As a critical aeroelastic issue for aircraft in the transonic flow regime, flutter is of great significance to flight safety and design verification. Conventional blowdown wind tunnels are constrained by test duration and model scaling, making them inadequate for flutter tests of large aircraft. In contrast, large-scale continuous transonic wind tunnels enable long-duration and large-scale model tests, thus serving as a key solution to this problem. This paper systematically reviewed the research progress on flutter test technologies in large-scale continuous transonic wind tunnels from five aspects: environment construction, model technology, support system, measurement method, and prediction method. Firstly, the safety protection system of continuous wind tunnels, the construction of heavy-gas medium environments and corresponding data-correction techniques were elaborated. Secondly, the dynamic similarity design, model manufacturing of composite materials, and ground test verification system for flutter test models were summarized. Thirdly, the development of flutter model support systems, as well as their limitations, was introduced. Finally, the development status of measurement techniques was reviewed. Accurate flutter boundary prediction methods based on subcritical response identification and artificial intelligence algorithms were discussed, and potential future directions were proposed.
Current research on aeroelastic bifurcation predominantly focuses on two-dimensional problems, and there is a lack of effective methods for studying bifurcation in three-dimensional aeroelastic systems, making it difficult to apply such research in practical engineering. This paper took a three-dimensional all-movable wing as an example and investigated systems with free-play nonlinearity and nonlinear energy sink. Both types of nonlinearities can lead to subcritical bifurcation, adversely affecting aeroelasticity. For free-play nonlinearity, an aeroelastic analysis method based on the CFD/CSD coupling approach was established using the fictitious mass method, combined with equivalent linearization for bifurcation analysis. Limit cycle oscillations, the influence of initial conditions, and hysteresis phenomenon were discussed. Notably, under transonic conditions, there is a dip in the limit cycle boundary, and this dip is significantly lower than that in the traditional linear flutter boundary. For nonlinear energy sink, the aeroelastic equations were established based on a reduced-order aerodynamic model, and an equivalent linearized system was constructed using the describing function method. Results show that when the damping coefficient of the energy sink is small, it can even reduce the aeroelastic stability boundary, triggering subcritical bifurcation. By increasing the damping coefficient of the energy sink, the subcritical bifurcation can be transformed into a supercritical bifurcation. Through bifurcation analysis, it is found that the system’s divergence speed is determined by a turning point, beyond which the energy sink cannot control flutter effectively.
The rapid development of multirotor aircraft in the field of urban air mobility has driven the demand for lightweight design, which however reduces structural stiffness and tends to induce instability due to structural-control coupling. To address this issue, a stability analysis method for structural-control coupling in multirotor aircraft was proposed. This method extended beyond the traditional rigid-body assumption by integrating structural elastic modes into dynamic modeling, and established a roll-angle correction model considering multi-order elastic modes based on the mode superposition method. A simulation platform for the lateral roll channel of a quadrotor was developed in Simulink. By combining frequency-domain analysis with time-domain simulation, the effects of elastic modes and control gains on the stability of the attitude control system were systematically investigated. The results show that the antisymmetric first bending mode of the support arms around the x-axis of the aircraft is prone to coupling with the roll-channel control system, manifested as a resonance peak at 4.8 Hz in the frequency domain with a gain margin below 1 dB, leading to divergence in the time-domain response. The control system gain is a key influencing factor of the coupling effect: low gain leads to insufficient control accuracy, while high gain triggers instability. The notch filter can effectively improve system stability by filtering out the elastic signals of key modes. This study reveals the structural-control coupling mechanism of multirotor aircraft and provides a theoretical foundation and practical engineering guidance for the design of control systems for multirotor aircraft.
Variable-geometry lip structures can effectively broaden the operating range and enhance the performance of hypersonic inlets. However, the presence of connection devices leads to a substantial reduction in structural stiffness, making the lip structure highly susceptible to fluid-structure interaction （FSI） problems. In this study, an in-house high-order CFD/CSD coupled solver was employed to investigate the buzz characteristics of a hypersonic inlet under lip FSI. The results demonstrate that lip FSI induces instability of the downstream shock train, leading to the premature onset of inlet buzz. Compared with the rigid inlet buzz, the buzz induced by FSI exhibits two novel buzz modes: namely violent buzz and mild buzz, which alternate in an irregular manner. During violent buzz, the inlet briefly starts, resulting in a fundamentally different evolution of the flowfield structure from that of the rigid inlet buzz. During mild buzz, the inlet remains unstarted, and thus the evolution of the flowfield structure is similar to that of the rigid inlet buzz. This study deepens the understanding of the buzz characteristics of hypersonic inlets under lip FSI and provides a new perspective on the mechanisms responsible for buzz initiation.
The pitching moment of inertia of flexible flying-wing aircraft is relatively low, making such aircraft prone to coupling with low-order elastic modes and inducing body-freedom flutter, which poses a serious threat to flight safety. For flexible flying-wing aircraft, an aeroelastic sub-scale model design method was proposed, which accurately reflected the body-freedom flutter characteristics of the flying-wing baseline model. Firstly, based on dimensional theory, the aeroelastic scaling similarity laws for flying-wing aircraft were derived to obtain the key parameters that the sub-scale model must satisfy. Then, a baseline model of the flying-wing aircraft was established, and the design of the sub-scale model of the flying-wing aircraft was completed based on the baseline model. An aeroelastic analysis was conducted to preliminarily validate its effectiveness. Finally, flight tests were carried out to further verify the similarity relationship between the sub-scale model and the baseline model. The test results indicate that the flying-wing test model designed based on the aeroelastic scaling similarity criteria can accurately reflect the body-freedom flutter characteristics of the baseline model.
