A method for designing a fixed leading edge structure of an integral detachable wing

Abstract

The application belongs to the field of aircraft structure design, and particularly relates to a kind of integral detachable wing fixed leading edge structure design method.The method comprises the following steps: step S1, determining fixed leading edge segmentation;Step S2, determine the form of the bulkhead structure;Step S3, determine whether to arrange the front wall;Step S4, determine the design of the fixed leading edge flap;Step S5, determine the docking form between the fixed leading edge and the peripheral structure and the adjacent two fixed leading edges;Step S6, determine the fixed leading edge structure parameters.The integral detachable wing fixed leading edge structure design method of the application is used to solve the problem of system installation and maintenance difficulty under the condition of low wing profile height and tight fixed leading edge space of small and medium-sized aircrafts;improve the system installation and maintenance efficiency, solidify the detachable fixed leading edge design process, so as to improve the design efficiency and design level.

Description

Technical Field

[0001] This application belongs to the field of aircraft structural design, and specifically relates to a design method for an integral detachable wing fixed leading edge structure. Background Technology

[0002] The fixed leading edge is the fixed section of the wing preceding the front spars and is attached to the front spars. For aircraft with slats, the fixed leading edge is located before the front spars and after the slats. Numerous systems, such as anti-icing systems, electrical systems, and hydraulic systems, are typically housed within the fixed leading edge. The fixed leading edge provides access for their installation and maintenance. For some small and medium-sized aircraft with low airfoil heights and extremely limited space on the fixed leading edge, system installation and maintenance are very difficult.

[0003] To address the aforementioned issues, this patent proposes a design method for an integrally detachable fixed leading edge structure for airfoils, which improves system installation and maintenance efficiency, solidifies the design process for detachable fixed leading edges, and thereby achieves the goal of improving design efficiency and design quality.

[0004] Therefore, there is an urgent need for a technical solution to overcome or mitigate at least one of the aforementioned defects in the existing technology. Summary of the Invention

[0005] The purpose of this application is to provide a design method for an integral detachable wing fixed leading edge structure to solve at least one problem existing in the prior art.

[0006] The technical solution of this application is:

[0007] A design method for an integrally detachable wing fixed leading edge structure includes:

[0008] Step S1: Determine the fixed leading edge segment;

[0009] Step S2: Determine the partition structure;

[0010] Step S3: Determine whether to install a front wall;

[0011] Step S4: Determine the design of the cap with a fixed leading edge;

[0012] Step S5: Determine the docking form between the fixed leading edge and the surrounding structure, as well as between two adjacent fixed leading edge segments;

[0013] Step S6: Determine the parameters of the fixed leading edge structure.

[0014] In at least one embodiment of this application, in step S1, the fixed leading edge segment is determined according to the size of the raw material, the position of the skeleton, and the position of the cap.

[0015] In at least one embodiment of this application, in step S1, the weight of each fixed leading edge segment does not exceed 13 kg and the length does not exceed 2.5 m.

[0016] In at least one embodiment of this application, in step S2, the end partitions at both ends of each fixed front edge are connected to the front beam, and the middle partition of each fixed front edge is not connected to the box segment.

[0017] In at least one embodiment of this application, in step S3, when the load is less than the load threshold, the front wall is not arranged, and the long truss is arranged at the original front wall position.

[0018] In at least one embodiment of this application, in step S4, the fixed leading edge is used as an integral cover, and a partial cover is arranged in the high maintenance frequency area.

[0019] In at least one embodiment of this application, in step S5, the fixed leading edge is not connected to the fuselage, and a 15mm gap is reserved between the fixed leading edge and the fuselage and sealed with rubber profiles.

[0020] In at least one embodiment of this application, in step S5, the fixed front edge is connected to the box segment by a mating belt plate, and the skin of the fixed front edge is connected to the mating belt plate by a support plate nut.

[0021] In at least one embodiment of this application, in step S5, the skins of two adjacent fixed leading edges are joined at the end partition, and the skins and the end partition are connected by a support plate nut.

[0022] In at least one embodiment of this application, step S6, determining the parameters of the fixed leading edge structure, includes:

[0023] Perform parameter design for the riveted skin of the long stringer;

[0024] Design the parameters of the partition.

[0025] In at least one embodiment of this application, the parameter design of the stringer riveted skin includes:

[0026] Determine the initial thickness of the skin, the initial cross-sectional parameters of the stringers, and the initial spacing of the stringers;

[0027] Calculate the local critical instability stress of the skin;

[0028] Calculate the critical stress for local buckling of the stringer;

[0029] Calculate the critical stress for stress loss of the stringer;

[0030] Calculate the critical stress for overall instability of the riveted skin of the stringer;

[0031] Calculate the safety margin based on each critical stress;

[0032] Adjust the skin thickness, stringer section parameters, and stringer spacing according to the safety margin, recalculate the safety margin based on the adjusted parameters, and iterate repeatedly until the requirements are met.

