Elevator failure may have fatal consequences for fighter aircraft that are unstable due to their high manoeuvrability requirements. Many studies have been conducted in the literature using active and passive fault-tolerant control structures. However, these studies mostly include sophisticated controllers with high computational load that cannot work in real systems. Considering the multi-functionality and broad operational prospects of fighter aircraft, computational load is very important in terms of applicability. In this study, an integrated fault-tolerant control strategy with low computational load is proposed without sacrificing the ability to cope with failures. This control strategy switches between predetermined controllers in the case of failure. One of these controllers is designed to operate in a non-failure condition. This controller is a basic controller that requires very little computational effort. The other controller operates when an asymmetric elevator failure occurs. This controller is a robust fault-tolerant controller that can fly the aircraft safely in case of elevator failure. The switching is decided by a failure detection system. The proposed integrated fault-tolerant control system is verified by non-linear F-16 flight simulations. These simulations show that the proposed method can cope with failures but requires less computational load because it uses a conventional controller in the case of no failure.

A novel synthetic air data estimation method without using air data sensors is presented, and the method only relies on the information from the Navigation System (NS) and Flight Control System (FCS). The aircraft's aerodynamic model is also required to make a connection between the FCS control parameters and the NS measurements. The airspeed, angle of attack and sideslip, angular velocity and wind speed are defined as state vectors, and state equations are established through the aircraft's aerodynamic model and dynamics. Linear velocity and angular velocity provided by the navigation system are considered as the measurement vector. To deal with variable wind fields, a novel Initialised Three-step Extended Kalman Filter (ITEKF), which considers the wind speed as an unknown input, is developed to track the variation of wind speed. Simulation results based on a Generic Hypersonic Vehicle (GHV) model are presented and compared with an existing method. Factors affecting the method's accuracy include the navigation system accuracy and the aerodynamic model error, are also discussed.

Helicopter downwash impact on ship airwake is addressed from a three-pronged approach: (1) Analysis of one-point statistics of autospectrum and two-point statistics of cross-spectrum and coherence from a Computational Fluid Dynamics database of flow velocities effected by helicopter downwash and shipboard airwake; (2) Development of a mathematical framework for extracting interpretive turbulence models in closed form from these autospectral statistics; and (3) Simulation through white-noise-driven filters for the extracted models. The framework begins with an earlier-exercised perturbation-type series expansion of the autocorrelation for all three velocity components, where the first term of the series has a form of the von Karman longitudinal or lateral correlation function. After transformation into equivalent series of autospectrum, the coefficients in the series are evaluated by satisfying theoretical constraints and fitting a curve on a set of selected autospectral data points generated from the database. The framework represents a sensible combination of series expansion, exploitation of a database, and theoretical constraints to provide a foothold on airwake-downwash phenomenon for engineering analysis. It ensures that the extracted model and the autospectral data points have the same time scale, mean square value, and asymptotic decay according to the Kolmogorov –5/3 Law. The framework's strengths and weaknesses, and its major advancement over the earlier series-expansion schemes are also addressed. Finally it is shown that downwash increases airwake energy (mean square value) by one order of magnitude, and almost all of this airwake-downwash energy is concentrated within the bandwidth (0.16 < f(Hz) < 1.6) that affects flight mechanics.

Efforts to develop the next generation of aircraft with ever-increasing levels of performance – higher, farther, faster, cheaper – face great technical challenges. One of these technical challenges is to reduce structural weight of the aircraft. Another is to look to aircraft configurations that have been unrealizable to date. Both of these paths can lead to a rigid flex coupling phenomenon that can result in anything from poor flying qualities to the loss of an aircraft due to flutter. This has led to a need to develop an integrated flight and aeroelastic control capability where structural dynamics are included in the synthesis of flight control laws. Studies have indicated that the application of an integrated flight and aeroelastic control approach to a SensorCraft high-altitude long-endurance vehicle would provide substantial performance improvement(1,2). Better flying qualities and an expanded flight envelope through multi-flutter mode control are two areas of improvement afforded by integrated flight and aeroelastic control. By itself, multi-flutter mode control transforms the flutter barrier from a point of catastrophic structural failure to a benign region of flight. This paper discusses the history and issues associated with the development of such an integrated flight and aeroelastic control system for the X-56A aircraft.

Artillery rockets are sensitive to disturbances (total impulse variation, wind, thrust misalignment, etc.). As the range of ground-to-ground rockets increases, the accuracy of free flight rockets decreases. Requirements for the increase of the range and minimisation of the impact point dispersions can be solved by adding guidance and control systems to rockets. Based on the differences between the measured flight parameters and the calculated parameters for the nominal trajectory, the flight path angle correction algorithm is obtained by adding the correction to the nominal value. The flight path steering guidance system with lateral acceleration autopilot in the inner loop, is used for the guidance of hypothetical artillery rockets. The guidance algorithm given in this paper eliminates the time correction due to rocket velocity variations. Efficiency of the proposed algorithm is illustrated by numerical simulation. There is no practical influence of the external disturbances on impact point dispersion. It is also shown that the measurement errors of the inertial measurement unit are the dominant factor affecting impact point dispersion of artillery rockets modified by adding a guidance system based on the flight path steering method.

Machine vision-based sensing enables automatic hover stabilization of helicopters. The evaluation of image data, which is produced by a camera looking straight to the ground, results in a drift free autonomous on-board position measurement system. No additional information about the appearance of the scenery seen by the camera (e.g. landmarks) is needed. The technique being applied is a combination of the 4D-approach with two dimensional template tracking of a priori unknown features.