This paper presents the first results from research into active control of structural load alleviation (SLA) for tiltrotor aircraft carried out in the European ‘critical technology’ RHILP project. The importance of and the need for SLA in tiltrotors are discussed, drawing on previous US experience reported in the open literature. The paper addresses the modelling aspects in some detail; hence forming the foundation for both the FLIGHTLAB simulated XV-15 and EUROTILT configurations. The primary focus of attention is the suppression of in-plane rotor yoke loads for pitch manoeuvres in airplane mode; without suppression these loads would result in a very high level of fatigue damage. Multi-variable control law design methods are used to develop controller schemes and load suppression of 80-90% is demonstrated using rotor cyclic control, albeit at a 20-30% performance penalty. However, rotor flapping transients tend to increase by the action of the SLA system. A dual-objective control design approach demonstrates the effectiveness of suppressing both loads and flapping simultaneously.

Operating micro-UAVs autonomously in complex urban areas requires that the guidance algorithms on-board are robust to changes in the operating environment. Limited endurance capability demands an optimal guidance algorithm, which will change as the environment does. All optimal path-planning routines are computationally intensive, with processor load a function of the environmental complexity. This paper presents a new algorithm, the reactive route selection algorithm, for storing a bank of optimal trajectories computed off-line and blending between these optimal trajectories as the operating environment changes. An example is presented using a mixed-integer linear program to generate the optimal trajectories.

Weapon target allocation (WTA) is an effective method to solve the battlefield fire optimisation problem, which plays an important role in intelligent automated decision-making. We researched the multitarget allocation problem to maximise the attack effectiveness when multiple interceptors cooperatively attack multiple ground targets. Firstly, an effective and reasonable fitness function is established, based on the situation between the interceptors and targets, by comprehensively considering the relative range, relative angle, speed, capture probability and radiation source matching performance and thoroughly evaluating them based on the advantage of the attack effectiveness. Secondly, the optimisation performance of the particle swarm optimisation (PSO) algorithm is adaptively improved. We propose an adaptive simulated annealing-particle swarm optimisation (SA-PSO) algorithm by introducing the simulated annealing algorithm into the adaptive PSO algorithm. The proposed algorithm can enhance the convergence speed and overcome the disadvantage of the PSO algorithm easily falling into a local extreme point. Finally, a simulation example is performed in a scenario where ten interceptors cooperate to attack eight ground targets; comparative experiments are conducted between the adaptive SA-PSO algorithm and PSO algorithm. The simulation results indicate that the proposed adaptive SA-PSO algorithm demonstrates great performance in convergence speed and global optimisation capabilities, and a maximised attack effectiveness can be guaranteed.

The issue of the Royal Aeronautical Society Structures Data Sheets 02.01.28 to 37* prompts the publication of a short survey of the theory on which these sheets are based and of the manner in which they may be applied. To save the need to refer to original sources the theory is developed from first principles; but for brevity the application is illustrated only by one simple example. Other more elaborate applications are illustrated by the example data sheets themselves. A recent paper by Chilver covers the same ground from the mathematical point of view and the purpose of the present paper is to illustrate the practical application of the method to familiar problems in the computation of initial buckling stress for sheet-stiffener combinations.

The current understanding of periodic transonic flow is reviewed briefly. The effects of boundary-layer transition, non-adiabatic wall conditions and modifications to the aerofoil surface geometry at the shock interactions on periodic transonic flow are discussed. Through the methods presented, it is proposed that the frequency of periodic motion can be predicted with reasonable accuracy, but there are limitations on the prediction of buffet boundaries associated with periodic transonic flows. Several methods have been proposed by which the periodic motion may be virtually eliminated, most relevantly by altering the position of transition fix, contouring the aerofoils surface or adding a porous surface and a cavity in the region of shock interaction. In addition, it has been shown that heat transfer can have a significant effect on buffet.

Unmanned aerial vehicles (UAV) have seen a rapid growth in utilisation for reconnaissance, mostly using single UAVs. However, future utilisation of UAVs for applications such as bistatic synthetic aperture radar and stereoscopic imaging, will require the use of multiple UAVs acting cooperatively to achieve mission goals. In addition, to de-skill the operation of UAVs for certain applications will require the migration of path-planning functions from the ground to the UAV. This paper details a computationally efficient algorithm to enable path-planning for single UAVs and to form and re-form UAV formations with active collision avoidance. The algorithm presented extends classical potential field methods used in other domains for the UAV path-planning problem. It is demonstrated that a range of tasks can be executed autonomously, allowing high level tasking of single and multiple UAVs in formation, with the formation commanded as a single entity.

The vorticity production by a single vortex generator jet in a turbulent boundary layer at high speed and zero pressure gradient was studied by analysing surface-flow patterns. The investigation involved a range of jet mass flow rates and skew angles while the freestream Mach number and jet pitch angle were held constant. A qualitative analysis of the skin friction patterns demonstrated the effective role of the skew angle as a bifurcation parameter. The vortex production is classified into three main regimes as the skew angle increases from zero (fully downstream) to 180° (fully upstream). Better understanding of the physical mechanism behind the transition from the first regime to the second is crucial in the design of the airjet vortex generator.

