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This study investigates the receptance method for the partial pole assignment of time-delay nonlinear systems using multi-input control from measurement output feedback (i.e., acceleration, velocity and displacement). The receptance method has a remarkable advantage compared to other methods in that there is no need to know the mass, damping and stiffness matrices of the system, which are typically obtained from the finite element method. We achieve partial assignment of the desired poles with no spillover using the assigned and unchanged poles and the corresponding eigenvectors of the closed-loop system. We used different types of generalised inverse matrices to obtain the realisable control gains. The modal constraints for the assigned eigenvectors are thus obtained. Because certain components of the measurement output were found to be unmeasurable, a numbering system is proposed for determining zero elements in the control gains. Then, realisable control gains are obtained after zero-column substitutions are made in the corresponding matrix with the numbering system. Our theoretical results show that having multi-input control from the measurement output feedback is effective for a partial pole assignment with time delay in structures. This theory is demonstrated by several numerical examples of a three-degree-of-freedom damped mass-spring system.

New output feedback-linearisation theory is presented for the treatment of nonlinear vibration problems by a receptance-based approach. An important aspect is a new formulation for investigating the stability of the zero dynamics. The overall methodology possesses the usual benefits of the receptance method, namely that the system matrices (with associated assumptions and approximations) do not have to be known. In addition, it has the distinction of not requiring the form and parameter values of the nonlinearity when the input and output degrees of freedom are away from the nonlinearity itself. This represents a valuable advance over the conventional time-domain feedback linearisation approach.

A method to augment range and endurance for a propeller-driven unmanned aerial vehicle (UAV) is presented in this paper. The basic idea is to find the optimal matching parameters among the flight conditions, the engine operation states, and the propeller characteristics. This can be achieved by minimizing fuel consumption per hour, and the fuel consumption per kilometer, respectively. The relationship between the required power and the available one is set up to satisfy the specific flight condition, and the calculation of parameters is discussed. A sequential quadratic programming (SQP) procedure is employed to solve the problem. Results show that the presented method can increase the endurance by 35% and the range by 11%, compared with the flight test.

The aeroelastic analysis of high-altitude, long-endurance (HALE) aircraft that features high-aspect-ratio flexible wings needs take into account structural geometrical nonlinearities and dynamic stall. For a generic nonlinear aeroelastic system, besides the stability boundary, the characteristics of the limit-cycle oscillation (LCO) should also be accurately predicted. In order to conduct nonlinear aeroelastic analysis of high-aspect-ratio flexible wings, a first-order, state-space model is developed by combining a geometrically exact, nonlinear anisotropic beam model with nonlinear ONERA (Edlin) dynamic stall model. The present investigations focus on the initiation and sustaining mechanism of the LCO and the effects of flight speed and drag on aeroelastic behaviors. Numerical results indicate that structural geometrical nonlinearities could lead to the LCO without stall occurring. As flight speed increases, dynamic stall becomes dominant and the LCO increasingly complicated. Drag could be negligible for LCO type, but should be considered to exactly predict the onset speed of flutter or LCO of high-aspect-ratio flexible wings.

An analytical model of hydraulic damper was presented in forward flight accounting for pitch/flap/lag kinematic coupling and its nonlinear force-velocity curve. The fourth order Runge-Kutta was applied to calculate the damper axial velocity in time domain. Fourier series based moving block analysis was applied to calculate equivalent linear damping in terms of transient responses of damper axial velocity. Results indicate that equivalent linear damping will be significantly reduced if pitch/flap/lag kinematic coupling introduced for notional model and flight conditions.

This paper presents a methodology to navigate a micro aerial vehicle (MAV) flying through a cluttered and unknown environment. Due to the limited payload and power availability of the MAV platform, the guidance algorithm should be simple enough to be performed onboard with low computation and power consumption. The approach proposed here employs a miniature scanning laser rangefinder as a sensor to learn continuously about the surroundings, and a reflexive method is used to plan a collision-free path based on the information given by the sensor. To track the path, a simple difference controller is implemented on the MAV platform. The proposed methodology was tested in several typical scenarios. Simulation results show that the designed system provides an effective and robust solution to the navigation problem. The vehicle reached the destination point without collision to encountered obstacles.

Air-breathing hypersonic vehicles (HSVs) are typically characterized by interactions of elasticity, propulsion and rigid-body flight dynamics, which may result in intractable aeroservoelastic problem. When canard is added, this problem would be even intensified by the introduction of low-frequency canard pivot mode. This paper concerns how the aeroservoelastic stability of a canard-configured HSV is affected by the pivot stiffnesses of all-moveable horizontal tail (HT) and canard. A wing/pivot system model is developed by considering the pivot torsional flexibility, fuselage vibration, and control input. The governing equations of the aeroservoelastic system are established by combining the equations of rigid-body motion, elastic fuselage model, wing/pivot system models and actuator dynamics. An unsteady aerodynamic model is developed by steady Shock-Expansion theory with an unsteady correction using local piston theory. A baseline controller is given to provide approximate inflight characteristics of rigid-body modes. The vehicle is trimmed for equilibrium state, around which the linearized equations are derived for stability analysis. A comparative study of damping ratios, closed-loop poles and responses are conducted with varying controller gains and pivot stiffnesses. Available bandwidth for control design is discussed and feasible region for pivot stiffnesses of HT and canard is given.