A method for designing an aircraft pylon structure
By optimizing the design of the aircraft sling structure through finite element analysis, the problems of relative deformation control and vibration response of the center of gravity of special equipment on the aircraft were solved, and the strength, stiffness and vibration matching of the sling structure were achieved, meeting the performance requirements of the aircraft.
Patent Information
- Authority / Receiving Office
- CN · China
- Patent Type
- Patents(China)
- Current Assignee / Owner
- AVIC XAC COMMERCIAL AIRCRAFT CO LTD
- Filing Date
- 2021-07-27
- Publication Date
- 2026-07-24
AI Technical Summary
When mounting special equipment on an aircraft, existing technologies struggle to control the relative deformation of the equipment's center of gravity to within 0.1 mm under stable level flight conditions, while simultaneously avoiding the aircraft's main excitation frequency, meeting the strength, stiffness, and vibration response requirements of the mounting structure, and without affecting aircraft performance.
Finite element method (FEM) analysis was used to design an aircraft suspension structure. By calculating the relative deformation of different fuselage angles of attack and sideslip angles, the main load-bearing truss structure and tie rod connections were optimized to establish the aircraft flight attitude range that meets the relative deformation requirements. Strength and vibration calculations were performed in conjunction with aerodynamic and inertial loads to optimize the design of a suspension structure that meets the requirements.
It achieves the control of the relative deformation of the equipment's center of gravity within 0.1mm under stable level flight conditions, avoids the main excitation frequency of the aircraft, meets the strength and vibration response requirements of the suspension structure, and at the same time reduces the impact of weight to ensure that the aircraft's performance is not damaged.
Smart Images
Figure CN115688500B_ABST
Abstract
Description
Technical Field
[0001] This invention relates to the field of aircraft structural strength design, specifically to a design method for aircraft sling structures based on finite element analysis, which allows for relative deformation requirements at the center of gravity of the load. Background Technology
[0002] With the development of the modern aviation industry, under the premise of meeting basic performance requirements, there is a desire to realize more uses for aircraft, especially special aircraft that are modified or equipped with special equipment, which greatly expands the application of aircraft.
[0003] In the development of a certain type of aircraft, special equipment needs to be mounted on both sides of the lower fuselage. Under stable level flight conditions, the relative deformation at the center of gravity of the equipment must not exceed 0.1 mm. If this control requirement is not met, the equipment will malfunction. Simultaneously, the low-order natural frequencies of the mounting structure must avoid the fundamental frequency (propeller passing frequency) and its higher harmonics of the aircraft's main excitation source, reduce the vibration response displacement amplitude of the mounting structure, and improve cabin comfort while meeting the relative deformation requirements. Therefore, a mounting structure (mounting device) needs to be designed to mount the equipment. While ensuring aircraft performance, the mounting structure must meet the strength and stringent stiffness design requirements for the mounted equipment, as well as vibration control requirements. It must be lightweight, not affect the aircraft's lift or drag, and the flight attitude of the aircraft must be clearly defined to satisfy the relative deformation requirements of the center of gravity of the mounted equipment on both sides. Summary of the Invention
[0004] The purpose of this invention is to provide a design method based on finite element analysis. Under stable level flight conditions, strength optimization is achieved through relative deformation and vibration calculations at the load center of gravity position for different fuselage angles of attack and sideslip angles. This results in a sling structure that perfectly matches strength, stiffness, and vibration. For the calculated parameters such as fuselage angle of attack and sideslip angle at different speeds, a correspondence between the fuselage angle of attack and sideslip angle is established to ensure that the relative deformation does not exceed the allowable requirements. Furthermore, based on the deformation results obtained from finite element analysis, the range of aircraft flight attitudes that meets the allowable relative deformation requirements is calculated for variations in the fuselage angle of attack and sideslip angle at different speeds.
