For boundary layer flows over curved surfaces at moderately high supersonic speeds the existence of normal pressure gradients within the boundary layer becomes important even for small curvatures and they cannot be ignored. The describing equations are basically parabolic in form so that the simplifications inherent in hyperbolic flows would not at first sight seem to be relevant. However, the equations of motion for a two-dimensional, supersonic, rotational, viscous flow are analysed along the lines of a hyperbolic flow and the individual effects of viscosity and vorticity are examined with regard to the isobar distributions. It is found that these two properties have compensating effects and the experimental evidence presented confirms the conclusion that inside the boundary layer the isobars follow much the same rules as those which determine the isobars in the external hyperbolic flow. Since for turbulent boundary layers the fullness of the Mach number profile produces almost linear Mach lines in the boundary layer, this provides a simple extension to the methods of analysis, and the momentum integral equation is reformulated using a swept element bounded by linear isobars. The final equation is similar in form to the conventional one except that the momentum and displacement thicknesses are now defined by integrals along the swept isobars, and all normal pressure gradients due to centrifugal effects are accounted for.

The equations of motion of an aircraft flying at a constant glide slope (Fig. 3), in the presence of arbitrary head- or tailwinds, and up- or downflows is considered. The equations are integrated analytically, in the case of an aircraft initially on a steady flight, perturbed by winds of ‘moderate’ strength, in the sense that the wind velocity is not negligible compared to the aircraft's steady speed, but the ratio of their squares is much smaller than unity. The case of an approach through a downburst (Fig. 1), leads to winds which can be simplified to a one-period sinusoidal wind along the flight path, changing from head- to tailwind, at the peak of a superimposed downflow (Fig. 2), of half-period sinusoidal shape, and this is discussed in some detail. Data sheets are presented for three combinations of amplitudes of the head-to-tailwind and downflow; each contains plots of the scheduling of incidence that exactly cancels windshear effects, and of the groundspeed a'nd airspeed profiles which will keep the aircraft flying along the original glide slope. Each plot is given for a range of values of the windshear susceptibility parameter, representing aircraft with small or large inertia, with high or low approach speeds, including light aircraft, jet fighters and large transports; the cancellation of windshear effects isachievable only if the incidence schedule lies wholly below the stall limit. Since the rotational inertia of the aircraft is neglected, the short period mode is absent, and the pitch control acts to cancel the ‘phugoid’ instability induced on the aircraft by the wind profiles typical of a microburst.

The method developed by G. N. Ward for the treatment of slender pointed bodies in a uniform supersonic stream is applied to three special cases.

• (i) Supersonic flow past a body of revolution with thin wings of symmetrical section and of small aspect ratio at zero incidence.

• (ii) Supersonic flow past a body of revolution with plane wings of small aspect ratio set at incidence to the body, the whole being at incidence to the stream.

• (iii) Supersonic flow past a body of revolution with a plane fin of small aspect ratio set at incidence, the whole being at incidence to the stream.

The pressure distribution on the wing has been calculated for a special case of (i) and is given in the Appendix.

Long before the legendary Queen Dido cleverly maximised the area of future Carthage by choosing the optimum perimeter shape, man probably consciously or subconsciously tried to optimise his designs as well as his actions.

In a more modern time, the post Renaissance period witnessed development of the principles of optimisation on a rigorous mathematical basis, as exemplified in works of Newton, Leibniz, Euler and many others. Eventually, those optimisation ideas were even transplanted by Leibniz and his school to the field of general philosophy, providing an attractive target for the merciless satire of Voltaire in Candide.

I should like to express my gratitude to the Society, on behalf of the Federal Aviation Agency of the United States, for the opportunity of participating in this meeting.

Before speaking of the lessons gained in Helicopter Air Traffiic Control some background information is necessary on the types of helicopter operations which are being conducted in the United States.

As in many countries, helicopter operations are conducted by both the military services and civil operators. The military operations are primarily those of a tactical training nature and search and rescue work. Civil helicopter operations are conducted mainly for transport and utility purposes.

