The skin panel structures of vehicles in supersonic flights are subjected to combined thermal, acoustic, and aerodynamic loads. These combined loads may cause a nonlinear dynamic response of the high-speed flight vehicle’s panel structures. Among these nonlinear dynamic responses, the snapthrough and limit cycle oscillation response seriously affect the fatigue failure of the panel structures. This work investigates the nonlinear dynamic responses using the numerical method, when combined supersonic aerodynamic, thermal, and random acoustic loads are applied to the panel structures simultaneously. To consider the thin and thick composite panels, the first-order shear deformation plate theory (FSDT) and the von Karman nonlinear displacement–strain relationship are applied. The aerodynamic load in the supersonic flow is modeled using the first-order piston theory. The thermal load distribution is assumed constant in the thickness direction of the composite panel. The random acoustic load is represented as stationary white-Gaussian random pressure with zero mean and uniform magnitude over the panels. The nonlinear equations of motion of the composite panel under combined loads are derived using the principle of virtual work and the finite element method. The static displacement, which is the solution of the nonlinear static equation, is calculated using the Newton–Raphson method, and the nonlinear dynamic equation is solved using the Newmark-β time integration method. Using these numerical methods, the nonlinear dynamic analyses are conducted under various loading conditions such as thermal–random acoustic loads, thermal–supersonic aerodynamic loads, and supersonic aerodynamic–thermal–random acoustic loads. Numerical results show the nonlinear dynamic response of the composite thin and thick panels such as snapthrough and limit cycle oscillation responses. Particularly, the snapthrough response is caused when the random acoustic load is applied appropriately to the thermally buckled composite plate when the aerodynamic load is not applied or applied with the relatively small magnitude of the dynamic pressure.

中文翻译：

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组合超音速气动、热和随机声载荷下剪切变形复合板的非线性动态响应

超音速飞行器的蒙皮板结构承受热、声和空气动力的综合载荷。这些组合载荷可能导致高速飞行器面板结构的非线性动态响应。在这些非线性动力响应中，突触和极限循环振荡响应严重影响面板结构的疲劳失效。这项工作使用数值方法研究了当组合超音速空气动力、热和随机声学载荷同时施加到面板结构时的非线性动态响应。为了考虑薄复合板和厚复合板，应用了一阶剪切变形板理论（FSDT）和冯卡曼非线性位移-应变关系。超音速流动中的气动载荷使用一阶活塞理论建模。假设热负荷分布在复合板的厚度方向上是恒定的。随机声学负载表示为固定的白高斯随机压力，面板上的均值为零且幅度均匀。利用虚功原理和有限元方法推导出复合板在复合荷载作用下的非线性运动方程。静态位移是非线性静力方程的解，采用 Newton-Raphson 法计算，非线性动力方程采用 Newmark-β 时间积分法求解。使用这些数值方法，在各种载荷条件下进行非线性动态分析，例如热随机声载荷，热-超音速气动载荷和超音速气动-热-随机声载荷。数值结果显示了复合薄板和厚板的非线性动态响应，例如快速穿透和极限循环振荡响应。特别地，当不施加空气动力载荷或施加相对较小的动态压力时，当随机声学载荷适当地施加到热屈曲复合板时，会引起快速响应。