure; 6. Rear spars.

[0027] The accompanying drawings illustrate specific embodiments of this application, which will be described in more detail below. These drawings and descriptions are not intended to limit the scope of the concept in any way, but rather to illustrate the concept of this application to those skilled in the art through reference to particular embodiments. Detailed Implementation

[0028] Exemplary embodiments will now be described in detail, examples of which are illustrated in the accompanying drawings. When the following description relates to the drawings, unless otherwise indicated, the same numbers in different drawings denote the same or similar elements. The embodiments described in the following exemplary embodiments do not represent all embodiments consistent with this application. Rather, they are merely examples of apparatuses and methods consistent with some aspects of this application as detailed in the appended claims.

[0029] The collection, storage, use, processing, transmission, provision, and disclosure of relevant data and information in the technical solution of this application all comply with the provisions of relevant laws and regulations and do not violate public order and good morals.

[0030] It should be noted that in the embodiments of this application, certain software, components, models and other existing solutions in the industry may be mentioned. These should be regarded as exemplary and are only intended to illustrate the feasibility of implementing the technical solution of this application. However, it does not mean that the applicant has used or necessarily used the solution.

[0031] The technical solution of this application and how the technical solution of this application solves the above-mentioned technical problems are described in detail below with specific embodiments. These specific embodiments can be combined with each other, and the same or similar concepts or processes may not be described again in some embodiments. The embodiments of this application will now be described with reference to the accompanying drawings.

[0032] Example 1: This application provides a trailing edge variable camber airfoil composed of a bistable composite material structure, its design, and a control method thereof. This embodiment aims to illustrate the basic implementation principle of the above-mentioned airfoil, its design, and the control method.

[0033] One aspect of this application provides a trailing-edge variable camber airfoil composed of a bistable composite material structure, such as... Figures 1 to 3 As shown, the variable camber airfoil with a trailing edge composed of a bistable composite material structure is entirely made of composite materials. Along the chord direction, it includes a leading edge section 1, a mid-wing section 2, and a trailing edge section 3. A front sparsity 4 is provided between the leading edge section 1 and the mid-wing section 2. The trailing edge section 3 is fixed to the airfoil's rear sparsity 6 and is the core functional section, composed of multiple layers of glass / carbon hybrid bistable composite material structures 5 arranged from top to bottom along the wingspan. Each glass / carbon hybrid bistable composite material structure 5 has a first stable state and a second stable state, corresponding to the airfoil's first and second flight states, respectively. One end of each glass / carbon hybrid bistable composite material structure 5 is fixedly connected to the airfoil's rear sparsity 6, and the other end is glued together. The first flight state, such as hovering, corresponds to a trailing edge deflection angle of 0°, and the second flight state, such as forward flight, corresponds to a trailing edge deflection angle of 0°. θ °.

[0034] In this embodiment, the entire airfoil requires no additional drive structure. The trailing edge segment 3 is the core of the airfoil, composed of multiple layers of glass / carbon hybrid bistable composite material structures arranged from top to bottom along the wingspan. Each structure has a first stable configuration and a second stable configuration, corresponding to the airfoil's first flight state (e.g., hovering, trailing edge deflection angle of 0°) and second flight state (e.g., forward flight, trailing edge deflection angle of 0°). θ °).

[0035] Unlike bistable composite structures made from a single composite material, this embodiment employs a glass / carbon hybrid bistable composite structure. The designability of the trailing edge bistable composite structure can be improved by adjusting the hybrid parameters, allowing the curvature variation under different steady states to be specifically designed based on its position in the wingspan direction. Furthermore, each piece of the glass / carbon hybrid bistable composite structure is fixedly connected at one end to the airfoil's rear sparsity 6, and the other end is glued together, ultimately forming a complete, gapless, and sealable trailing edge assembly.

[0036] In some embodiments, the first stable state of the glass / carbon hybrid bistable composite material structure 5 is a circular arc structure, the second stable state is a straight line structure, and the curvature variation of the glass / carbon hybrid bistable composite material structure 5 at different wingspan positions is customized according to its installation position.

