Non-linear behaviour is usually undesirable for the operation of any system and it needs to be minimised in flight control systems in order to provide the pilot with a predictable and well-behaved aircraft. The non-linear response characteristics exhibited by all aircraft and their flight control systems can be characterised by using nonlinear functions to simulate and analyse their behaviour. This paper explores the types of physical non-linearities that exist in aircraft flight control systems and provides some examples. It describes how the non-linear stability might be analysed and addresses the subject of actuation systems modelling. The final section provides recommendations for minimising non-linear behaviour and gives some general guidelines on how to deal with non-linear characteristics.

This paper describes a method for using inverse simulation to obtain a preliminary assessment of helicopter handling qualities. Formal descriptions of standard manoeuvres, defined to establish the handling qualities of military helicopters, are used to drive an inverse simulation of a subject helicopter. The simulation generates the controls and states of the helicopter as it executes the manoeuvre and the results may be used to calculate values of quickness, a parameter defined to measure responsiveness. Initial results reveal that in the context of inverse simulation quickness is independent of vehicle configuration when, as specified in the requirements, the quickness is based on the helicopter's kinematic states. An alternative quickness parameter, associated with the control displacements required to fly the manoeuvre is shown to be capable of discriminating between the pilot workload involved in flying two different configurations through the same manoeuvre.

The “post-stall” manoeuvre capabilities of current and projected combat aircraft have prompted a resurgence of interest in the dynamic response to wing motion of the vortex burst at high angles of attack. A large number of experimental studies have been reported, but these have tended to be restricted to a limited range of motion profiles and a single specific wing configuration. Presumably due to the complexity of the phenomenon, no attempt appears to have been made to undertake a systematic study of the burst response for a range of wing planforms and motion profiles. Hence, a survey was made of the published data and a semi-empirical correlation derived using a system identification methodology. For preliminary performance studies, the dynamic response of the vortex burst is modelled adequately by a second order lag transfer function. The physical significance of this system model and the implications for the design of a systematic experimental study are discussed.

Measurements in two-dimensional flow on rectangular cylinders confirm earlier work of Nakaguchi et al in showing a maximum drag coefficient when the height h of the section (normal to the stream) is about 1.5 times the width d. Reattachment on the sides of the cylinder occurs only for h/d < 0.35.

For cylinders of D-shaped section (Fig 1) the pressure distribution on the curved surface and the drag are considerably affected by the state of the boundary layer at separation, as for a circular cylinder. The lift is positive when the separation is turbulent and negative when it is laminar. It is found that simple empirical expressions for base pressure or drag, based on known values for the constituent half-bodies, are in general not satisfactory.

Design optimisation of a helicopter rotor blade is performed. The objective is to reduce helicopter vibration and constraints are put on frequencies and aeroelastic stability. The ply angles of the D-spar and skin of the composite rotor blade with NACA 0015 aerofoil section are considered as design variables. Polynomial response surfaces and space filling experimental designs are used to generate surrogate models of the objective function with respect to cross-section properties. The stacking sequence corresponding to the optimal cross-section is found using a real-coded genetic algorithm. Ply angle discretisation of 1°, 15°, 30° and 45° are used. The mean value of the objective function is used to find the optimal blade designs and the resulting designs are tested for variance. The optimal designs show a vibration reduction of 26% to 33% from the baseline design. A substantial reduction in vibration and an aeroelastically stable blade is obtained even after accounting for composite material uncertainty.

A detailed and accurate three-dimensional finite element stress analysis was conducted on countersunk rivet holes in a plate subjected to tension loading. The analysis included a wide range of countersunk depths, plate thicknesses, countersunk angles and plate widths. The study confirmed some of the previous results, addressed their differences, provided many new results, and investigated countersunk angle and width effects. Using the detailed FE results and the limiting conditions, a design equation for stress concentration was developed and verified.