[0033] In at least one embodiment of this application, the parameters of the partition are designed, including:

[0034] Determine the initial dimensions of the partition;

[0035] Calculate the critical buckling stress of the web;

[0036] Calculate the critical buckling stress of the flange under compression;

[0037] Calculate the safety margin based on each critical stress;

[0038] Adjust the partition size according to the safety margin, recalculate the safety margin based on the adjusted parameters, and iterate repeatedly until the requirements are met.

[0039] The invention has at least the following beneficial technical effects:

[0040] The present application provides a design method for an integral detachable fixed leading edge structure for airfoils, which addresses the challenges of system installation and maintenance in small and medium-sized aircraft with low airfoil height and limited space at the fixed leading edge. This method improves system installation and maintenance efficiency, solidifies the design process for detachable fixed leading edges, and ultimately enhances design efficiency and quality. Attached Figure Description

[0041] Fig. 1 This is a flowchart of a design method for an integrally detachable wing fixed leading edge structure according to one embodiment of this application;

[0042] Fig. 2 This is a schematic diagram of an integral detachable wing fixed leading edge structure according to one embodiment of this application.

[0043] in:

[0044] 1-Skin; 2-Intermediate partition; 3-End partition; 4-Stringer; 5-Matching strip. Detailed Implementation

[0045] To make the objectives, technical solutions, and advantages of this application clearer, the technical solutions in the embodiments of this application will be described in more detail below with reference to the accompanying drawings. In the drawings, the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout. The described embodiments are some, but not all, embodiments of this application. The embodiments described below with reference to the accompanying drawings are exemplary and intended to explain this application, and should not be construed as limiting this application. All other embodiments obtained by those skilled in the art based on the embodiments of this application without creative effort are within the scope of protection of this application. The embodiments of this application will be described in detail below with reference to the accompanying drawings.

[0046] In the description of this application, it should be understood that the terms "center", "longitudinal", "lateral", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing this application and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limiting the scope of protection of this application.

[0047] The following is in conjunction with the appendix Figs. 1-2 This application will be described in further detail.

[0048] This application provides a design method for an integrally detachable wing fixed leading edge structure, such as... Fig. 1 As shown, it includes the following steps:

[0049] Step S1: Determine the fixed leading edge segment;

[0050] Step S2: Determine the partition structure;

[0051] Step S3: Determine whether to install a front wall;

[0052] Step S4: Determine the design of the cap with a fixed leading edge;

[0053] Step S5: Determine the docking form between the fixed leading edge and the surrounding structure, as well as between two adjacent fixed leading edge segments;

[0054] Step S6: Determine the parameters of the fixed leading edge structure.

[0055] Specifically, in step S1, the segmentation of the fixed leading edge is determined based on the material size, frame position, and cap position. When segmenting the fixed leading edge in a conventional manner, the material size and frame position are the primary considerations. For a fully detachable wing, in addition to the above factors, the ease of use of the fixed leading edge as a cap must also be considered. If a single segment of the fixed leading edge is too long, it will increase the weight of each segment and the number of fasteners that need to be removed during disassembly, undoubtedly increasing the workload and construction difficulty. Conversely, if there are too many segments, it will increase the mating surfaces, thus increasing the overall weight. Therefore, an optimal solution must be found between weight and ease of use.

[0056] In a preferred embodiment of this application, the weight of each fixed front edge segment that is disassembled individually does not exceed 13 kg and the length does not exceed 2.5 m, thereby improving ease of use while reducing weight.

[0057] In step S2, the end slats at both ends of each fixed leading edge segment are connected to the front spar, while the middle slat of each fixed leading edge segment is not connected to the box segment. In a conventional fixed leading edge, the slats are connected to the front spar, which improves skin stability and transfers the load on the fixed leading edge to the front spar. In the integrally detachable wing fixed leading edge, for ease of disassembly, only the end slats at both ends of each fixed leading edge segment are connected to the front spar; the middle slat is not connected to the box segment and only serves to improve skin stability.