[0037] The glass / carbon hybrid bistable composite structure 5 possesses a first steady-state circular arc configuration and a second steady-state linear configuration, which are key structural features for achieving subsequent adaptive deformation of the airfoil. This bistable characteristic originates from the residual stress distribution caused by the material layup design, enabling the structure to stably maintain two distinctly different geometric shapes when no external forces are applied. More importantly, the curvature variation of structural units at different wingspan positions is not uniformly set, but customized according to the aerodynamic environment and structural stress characteristics at each position. For example, in the high dynamic pressure region near the blade tip, the curvature variation of the structural unit is designed to be larger to provide sufficient camber change and resist stronger aerodynamic loads; while in the inner region, the curvature variation is relatively smaller. This differentiated design ensures that all structural units can still coordinately complete the steady-state transition under the constraint that both ends are fixed, thereby achieving seamless and sealed integration of the entire structure, effectively avoiding the aerodynamic performance degradation and process complexity problems caused by reserved movement gaps in traditional solutions.

[0038] In some embodiments, the curvature change of the trailing edge segment 3 is triggered by the trailing edge distributed aerodynamic load. When the aerodynamic load exceeds the bistable jump critical load, all glass / carbon hybrid bistable composite material structures 5 automatically jump to the corresponding steady state synchronously, and can stably bear the maximum aerodynamic load in the corresponding steady state in both steady states.

[0039] The camber change mechanism of subsequent segment 3 relies entirely on distributed aerodynamic load triggering, embodying the passive adaptive control principle of the airfoil structure. When the flight state changes, the distribution of aerodynamic loads acting on the subsequent surface changes accordingly. Once the local load exceeds the jump threshold of a certain glass / carbon hybrid bistable composite material structure 5, that unit will be the first to undergo a steady-state transition. This transition immediately changes the local flow field and load distribution, triggering a chain reaction in adjacent units, ultimately driving all structural units to synchronously and automatically jump to another stable state in a very short time. The entire transition process requires no external actuator or complex control system intervention, achieving truly driverless, adaptive deformation. Furthermore, after completing the transition, the structure, thanks to the inherent self-locking properties of the bistable material, will stably maintain the new geometric configuration and be able to withstand the maximum aerodynamic load that may occur in this steady state. Thus, while ensuring lightweight design, it possesses superior load-bearing capacity and structural reliability compared to traditional flexible airfoils.

[0040] One aspect of this application provides a control method for a trailing-edge variable camber airfoil composed of a bistable composite material structure. The control method includes: automatically switching between two steady states through changes in distributed aerodynamic loads at the trailing edge, wherein: in a first flight state, all glass / carbon hybrid bistable composite material structures are in the first steady state, and the airfoil exhibits a camber corresponding to a trailing-edge deflection angle of 0°; when a change in flight state causes a change in the aerodynamic load distribution acting on the trailing-edge segment, and the aerodynamic load exceeds the critical load for bistable state transition, all glass / carbon hybrid bistable composite material structures synchronously and automatically transition to the second steady state, and the airfoil exhibits a camber corresponding to a trailing-edge deflection angle. θ The structure maintains a 30° bend and remains stable. When the flight state changes in the opposite direction, or the aerodynamic load direction or magnitude changes to another critical load, all glass / carbon hybrid bistable composite material structures automatically jump back to the first stable state, completing the reverse conversion.

[0041] Specifically, the airfoil at 0° and θ The airfoil remains stable at trailing edge deflection angles of °, corresponding to two flight states. The trailing edge camber change is triggered by distributed aerodynamic loads on the trailing edge. In the first flight state, all bistable structures on the trailing edge are in the first stable state, and the airfoil exhibits the first type of camber. When a change in flight state causes a change in the distribution of aerodynamic loads acting on the trailing edge, exceeding the critical load for bistable transition, all bistable structures simultaneously and automatically transition to the second stable state, causing the airfoil to exhibit the second type of camber, and stabilize in this new state. The reverse transition is similar. The entire process requires no external actuator intervention and is a passive adaptive process.

[0042] One aspect of this application provides a design method for a trailing-edge variable camber airfoil composed of a bistable composite material structure, such as... Figure 4As shown, the design method includes the following steps S10-S40.

[0043] S10: Calculate the aerodynamic loads on the trailing edge section structure.