Morgan has described a spatial instability in the flow through screens or grids of small open-area ratio. Head and Rechenberg and others have observed large span-wise variations in the thickness and shear stress of nominally two-dimensional boundary layers on flat plates and aerofoils in wind tunnels. It now appears that these spanwise variations are caused by the instability of flow through the screens. The jets of air issuing from the pores of the screen attempt to entrain more air by the usual mixing processes, but can only entrain it from each other, so that groups of jets coalesce in rather random (steady) patterns determined by small irregularities in the weave. The resulting variations in axial velocity are virtually eliminated by the wind tunnel contraction, but variations in flow direction are not so greatly reduced: a theoretical analysis shows that the observed variations of boundary-layer thickness, which often reach ± 10 per cent of the mean, can be produced by directional variations in the working section of the order of ± 1/20 deg, with a spanwise wavelength of the same order as the boundary-layer thickness.

A method for describing the action of a pilot in mathematical terms is presented. The method applies for tracking tasks and gives a pilot describing function which can be used in conventional servo-analysis. Flight simulator data obtained in this way for a wide variety of aircraft longitudinal response characteristics are presented and the effect of these characteristics on the pilot's action is discussed fully. In order to illustrate that work of this kind can lead to a quantitative understanding of pilot's “handling qualities” opinion, the recorded pilot opinions are “predicted” from a knowledge of three simple aircraft characteristics.

In many important problems of supersonic flow, either for the whole field of flow or a part of it, the velocity components are constant on straight lines through a fixed point. Such velocity fields are called conical fields. In the usual theory of the linearised perturbations of a steady supersonic flow, the velocity is assumed to differ only slightly from a uniform undisturbed velocity, and in the defining equations and boundary conditions all non-linear terms in the components of the perturbation velocity (and their space derivatives) are neglected. In this paper the equations of linearised supersonic conical fields, and their general solution, are set out both for the region inside and for the region outside the Mach cone of the origin in the conical field. The results are applied to flow past plane triangular aerofoils with straight edges downstream of the vertex (of which there are six cases), to flow past those plane aerofoils of more extended shape and finite span for which the solution may be obtained by combining a finite number of conical fields, and to the problem of plane triangular vanes in semi-infinite free jets. In the applications, the velocity fields are determined in considerable detail, but a main purpose in setting them out was to exhibit the mathematical methods used and the physical considerations that enter in determining the mathematical solutions.

This paper is concerned with the local buckling of long thin flat-walled structures, such as integrally stiffened panels or corrugated core sandwich panels, loaded in such a way that the individual flats are in combined uniform longitudinal compression and shear. When buckling occurs the line junctions between adjoining flats remain straight, and the flats are subjected on their long edges to sinusoidally varying edge moments. These produce sinusoidally varying edge rotations which, when shear is present, are in general out of phase with each other and with the moments. Relations between the edge moments and rotations are obtained in terms of two stability functions, one of which is real and the other complex, to take account of phase differences. Explicit expressions are derived for these stability functions and tables are included, giving their values for the case of pure shear.

Fatigue of aircraft parts due to small but numerous vibrations has been known about for some time, but in 1944–5 cases of structural fatigue were found which were thought to have been caused by loads greater, but less frequent, than the usual vibration loads. Such loads could be ground loads due to taxi-ing, take-off and landing or they could be air loads caused by gusts and manœuvres.

A general study of the fatigue properties of typical structures was made in the laboratory and it became apparent that the loads likely to do the most damage were those occurring about a million times in the life of the aircraft.

The boundary element method is combined with the method of compatible deformations to obtain stress intensity factors for a cracked sheet reinforced with a repair patch. The method is applied to the analysis of a circular patch over a central crack in a rectangular uniaxially stressed sheet. It is shown that the proximity of the edges of the sheet to the patch edge has a negligible effect on the stress intensity factor of a crack completely under the patch.

This paper presents a rational method of stressing single or many cell tubes particularly of the type encountered in wing structures. The theory has been developed for conical or cylindrical tubes of arbitrary cross-section the shape of which is maintained by a closely spaced system of diaphragms rigid in their own planes and parallel to the root section. Within the limits of the assumptions the theory is exact for right cylindrical tubes, and is applicable with adequate accuracy to cylindrical or conical tubes in which the inclination of any generator to the normal to the root plane does not exceed 10°.

The analysis given unifies the theories of bending and torsion and shows that the commonly used method of separating the bending and torsion loads by means of a shear centre is in general incorrect.