The design, manufacture and testing of a 480 cubic metre helium-filled airship, or dirigible balloon, intended for flight above tropical rain forest canopy is outlined. The requirements for the airship and the preliminary sizing method are set-out. Details of the final design and flight test results are included, as well as suggestions concerning future work. Particular attention is paid to the issue of accurate station-keeping and operation close to the canopy with winds present. It is concluded that the development of a truly useful aerial platform for canopy biological studies requires further research in aerodynamics, dynamic simulation and ultra-quiet propulsion systems.

The paper presents the aerodynamic characterization of a low-speed unmanned aerial vehicle, with cropped delta planform and rectangular cross section, at and around high angles-of-attack using flight test methods. Since the linear models used for identification from flight data at low and moderate angles of attack become unsuitable for accurate parameter estimation at high angles of attack, a non-linear aerodynamic model has to be considered. Therefore, the Kirchhoff's flow separation model was used to incorporate the non-linearity in the aerodynamic model in terms of flow separation point and stall characteristic parameters. The Maximum Likelihood (ML) and Neural Gauss-Newton (NGN) methods were used to perform the parameter estimation on one set of low angle-of-attack and one set of near-stall flight data. It is evident from the estimates that the NGN method, which does not involve solving equations of motion, performs on a par with the classical ML method. This may be attributed to the reason that NGN method uses a neural network which has been trained by performing point to point mapping of the measured flight data. This feature of NGN method enhances its application over a wider envelope of high angles of attack flight data.

Structural integrity is a key ingredient of aircraft design. It takes its place alongside flight performance, operational effectiveness, reliability, maintainability and economy in determining, or contributing in very large measure to, the success of a project in service.

What is expected from the contractor for a new project in respect of these qualities is set down by the customer in a specification which includes reference to a set of mandatory design requirements. Thus, flight performance parameters such as speed, range and payload, are defined in the specification : there too is to be found a description of the sorties to be flown, numbers and types of manoeuvres and their severity (conventionally including the levels of normal acceleration) and the regions in which the aircraft is required to operate. Sufficient detail is given to establish, as the process of design evolves, the loading and environmental conditions to which the structure will be subjected.

An initial value method is introduced in this paper for the solution of the two-point non-linear ordinary differential equations resulting from an analysis of the MHD boundary-layer flow originally treated by Greenspan and Carrier. By using this method, the iteration process is eliminated. The method is seen to be applicable to the solution of similar two-point boundary value problems where certain physical parameters appear either in the differential equation or in the boundary conditions and solutions for a range of the parameter are sought.

A simple circuit is described which gives a direct meter reading of the intermittency factor in a turbulent flow. The operation and setting-up procedure of the instrument is discussed.

While expansion deflection (ED) nozzles have traditionally been considered primarily for use as altitude compensating devices to improve the performance of single stage to orbit vehicles, they also offer the potential for enhancing high altitude propulsion systems. If intended to only operate in near vacuum conditions, the complexity of analysis and inherent risks involved in the ED concept are greatly reduced. An integrated approach to the design and performance analysis of such nozzles is presented, comprising a mixture of computational fluid dynamics, the method of characteristics, and a semi-empirical model to allow full analysis of the closed wake flow-field of an ED nozzle. While it is demonstrated that the influence of the parameters used to define the throat region is critical to the successful application of the ED nozzle, it is also shown that with careful design the weight savings possible are significant. The analysis method itself is flexible and rapid, and lends itself well to incremental improvements in accuracy as the flow under consideration becomes better understood.

We consider the effects of atmospheric disturbances, such as wind, shears, turbulence and wakes, on the aerodynamics of aircraft, in order (§1) to quantify the phenomena in question, and provide a basis for future application to problems of performance and stability. The present method of description of the effects of non-uniformity and unsteadiness of the incident flow on aircraft aerodynamics, is presented in some detail for a head- or tailwind sheared vertically (§2), and its extension to all three components of velocity and six shear derivatives is outlined (§11). The starting point (§3) is the relative life change due to a uniform or sheared wind; in the latter case we introduce (§6) a dimensionless shear number S, which plays with regard to vorticity, a role similar (§6) to the Mach and Reynolds numbers respectively in connection with compressibility and viscosity. The aerodynamics of a body or aerofoil in an incident stream containing vorticity, is specified (§7) by a shear coefficient Cs, playing a role similar to the lift, drag and moment coefficients for aerodynamic forces and torques. The present theory is consistent with the formulas of theoretical aerodynamics, the results of wind tunnel tests and observation of atmospheric disturbances involving shear flows. The use of the concepts of shear number S and shear coefficient Cs allows the calculation (§10) of the uniform wind equivalent to a given shear, ie, the wind velocity which would cause the same lift change as the shear. This equivalent wind may be useful in the calculation of the aerodynamic effects of sheared flows, eg, forebody vortices on a wing, wing in a propeller slipstream, tailplane in a wing’s wake, an aircraft landing, behind another, or flight through storms or microbursts.

