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The aerothermoelastic behavior of a conical shell in supersonic flow is studied in the paper. According to Love’s first approximation shell theory, the kinetic energy and strain energy of the conical shell are expressed and the aerodynamic model is established by using the linear piston theory with a curvature correction term. By taking the characteristic orthogonal polynomial series as the admissible functions, the mode function of conical shell under different boundary conditions can be obtained using the Rayleigh–Ritz method. Then, the dynamic model of the conical shell is derived by using the Lagrange equation. Based on the model, variations in the natural frequencies with respect to temperature and free-stream static pressure are analyzed. Additionally, the effects of the length-to-radius ratio, the thickness-to-radius ratio, and semi-vertex angle, as well as the thermal and aerodynamic loads on the aerothermoelastic stability of the structure are investigated in detail.

Hypersonic vehicles are susceptible to considerable aerodynamic heating and noticeable aerothermoelastic effects during flight due to their high speeds. Functionally graded materials (FGMs), which enable continuous changes in material properties by varying the ratio of different materials, provide both thermal protection and load-bearing capabilities. Therefore, they are widely used in thermal protection structures for hypersonic vehicles. In this work, the aerothermoelastic behaviors of functionally graded (FG) plates under arbitrary temperature fields are analyzed via a semianalytical method. This research develops a method considering the influence of thermal loading, specifically the decrease in stiffness due to thermal stresses, as well as the correlation between material properties and temperatures under arbitrary temperature fields, based on Ritz’s method. The classical plate theory, von–Karman’s large defection plate theory and piston theory are employed to formulate the strain energy, kinetic energy and external work functions of the system. This paper presents a novel analysis of static aerothermoelasticity of FG plates, in addition to the linear/nonlinear flutter under arbitrary temperature fields, such as uniform, linear and nonlinear temperature fields. In addition, the effects of the volume fraction index, dynamic pressure, and temperature increase on the aerothermoelastic characteristics of FG plates are analyzed.

A body flap/RCS-integrated configuration is often used to achieve pitch trimming and controlled flight in near space for hypersonic vehicles. Under the high temperature and pressure load induced by the expansion wave at the nozzle exit, the body flap is prone to significant structural deformation, which leads to a change in the resulting moment, even comparable to the control ability, and bring additional challenges to the control system. Based on the CFD/CTD/CSD coupling method, the aerothermoelastic effect on the aerodynamic characteristics and structural deformation of the body flap under jet interaction is systematically studied. Numerical results indicate that the pitching moment coefficients show an increasing trend for all the models, rigid, elastic and thermoelastic, while the increment significantly decreases with the increase in trajectory altitudes. With the increase in deflection angle, the pitching moment coefficients of the three models decrease nonlinearly at high altitude, and the aerothermoelastic effect significantly decreases. At a middle-lower altitudes, the pitching moment coefficient is reversed at a lager deflection angle, and the trailing edge of the body flap presents the deformation characteristics of upward bending, which makes the aerothermoelastic phenomenon degenerate into an aeroelastic problem. From the station along the chord and spanwise direction, the change in displacement increment of the thermoelastic model reflects the competitive relationship between normal stress and thermal stress imposed by jet interaction.

The stability of a nonlinear aero-thermo-elastic panel in supersonic flow is analyzed numerically. In light of Hamilton’s principle, the governing equation of motion for a two-dimensional aero-thermo-elastic panel is established taking geometric nonlinearity and curvature effect into account. Coupling with the panel vibration, aerodynamic pressure is evaluated by first order supersonic piston theory and aerothermal load is approximated by the quasi-steady theory of thermal stress. A Galerkin method based on approximate inertial manifolds is deduced for low-dimensional dynamic modeling. The efficiency of the method is discussed. Finally, the complex stability regions of the system are presented within the parametric space. The Hopf bifurcation is found during the onset of flutter as the dynamic pressure increases. The temperature rise imposes a significant effect on the stability region of the panel. Since the material parameters of the panel (elastic modulus and thermal expansion coefficient in this case) are the function of temperature, the panel tends to lose its stability as the temperature gets higher.