The formulae have been developed in such forms that attention is concentrated on the necessary corrections to the stresses as indicated by the ordinary engineers' theory. These correction terms include all effects of shear lag, diffusion, and end effects hitherto taken into account only in some very special cases.

Particularly important for practical applications is the structure consisting of a number of direct stress carrying members (booms) connected by walls effective only in shear. The simplest structure of the latter type is the four-boom tube with or without nose and trailing cells, and in this case explicit formulae are given which are immediately applicable to practical calculations. Formulae are also given for the more complex case of a six-boom tube in which the two extra booms are introduced to represent more accurately the direct stress carrying capacity of the top and bottom panels between the two spars.

The US Naval Research Laboratory (NRL) is developing technologies that will enable Navy-relevant missions with the smallest practical Micro Air Vehicles (MAVs). The NRL Micro Tactical Expendable (MITE) air vehicle is a result of this research. MITE is a hand-launched, dual-propeller, fixed-wing air vehicle, with a 25cm chord and a wingspan of 25–47cm, depending on payload weight. Vehicle gross weight is 130–350g. Miniature autopilot systems, based on visual imaging techniques, are being developed for MITE. These will be used in conjunction with conventional autopilot sensors to allow the MITE to fly autonomously. This paper provides an overview of the MITE development, including aerodynamic design considerations, electric propulsion, and vision-based autopilot research. Also presented is a rationale for the development of control laws that can direct the behavior of large groups of MAVs or other vehicle agents. Dubbed ‘physicomimetics,’ this process can bring about the self-assembly of complex MAV formations, though individual MAVs have minimal onboard processing power and limited local sensing capabilities.

Current state-of-the-art trajectory prediction tools typically model aircraft as three-dimensional point-masses, and make a number of simplifying assumptions about the actual and anticipated dynamics states of the aircraft. They are typically based on predefined settings obtained from existing databases such as Eurocontrol’s Bada rather than real-time information, including on the environment, available onboard the aircraft. This significantly limits trajectory prediction performance. This paper proposes a high-accuracy four-dimensional trajectory prediction model for use onboard civil aircraft, as well as by ground-based systems, which addresses these limitations. It is designed for strategic traffic capacity optimisation and conflict-detection and resolution over time-horizons covering the entire duration of a flight. The model incorporates a number of features including a novel flight-control-system and an enhanced flight-script that incorporates new taxonomy and content thereby enabling better definition of aircraft intent. The accuracy of the model is characterised using operational data acquired during a real flight trial. Results show that the performance of the proposed model is significantly better than the current models. Its accuracy is better than the required navigation performance for departure, en route and Non-Precision-Approach phases of flight.

The electrification of the commuter aircraft is instrumental in the development of novel propulsion systems. The scope of this work aims to explore the design space of a parallel hybrid-electric configuration with an entry into service date of 2030 and beyond and determine the impact of other disciplines on conceptual design, such as components positioning, aircraft stability and structural integrity. Three levels of conceptual sizing are applied and linked with a parametric aircraft geometry tool, to generate the aircraft’s geometry and position the components. Subsequently, the structural optimisation of the wing box is performed, providing the centre of gravity of the components placed inside the wing, that minimise the induced stresses. Furthermore, the stability and trim analysis follow, with the former being highly affected by the positioning of components. Results are compared to a similar aircraft with entry into service technology of 2014 and it is indicated that in terms of block fuel reduction the total electrification benefit increases with the increase of degree of hybridisation, if aircraft mass is kept constant. On the other hand, if battery specific energy is kept constant, similar block fuel reduction is possible with lower hybridisation degrees. The potential for improvement in terms of carbon dioxide emissions and block fuel reduction ranges from 15.73% to 21.44% compared to the conventional aircraft, for levels of battery specific energy of 0.92 and 1.14 kWh/kg respectively. Finally, the component positioning evaluation indicates a maximum weight limitation of 240 kg for the addition of an aft boundary layer ingestion fan to a tube and wing aircraft configuration, without compromising the aircraft static stability.