[0044] In some embodiments, the aerodynamic loads on the trailing edge structure can be calculated in the following manner. Obtaining accurate aerodynamic loads under two flight conditions is the starting point and foundation of the entire design process. (The last sentence appears to be incomplete and possibly refers to a separate topic: trailing edge structure surface aerodynamic loads.) The expression is: In the formula, p For pressure components; τ It is the viscous shear stress tensor; n It is the surface normal vector; S The surface area of ​​the trailing edge segment. p ( x , y ) is a surface structure point ( x , y The pressure component of ) τ ( x , y ) is a surface structure point ( x , y The viscous shear stress tensor of ) x , y These are the coordinates of surface structure points in the trailing edge segment structure. n For the surface normal vector, S Let be the surface area of ​​the trailing edge segment.

[0045] S20: Based on the established overall stiffness calculation model of the trailing edge, calculate the equivalent bending stiffness, and calculate the trailing edge deflection angle based on the equivalent bending stiffness.

[0046] In some embodiments, the overall deflection angle of the trailing edge structure can be calculated as follows: First, a calculation model for the overall stiffness of the trailing edge, consisting of multiple layers of parallel-arranged bistable composite material structures, is established. This model includes the relationship between the overall bending stiffness and the stiffness of each layer, as well as the overall deflection angle of the trailing edge structure. Relationship with stiffness and load.

[0047] The relationship between the overall bending stiffness and the stiffness of each layer is expressed as follows: In the formula: The equivalent bending stiffness of the trailing edge segment structure; Number of floors; : No. i The equivalent elastic modulus of the composite material in the longitudinal (chordal) direction; : No. i Moment of inertia of a section about its own centroidal axis (per unit width in the width direction); : No. i Cross-sectional area per unit width; : No. i The distance from the centroid of the layer to the reference axis; : The position of the neutral axis of the entire cross section.

[0048] For the overall deflection angle of the trailing edge segment structure In relation to stiffness and load, the trailing edge can be simplified as a cantilever beam. Under aerodynamic loads, its tip deflection angle (i.e., the overall deflection angle) is... )for: (Simplified for uniformly distributed loads) In the formula: : Arc length of the centerline of the trailing edge section; : Bending moment distribution at each section of the trailing edge; Aerodynamic loads on the surface of the trailing edge section structure; ( EI ) eq : Equivalent bending stiffness of the trailing edge segment structure.

[0049] S30: Calculate the radius of curvature of each layer of the bistable composite material structure. Based on the assumption that the center line of the second steady-state section of each layer is a straight line, the center line of the first steady-state section is a circular arc with constant curvature, and the displacement and deflection angle of the free ends of each layer are the same, the radius of curvature of the first steady state is obtained by inverse equations.

[0050] In some embodiments, step S30 is implemented as follows: Calculate the radius of curvature of each layer of the bistable composite material structure. First, each layer of the bistable composite material structure is equivalent to a curve (i.e., the centerline of the cross section, where the second stable state is a straight line and the first stable state is a circular arc). Based on this assumption, a calculation model for the first stable state radius of curvature of each layer of the bistable composite material structure in the trailing edge segment is established. Then, based on the overall deflection angle of the trailing edge segment structure, the radius of curvature of the circular arc of each bistable composite material structure in the first stable state is calculated.

[0051] Basic assumptions: (1) Each layer is a straight line (curvature) in the second steady state. (2) The arc of each layer with constant curvature in the first steady state (curvature) (3) Each layer is bonded to a single point at the beam end (free end), therefore each layer's free end has the same displacement and deflection angle, i.e. , .

[0052] Using the neutral layer of the trailing edge segment as a reference, the calculation model (i.e., the system of equations) for the first steady-state radius of curvature of each layer of the bistable composite material structure in the trailing edge segment is expressed as follows: The displacement of the free end of each layer of the bistable composite material structure in the trailing edge segment; This represents the displacement of the free end of the neutral layer in the trailing segment; θ i The free end deflection angle of each layer of the bistable composite material structure in the trailing edge segment; L i Let be the arc length of the centerline of the cross-section of each layer of the bistable composite material structure in the rear edge segment, which is a known term in the second steady state; R i The radius of curvature of the centerline of the first steady-state section of each layer of the bistable composite material structure in the trailing edge segment; R 0 represents the radius of curvature of the centerline of the first steady-state section of the neutral layer in the trailing edge segment; θ The overall deflection angle of the trailing edge segment structure.