[0058] In step S3, when the load is less than the load threshold, no front wall is installed; instead, a stringer is installed at the original front wall location. For a conventional fixed leading edge, a front wall can be installed depending on the stress conditions. As a spanwise load-bearing component, the front wall provides support for the fixed leading edge skin and bulkhead, and transfers the aerodynamic load on the fixed leading edge to the fixed leading edge end bulkhead in the form of shear force after being converted through the skin, bulkhead, and stringer. However, for a fixed leading edge of an integrally detachable wing, installing a front wall would greatly increase the difficulty of disassembly. Therefore, when the load is not high, a stringer is installed at the original front wall location to improve the stiffness at that point.

[0059] In step S4, the fixed leading edge is used as a single access panel, and partial access panels are installed in areas with high maintenance frequency. For the integrally detachable wing fixed leading edge, the entire fixed leading edge is disassembled as an access panel. Small access panels can be installed in areas with high maintenance frequency for inspection or observation purposes, facilitating disassembly and improving maintenance efficiency. When disassembling the fixed leading edge as a single access panel, only the support plates connecting the skin to the end bulkheads and the skin to the strip plate need to be removed. The skin, stringers, and intermediate bulkheads are typically disassembled as a single unit; the end bulkheads are not removed.

[0060] In step S5, the fixed front edge is not connected to the body, and a 15mm gap is reserved between the fixed front edge and the body and sealed with rubber profiles; the fixed front edge is connected to the box section through a mating strip plate, and the skin of the fixed front edge is connected to the mating strip plate through a support plate nut; the skins of two adjacent fixed front edge sections are mated at the end partition, and the skins are connected to the end partition through a support plate nut.

[0061] The integral, detachable wing's fixed leading edge is not connected to the fuselage. A 15mm gap is left in the design, sealed with a special rubber profile. This design meets aerodynamic requirements, waterproofing requirements, and prevents structural damage due to deformation inconsistencies, while also allowing for the removal of the fixed leading edge. The fixed leading edge is connected to the box section via a mating plate, and a support plate nut connects the fixed leading edge to the plate, facilitating disassembly. The two fixed leading edge skin sections are mated at the end bulkheads and connected to the end bulkheads with support plate nuts, also facilitating disassembly.

[0062] In step S6, the parameters of the fixed leading edge structure are determined, including:

[0063] Perform parameter design for the riveted skin of the long stringer;

[0064] Design the parameters of the partition.

[0065] When designing riveted skin for long girder joints, multiple failure modes must be considered, making it difficult to effectively determine the structural dimensions. Therefore, it is necessary to comprehensively consider various failure modes, calculate their critical stresses, and compare them with the working stresses.

[0066] In a preferred embodiment of this application, the parameter design process for the stringer riveting skin is as follows:

[0067] Determine the initial thickness of the skin, the initial cross-sectional parameters of the stringers, and the initial spacing of the stringers;

[0068] Based on connection requirements and bird strike prevention requirements, the initial skin thickness, initial stringer cross-section parameters, and stringer spacing are determined with reference to parameters of other models.

[0069] Calculate the local critical instability stress of the skin;

[0070] The foremost part of the fixed leading edge is calculated using the cylindrical method:

[0071] ;

[0072] Where E is the elastic modulus of the material, and R is the radius of the cylinder. Thickness (cylinder wall thickness). The elastic Poisson's ratio of the material, This is a correction factor;

[0073] The remaining areas are calculated using the curved plate method:

[0074] ;

[0075] Where b is the width (width of the curved plate). The elastic Poisson's ratio of the material, The critical compressive stress coefficient, This is the plasticity correction factor. ;

[0076] Calculate the critical stress for local buckling of the stringer;

[0077] ;

[0078] in, The width of the loaded edge. For the thickness of the plate, The elastic Poisson's ratio of the material, The critical compressive stress coefficient, The elastic modulus of the material;

[0079] Calculate the critical stress for stress loss of the stringer;

[0080] ;

[0081] Among them, A i This refers to the crimping area;

[0082] Calculate the critical stress for overall instability of the riveted skin of the stringer;

[0083] ;

[0084] in, Let be the allowable column stress of the structural unit. For the compressive stress of the stringer (profile), The compressive elastic modulus of the stringer material. The effective length of the structural unit. , The length of the structural unit. For simply supported conditions, the end support coefficient is... , Let be the radius of gyration of the structural unit. , The moment of inertia of the panel section including the effective skin. The cross-sectional area of ​​the panel section including the effective skin;

[0085] Calculate the safety margin based on each critical stress;

[0086]

[0087] in, Allowable stress;

[0088] Adjust the skin thickness, stringer section parameters, and stringer spacing according to the safety margin, recalculate the safety margin based on the adjusted parameters, and iterate repeatedly until the requirements are met.

[0089] From the perspective of stress mode analysis, the failure mode of the web of a common diaphragm is instability failure. The failure modes of the flange are compressive instability and tensile fracture failure.