The work described here concentrates on under-expanded, axisym-metric turbulent jets issuing into quiescent conditions. Under-expanded turbulent jets are applicable to most aircraft propulsion applications that use convergent nozzles. Experimental studies used laser doppler velocimetry (LDV) and pitot probe measurements along the jet centreline. These measurements were made for two nozzle pressure ratios (2·5 and 4·0) and at various streamwise positions up to 10 nozzle diameters downstream of the nozzle exit plane. A computational fluid dynamics (CFD) model was developed using the Fluent code and utilised the RNG K-ε two-equation turbulence model. A mesh resolution of approximately one hundredth of nozzle exit diameter was found to be sufficient to establish a mesh independent solution.

Comparison of the jet centreline axial velocity (LDV data) and pressure ratio (pitot probe data) showed good agreement with the CFD model. The correct number of shock cells had been predicted and the shock strength agreed well between the experiments and numerical model. The CFD model was, however, found to over-predict the shock cell length resulting in a longer supersonic core. There was some evidence, based on analysis of the LDV measurements that indicates the presence of swirl and jet unsteadiness, which could contribute to a shortening of the shock cell length. These effects were not modelled in the CFD. Correlation between the LDV and pitot probe measurements was generally good, however, there was some evidence that probe interference may have caused the premature decay of the jet. Overall, this work has indicated the good agreement between a CFD simulation using the RNG k-ε turbulence model and experimental data when applied to the prediction of the flowfield generated by under-expanded turbulent jets. The suitability of the LDV technique to jet flows with velocities up to 500ms-1 has also been demonstrated.

The aerodynamic pitching moment, lift and drag forces acting on a ski-jumper during the free flight and landing phases were measured using scaled models in a wind tunnel. Comparisons are made between those meaurements and observations of actual ski-jumps.

It is shown that the high angles of incidence (typically α varies from 35° to 55°) adopted by ski-jumpers during free flight is determined by the need to prevent tumbling. Over this range of incidence angles, the opposing moments due to weight distribution and aerodynamic pressure distribution are approximately in equilibrium, and small out-of-balance moments can be accommodated by arm movements. The maximum lift/drag ratio on a ski-jumper in free flight occurs at α ≃ 25°. If it were possible to avoid tumbling at low incidence angles, maximum jump lengths would be achieved at α ≃ 8°.

Experiments were conducted to study the types of flow that occur on the lee surface of delta wings at supersonic speeds. Two sets of flat topped delta wings of different thickness (wedges with 10° and 25° normal angle respectively), each with leading edge sweep angles of 45°, 50°, 60° and 70°, were tested. The measurements, carried out at Mach numbers of 1·4, 1·6, 1·8, 2·0, 2·5 and 3·0, included oil flow visualisations (on both sets of wings) and static pressure distributions (on the thicker wing only). In addition, a 60° sweptback delta wing with a normal angle of 40° was also tested. The tests on this wing included both oil flow visualisations and static pressure measurements. From these and other existing measurements, the leeside flows have been classified into nine distinct types, namely (i) leading edge separated flow with secondary separation, (ii) leading edge separated flow with secondary and tertiary separation, (iii) leading edge separated flow with a shock wave beneath the primary vortex, (iv) leading edge separated flow with shock-induced secondary separation, (v) fully attached flow, (vi) flow attached at the leading edge with inboard shock-induced separation, (vii) mixed type of flow, (viii) flow with a leading edge separation bubble and (ix) leading edge separated flow with a shock wave lying on the lee surface in between the leading edge vortices. These types of flow have been displayed in a plane of Mach number and angle of attack normal to the leading edge. The experimental results indicate that increasing wing thickness has no qualitative effect on the types of flow observed but does shift the boundaries between some of the types of flow.

Computer based mathematical models describing the aircraft evacuation process have a vital role to play in the design and development of safer aircraft, the implementation of safer and more rigorous certification criteria, in cabin crew training and post-mortem accident investigation. As the risk of personal injury and the costs involved in performing large-scale evacuation experiments for the next generation ultra high capacity aircraft (UHCA) are expected to be high, the development and use of these evacuation modelling tools may become essential if these aircraft are to prove a viable reality. This paper describes the capabilities and limitations of the airEXODUS evacuation model and some attempts at validation, including its successful application to the prediction of a recent certification trial, prior to the actual trial taking place. Also described is a newly defined performance parameter known as OPS that can be used as a measure of evacuation trial optimality. In addition, sample evacuation simulations in the presence of fire atmospheres are described.