[0053] S40: Calculate the interlaminar hybrid parameters of glass / carbon fiber hybrid bistable composite structures.

[0054] Specifically, firstly, a steady-state configuration prediction model for the glass / carbon fiber hybrid bistable composite structure is established; then, for each bistable composite structure, based on the change in radius of curvature under the two steady states and material properties (elastic modulus of glass fiber and carbon fiber, Poisson's ratio, etc.), the interlaminar hybridization parameters are calculated and determined. These interlaminar hybridization parameters include: the thickness ratio of carbon fiber layup to glass fiber layup and the layup sequence. By differentially designing the hybridization parameters of each bistable composite structure, its bistable transition behavior is made to precisely match the target radius of curvature change calculated in step S30. Step S40 specifically includes the following steps S401-S403: S401: Determine the formula for the residual bending moment of an asymmetric orthogonal ply laminate. For asymmetric plies, the thermal expansion coefficients of the different plies after curing and cooling are significant. Different, resulting in residual bending moment : After simplification, for two layers of different materials (thickness) , Modulus , coefficient of thermal expansion , The asymmetric orthogonal ply of ) with stable curvature It can be approximated as: In the formula: The difference between the curing temperature and the working temperature; w i The type of layup order affects the weight; : Thickness of carbon fiber and glass fiber layup; : Elastic modulus in the fiber direction of unidirectional layup of carbon fiber and glass fiber; : The coefficient of thermal expansion in the fiber direction of unidirectional layup of carbon fiber and glass fiber.

[0055] S402: Design equation.

[0056] For the Layer, its first steady-state curvature Second steady-state curvature .

[0057] The design objective is to select material parameters and thickness ratio. And the type of layup sequence, so that the curvature predicted by the model equals the target value. .

[0058] S403: Determine hybridization parameters: thickness ratio The layup sequence type is obtained by solving the above equations.

[0059] Example 2: Based on Example 1, this application provides a specific application case of a trailing edge variable camber airfoil composed of a bistable composite material structure, as well as its design and control method, which is claimed in this application.

[0060] Using a medium-span composite main rotor blade as the implementation object, based on a trailing edge variable camber airfoil composed of a bistable composite material structure provided in Example 1, the trailing edge portion of the outer side of the blade (approximately 30% of the span from the blade tip inward) is modified.

[0061] The trailing edge variable camber section is designed as follows: This trailing edge consists of an eight-layer bistable composite material structure, with each layer consisting of a glass / carbon hybrid bistable composite material structural piece as a deformable element. These structural pieces are arranged side by side along the spanwise direction. One end of each glass / carbon hybrid bistable composite material structure is fixedly connected to the airfoil's rear sparsity, and the other end is glued together to form a sealed trailing edge section integrated with the main blade.

[0062] The bistable design is as follows: First steady state (hovering): All bistable structural pieces are in a low curvature state, so that the entire trailing edge section is smoothly connected to the arc of the main blade, forming a low-drag, high-efficiency airfoil (equivalent trailing edge deflection angle is about 0°).

[0063] Second steady state (forward flight): All bistable structural pieces transition to a high curvature state, causing the trailing edge section to deflect downwards as a whole, forming an airfoil with a high lift coefficient (trailing edge deflection angle is...). θ °).

[0064] Unlike existing bistable trailing edge designs that assume uniform curvature, in this embodiment, the target curvature change (Δκ) of the eight bistable structures from the propeller root to the propeller tip is determined through rigorous aeroelastic calculations rather than a uniform design. For example, the unit closer to the propeller tip, with higher linear velocity, has a larger Δκ ​​to overcome stronger aerodynamic loads and generate the required camber; while the inner units have a relatively smaller Δκ. It is this customized curvature design that allows all units to coordinately transition from the first steady state to the second steady state under boundary conditions with fixed ends, solving sealing and manufacturing challenges.

[0065] The specific application of the design methodology in rotorcraft scenarios includes the following four steps.

[0066] Step 1: Calculate the aerodynamic loads on the trailing edge section of the helicopter main rotor blades. Calculated according to the following expression: .