In addition to numerous civil air taxi operators providing transport by helicopter, we have three commercial scheduled helicopter airlines; New York Airways, Chicago Airways, and Los Angeles Airways. These three helicopter airlines serve the aerodromes and residential communities in the metropolitan areas of the cities from which they derive their names. Utility helicopter flying includes such services as forest fire fighting, insecticide spraying of farm crops, power line inspection, and construction work, just to name a few.

The application of electrical analogue methods to the analysis of the extension and flexure of flat plates is reviewed and the difficulties encountered in the satisfaction of the various boundary conditions are discussed. A new method for treating certain boundary conditions and the operation of the electrical analogue is described. New experimental results for two cases which present great analytical difficulty, the flexure of a plate with a free edge and a plate supported on columns, are given.

A complete unsteady Newton-Busemann flow theory, including centrifugal force corrections, was recently given by Hui and Tobak and successfully applied to study the dynamical stability of oscillating aerofoils (Part I) and bodies of revolution (Part II). The present paper extends these results, which are restricted to the first order in frequency, to general frequencies that may be applicable to flutter analysis as well. Furthermore, the new results are used to investigate the behaviour of the indicial response functions in unsteady flow at very high Mach numbers. In particular, it is found that for a family of body shapes in the Newtonian flow, including the cone, the wedge, the delta wing and slender aerofoils, the aerodynamic response to a step change in angle of attack or pitching velocity contains an impulse at the initial instant. This is followed by a rapid adjustment to the new steady-flow conditions, the adjustment time being equal to that required for a fluid particle to travel a distance of the body length. The impulse component appearing at the infinite Mach number limit (Newtonian flow) is in effect an apparent mass term, analogous to that which occurs initially in the aerodynamic indicial response at the zero Mach number limit (incompressible flow). In the latter case the initial impulse is followed by an asymptotic adjustment to the new steady-flow conditions, theoretically taking infinite time, in direct contrast to the rapid adjustment in Newtonian flow.

A method was developed to predict numerically the damage of composite laminates with multiple plies under low-velocity impact loading. The Puck criterion for 3D stress states was adopted to model the intralaminar damage including matrix cracking and fibre breakage, and to obtain the orientation of the fracture plane due to matrix failure. According to interlaminar delamination mechanism, a new delamination criterion was proposed. The influence of transverse and through-thickness normal stress, interlaminar shear stress and damage conditions of adjacent plies on delamination was considered. In order to predict the impact-induced damage of composite laminates with more plies quickly and efficiently, an approach, which can predict the specific damage of several plies in a single solid element, was proposed by interpolation on the strains of element integration points. Moreover, the proposed model can predict specific failure modes. A good agreement between the predicted delamination shapes and sizes and the experimental results shows correctness of the developed numerical method for predicting low-velocity impact damage on composite laminates.

This paper aims to build the finite element model of the composite sinusoidal specimens and to carry out the parametric analysis. In this paper, the damage behaviour and the energy-absorbing results of composite sinusoidal specimens have been studied by quasi-static crushing experiments. The failure mechanisms of specimens under quasi-static crushing is further analysed. A numerical simulation has been performed by using the finite element model code LS-DYNA. The numerical results, in terms of load -displacement data, have been compared against experimental data, and good agreement has been found. Moreover, a sensitivity study has been carried out by varying material properties in order to assess their influence on the numerical results, and the material parameter selection scheme is optimised based on the constructed corresponding response surfaces. The results show that the response surface model has passed the test of goodness of fit, and the optimisation method can effectively assist the finite element modelling, and greatly decrease the numbers of trial and error.

In the past decade considerable attention has been given to the possibility of using self-adaptive control systems to cope with the problem of providing stability augmentation to aircraft over extended flight envelopes. On the other hand, it has sometimes been considered that insufficient attention has been given to the design of autostabilisers with fixed parameters. The authors have applied recent developments In dynamic programming to the design of aircraft autostabilisers. The methods provide a highly systematic approach to the design of multi-loop systems which is especially useful for control of the lateral mode. Furthermore, the resulting optimum multiple feedback coefficients vary little throughout the flight envelope and the use of fixed parameters for the whole of the flight envelope appears to be possible. Aerodynamic data for two aircraft were used: a supersonic VTOL aircraft and a supersonic transport.

