Calculation method of airplane under action of steady-state non-uniform temperature field
By using the finite element model and the heat conduction-corrected Kriging interpolation method, a steady-state non-uniform temperature field of the aircraft is constructed, which solves the problem that high-stress regions are difficult to capture in traditional methods. This enables high-precision calculation of the real physical field coupling between temperature stress and mechanical stress, thereby improving the safety and reliability of aircraft design.
Patent Information
- Application Number
- CN202511718228.9
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2025-11-21
- Publication Date
- 2026-03-17
AI Technical Summary
In existing aircraft design, the traditional assumption of a uniform temperature field makes it difficult to capture high-stress areas, and the deployment of a large number of temperature sensors increases system weight and cost. At the same time, the limited sensor data cannot reconstruct the temperature field, resulting in insufficient calculation accuracy and safety.
A finite element model combined with a heat conduction-corrected Kriging interpolation method is used to construct an initial non-uniform temperature field based on effective temperature measurement data. Through heat transfer analysis and verification data correction, the non-uniform temperature field is accurately calculated, and material/geometric nonlinear simulation calculations are performed to superimpose mechanical stress.
It improves the calculation accuracy of high-stress-risk locations, enhances calculation reliability, reduces aircraft safety risks, and supports aircraft fatigue life assessment and airworthiness certification.
Smart Images
Figure CN121683008A_ABST
Abstract
Description
TECHNICAL FIELD
[0001] The application belongs to the technical field of aircraft design, and relates to a calculation method of an aircraft under the action of a steady-state non-uniform temperature field. BACKGROUND
[0002] Finite element simulation calculation is a commonly used technical means in the field of aircraft design. By applying boundary conditions and loads to the model, various conditions in the actual flight process of the aircraft are simulated to achieve the purpose of prediction, monitoring and analysis. Due to the need for ice prevention, the leading edge of the aircraft wing and other structures have a certain temperature distribution, forming a temperature field. For stress calculation of the structure in this area, the traditional calculation method is to apply a uniform temperature field with the highest temperature value in the interval temperature field for linear calculation to solve the thermal stress, and linearly superimpose the mechanical stress generated by the aircraft aerodynamic load.
[0003] Although this simplified processing method is simple and easy to understand, it saves work and is convenient for engineers to quickly carry out corresponding calculations, but it ignores the temperature difference between the internal structures of the aircraft. For aircraft structural components with large temperature differences, this simplification of temperature calculation will make it difficult to capture the real high stress area, and only reflect the average stress level of the interval structure, which poses a risk to the safety of the aircraft in flight.
[0004] In addition, when conducting flight tests, if a large number of temperature sensors are arranged to obtain sufficient temperature data, the system installation weight will undoubtedly increase, resulting in high test and maintenance costs, and the reliability of the temperature data will also be reduced due to the complexity of the test system. The number of sensors arranged on the aircraft is usually limited, and the temperature measurement point data of the test is insufficient to meet the needs of reconstructing the temperature field.
[0005] Therefore, it is necessary to study a new method to overcome the shortcomings of the prior art to solve the above problems. SUMMARY
[0006] The purpose of the present application is to provide a calculation method of an aircraft under the action of a steady-state non-uniform temperature field. The present application can improve the calculation accuracy of high stress risk positions, improve the calculation reliability, and reduce the safety risk.
[0007] The technical solution of the present application is: a calculation method of an aircraft under the action of a steady-state non-uniform temperature field, comprising: Step 1: According to the aircraft structural components, screen the effective temperature measurement point data in the flight test as field data; Step 2: Form an initial non-uniform temperature field by a thermal conduction correction Kriging interpolation method according to the spatial position; Step 3: Use the finite element model of the aircraft to perform heat transfer analysis and calculation of the aircraft in a thermal convection environment under the boundary condition of the initial non-uniform temperature field to obtain a steady-state non-uniform temperature field; Step 4, in addition to the field data, arrange temperature measuring points for verifying the steady-state non-uniform temperature field and obtain temperature data as verification data, complete the correction of the steady-state non-uniform temperature field, and the verification data measuring points are preferentially arranged on the typical section; Step 5, load the temperature field and the material / geometric nonlinear simulation calculation of the severe working condition aerodynamic load on the finite element model of the aircraft.