[0067] Step 2: Calculate the overall trailing edge deflection angle of the helicopter main rotor blades in the second flight state. θ Calculated according to the following expression: In the formula, : 8 floors; : No. i The equivalent elastic modulus of the composite material in the longitudinal (chordal) direction; : No. i Moment of inertia of a section about its own centroidal axis (per unit width in the width direction); : No. i Cross-sectional area per unit width; : No. i The distance from the centroid of the layer to the reference axis; : The position of the neutral axis of the entire cross section; : Arc length of the centerline of the trailing edge section; : Bending moment distribution at each section of the trailing edge; Aerodynamic load on the surface of the trailing edge section structure.

[0068] Step 3: Calculate the first steady-state radius of curvature of each layer of bistable composite material structure at the trailing edge of the helicopter main rotor blade. Calculated according to the following expression: Step 4: Calculate the hybrid parameters of the glass / carbon fiber hybrid bistable composite structure according to the following expression: In the formula, The difference between the curing temperature and the operating temperature; w i The weighting is affected by the layer sequence type; For the thickness of carbon fiber and glass fiber layup; It represents the elastic modulus in the fiber direction of the unidirectional layup of carbon fiber and glass fiber; is the coefficient of thermal expansion in the fiber direction of the unidirectional layup of carbon fiber and glass fiber; k The curvature of the centerline of the first steady-state section of the bistable composite material structure; R The radius of curvature of the centerline of the first steady-state section of the bistable composite material structure; The airfoil designed using the above method is applicable to helicopter rotors, and its feasibility and unique advantages are mainly reflected in its ability to solve special rotor problems. Specifically, rotor blades operate in extreme environments, bearing enormous centrifugal forces, alternating aerodynamic loads, and vibrations. The reliability, sealing, and maintainability of traditional mechanically articulated trailing edge flaps under these conditions pose significant challenges. This embodiment employs an integrated composite material structure with fixed supports at both ends, fundamentally eliminating movement gaps, providing excellent sealing to prevent rainwater and dust intrusion, and withstanding high centrifugal forces. The self-locking characteristic of the bistable structure after a jump gives it high stiffness and strong load-bearing capacity in steady state, meeting the rotor's load requirements.

[0069] Targeted beneficial effects include: (1) Vibration and noise suppression: The smooth and continuous trailing edge camber variation avoids the pulse noise and vibration sources generated by airflow separation at the traditional flap hinge. This is of great significance for improving the helicopter's ride comfort and stealth capabilities.

[0070] (2) Adaptive performance enhancement: The trailing edge camber can be passively and adaptively adjusted according to the flight state (such as the need for high thrust for hovering and high efficiency for forward flight), which can optimize the aerodynamic efficiency of the entire flight envelope and significantly improve hovering efficiency or extend the cruise range.

[0071] (3) Extreme lightweight and high reliability: The complex flap actuator, rocker arm, hydraulic lines and other heavy mechanisms are completely eliminated, which significantly reduces the weight of the rotor hub and blades, which is extremely valuable for the rotor system. At the same time, the structure is greatly simplified, the potential failure points are reduced and the reliability is improved.

[0072] To verify the above embodiments, the following three-level experiment was designed: (1) Ground static jump and load test Procedure: Eight bistable specimens with different layup designs were prepared to simulate fixed boundary conditions at both ends. Loads were applied at specific locations using an actuator, and the critical jump load, jump process, and shape under both steady states were measured. Subsequently, a distributed static load simulating aerodynamic pressure was applied under both steady states to test the stiffness and ultimate bearing capacity.

[0073] Expected results: The data will confirm that each specimen can jump under different critical loads and withstand aerodynamic pressures exceeding the design value under two steady states, verifying the non-uniform design and high load-bearing capacity.

[0074] (2) Dynamic function test of the rotary table Procedure: A blade segment with an integrated bistable trailing edge was mounted on a rotating test bench. At different rotational speeds (simulating centrifugal force), the trailing edge was excited to jump using airflow impact or a small triggering device. The synchronicity, reliability, and structural integrity of the jump under centrifugal force were observed using a high-speed camera and strain gauges.

[0075] Expected results: Verify that the structure can still stably complete the jump and maintain its shape under high centrifugal force, proving that it is suitable for real rotor operating conditions.

[0076] (3) Comparative test of wind tunnel aerodynamic performance Procedure: A full-size blade segment model with a bistable trailing edge was constructed and tested in a low-speed wind tunnel. First, the lift and drag coefficients as a function of angle of attack were measured under two steady-state configurations. Second, during wind tunnel operation, a dynamic transition was triggered by changing the incoming flow angle of attack, and aerodynamic data during the transition process were measured. A comparative model with conventional articulated flaps was also set up.

