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Understanding the performance and flight envelope of a damaged aircraft is a preliminary requirement to recover the aircraft after damage. This paper aims to provide a comprehensive understanding of wing damage effect on airplane performance, local stability, and flying quality of each trim state inside the achievable flight envelope.

This paper demonstrates the use of attainable equilibrium points which are referred as trim states in order to estimate a damaged airplane manoeuvring flight envelope using a numerical computation method.

Wing damaged airplane manoeuvring flight envelope is estimated for different portions of the wing tip loss. Local stability at each trim condition inside the estimated flight envelope is analysed, and also motion flight modes and flying quality sensitivity to the wing damage are explored.

Local stability and flying quality analysis at each trim condition inside the flight envelope which demonstrate the effect of damage provides a criterion to prioritize the choice of trimmed flight condition as motion primitives for the airplane post‐damage flight and safe landing.

The purpose of this paper is to describe the longitudinal dynamics of a hover‐capable rigid‐winged unmanned aerial vehicles (UAV) under various equilibrium flight conditions. The effects of the variable‐incidence wing in comparison with the fixed in‐incidence wing on the dynamics of UAV are also discussed.

The aerodynamic modeling of the vehicle covers both pre‐stall and post‐stall regimes using a three‐dimensional vortex lattice method incorporating viscous corrections. The trim states across a velocity spectrum are evaluated using a nonlinear constrained optimization scheme based on sequential quadratic programming. Then linearized dynamic analysis around trim states is carried out in order to compare the characteristics of the conventional platform with the modified platform incorporating variable‐incidence wing.

It is found that with the variable‐incidence wing, the longitudinal equilibrium flights can be achieved with reduced thrust‐to‐weight ratio demands and lower elevator deflection. However, the use of the variable‐incidence wing changes the dynamic characteristics of the vehicle considerably as indicated through the linear dynamic analysis.

The results presented in this paper are based on linear dynamic analysis about static trim point data. Further analysis taking into account nonlinearity, the unsteady aerodynamic effects and associated cross‐coupling because of asymmetric forces may be needed to reveal the true dynamics of the vehicle under unsteady maneuvers.

The variable‐incidence wing is a useful design feature to reduce the thrust‐to‐weight ratio requirements and to increase elevator control authority, however its effect on the dynamics warrants further investigation.

This is the first paper highlighting the effects of variable‐incidence wing on an agile hover‐capable UAV.

The purpose of this paper is to examine the cross‐coupled responses of a coupled rotor‐fuselage flight dynamic simulation model, including a finite‐state inflow aerodynamics and a coupled flap‐lag and torsion flexible blade structure.

The methodology is laid out based on model development for an articulated main rotor, using the theories of aeroelastisity, finite element and finite‐state inflow formulation. The finite‐state inflow formulation is based on a 3D unsteady Euler‐based concepts presented in the time domain. The most advantages of the model are the capability of modeling dynamic wake effects, tip losses and skewed wake aerodynamics. This is, in fact, a special type of the inflow model relating inflow states, to circulatory blade loadings through a set of first‐order differential equations. A non‐iterative solution of the differential equations has practically altered the model into a simple and direct formulation appending properly to the rest of the helicopter mathematical model. A non‐linear distribution of the induced velocity over the rotor disc is finally obtained by the use of both Legendre polynomials and higher‐harmonic functions. Ultimately, validations of the theoretical results show that the on‐axis response, direct reaction to the pilot input, has a good accuracy both quantitatively and qualitatively against flight test data, and the off‐axis response, cross‐coupled or indirect reaction to the pilot input are improved by this approach of modeling.

Improvements in dynamic prediction of both trim control settings and dynamic cross‐coupled responses of helicopter to pilot inputs are observed.

Further work is required for investigation of the augmented finite state inflow model, including the wake rotation correction factors to describe helicopter maneuvering flight characteristics.