To solve the on-ground lateral direction control problem of the unswept flying-wing unmanned aerial vehicle (UAV) without rudder, steering system or breaking system, a control approach which uses differential propeller thrust to control the lateral direction is proposed. First, a mathematical model of the unswept flying-wing UAV on-ground moving is established. Second, based on the active disturbance rejection control (ADRC) theory, a yaw angle controller is designed by using the differential propeller thrust as the control output. Finally, a straight line trajectory tracking control law is designed by improving the vector field path following method. Experiment results show that the proposed control laws have a shorter response time, better robustness and better control precision compared with proportional integral derivative (PID) controller. The proposed controller has small computational complexity, simple parameter setting process, and uses practical measurable physical quantities, providing a reference solution for further engineering applications.

This paper presents the results of a research performed at the Research Laboratory in Active Controls, Avionics and Aeroservoelasticty (LARCASE), at ÉTS, concerning optimisation strategies for cruise flight segments with imposed flight time (delimited by waypoints with required time of arrival constraints). Specifically, a new algorithm is presented that identifies the optimal vertical navigation profile (flight altitude and speed optimisation) for a cruise segment with imposed lateral navigation profile, bounded by two waypoints with required time of arrival constraints. The set of evaluated vertical navigation profiles are characterised by identical altitudes and speeds at their initial and final waypoints (at the beginning and the end of the cruise segment under optimisation), a maximum of one altitude step (relative to the initial altitude), and are flown at constant speed. This study investigates the flight performance increase (total cost reduction) for a flight along the optimal vertical navigation profile, relative to a flight at the optimal speed and initial cruise altitude. The evaluation was performed using a medium haul transport aircraft flight performance model, for three lateral navigation profiles and three wind profiles. The algorithm is targeted for Flight Management Systems platforms, to provide the optimal flight trajectory for the imposed lateral flight profile and time constraints.

An approximate analytical method is presented for determining the swirling compressible flow through a nozzle. The method is developed from the techniques introduced by Carpenter and Johannesen. It is perfectly general in that the swirl distribution need not be specified a priori but the radial gradients of the stagnation enthalpy and entropy are assumed to be small. The principal assumption is that changes in the nozzle cross-sectional area are sufficiently gradual and smooth for the radial velocity component to be neglected at each section, i e the usual assumption of one-dimensional compressible flow theory. Analytical expressions are derived for various flow characteristics, e g mass-flux coefficient, impulse function and back-pressure ratio required for choking. The analytical results are compared to numerical results for two main classes of swirling flows. On the whole the analytical results are found to be good approximations for moderate swirl levels. The approximate numerical method is found to be reasonably successful at predicting flow reversal at the nozzle wall. The main result is that a swirl parameter is found such that, with the sole exception of the free-vortex case, the mass-flux-coefficient data for all swirling flows investigated collapse onto a single curve.

An approximate theory, which yields the minimum induced drag of a planar wing with end-plates in ground effect, is presented. As a check on the accuracy of the method, it is compared with the exact linearised theory in the case of end-plates of vanishing depth (planar configuration). This shows that, within the limitations of the linearised lifting-surface theory, the present method is accurate in the range of wing heights of interest. In Section 5 the theory is compared with experiment. In this comparison two interpretations of the theory are employed. The first assumes that the vortex trace is the projection of the trailing edge of the configuration onto a plane normal to the direction of motion when the wing incidence is zero. The second is different from this only in that no constraint is placed on the wing incidence. Of the two interpretations the first is found to be in better agreement with experiment than the second for the planar configuration. However, the opposite is found to be true when end-plates of sufficient depth are fitted. A reason for this is proposed in Section 5.

Aspects of the turbulent wind environment Micro Air Vehicles (MAVs) experience when flying outdoors were replicated in a large wind tunnel. An overview of the facility, instrumentation and initial flight tests is given. Piloting inputs and aircraft accelerations were recorded on fixed and rotary wing MAVs and for some tests, measurements of the approach flow (u,v,w sampled at 1,250Hz at four laterally disposed upstream locations) were made. The piloting aim was to hold straight and level flight in the 12m wide × 4m high × ~50m long test section, while flying in a range of turbulent conditions. The Cooper-Harper rating system showed that a rotary craft was less sensitive to the effects of turbulence compared to the fixed wing craft and that while the fixed wing aircraft was relatively easy to fly in smooth air, it became extremely difficult to fly under high turbulence conditions. The rotary craft, while more difficult to fly per. se., did not become significantly harder to fly in relatively high turbulence levels. However the rotary craft had a higher mass and MOI than the fixed wing craft and further work is planned to understand the effects of these differences.