[0008] In the foregoing calculation method of the aircraft under the action of the steady-state non-uniform temperature field, in step 1, the method for screening the temperature measuring point data in the flight test for the same aircraft part is the highest temperature of the same temperature measuring point under different flight states.
[0009] In the foregoing calculation method of the aircraft under the action of the steady-state non-uniform temperature field, in step 2, the interpolation method of the heat conduction correction Kriging method is as follows: Let the heat flow into the actual temperature field be equal to the heat flow out, then:
[0010] In the formula, indicates the weight coefficient, indicates the variogram model between the known temperature measuring points, indicates the variogram model between the known temperature measuring points and the unknown point, indicates the thermal conductivity field function, indicates the temperature of the unknown point, indicates the known temperature, indicates the degree of change of the temperature field function in space, indicates the heat flow density, indicates the steady-state heat conduction equation, indicates the Lagrange multiplier.
[0011] In the foregoing calculation method of the aircraft under the action of the steady-state non-uniform temperature field, in step 3, when setting the parameters for heat transfer analysis calculation, the heat conduction coefficient, the heat convection coefficient under the specified airspeed, the convection temperature, and the heat flow density parameter need to be considered.
[0012] In the foregoing calculation method of the aircraft under the action of the steady-state non-uniform temperature field, in step 3, the finite element model of the aircraft is the finite element model corrected by the full aircraft static test data.
[0013] In the foregoing calculation method of the aircraft under the action of the steady-state non-uniform temperature field, in step 4, the verification data for correcting the steady-state non-uniform temperature field are taken from the same flight test.
[0014] In the foregoing calculation method of the aircraft under the action of the steady-state non-uniform temperature field, in step 4, the selection criteria of the typical section include load severity, force transmission path representativeness, and structural discontinuity.
[0015] In the calculation method of the foregoing aircraft under the action of a steady-state non-uniform temperature field, in step 5, the basis for screening the aerodynamic load severe working condition is the wing internal force, and the internal force includes the shear force and the bending moment of the wing.
[0016] Beneficial effects: the present application provides a calculation method of an aircraft under a steady-state non-uniform temperature field, which can realize the construction of a full-aircraft three-dimensional temperature field and the real physical field coupling calculation of the superposition of aircraft temperature stress and mechanical stress under the premise of limited temperature measuring points, improve the calculation accuracy of high stress risk positions, improve the calculation reliability, reduce the safety risk, and solve the problem that the traditional linear superposition of thermal stress and mechanical stress cannot accurately predict the post-buckling characteristics of the aircraft.
[0017] The present application initiates a high-precision construction method of a steady-state non-uniform temperature field of an aircraft and a superposition calculation method of a mechanical field, breaks through the limitations of the traditional single-point temperature measurement and mean value loading calculation method, and can realize the construction of a full-aircraft three-dimensional temperature field through a thermal conduction Kriging interpolation algorithm and a simulation calculation method; the technology realizes the real physical field coupling calculation of the temperature stress of an aircraft and the mechanical stress for the first time, greatly improves the accuracy of aircraft structure risk positioning, and effectively supports the aircraft fatigue life evaluation and airworthiness certification work. BRIEF DESCRIPTION OF DRAWINGS
[0018] Fig. 1 is a flight temperature measuring point data time history curve schematic diagram in the example of the present application.
[0019] Fig. 2 is a steady-state non-uniform temperature field schematic diagram in the example of the present application. DETAILED DESCRIPTION
[0020] The features and illustrative embodiments of various aspects of the present application will be described below in detail. In the following detailed description, numerous specific design details are set forth in order to provide a more thorough understanding of the present application. However, it will be apparent to one of ordinary skill in the art that the present application can be practiced without some or all of these specific details. The description of the embodiments is merely to provide a better understanding of the present application by showing examples of the present application. The present application is in no way limited to any specific settings and methods presented below, but covers any improvements, replacements and modifications of structures, methods and devices without departing from the spirit of the present application. In the drawings and the following description, the parts not described in detail are considered to be common knowledge or conventional implementations in the art.