[0005] A design method for an aircraft sling structure, comprising a main load-bearing truss, fixed docking joints, tie rods, and fairings, wherein the central symmetry plane of the main load-bearing truss coincides with the symmetry plane of the aircraft, and an equipment pod is symmetrically mounted at each end of the main load-bearing truss. The width and length of the main load-bearing truss, the position of the load center of gravity, the allowable relative deformation at the load center of gravity positions on both sides of the aircraft sling structure, and the vibration stiffness requirements of the aircraft sling structure are known. The method is characterized by the following: 1) proposing a preliminary design scheme for the aircraft sling structure, in which the shape and dimensions of the main load-bearing truss structure, as well as the connection positions and angles of the tie rods on the main load-bearing truss, are preliminarily designed; 2) establishing a finite element model based on the preliminary design scheme of the aircraft sling structure, loading aerodynamic and inertial loads under different flight attitudes and cruise states of the aircraft onto the finite element model, calculating the strength of the aircraft sling structure and the relative deformation at the load center of gravity position, obtaining the deformation and relative deformation at the load center of gravity position, and simultaneously performing vibration calculations on the aircraft sling structure;
[0006] 3) Based on the deformation and relative deformation of the load center of gravity, and the influence of vibration on the stiffness of the aircraft sling structure, optimize the shape and size of the main load-bearing truss structure in the preliminary scheme, as well as the connection position and angle of the tie rod in the main load-bearing truss, to obtain the optimized design scheme of the aircraft sling structure; 4) Repeat steps 2) and 3) above for the optimized design scheme of the aircraft sling structure until the final design scheme of the aircraft sling structure that meets the allowable relative deformation of the load center of gravity on both sides and the vibration stiffness requirements of the aircraft sling structure is obtained.
[0007] Determining the flight attitude range of the aircraft using the sling structure includes the following: 1) On the finite element model of the final design scheme of the aircraft sling structure, aerodynamic and inertial loads of the aircraft cruise state at different flight speeds and different flight attitudes are applied, and the deformation of the center of gravity of the sling on both sides is calculated to obtain the deformation of the aircraft sling structure and the relative deformation of the center of gravity of the sling; 2) The deformation of the center of gravity of the sling on both sides calculated in step 1) is compared with the deformation of the center of gravity of the sling on both sides under the allowable stable level flight attitude of the aircraft at the same flight speed and flight attitude, and the relative deformation value of the center of gravity of the sling on both sides under each flight speed and flight attitude is calculated; 3) The relative deformation value of the center of gravity of the sling on both sides under each flight speed and flight attitude is obtained by linear interpolation to obtain the feasible flight attitude range, and is made explicit by curves.
[0008] The advantages of this application are:
[0009] 1) The optimal suspension structure was selected through finite element strength optimization and vibration technology requirements and their impact on stiffness. The variable cross-section truss structure ensured that the suspension had the least impact on aircraft vibration, and also ensured the lightest weight and met the strength and stiffness requirements.
[0010] 2) For different combinations of fuselage angle of attack and sideslip angle at different speeds, the relative deformation at the center of gravity of the load on both sides in a stable level flight state can be calculated;
[0011] 3) Based on the relative deformation at the center of gravity of the two pods, the relative deformation at the center of gravity of the two pods under a certain range of fuselage angle of attack and sideslip angle can be calculated by interpolation.
[0012] 4) The calculation results from 2) and 3) can be easily filtered to find the range of fuselage angle of attack and sideslip angle that meet the mission requirements, and the flight attitude of the aircraft can be determined.
[0013] This application selects a suitable combination of fuselage angle of attack and sideslip angle, and uses finite element strength optimization design to create a pylon structure that adapts to the aircraft performance and mission requirements. The relative deformation at the center of gravity of the two pods can be calculated by using finite element calculation and interpolation calculation, thereby selecting the aircraft flight attitude that meets the mission requirements. The design method is convenient, quick, and highly practical.
[0014] The present invention will now be described in detail with reference to the embodiments in the accompanying drawings. Attached Figure Description
[0015] Figure 1 This is a schematic diagram of the unoptimized suspension structure.
[0016] Figure 2 This is a schematic diagram of the optimized suspension structure.
[0017] Figure 3 This is a schematic diagram of the optimized main load-bearing truss structure.
[0018] Figure 4 This is a schematic diagram of the aircraft's flight attitude range.
