A linearization method and an engineering approach for the geometric nonlinear aeroelastic stability analysis of the very flexi- ble aircraft with high-aspect-ratio wings are established based on the little dynamic perturbation assumption.The engineering practicability of the method is validated by a complex example.For a high-altitude long-endurance unmanned aircraft,the nonlinear static deformations under straight flight and the gust loads are calculated.At the corresponding nonlinear equilibrium state,the complete aircraft is linearized dynamically and the vibration modes are calculated considering the large deformation effects.Then the unsteady aerodynamics are calculated by the double lattice method.Finally,the aeroelastic stability of the complete aircraft is analyzed.The results are compared with the traditional linear calculation.The work shows that the geometric nonlinearity induced by the large structural deformation leads to the motion coupling of the wing chordwise bending and the torsion,which changes the mode frequencies and mode shapes.This factors change the aeroelastic coupling relationship of the flexible modes leading to the decrease of the flutter speed.The traditional linear method would give not only an imprecise flutter speed but also a possible dramatic mistake on the stability.Hence,for a high-altitude long-endurance unmanned aircraft with high-aspect-ratio wings,or a similar very flexible aircraft,the geometric nonlinear aeroelastic analysis should be a necessary job in engineering practice.
The effectiveness of the sliding mode control(SMC) method for active flutter suppression(AFS) and the issues concerning control system discretization and control input constraints were studied using a typical two-dimensional airfoil.The airfoil has a trailing-edge flap for flutter control.The aeroelastic system involves a two-degrees-of-freedom motion(pitch and plunge),and the equations were constructed by utilizing quasi-steady aerodynamic forces.The control system,designed by the output feedback SMC method,was incorporated to suppress the pitch-plunge flutter.Meanwhile,the system discretization and the flap deflection constraints were implemented.Then,a classical Runge-Kutta(RK) algorithm was utilized for numerical calculations.The results indicated that the close-loop system with the SMC system could be stable at a speed above the flutter boundary.However,when the flap deflection limits are reached,the close-loop system with the simple discretized control system loses control.Furthermore,control compensation developed by theoretical analysis was proposed to make the system stable again.The parameter perturbations and the time delay effects were also discussed in this paper.
YANG Chao,SONG Chen,WU ZhiGang & XIE ChangChuan School of Aeronautic Science and Engineering,Beijing University of Aeronautics & Astronautics,Beijing 100191,China
Model uncertainty directly affects the accuracy of robust flutter and limit-cycle-oscillation (LCO) analysis. Using a data-based method, the bounds of an uncertain block-oriented aeroelastic system with nonlinearity are obtained in the time domain. Then robust LCO analysis of the identified model set is performed. First, the proper orthonormal basis is constructed based on the on-line dynamic poles of the aeroelastic system. Accordingly, the identification problem of uncertain model is converted to a nonlinear optimization of the upper and lower bounds for uncertain parameters estimation. By replacing the identified memoryless nonlinear operators by its related sinusoidal-input describing function, the Linear Fractional Transformation (LFT) technique is applied to the modeling process. Finally, the structured singular value (μ) method is applied to robust LCO analysis. An example of a two-degree wing section is carried out to validate the framework above. Results indicate that the dynamic characteristics and model uncertainties of the aeroelastic system can be depicted by the identified uncertain model set. The robust LCO magnitude of pitch angle for the identified uncertain model is lower than that of the nominal model at the same velocity. This method can be applied to robust flutter and LCO prediction.
Missiles may be damaged when aeroservoelastic problem occurs,which is caused by the interaction of structure flexibility and flight control system.Because of the limit of wind tunnel test condition,numerical methods are mostly used in previous aeroservoelastic studies.However,series of assumptions and simplification on structures,aerodynamics and flight control systems are unavoidably introduced,and various nonlinear factors are also ignored,therefore,they result in considerable errors.A novel method called aeroservoelasticity semi-physical simulation test is proposed in this paper,which takes the flexible missile with control system as the test object.Vibration signals at several locations of the missile are measured by accelerometers,then corresponding unsteady aerodynamics is computed based on the fact that airflow at high Mach is nearly quasi-steady,and finally unsteady aerodynamics is exerted simultaneously by shakers at certain locations of the missile.The aeroservoelasticity semi-physical simulation test system can be constructed after the control system is closed.Open loop transfer function test and closed loop stability test are carried out in sequence.The test principle and method proposed in this paper are verified by the concordance between the results of numerical simulation and experiment.
A high-altitude long-endurance aircraft with high-aspect-ratio wing usually generates large deformation,which brings the geometric nonlinear aeroelastic problems.In recent decades,it has become a key focus of the international researchers of aeroelasticity.But some critical technologies are not developed systematically,such as aerodynamic calculation methods of the curved wing with deformation,moreover,there are few experimental validations of these technologies.In this paper,we established the steady aerodynamic calculating method of the curved wing with quite large deformation based on the extended lifting line method,and calculated the unsteady aerodynamics using the strip theory considering curved surface effects.Combining the structure geometrical nonlinear finite element method,we constructed a systematic analytic approach for the static aeroelasticity and flutter of very flexible wing,and further designed the ground vibration and wind tunnel test to verify this approach.Through the test and the theoretic results comparison,we concluded that the extended lifting line method has adaptable precision for the static aeroealsticity and the strip theory considering curved surface effects for flutter analysis can give exact critical speed and flutter mode when the dynamic stall does not happen.The work in this paper shows that the geometric nonlinear aeroelastic analytic approach for very flexible wing has very high efficiency and adaptable precision.It can be used in the engineering applications,especially the iterated design in preliminary stage.
A new hybrid approach to constructing reduced-order models(ROM)of unsteady aerodynamics applicable to aeroelastic analysis is presented by using proper orthogonal decomposition(POD)in combination with observer techniques.Fluid modes are generated through POD by sampling observations of solutions derived from the full-order model.The response in the POD training is projected onto the fluid modes to determine the time history of the modal amplitudes.The resulting data are used to extract the Markov parameters of the low-dimensional model for modal amplitudes using a related deadbeat observer.The state-space realization is synthesized from the system's Markov parameters that are processed with the eigensystem realization algorithm.The POD-observer method is applied to a two-dimensional airfoil system in subsonic flow field.The results predicted by the ROM are in general agreement with those from the full-order system.The ROM obtained by the hybrid approach captures the essence of a fluid system and produces vast reduction in both degrees of freedom and computational time relative to the full-order model.
YANG Chao,LIU XiaoYan &WU ZhiGang Beijing University of Aeronautics and Astronautics,Beijing 100191,China