[0077] The expected results and analysis are shown in Table 1 below.

[0078] Table 1 Expected Results and Analysis

[0079] The comparative test results shown in Table 1 clearly verify the multi-faceted performance advantages of the bistable trailing edge structure described in this application compared to the traditional hinged flap model. In cruise configuration, thanks to the smooth and continuous aerodynamic shape of the trailing edge, this application effectively reduces drag caused by surface friction and airflow interference, resulting in a lift-to-drag ratio comparable to or even slightly improved compared to traditional schemes. When the airfoil switches to a high-lift configuration, the continuously changing camber significantly delays airflow separation, thereby increasing the maximum lift coefficient and enhancing the aerodynamic performance of the aircraft during takeoff and landing. More importantly, this application achieves a smoother transition between the two flight states, with significantly reduced aerodynamic force and torque fluctuations. This is due to its continuous deformation mechanism, which avoids the sudden interruption and reattachment of local airflow caused by traditional hinged structures. Flow observation results provide direct evidence that the airflow adhesion on the trailing edge surface of this application is good, the separation point is pushed back, and the obvious vortices and flow separation zones present at the traditional flap hinge are completely eliminated, resulting in a substantial improvement in overall flow quality. In summary, these expected results not only confirm the excellent characteristics of the proposed method in terms of aerodynamic efficiency, smoothness of state transitions, and flow control, but also provide strong support for its application in the design of high-performance and high-reliability aircraft.

[0080] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of this application, and are not intended to limit them. Although this application has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that modifications can still be made to the technical solutions described in the foregoing embodiments, or equivalent substitutions can be made to some or all of the technical features therein. Such modifications or substitutions do not cause the essence of the corresponding technical solutions to deviate from the scope of the technical solutions of the embodiments of this application.

Claims

1. A trailing edge variable camber airfoil consisting of a bistable composite material structure, characterized in that, The chord direction includes a leading edge section (1), a wing middle section (2) and a trailing edge section (3) in sequence, the leading edge section (1) is provided with a front beam (4) between the wing middle section (2), the trailing edge section (3) is fixed to the airfoil rear beam (6), and the trailing edge section (3) is a core function section, which is composed of a plurality of glass / carbon hybrid bistable composite material structures (5) arranged from top to bottom along the span direction; Each glass / carbon hybrid bistable composite material structure (5) has a first stable state and a second stable state, corresponding to a first flight state and a second flight state of the airfoil respectively, and one end of each glass / carbon hybrid bistable composite material structure (5) is fixedly connected to the airfoil rear beam (6), and the other end is fixedly connected as a whole through gluing; The first flight state, such as hovering, corresponds to a trailing edge deflection angle of 0°, and the second flight state, such as forward flight, corresponds to a trailing edge deflection angle of The first stable state section center line of the glass / carbon hybrid bistable composite material structure (5) is a circular arc, the second stable state section center line is a straight line, and the curvature of the glass / carbon hybrid bistable composite material structure (5) at different positions is customized according to the installation position. °.

2. A trailing edge variable camber airfoil composed of bistable composite structures according to claim 1, characterized in that, The curvature change of the trailing edge section (3) is triggered by the trailing edge distributed aerodynamic load, when the aerodynamic load exceeds the bistable jump critical load, all glass / carbon hybrid bistable composite material structures (5) jump to the corresponding stable state synchronously and automatically, and can stably bear the maximum aerodynamic load in the two stable states.

3. The bistable composite structure composed trailing edge variable camber airfoil of claim 1, wherein, The control method comprises:

4. A method of controlling a trailing edge variable camber airfoil composed of the bistable composite structure according to any one of claims 1 to 3, characterized in that, The automatic conversion between the two stable states is realized by the change of the trailing edge distributed aerodynamic load, wherein: In the first flight state, all glass / carbon hybrid bistable composite material structures (5) are in the first stable state, and the airfoil presents a curvature corresponding to a trailing edge deflection angle of 0°. When the flight state reverses, the aerodynamic load direction or size changes to another critical load, all glass / carbon hybrid bistable composite material structures (5) jump back to the first stable state synchronously and automatically, and the reverse conversion is completed. When the flight state changes cause the aerodynamic load distribution acting on the trailing edge section (3) to change, and the aerodynamic load exceeds the critical load of the bistable jump, all the glass / carbon hybrid bistable composite structures (5) automatically jump to the second stable state simultaneously, the airfoil presents the camber corresponding to the trailing edge deflection angle The critical load of the bistable jump is preset by customizing the curvature change of each glass / carbon hybrid bistable composite material structure (5). . The design method comprises:

5. The control method according to claim 4, characterized by Calculate the aerodynamic load of the trailing edge section structure; 6. A method of designing a trailing edge variable camber airfoil consisting of a bistable composite structure according to any one of claims 1 to 3, characterized in that, Based on the established overall stiffness calculation model of the trailing edge, calculate the equivalent bending stiffness, and calculate the trailing edge deflection angle based on the equivalent bending stiffness; Calculate the curvature radius of each layer of bistable composite material structure, based on the assumption that the second stable state section center line of each layer is a straight line, the first stable state section center line is a constant curvature circular arc, and the free end displacement and deflection angle of each layer are the same, the first stable state curvature radius is solved by equation set; Calculate the interlaminar hybrid parameter of the glass / carbon fiber hybrid bistable composite material structure. The surface aerodynamic load of the trailing edge section structure is calculated by the following formula: Based on the established overall stiffness calculation model of the trailing edge, the equivalent bending stiffness is calculated by the following formula:

7. The method of designing according to claim 6, wherein, The equation set is represented as: wherein F a is the pressure component at the surface structure point p x y x y The interlaminar hybrid parameter of the glass / carbon fiber hybrid bistable composite material structure is calculated by the following method: x y x y x y is the coordinate of the surface structure point in the trailing edge segment structure n is the surface normal vector S is the surface area of the trailing edge segment​​​​​​​​​​​ 8. The method of claim 6, wherein, A stable state configuration prediction model is established, and the interlaminar hybrid parameter is solved according to the following formula: wherein, is the equivalent bending stiffness of the trailing edge segment structure; is the number of layers; is the equivalent elastic modulus of the composite material in the length direction of the layer; i is the equivalent elastic modulus of the composite material in the length direction of the layer; is the moment of inertia of the cross section of the layer about its centroidal axis; i is the moment of inertia of the cross section of the layer about its centroidal axis; is the cross-sectional area of the layer; i is the cross-sectional area of the layer; is the distance from the centroid of the layer to the reference axis; i is the distance from the centroid of the layer to the reference axis; is the position of the neutral axis in the entire cross section; Based on the equivalent bending stiffness, the trailing edge skew angle is calculated by the following equation ​ : wherein, is the center line arc length of the structural cross section of the trailing edge segment; is the bending moment distribution of each cross section of the trailing edge segment; is the structural surface aerodynamic load of the trailing edge segment; ​ eq is the equivalent bending stiffness of the trailing edge segment structure.​ 9. The method of claim 6, wherein, ​ wherein is the free end displacement of the bistable composite structure of the trailing edge segment for each layer; is the free end displacement of the neutral layer of the trailing edge segment; ​ i is the free end deflection angle of the bistable composite structure of the trailing edge segment for each layer; L i is the centerline arc length of the cross section of the bistable composite structure of the trailing edge segment for each layer, which is a known quantity in the second stable state; R i is the first stable state centerline radius of curvature of the cross section of the bistable composite structure of the trailing edge segment for each layer; R 0 is the first stable state centerline radius of curvature of the cross section of the neutral layer of the trailing edge segment; ​ is the overall deflection angle of the trailing edge segment structure.

10. The method of claim 6, wherein, ​ ​ wherein is the difference between the curing temperature and the working temperature; w i is the ply sequence type influence weight; is the carbon fiber and glass fiber ply thickness; is the elastic modulus of the carbon fiber and glass fiber unidirectional ply fiber direction; is the thermal expansion coefficient of the carbon fiber and glass fiber unidirectional ply fiber direction; k is the bistable composite structure first stable state cross section centerline curvature; R is the bistable composite structure first stable state cross section centerline curvature radius; The interlaminar hybrid parameters include a thickness ratio of carbon fiber plies to glass fiber plies and ply sequence type.

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