The results of this work support the future researches on design and development of advanced flight control system, incorporating a high bandwidth with low‐phase delay to control inputs and also high levels of dynamic stability within minimal controls cross coupling.

This paper provides detailed characteristics on the mathematical integration problems associated with the advanced helicopter flight dynamics research.

This paper aims to propose output feedback-based control algorithms for the flight control system of a scaled, un-crewed helicopter in its hover flight mode.

The proposed control schemes are based on H_∞ control and composite nonlinear control. The gains of the output feedback controllers are obtained as the solution of a set of linear matrix inequalities (LMIs).

In the proposed schemes, the finite-time convergence of system states to trim condition is achieved with minimum deviation from the steady-state. As the proposed composite nonlinear output feedback design improves the transient response, it is well suited for a scaled helicopter flight. The use of measured output vector instead of the state vector or its estimate for feedback provides a simple control structure and eliminates the need for an observer in real-time application. The proposed control strategies are relevant to situations in which a simple controller is essential due to economic factors, reliability and hardware implementation constraints.

The proposed control strategies are relevant to situations in which a simple controller is essential due to economic factors, reliability and hardware implementation constraints. They also have significance in applications where the number of measurement quantities needs to be minimized such as in a fully functional rotor-craft unmanned aerial vehicle.

The developed output feedback control algorithms can be used in small-scale helicopters for numerous civilian and military applications.

This work addresses the LMI-based formulation and solution of an output feedback controller for a hovering un-crewed helicopter. The stability and robustness of the closed-loop system are proved mathematically and the performance of the proposed schemes is compared with an existing strategy via simulation studies.

This paper aims to provide a mathematical modeling and design of H-infinity controller for an autonomous vertical take-off and landing (VTOL) Quad Tiltrotor hybrid unmanned aerial vehicles (UAVs). The variation in the aerodynamics and model dynamics of these aerial vehicles due to its tilting rotors are the key issues and challenges, which attracts the attention of many researchers. They carry parametric uncertainties (such as non-linear friction force, backlash, etc.), which drives the designed controller based on the nominal model to instability or performance degradation. The controller needs to take these factors into consideration and still give good stability and performance. Hence, a robust H-infinity controller is proposed that can handle these uncertainties.

A unique VTOL Quad Tiltrotor hybrid UAV, which operates in three flight modes, is mathematically modeled using Newton–Euler equations of motion. The contribution of the model is its ability to combine high-speed level flight, VTOL and transition between these two phases. The transition involves the tilting of the proprotors from 90° to 0° and vice-versa in 15° intervals. A robust H-infinity control strategy is proposed, evaluated and analyzed through simulation to control the flight dynamics for different modes of operation.

The main contribution of this research is the mathematical modeling of three flight modes (vertical takeoff–forward, transition–cruise-back, transition-vertical landing) of operation by controlling the revolutions per minute and tilt angles, which are independent of each other. An autonomous flight control system using a robust H-infinity controller to stabilize the mode of transition is designed for the Quad Tiltrotor UAV in the presence of uncertainties, noise and disturbances using MATLAB/SIMULINK. This paper focused on improving the disturbance rejection properties of the proposed UAV by designing a robust H-infinity controller for position and orientation trajectory regulation in the presence of uncertainty. The simulation results show that the Tiltrotor achieves transition successfully with disturbances, noise and uncertainties being present.

A novel VTOL Quad Tiltrotor UAV mathematical model is developed with a special tilting rotor mechanism, which combines both aircraft and helicopter flight modes with the transition taking place in between phases using robust H-infinity controller for attitude, altitude and trajectory regulation in the presence of uncertainty.