[0019] The numbers in the diagram are explained as follows: 1. Main load-bearing truss; 2. Fixed docking joint; 3. Tie rod; 4. Fairing; 5. Equipment pod. Detailed Implementation
[0020] Referring to the accompanying drawings, this embodiment describes the design of a sling structure for a certain type of aircraft, used to mount special equipment on both sides of the lower fuselage. The requirement is that, under stable level flight conditions, the relative deformation at the center of gravity of the equipment should not exceed 0.1 mm during operation. Based on the aircraft data, the width and length of the main load-bearing truss and the position of the sling center of gravity can be determined. It is known that the allowable relative deformation at the sling center of gravity on both sides of the aircraft sling structure should not exceed 0.1 mm, and the vibration stiffness requirements of the aircraft sling structure are also known.
[0021] The specific design steps are as follows:
[0022] 1) A preliminary design scheme for the suspension structure is proposed. The aircraft suspension structure includes a main load-bearing truss 1, a fixed docking joint 2, a tie rod 3, and a fairing 4. The central symmetry plane of the main load-bearing truss coincides with the symmetry plane of the aircraft. An equipment pod 5 is symmetrically suspended at both ends of the main load-bearing truss 1. In this preliminary scheme, the shape and size of the main load-bearing truss structure, as well as the connection position and connection angle of the tie rod on the main load-bearing truss are preliminarily designed.
[0023] 2) A finite element model was established based on the preliminary design of the suspension structure. Aerodynamic and inertial loads for six flight attitudes under the aircraft's cruise state were applied to the finite element model. The strength of the suspension structure and the relative deformation at the center of gravity of the load were calculated, yielding the deformation and relative deformation at the center of gravity. Vibration calculations for the suspension structure were also performed simultaneously. The six flight attitudes were: 0° angle of attack and 0° sideslip; 0° angle of attack and 2° sideslip; 2° angle of attack and 0° sideslip; 2° angle of attack and 2° sideslip; 3° angle of attack and 0° sideslip; 3° angle of attack and 0° sideslip; and 3° angle of attack and 2° sideslip. The equivalent airspeed for the aircraft's cruise state was 420 km / h.
[0024] 3) Based on the deformation and relative deformation of the load center of gravity, the relative deformation values of the load center of gravity on both sides of the load center of gravity relative to the stable level flight state (the flight attitude of the aircraft in a level flight state with an angle of attack of 3° and a sideslip angle of 0° is defined as the stable level flight state), and the influence of the vibration of the suspension structure on the stiffness requirements, the shape and size of the main load-bearing truss structure in the preliminary scheme, as well as the connection position and connection angle of the tie rod in the main load-bearing truss, are optimized to obtain the optimized design scheme of the aircraft suspension structure.
[0025] 4) For the optimized design of the aircraft sling structure, repeat steps 2) and 3) above until the final design scheme that satisfies the allowable relative deformation of the center of gravity of the load on both sides and the vibration stiffness requirements of the aircraft sling structure when the sideslip angle is within ±1° is obtained.
[0026] Based on the optimized design of the aircraft's pylon structure, the flight attitude range of the carrier aircraft needs to be determined, specifically including the following:
[0027] 1) On the finite element model of the final design scheme of the aircraft sling structure, aerodynamic and inertial loads were applied to the aircraft in stable level flight state under five different flight speeds (equivalent airspeeds of 420km / h, 400km / h, 350km / h, 300km / h, and 280km / h) and six different flight attitudes (0° fuselage angle of attack and 0° sideslip angle, 0° fuselage angle of attack and 2° sideslip angle, 2° fuselage angle of attack and 2° sideslip angle, 3° fuselage angle of attack and 0° sideslip angle, and 3° fuselage angle of attack and 2° sideslip angle). The deformation of the center of gravity of the sling structure on both sides was calculated. The deformation of the aircraft sling structure and the relative deformation of the center of gravity of the sling structure were obtained.
