In this study, an experimental investigation is performed on a scaled, all-movable horizontal tail to study the aeroelastic behaviors induced by multiple free-plays. The dynamic response in wind tunnel tests is measured by strain gauges, an accelerometer, and a binocular vision measurement system. The obtained results indicate that the present aeroelastic system exhibits highly nonlinear characteristics and undergoes two independent limit cycle oscillations (LCOs) induced by bending free-play and torsion free-play, respectively. Further, various parametric studies are conducted to evaluate the effects of the free-play angles, angle of attack, flow velocity, and gust excitation on the LCOs. It is found that the value of free-play angle has no significant effect on the critical flow velocity which leads to the occurrence of LCOs. The amplitude and frequency of LCOs increase with the increasing free-play angle and flow velocity. Moreover, the horizontal tail experiences high-order harmonic resonances when LCOs appear. Finally, the stability of limit cycles is analyzed based on the gust excitation experiment. Overall, compared to an all-movable horizontal tail with single free-play, the multiple free-plays system exhibits more complex dynamic behaviors. In this paper, the measured results of the scaled model, which has a similar mass distribution and stiffness distribution as actual aircraft, may be valuable for predicting such LCOs induced by multiple free-plays, and providing a reference for the design of all-movable horizontal tail to prevent LCOs.

Free-play-induced nonlinear dynamic behavior has been one of the most important topics of aeroelastic research in recent decades. In this paper, the describing function (DF) method is developed to investigate the complex dynamic response of a popular all-movable horizontal tail with free-play. Piecewise expressions for the time history and phase portrait of limit cycle oscillation (LCO) are derived by the developed DF method, which is conducive to understand the mechanism of free-play-induced LCO. Another advantage of the developed DF method is the ability to predict the high-order harmonics, which cannot be realized by the classic DF method. A three-dimensional (3D) all-movable horizontal tail model with torsional free-play was designed and manufactured to implement wind tunnel tests via various initial parameters. A good agreement was found between the numerical and experimental results, which can demonstrate the effectiveness of the proposed method. The influence of the initial parameters of the all-movable horizontal tail on the LCO characteristics is analyzed by both numerical calculations and wind tunnel tests. The method and results in this paper can provide a significant reference for the design of all-movable horizontal tail versus the free-play-induced LCO.

The V-tail configuration has excellent stealth performance and has been using widely in the aerodynamic shape design of advanced aircraft. Many recent studies have focused on numerical simulation about V-tail configuration flight performance. The relative wind tunnel tests still need to be developed. This challenge is a focused aspect in such research. In the present experimental study, the role of flight control law was investigated in order to keep the test model in the target attitude and height. An effective design method of a full model of the aircraft with twin V-tails is proposed based on CFD evaluation. This model was manufactured based on the design of a two degrees of freedom support system via a Chinese wind tunnel. A longitudinal flight control law was proposed and simulated. Wind tunnel tests were employed to find the effectiveness of the model design and the control law. It is seen from the results that the proposed experimental method via a full model of the aircraft with twin V-tails and a novel longitudinal flight control law is effective. These test results can provide appliable contributions on the development of the support system for wind tunnel experiments. The proposed model design and test methods can be useful for applications in the aeroelastic wind tunnel tests of the full model aircrafts.

This paper presents an innovative aeroelastic solution method called the implicit dynamic approach (IDA). The IDA is a novel technique implemented to address the complexities of structural displacement and fluid–structure interactions. At the core of this method lie the Navier–Stokes equations, which are pivotal for resolving the unsteady aerodynamic forces for studying aeroelasticity. This paper develops a time-step-coupling data-solving interface that bridges the structural dynamic and fluid dynamic domains, ensuring a seamless integration of these two critical aspects of aeroelastic analysis. To substantiate the credibility of the IDA, the aeroelastic responses of a two-dimensional wing and the benchmark supercritical wing (BSCW) are analyzed. The results obtained from these models demonstrate the IDA’s credibility and stability. The IDA proves to be reliable in predicting aeroelastic responses and offers a powerful tool for analyzing aeroelastic problems.

The aerodynamic configuration of large swept-back tri-wings is generally adopted for hypersonic vehicles, but the structural stiffness of the ailerons is weak, which may lead to damage due to the flutter behavior. In the initial stage of structural design, studying the flutter characteristics of tri-wing flight vehicles is necessary and can provide the stiffness index of the tri-wing structural design. To assess the flutter characteristics of tri-wing flight vehicles efficiently, a rapid modeling technique of the finite element method was used in this paper. For the structural scheme of large swept-back tri-wing flight vehicles, a structural dynamic model was modeled using the rapid modeling technique, the unsteady aerodynamic was computed using the double-lattice method, and the flutter characteristics were analyzed using the P-K method. Variable parametric studies were conducted to evaluate the effects of the stiffness of the aileron skin, the stiffness of the control mechanism, and the mass distribution of the aileron on the flutter characteristics of large swept-back tri-wing flight vehicles. The results showed that the key flutter coupling modes of such vehicles are symmetric and anti-symmetric combinations of aileron rotation and torsion. Additionally, optimizing the control mechanism stiffness and mass distribution of the aileron could improve the flutter boundary, which can be helpful in the structural design of such vehicles. The flutter optimization technique effectively improved the flutter boundary, significantly enlarged the flight envelope, and accurately provided the stiffness index for the structural design of large swept-back tri-wing flight vehicles.

Large transport aircraft tend to adopt a wing layout with a high aspect ratio and swept-back angle due to the requirement of a high lift-to-drag ratio. Composite material is typically employed to ensure the light weight of the structure, causing serious static aeroelasticity problems to the aircraft. When the airplane is flying in the transonic region, its aerodynamic load is very complex, and the large load leads to large deformation of the wing, triggering geometric nonlinear effects, which further affects the static aerodynamic elasticity characteristics of the wing. In this study, in order to study the static aeroelastic characteristics of the transonic flow of a high-aspect-ratio airfoil, a new design method of the scaled similar optimization model is described, and the change in the model lift coefficient due to geometrically nonlinear static aeroelasticity effects when the angle of attack is changed was investigated by using simulation and wind tunnel test methods. In order to ensure the accuracy of the wing shape when the model was deformed greatly, this study employed the structural design scheme of the wing with the skin as the main stiffness component, and the thicknesses of different regions of the skin were used as the design variables for the stiffness optimization design. The engineering algorithm of nonlinear finite elements was used in this study to calculate the curve of lift with the angle of attack considering the geometric nonlinear static aeroelasticity effect. The results show that the similarity optimization process employed in this study can be used to complete the design of the high-speed aerostatic wing test model, and the wind tunnel test results show that geometric nonlinearity has a large impact on the lift coefficient of the wing.