The use of offsets is one of the main characteristics of international defence trade. The rising costs of defence equipment and the significant contraction of defence spending have resulted in an environment that favoured the use of offset policies, the latter becoming increasingly demanding in both quantitative and qualitative terms. The chapter analyses the role of offsets on the process of integration of defence equipment markets, with a specific focus on the EU. Particular attention is given to the offset-relevant regulation and practice and to their recent evolution in the EU following the adoption of European Directive on defence and security procurement (81/81/EC). Offsets play a dual role with regard to the integration of defence industries: on one hand they can be trade-distorting and contribute to the survival of inefficient suppliers in arms importing countries; on the other hand, they can contribute in overcoming barriers that may otherwise prevent some potentially efficient suppliers from accessing the supply chains of the big system integrators. The chapter draws the attention on the need to complement the regulatory evolution by further initiatives aiming at improving the access of non-incumbent suppliers to the supply chains of the large defence system integrators.

This chapter develops robust panel estimation in the form of trimmed mean group estimation for potentially heterogenous panel regression models. It trims outlying individuals of which the sample variances of regressors are either extremely small or large. The limiting distribution of the trimmed estimator can be obtained in a similar way to the standard mean group (MG) estimator, provided the random coefficients are conditionally homoskedastic. The authors consider two trimming methods. The first one is based on the order statistic of the sample variance of each regressor. The second one is based on the Mahalanobis depth of the sample variances of regressors. The authors apply them to the MG estimation of the two-way fixed effects model with potentially heterogeneous slope parameters and to the common correlated effects regression, and the authors derive limiting distribution of each estimator. As an empirical illustration, the authors consider the effect of police on property crime rates using the US state-level panel data.

The purpose of this paper is to model the aircraft-cargo’s coupling dynamics during ultra-low altitude heavy cargo airdrop and to design the aircraft’s robust flight control law counteracting its aerodynamic coefficients perturbation induced by ground effect and the disturbance from the sliding cargo inside.

Aircraft-cargo system coupling dynamics model in vertical plane is derived using the Kane method. Trimmed point is calculated when the cargo fixed in the cabin and then the approximate linearized motion equation of the aircraft upon it is derived. The robust stability and robust H_∞ optimal disturbance restraint flight control law are designed countering the aircraft’s aerodynamic coefficients perturbation and the disturbance moment, respectively.

Numerical simulation shows the effectiveness of the proposed control law with elevator deflection as a unique control input.

The model derived and control law designed in the paper can be applied to heavy cargo airdrop integrated design and relevant parameters choice.

The dynamics model derived is closed, namely, the model can be called in numerical simulation free of assuming the values of parachute’s extraction force or cargo’s relative sliding acceleration or velocity as seen in many literatures. The modeling is simplified using Kane method rather than Newton’s laws. The robust control law proposed is effective in guaranteeing the aircraft’s flight stability and disturbance restraint performance in the presence of aerodynamic coefficients perturbation.

The purpose of this paper is to study the control strategy of transition mode of the stopped-rotor (SR) aircraft under the condition of redundant control and complex aerodynamic characteristics.

This paper first proposes a transition strategy for the conversion between helicopter mode and fixed-wing mode. Then, aiming at the redundancy of the two control systems in the transition process, a control model is proposed, which greatly simplifies the control in conversion mode. Then, to facilitate the design of the control system, the Takagi-Sugeno model of the SR aircraft in transition mode is established. Finally, an explicit model tracking and tuning parameter stability augmentation control system is designed, so that the SR aircraft has a good stability during the transition process. Then, the outer loop control system of transition flight is designed.

The simulation results show that the control strategy proposed in this paper can realize the mode conversion well. It lays a solid foundation for the subsequent engineering flight test for the SR aircraft.

The work done in this paper provides ideas and methods for the flight control system design of SR aircraft in transition mode. The method of designing control model to solve the coordination of redundant control system is also applicable for other multimode aircraft, which provides a simple and convenient method for the multimode aircraft control.

Whereas the Minister of Labour (hereafter in this Order referred to as the “Minister”) has received from the Brush and Broom Wages Council (Great Britain) the wages regulation proposals set out in Schedules 1 and 2 hereof;