[0028] 2) Compare the deformation of the center of gravity of the two loads calculated in step 1) with the deformation of the center of gravity of the two loads under the allowable stable level flight attitude of the aircraft foundation under the same flight speed and flight attitude, and calculate the relative deformation value of the center of gravity of the two loads under each flight speed and flight attitude. The flight attitude corresponding to the relative deformation value is an isolated sideslip angle.
[0029] 3) The relative deformation values of the center of gravity of the load on both sides under each flight speed and flight attitude are obtained by linear interpolation to obtain the feasible flight attitude range, and then made explicit by curves.
[0030] As can be seen from the above embodiments, by using the method of the present invention, the relative displacement of the calculation points required by the mission can be obtained by analyzing and screening the finite element calculation results for different flight attitudes, and the flight attitude of the aircraft can be clearly defined. The calculation process is simple and direct and has strong practicality.
Claims
1. A design method for an aircraft sling structure, the aircraft sling structure comprising a main load-bearing truss, fixed docking joints, tie rods, and a fairing, wherein the central symmetry plane of the main load-bearing truss coincides with the symmetry plane of the aircraft, and an equipment pod is symmetrically mounted at each end of the main load-bearing truss, wherein the width and length of the main load-bearing truss and the position of the sling center of gravity are known, the allowable relative deformation at the sling center of gravity positions on both sides of the aircraft sling structure is known, and the vibration stiffness requirements of the aircraft sling structure are known, characterized in that... This includes the following: 1) Proposing a preliminary design scheme for the aircraft sling structure, which preliminarily designs the shape and dimensions of the main load-bearing truss structure, as well as the connection positions and angles of the tie rods on the main load-bearing truss; 2) Establishing a finite element model based on the preliminary design scheme of the aircraft sling structure, loading aerodynamic and inertial loads under different flight attitudes and cruise states of the aircraft onto the finite element model, calculating the strength of the aircraft sling structure and the relative deformation of the load center of gravity, obtaining the deformation and relative deformation amount at the load center of gravity, and simultaneously performing vibration calculations on the aircraft sling structure; determining the flight attitude range of the aircraft using the sling structure, loading aerodynamic and inertial loads under different flight speeds and flight attitudes and cruise states of the aircraft onto the finite element model of the final design scheme of the aircraft sling structure, calculating the deformation of the load center of gravity on both sides, obtaining the deformation of the aircraft sling structure and the relative deformation amount at the load center of gravity; and calculating the deformation of both sides. The deformation at the center of gravity of the load is compared with the deformation at the center of gravity of the load on both sides under the allowable stable level flight attitude of the aircraft foundation at the same flight speed and flight attitude. The relative deformation values of the center of gravity of the load on both sides under each flight speed and flight attitude are calculated. The relative deformation values of the center of gravity of the load on both sides under each flight speed and flight attitude are obtained by linear interpolation to obtain the feasible flight attitude range and made explicit by curve. 3) Based on the deformation and relative deformation of the center of gravity of the load, and the influence of the vibration of the aircraft sling structure on the stiffness requirements, the shape, size and the connection position and connection angle of the tie rod in the main load truss structure in the preliminary scheme are optimized to obtain the optimized design scheme of the aircraft sling structure. 4) For the optimized design scheme of the aircraft sling structure, repeat the above steps 2) and 3) until the final design scheme of the aircraft sling structure that meets the allowable relative deformation of the center of gravity of the load on both sides and the vibration stiffness requirements of the aircraft sling structure is obtained.
2. The aircraft sling structure design method as described in claim 1, characterized in that, The different flight attitudes can be selected from the following six types: the first type is 0° fuselage angle of attack and 0° sideslip; the second type is 0° fuselage angle of attack and 2° sideslip; the third type is 2° fuselage angle of attack and 0° sideslip; the fourth type is 2° fuselage angle of attack and 2° sideslip; the fifth type is 3° fuselage angle of attack and 0° sideslip; and the sixth type is 3° fuselage angle of attack and 2° sideslip.
3. The aircraft sling structure design method as described in claim 1, characterized in that, In step 1), the different flight speeds are selected from the following five types, with equivalent airspeeds of 420, 400, 350, 300, and 280 km / h respectively.