[0021] It should be noted that the embodiments of the present application and the features in the embodiments can be combined with each other without conflict, and each embodiment can be mutually referenced and quoted. The present application will be described in detail below with reference to the drawings and in combination with the embodiments.
[0022] Embodiment 1. A method for calculating an aircraft under the action of a steady-state non-uniform temperature field, referring to Figs. 1-2 , an initial non-uniform temperature field is obtained by interpolation based on a limited number of temperature measurement points using a heat conduction modified Kriging method; a steady-state non-uniform temperature field is obtained by heat transfer analysis and calculation of the aircraft under a steady heat source and a convective environment using an aircraft finite element model with the initial non-uniform temperature field as the boundary condition; temperature measurement points for verifying the steady-state non-uniform temperature field are arranged outside the aforementioned temperature measurement points and temperature data is obtained to complete the correction of the steady-state non-uniform temperature field; the temperature field and severe working condition aerodynamic load are loaded on the aircraft finite element model for material / geometric nonlinear simulation calculation. Specifically, the following steps are included: Step 1. According to the aircraft wing skin, ribs, longitudinal walls and other components, the data of each measurement point is selected as the highest temperature in the time domain temperature curve of the measurement point under different flight conditions.
[0023] Step 2. According to the spatial position, the heat conduction modified Kriging method is used for interpolation, assuming that the actual temperature field at this time the heat inflow is equal to the heat outflow, i.e. no heat accumulation, the heat conduction Kriging interpolation method is:
[0024] In the formula, represents the weight coefficient, represents the variogram model between the known temperature measurement points, represents the variogram model between the known temperature measurement points and the unknown point, represents the thermal conductivity field function, represents the unknown point temperature, represents the known temperature, represents the degree of change of the temperature field function in space, represents the heat flux density, represents the steady-state heat conduction equation, which is mathematically referred to as the divergence of the heat flux density, and physically represents that the heat flowing in and out of the temperature field is consistent, i.e. the total heat is constant, which is used to constrain and solve , represents the Lagrange multiplier.
[0025] Step 3. According to the performance of the metal material, the thermal conductivity coefficient, the heat convection coefficient between the structure and the air, and the convective temperature value are set for the materials of the components of the aircraft wing detail finite element model. Using the aircraft finite element model corrected by the full aircraft static test data, the steady-state non-uniform temperature field is obtained by heat transfer analysis and calculation of the aircraft in the thermal convection environment with the initial non-uniform temperature field as the boundary condition; Step 4. Selecting a high-load section located in the main transmission path as a typical section to arrange verification data measurement points and obtain verification data to complete the correction of the steady-state non-uniform temperature field; Step 5, load the temperature field and the aerodynamic load of the severe working condition to the finite element model of the airplane to perform material / geometric nonlinear simulation calculation.
[0026] Taking the simulation calculation of the steady-state non-uniform temperature field of the wing leading edge of a certain domestic large amphibious airplane for example, the method is described.
[0027] 1) According to the wing skin, rib, longitudinal wall and other components of the amphibious airplane, the data of each measuring point is selected as the highest temperature in the time domain temperature curve of the measuring point in different flight states.
[0028] 2) According to the spatial position, the thermal conduction correction Kriging interpolation is performed, and it is assumed that the heat inflow is equal to the heat outflow at this time, that is, there is no heat accumulation, and the thermal conduction Kriging interpolation method is:
[0029] In the formula, The weight coefficient is represented by W; The variation function model between the known temperature measuring points is represented by C; The variation function model between the known temperature measuring points and the unknown point is represented by C; The thermal conductivity field function is represented by K, and in this example, the thermal conductivity of the material is taken as the inherent property parameter of the material, The unknown point temperature is represented by T; The known temperature is represented by T; 3) According to the performance of the metal material, the thermal conductivity coefficient, the heat convection coefficient between the structure and the air, and the temperature value of the convection air of the finite element model of the wing of the amphibious airplane are set.