[0090] The parameter design process for the partition is as follows:

[0091] Determine the initial dimensions of the partition;

[0092] Determine the initial dimensions of the partition by referring to the dimensions of similar models, including the partition thickness and partition width.

[0093] Calculate the critical buckling stress of the web;

[0094] Calculate the critical buckling stress of the flange under compression;

[0095] Calculate the safety margin based on each critical stress;

[0096] The critical buckling stress of the web is:

[0097] ;

[0098] The safety margin is:

[0099] ;

[0100] The critical buckling stress of the flange under compression is:

[0101] ;

[0102] ;

[0103] The safety margin is:

[0104] ;

[0105] Let the maximum working tensile stress of the partition flange be... ,

[0106] The safety margin is calculated as follows:

[0107] ;

[0108] Adjust the partition size according to the safety margin, recalculate the safety margin based on the adjusted parameters, and iterate repeatedly until the requirements are met.

[0109] The integral detachable wing fixed leading edge obtained according to the above design method, such asFig. 2 As shown, the fixed leading edge includes a skin 1, a middle partition 2, an end partition 3, a stringer 4, and a docking strip 5. The structural form, connection relationship, and parameter design of each component refer to the design method of the fixed leading edge structure of a detachable wing.

[0110] This application presents a design method for an integrally detachable fixed leading edge structure for airfoils, designed to address the challenges of system installation and maintenance in certain small and medium-sized aircraft with low airfoil height and limited space at the fixed leading edge. It improves system installation and maintenance efficiency, solidifies the design process for detachable fixed leading edges, and thereby enhances design efficiency and overall design quality.

[0111] The above description is merely a specific embodiment of this application, but the scope of protection of this application is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in this application should be included within the scope of protection of this application. Therefore, the scope of protection of this application should be determined by the scope of the claims.

Claims

1. A design method for an integrally detachable wing fixed leading edge structure, characterized in that, include: Step S1: Determine the fixed leading edge segment; Step S2: Determine the partition structure; Step S3: Determine whether to install a front wall; Step S4: Determine the design of the cap with a fixed leading edge; Step S5: Determine the docking form between the fixed leading edge and the surrounding structure, as well as between two adjacent fixed leading edge segments; Step S6: Determine the parameters of the fixed leading edge structure; In step S1, the fixed front edge segments are determined according to the size of the raw material, the position of the skeleton, and the position of the cap; In step S1, the weight of each fixed leading edge segment shall not exceed 13kg and the length shall not exceed 2.5m; In step S2, the end partitions at both ends of each fixed front edge are connected to the front beam, while the middle partition of each fixed front edge is not connected to the box segment. In step S3, when the load is less than the load threshold, the front wall is not arranged, and the long truss is arranged at the original front wall position. In step S4, the fixed leading edge is used as the overall cover, and local covers are arranged in areas with high maintenance frequency. In step S5, the fixed front edge is not connected to the body, and a 15mm gap is reserved between the fixed front edge and the body and the gap is sealed with rubber profiles. In step S5, the fixed front edge is connected to the box segment by a mating belt plate, and the skin of the fixed front edge is connected to the mating belt plate by a support plate nut. In step S5, the two adjacent fixed front edge skins are joined at the end partition, and the skins and the end partition are connected by a support plate nut. In step S6, the parameters of the fixed leading edge structure are determined, including: Perform parameter design for the riveted skin of the long stringer; Design the parameters of the partition.

2. The design method for the integral detachable wing fixed leading edge structure according to claim 1, characterized in that, The parameter design for the riveted skin of the stringer includes: Determine the initial thickness of the skin, the initial cross-sectional parameters of the stringers, and the initial spacing of the stringers; Calculate the local critical instability stress of the skin; Calculate the critical stress for local buckling of the stringer; Calculate the critical stress for stress loss of the stringer; Calculate the critical stress for overall instability of the riveted skin of the stringer; Calculate the safety margin based on each critical stress; Adjust the skin thickness, stringer section parameters, and stringer spacing according to the safety margin, recalculate the safety margin based on the adjusted parameters, and iterate repeatedly until the requirements are met.

3. The design method for the integral detachable wing fixed leading edge structure according to claim 2, characterized in that, The parameters of the partition are designed, including: Determine the initial dimensions of the partition; Calculate the critical buckling stress of the web; Calculate the critical buckling stress of the flange under compression; Calculate the safety margin based on each critical stress; Adjust the partition size according to the safety margin, recalculate the safety margin based on the adjusted parameters, and iterate repeatedly until the requirements are met.

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