To study the influence of random vibration on the fatigue life of airborne equipment, an aircraft electrical control box was selected as the research object. First, finite element software was used to model the dynamics of the airborne electrical control box to investigate its mode frequencies. The accuracy of finite element simulations was verified by performing mode experiments. Second, the mode superposition method was used to analyze the flight direction (X axis), side direction (Y axis), and altitude direction (Z axis) random vibration responses of the electrical control box. The analysis results were combined with the Miner linear cumulative damage criterion and the Gaussian-distribution Steinberg method to estimate the fatigue life of the electrical control box in the three directions. Finally, the calculation results were verified by performing the random vibration durability test on the electrical control box. The finite element mode analysis results show good agreement with the vibration experiment results, and the maximum error is 13.4%, indicating that the finite element model established in this paper is acceptable. The fatigue life of the electrical control box in the three axes meets the user requirements, and random vibration along the side direction (Y axis) has the greatest impact on the fatigue life, which is consistent with the results of the actual experimental data. The research method can be extended to predict the fatigue life of other airborne equipment and thus has practical significance for structural design and reliability analysis of airborne equipment.

To achieve miniaturization and lightweight design of a flapping-wing aircraft, a high-performance biomimetic butterfly flyer was designed based on an analysis of the butterfly’s body structure and flight principles. The aircraft has a mass of 20.6 g and a wingspan of 0.295 m. To validate the rationality of the design, sensitivity analysis of the flapping-wing drive mechanism was first conducted using MATLAB 2022B software, and the length of the driving rod was optimized. Subsequently, a dynamic model was established to calculate the aerodynamic performance of the flapping wing. Then, the aerodynamic performance of the aircraft was verified using simulation software (XFLOW 2022). Finally, the flight stability of the aircraft was validated using the SIMULINK toolbox. Flight test results show that the biomimetic butterfly flyer achieves a maximum flight speed of 0.9 m/s, a climb rate of 0.12 m/s, and a flight endurance of up to 3 min, with good flight stability. This design provides a new approach for the development of small and lightweight flapping-wing aircraft.

At the initial stage of design, the structural strength and stability of the load-bearing structural members of the aircraft are mainly considered. In this paper, the structural strength and stability of a support of aircraft assembly are analyzed. First, structural strength analysis, under the design load of 30.151 KN, and numerical analysis shows that, in the stress concentration area, the stress peak value is 571.9 Mpa. The experimental value of the strain sensor S6 position is compared with the numerical results, and the error is within 2.5%, which verifies the validity of the numerical model and the structural strength to meet the design requirements. Second, structural stability adopts the method of buckling analysis. Linear buckling analysis shows that the first six order buckling modes and critical loads. The first order critical load is 464.9 KN. Risk arc length method is used in the nonlinear buckling analysis. Considering the geometric imperfection of the support, 1% and 10% initial imperfection values are used. Nonlinear buckling critical load values—without imperfection and considering the initial imperfection—are obtained, which are 110.6 KN, 108.4 KN, and 106.2 KN, respectively. There is little difference between the three values, indicating that the geometric imperfection of the support is not obvious to the nonlinear buckling, and they have structural stability. In the case of no imperfection, the maximum deformation of the main area reaches 21.4 mm, and the support enters the instability state. This paper comprehensively discusses the results of structural strength and structural stability analysis, which can provide reference for the design and optimization of bearing components.

The aircraft electronic devices are mainly composed of plate type, the structures with so many electronic components are complex. Most electronic components are standard parts. Reliability analysis of aircraft electronic equipment is mainly focused on the circuit of electronic equipment and components and so on. However, the investigation about reliability of electronic equipment structure research is seldom. This article introduces the principle and method of structural reliability of aircraft electronic equipment. Taking the computer box of air data processing for a type of aircraft as an example, reliability analysis process is described based on Monte Carlo Method. the result of calculating the strength and stiffness reliability of the box are 99.9746% and 99.9833% respectively, while the experimental reliability are 99.9986% and 99.9918%. The good consistency between simulation and experiment results indicates that the analysis method is reasonable and effective.

Airplane oil pressure reservoirs are major fuel supplying parts of airplanes to have flight maneuvers. Since they are always working under alternating loadings, they are prone to produce fatigue fractures. Therefore, they will undergo vibration endurance tests in general. And the fixed-frequency vibration endurance test is commonly used in engineering. This paper has established a finite element model for the vibration test system, taken into consideration the influence of liquid fuel and introduced the fluid-solid coupling method to conduct harmonic response analysis on different excitation conditions with MSC.Nastran, thus obtaining the stress distribution under different loading conditions. Moreover, the results have been compared with the test results to verify correctness of the method, thus providing references for the improvement in structural design.

Modal test is one of the basic works of studying on dynamic characteristics of aircraft. The ideal state is to make the specimen in the "free-free" state when doing modal test. In actual, specimen is supported by an elastic supporting system, the stiffness, added mass and supporting position of the supporting system will affect the rigid-body modal frequency of the system directly. In this paper, we took a certain aerospace products as an example and used ANSA, NASTRAN software, studied on the influence of stiffness, added mass and supporting position of the supporting system on modal test results, and put forward the design method of modal test supporting system.