The high operational cost of aircraft, limited availability of air space, and strict safety regulations make training of fighter pilots increasingly challenging. By integrating Live, Virtual, and Constructive simulation resources, efficiency and effectiveness can be improved. In particular, if constructive simulations, which provide synthetic agents operating synthetic vehicles, were used to a higher degree, complex training scenarios could be realised at low cost, the need for support personnel could be reduced, and training availability could be improved. In this work, inspired by the recent improvements of techniques for artificial intelligence, we take a user perspective and investigate how intelligent, learning agents could help build future training systems. Through a domain analysis, a user study, and practical experiments, we identify important agent capabilities and characteristics, and then discuss design approaches and solution concepts for training systems to utilise learning agents for improved training value.

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.

The rotorcraft industry needs Virtual Engineering first to ensure decisions made early in the life-cycle, at the requirements capture and preliminary design phases for example, are reliably informed. Later, in design, development and qualification, Virtual Prototypes can become the centre of attention for critical reviews and, ultimately, certification itself. A significant challenge is to ensure that model fidelity is good enough, not only for supporting design decisions but also in establishing requirements based on sufficiently mature technologies. This international conference, Rotorcraft virtual engineering; supporting life-cycle engineering through design and development, test and certification and operations and co-sponsored by the RAeS/AHS/A3F/DGLR/AIDAA addressed these themes and this paper reviews and assesses the value of the various contributions.

As simulation has become an integral part of the overall life-cycle support of aircraft, the need for effective Virtual Engineering (VE) tools to support these activities has increased. FLIGHTLAB is a state-of-the-art, aircraft modelling and simulation software tool, that has been designed to address this need and is widely used in rotorcraft design, analysis, test and evaluation, and full-flight simulation applications. This VE tool supports the development and analysis of both fixed and rotary wing aircraft with an extensive library of modelling components which have been successfully used and validated in numerous, real-world applications. These components provide comprehensive modelling of aerodynamic, structural, control and propulsion disciplines. Analyses include performance, dynamic response, stability and control, airloads, and structural loads. Graphical User Interfaces and an interactive scripting language provide user-friendly operation. This paper describes the capabilities and validation activities that have been undertaken to support the development of the commercial VE toolset FLIGHTLAB over the last 20 years and discusses future rotorcraft challenges that could be addressed by enhancements to current generation VE tools.

This paper adds data to help establish fidelity criteria to accompany the simulator motion system diagnostic test specified by the International Civil Aviation Organisation. Twelve airline transport pilots flew three tasks in the NASA Vertical Motion Simulator under four different motion conditions. The experiment used three different hexapod motion configurations, each with a different trade-off between motion filter gain and break frequency, and one large motion configuration that utilised as much of the simulator's motion space as possible. The motion condition significantly affected (1) pilot motion fidelity ratings, and sink rate and lateral deviation at touchdown for the approach and landing task, (2) pilot motion fidelity ratings, roll deviations, maximum pitch rate, and number of stick shaker activations in the stall task, and (3) heading deviation after an engine failure in the take-off task. Significant differences in pilot-vehicle performance were used to define initial objective motion cueing criteria boundaries. These initial fidelity boundaries show promise but need refinement.

This design study applied parameterisation to rotor blade for improved performance. In the design, parametric equations were used to represent blade planform changes over the existing rotor blade model. Design variables included blade twist, sweep, dihedral and the radial control point. Updates to the blade structural properties with changes in the design variables allowed accurate evaluation of performance objectives and realistic structural constraints – blade stability, steady moments (flap bending, chord bending and torsion) and the high-g manoeuvre pitch link loads. Performance improvement was demonstrated with multiple parametric designs. Using a parametric design with advanced aerofoils, the predicted power reduction was 1.0% in hover, 10.0% at μ = 0.30 and 17.0% at μ = 0.40, relative to the baseline UH-60A rotor, but these were obtained with a 35% increase in the steady chord bending moment at μ = 0.30 and a 20% increase in the half peak-to-peak pitch link load during the UH-60A UTTAS manoeuvre. Low vibration was maintained for this design. More rigorous design efforts, such as chord tapering and/or structural redesign of the blade cross section, would enlarge the feasible design space and likely provide significant performance improvement.

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.

Since the introduction of hexapod-type motion systems for flight simulation in the 1970s, Motion Drive Algorithm tuning has been primarily based on the subjective judgement of experienced pilots. This subjective method is often not transparent and often leads to ambiguous process of adjustment of the tuning parameters. Consequently, there are large variations in the motion cueing characteristics of flight training devices, a variability that subsequently raises questions regarding the value of motion cueing for pilot training itself. The third revision of ICAO 9625 Manual of Criteria for the Qualification of Flight Simulation Training Devices offered the opportunity to take a closer look at simulator motion cueing requirements in general. This led to the concept of the objective motion cueing test (OMCT), which was reported in 2006. After the method was evaluated on three research flight simulators, the results were published in 2007, demonstrating a larger spread in dynamic behaviour of cueing algorithms than expected. After discussions with the simulator industry regarding the form and methodology of the OMCT, an evaluation of the test in cooperation with the industry started in 2011. This led to the final form of the OMCT and cueing parameter criterion for the in-flight mode of transport aircraft. This paper describes the OMCT, the evaluation results and the criterion.

In this paper, we present a new force-feedback side-stick which has been developed and integrated into a research flight simulator. The developed 2 Degrees of Freedom (DOF) force-feedback joystick, as a kind of haptic device, provides two-way communication in both position and force, and allows users to interact with the simulation system. It has been designed by considering the main factors in designing a general use force-feedback device. Thus, the design must allow the restitution of aerodynamic forces onto the hand of the pilot. This is an important feature, which gives the pilot the ‘natural feel’ of traditional mechanical aircraft control. In order to provide the force feedback to enhance the realism of the simulation, we added the necessary software using Commercial-Off-the-Shelf (COTS) solutions (Microsoft Flight Simulator Software (MSFS)) and built-in data structure and methods. Thus, the main contribution of this paper concerns the design and implementation of an automatic controller based on fuzzy logic systems. It is not simply designing a force-feedback stick for flight simulation: we proposed a novel control principles and more importantly completely new approach to compute in real-time force feedback on the stick based on pilot knowledge that avoids the use of complex aerodynamics equations with unknown parameters. To our best knowledge, this work is the first to propose the integration of fuzzy logic force controller in flight simulation for creating force feedback. Results using the overall simulation are presented and evaluated and interesting sensations have been recorded.

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.