This work develops a framework for SIMP-based topology optimization of a metallic panel structure subjected to design-dependent aerodynamic, inertial, elastic, and thermal loads. Multi-physics eigenvalue-based design metrics such as thermal buckling and dynamic flutter are derived, along with their adjoint-based design derivatives. Locating the flutter point (Hopf-bifurcation) in a precise and efficient manner is a particular challenge, as is outfitting the optimization problem with sufficient constraints such that the critical flutter mode does not switch during the design process. Results are presented for flutter-optimal topologies of an unheated panel, thermal buckling-optimal topologies, and flutter-optimality of a heated panel (where the latter case presents a topological compromise between the former two). The effect of various constraint boundaries, temperature gradients, and (for the flutter of the heated panel) thermal load magnitude are assessed. Off-design flutter and thermal buckling boundaries are given as well.

To fulfill the design objective of a structure and thermal protection system, accurate load environment prediction is very important, so we present a high-fidelity aerothermoelastic load calculation method based on a partitioned computational fluid dynamics/computational structural dynamics/computational thermal dynamics (CFD/CSD/CTD) coupling analysis. For the data transformation between the CFD/CSD/CTD systems, finite element interpolation (FEI) is explored, and a shape-preserving grid deformation strategy is achieved via radical basis functions (RBFs). Numerical results are presented for validation of the proposed CFD/CSD/CTD coupling analysis. First, a simply supported panel in hypersonic flow is investigated for results comparison of the proposed coupling method and previous work. Second, a hypersonic forebody is investigated to explore the aerothermoelastic effects while considering the feedback between deformation and aerodynamic heating. The results show that the CFD/CSD/CTD coupling method is accurate for analysis of aerothermoelasticity. In addition, considering the aerothermoelastic effect, the shear force and bending movement increase with time before 900s and decrease after 900s, and at 900s increased percentages of 5.7% and 4.1% are observed, respectively. Therefore, it is necessary to adopt high-fidelity CFD/CSD/CTD coupling in the design of a structure and thermal protection system for hypersonic vehicles.

Summary In this paper, the optimal vibration control of a rotating, pre‐twisted, single‐celled box thin‐walled beam made of functionally graded material is discussed. This beam is under aerothermoelastic loading and warping restraint. A first‐order shear deformation theory enabling satisfaction of traction‐free boundary conditions is employed to achieve governing dynamical model. These equations include the effects of the presetting angle, the secondary warping, temperature gradient through the wall thickness of the beam, and also the rotational speed. Moreover, quasi‐steady aerodynamic pressure loadings are determined using first‐order piston theory, and steady beam surface temperature is obtained from gas dynamics theory. The extended Galerkin method is then used to transform the blade partial differential equations into a set of ordinary differential equations. Transversely isotropic sensor‐actuator piezoelectric pairs that are surface embedded along the blade are also considered for the purpose of closed‐loop control. An optimal observer‐based output feedback control scheme is used to stabilize the closed‐loop system. Simulation studies demonstrate the effectiveness of the proposed method.

Aerothermoelasticity plays a vital role in the design and optimisation of hypersonic aircraft. Furthermore, the transient and nonlinear effects of the harsh thermal and aerodynamic environment a lifting surface is in cannot be ignored. This article investigates the effects of transient temperatures on the flutter behavior of a three-dimensional wing with a control surface and compares results for transient and steady-state temperature distributions. The time-varying temperature distribution is applied through the unsteady heat conduction equation coupled to nonlinear aerodynamics calculated using 3rd order piston theory. The effect of a transient temperature distribution on the flutter velocity is investigated and the results are compared with a steady-state heat distribution. The steady-state condition proves to over-compensate the effects of heat on the flutter response, whereas the transient case displays the effects of a constantly changing heat load by varying the response as time progresses.

Natural modes and flutter of the sweptback wing are analyzed at different temperatures. Natural frequencies are reduced as temperature increasing, the temperature gradient in the structure tend to reduce model stiffness due to changing in material properties and development of thermal stress, which degrade the structure stability. Results in this paper show that the 1st~4th order natural frequencies of sweptback wing in this paper decrease by about 12.09%, 15.06%, 14.31% and 12.57%respectively from room temperature to 450°C.The flutter frequency and velocity are also reduced as the decreasing of natural frequency due to the elevating of temperature by aero-heating; and the flutter frequency and velocity decrease by about 12.42% and 16.69% respectively as the temperature elevating from normal up to 450°C at Ma=3.0.