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The aerodynamic layout of the Canard Rotor/Wing(CRW) aircraft in helicopter flight mode differs significantly from that of conventional helicopters. In order to study the flight dynamics characteristics of CRW aircraft in helicopter mode, first, the aerodynamic model of the main rotor system is established based on the blade element theory and wind tunnel test results. The aerodynamic forces and moments of the canard wing, horizontal tail, vertical tail and fuselage are obtained via theoretical analysis and empirical formula. The flight dynamics model of the CRW aircraft in helicopter mode is developed and validated by flight test data. Next, a method of model trimming using an optimization algorithm is proposed. The flight dynamics characteristics of the CRW are investigated by the method of linearized small perturbations via Simulink. The trim results are consistent with the conventional helicopter characteristics, and the results show that with increasing forward flight speed, the canard wing and horizontal tail can provide considerable lift,which reflects the unique characteristics of the CRW aircraft. Finally, mode analysis is implemented for the linearized CRW in helicopter mode. The results demonstrate that the stability of majority modes increases with increasing flight speed. However, one mode that diverges monotonously,and the reason is that the CRW helicopter mode has a large vertical tail compared to the conventional helicopter. The results of the dynamic analysis provide optimization guidance and reference for the overall design of the CRW aircraft in helicopter mode, and the model developed can be used for control system design. 相似文献
354.
This paper addresses the fixed-time adaptive model reference sliding mode control for an air-to-ground missile associated with large speed ranges, mismatched disturbances and un-modeled dynamics. Firstly, a sliding mode surface is developed by the tracking error of the state equation and the model reference state equation with respect to the air-to-ground missile. More specifically,a novel fixed-time adaptive reaching law is presented. Subsequently, the mismatched disturbances and the un-modeled dynamics are treated as the model errors of the state equation. These model errors are estimated by means of a fixed-time disturbance observer, and they are also utilized to compensate the proposed controller. Therefore, the fixed-time controller is obtained by an adaptive reaching law and a fixed-time disturbance observer. Closed-loop stability of the proposed controller is established. Finally, simulation results including Monte Carlo simulations, nonlinear six-DegreeOf-Freedom(6-DOF) simulations and different ranges are presented to demonstrate the efficacy of the proposed control scheme. 相似文献
355.
《中国航空学报》2020,33(10):2679-2693
In recent years, the Active Flutter Suppression (AFS) employing Linear Parameter-Varying (LPV) framework has become a hot spot in the research field. Nevertheless, the flutter suppression technique is facing two severe challenges. On the one hand, due to the fatal risk of flight test near critical airspeed, it is hard to obtain the accurate mathematical model of the aeroelastic system from the testing data. On the other hand, saturation of the actuator may degrade the closed-loop performance, which was often neglected in the past work. To tackle these two problems, a new active controller design procedure is proposed to suppress flutter in this paper. Firstly, with the aid of LPV model order reduction method and State-space Model Interpolation of Local Estimates (SMILE) technique, a set of high-fidelity Linear Time-Invariant (LTI) models which are usually derived from flight tests at different subcritical airspeeds are reduced and interpolated to construct an LPV model of an aeroelastic system. And then, the unstable aeroelastic dynamics beyond critical airspeed are ‘predicted’ by extrapolating the resulting LPV model. Secondly, based on the control-oriented LPV model, an AFS controller in LPV framework which is composed of a nominal LPV controller and an LPV anti-windup compensator is designed to suppress the aeroelastic vibration and overcome the performance degradation caused by actuator saturation. Although the nominal LPV controller may have superior performance in linear simulation in which the saturation effect is ignored, the results of the numerical simulations show that the nominal LPV controller fails to suppress the Body Freedom Flutter (BFF) when encountering the actuator saturation. However, the LPV anti-windup compensator not only enhances the nominal controller’s performance but also helps the nominal controller to stabilize the unstable aeroelastic system when encountering serious actuator saturation. 相似文献
356.
K.J. Badcock S. TimmeS. Marques H. KhodaparastM. Prandina J.E. MottersheadA. Swift A. Da RonchM.A. Woodgate 《Progress in Aerospace Sciences》2011,47(5):392-423
In this paper the use of eigenvalue stability analysis of very large dimension aeroelastic numerical models arising from the exploitation of computational fluid dynamics is reviewed. A formulation based on a block reduction of the system Jacobian proves powerful to allow various numerical algorithms to be exploited, including frequency domain solvers, reconstruction of a term describing the fluid-structure interaction from the sparse data which incurs the main computational cost, and sampling to place the expensive samples where they are most needed. The stability formulation also allows non-deterministic analysis to be carried out very efficiently through the use of an approximate Newton solver. Finally, the system eigenvectors are exploited to produce nonlinear and parameterised reduced order models for computing limit cycle responses. The performance of the methods is illustrated with results from a number of academic and large dimension aircraft test cases. 相似文献
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《中国航空学报》2020,33(12):3073-3081
Transition prediction is a hot research topic of fluid mechanics. For subsonic and transonic aerodynamic flows, eN method based on Linear Stability Theory (LST) is usually adopted reliably to predict transition. In 2013, Coder and Maughmer established a transport equation for Tollmien-Schlichting (T-S) instability so that the eN method can be applied to general Reynolds-Average-Navier-Stokes (RANS) solvers conveniently. However, this equation focuses on T-S instability, and is invalid for crossflow instability induced transition which plays a crucial role in flow instability of three-dimensional boundary layers. Subsequently, a transport equation for crossflow instability was developed in 2016, which is restricted to wing-like geometries. Then, in 2019, this model was extended to arbitrarily shaped geometries based on local variables. However, there are too many tedious functions and parameters in this version, and it can only be used for incompressible flows. Hence, in this paper, after a large amount of LST analyses and parameter optimization, an improved version for subsonic and transonic boundary layers is built. The present improved model is more robust and more concise, and it can be applied widely in aeronautical flows, which has great engineering application value and significance. An extensive validation study for this improved transition model will be performed. 相似文献