[0030] 4) Taking the initial non-uniform temperature field as the boundary condition, the heat transfer analysis is performed on the wing finite element model verified by the full machine static test, and the steady-state non-uniform temperature field of the wing leading edge skin, longitudinal wall, wing rib and other structures is obtained.
[0031] 5) Selecting the high-load section located in the main transmission path as a typical section to arrange verification data measuring points, obtaining verification data, and completing the correction of the steady-state non-uniform temperature field; 6) Load the steady-state non-uniform temperature field and the aerodynamic load in the flight test to the finite element model of the wing to perform material / geometric nonlinear simulation calculation.
[0032] The above only describes the preferred embodiments of the present application, and it should be noted that those skilled in the art can make some improvements without departing from the principles of the present application, and these improvements should also be considered as the protection scope of the present application.
Claims
1. A calculation method for an aircraft under steady-state non-uniform temperature field, characterized in that, include: Step 1: Select effective temperature measurement point data from the flight test according to the aircraft structural components, and use them as field data; Step 2: Generate an initial non-uniform temperature field according to spatial location using the heat conduction-corrected Kriging interpolation method; Step 3: Using the aircraft finite element model and the initial non-uniform temperature field as the boundary condition, perform heat transfer analysis and calculation of the aircraft in the thermal convection environment to obtain the steady-state non-uniform temperature field. Step 4: In addition to the field data, set up temperature measurement points to verify the steady-state non-uniform temperature field and acquire temperature data as verification data to complete the correction of the steady-state non-uniform temperature field. The verification data measurement points are preferentially arranged on typical cross-sections. Step 5: Perform material / geometric nonlinear simulation calculations on the aircraft finite element model to apply the temperature field and severe aerodynamic loads.
2. The calculation method for an aircraft under steady-state non-uniform temperature field according to claim 1, characterized in that, In step 1, the method for selecting temperature measurement data from various temperature measurement points during flight tests for the same aircraft component is the highest temperature of the same temperature measurement point under different flight conditions.
3. The calculation method for an aircraft under steady-state non-uniform temperature field according to claim 1, characterized in that, In step 2, the heat conduction-corrected Kriging interpolation method is as follows: Assuming that the heat inflow in the actual temperature field equals the heat outflow, then: In the formula, Indicates the weighting coefficient. This represents a variogram model between known temperature measurement points. This represents the variogram model between known temperature measurement points and unknown points. Represents the thermal conductivity field function. Indicates the temperature at an unknown point. Indicates a known temperature. This indicates the degree of drastic change of the temperature field function in space. Indicates heat flux density, The steady-state heat conduction equation is expressed as follows: It represents the Lagrange multiplier.
4. The calculation method for an aircraft under steady-state non-uniform temperature field according to claim 1, characterized in that, In step 3, when setting the parameters for heat transfer analysis calculations, the thermal conductivity coefficient, the thermal convection coefficient at a specified space velocity, the convection temperature, and the heat flux density parameters need to be considered.
5. The calculation method for an aircraft under steady-state non-uniform temperature field according to claim 1, characterized in that, In step 3, the aircraft finite element model is a finite element model corrected from the full-aircraft static test data.
6. The calculation method for an aircraft under steady-state non-uniform temperature field according to claim 1, characterized in that, In step 4, the verification data and field data used to correct the steady-state non-uniform temperature field are taken from the same flight test.
7. The calculation method for an aircraft under steady-state non-uniform temperature field according to claim 1, characterized in that, In step 4, the selection criteria for typical cross-sections include load severity, representativeness of the force transmission path, and structural discontinuity.
8. The calculation method for an aircraft under steady-state non-uniform temperature field according to claim 1, characterized in that, In step 5, the basis for screening severe aerodynamic load conditions is the magnitude of the internal forces of the wing, which include the shear force and bending moment of the wing.