A velocimeter is described based upon the same principle as the five-hole Conrad tube yawmeter. However it uses a single tube, cut obliquely at 45°, that rotates about its own axis together with an outer closely-fitting concentric tube, cut normal to its axis, that can be slid forward over the mouth of the chamfered tube. The small measuring volume, about 1 mm3, and well-defined measuring point enable the probe to be used with some confidence in regions of high shear. Some typical results obtained with the probe in a highly-sheared, three-dimensional flow are given.

For the turbulent boundary layer it is shown that, if an initial velocity profile is given, along with the local pressure gradient and shear-stress distribution through the layer, then the shape of the velocity profile a short distance downstream is unaffected by flow convergence or divergence, provided this is constant through the layer. For flow approaching an obstacle, increased divergence close to the surface is shown to account for the marked changes in profile shape that have been observed.

Up to the present most of the methods used to calculate the flow over delta wings with leading edge separations have been based on slender wing theory. In this paper one of the simplest of these methods is carried over to linear theory and the overall lift calculated. In general the results are in agreement with experiment, in particular both show a fall in nonlinear lift as the leading edges of the wing approach the free stream Mach angles. The discrepancies which exist appear to arise mainly from the use of linear theory rather than from the mathematical model introduced to treat the separated flow.

There is some evidence to indicate that injuries associated with parachute opening shock are much less frequent than injuries associated with the acceleration loads imposed by ejection seats, even though the stresses presumed to be induced in the human body may often be of the same order. Since all such data is based on operational experience, the details are necessarily vague, and it is hard to draw firm conclusions.

The Aeromedical Research Laboratory, USAF, as part of its studies of aircrew injury, decided to see whether there was any difference between the dynamics of parachute opening when used with a rigid anthropometric dummy of the type normally used to obtain opening shock data, and a human parachutist whose body is naturally more resilient.

Within an immunity-based architecture for aircraft fault detection, identification and evaluation, a structured, non-self approach has been designed and implemented to classify and quantify the type and severity of different aircraft actuators, sensors, structural components and engine failures. The methodology relies on a hierarchical multi-self strategy with heuristic selection of sub-selves and formulation of a mapping logic algorithm, in which specific detectors of specific selves are mapped against failures based on their capability to selectively capture the dynamic fingerprint of abnormal conditions in all their aspects. Immune negative and positive selection mechanisms have been used within the process. Data from a motion-based six-degrees-of-freedom flight simulator were used to evaluate the performance in terms of percentage identification rates for a set of 2D non-self projections under several upset conditions.

Unmanned aerial vehicle (UAV) design tends to focus on sensors, payload and navigation systems, as these are the most expensive components. One area that is often overlooked in UAV design is airframe and aerodynamic shape optimisation. As for manned aircraft, optimisation is important in order to extend the operational envelope and efficiency of these vehicles. A traditional approach to optimisation is to use gradient-based techniques. These techniques are effective when applied to specific problems and within a specified range. These methods are efficient for finding optimal global solutions if the objective functions and constraints are differentiable. If a broader application of the optimiser is desired, or when the complexity of the problem arises because it is multi-modal, involves approximation, is non-differentiable, or involves multiple objectives and physics, as it is often the case in aerodynamic optimisation, more robust and alternative numerical tools are required. Emerging techniques such as evolutionary algorithms (EAs) have been shown to be robust as they require no derivatives or gradients of the objective function, have the capability of finding globally optimum solutions among many local optima, are easily executed in parallel, and can be adapted to arbitrary solver codes without major modifications. In this paper, the formulation and application of a evolutionary technique for aerofoil shape optimisation is described.

Initially, the paper presents an introduction to the features of the method and a short discussion on multi-objective optimisation. The method is first illustrated on its application to mathematical test cases. Then it is applied to representative test cases related to aerofoil design. Results indicate the ability of the method for finding optimal solutions and capturing Pareto optimal fronts.

This article presents the findings of a study to investigate the impact perforation resistance of sandwich structures. The dynamic response of sandwich panels based on PVC foam cores has been evaluated by determining the energy to perforate the panels. The impact response of the sandwich structures was predicted using the finite element analysis package Abaqus/Explicit. The validated FE models were also used to investigate the effect of oblique loading and to study the impact response of sandwich panels subjected to a pressure differential equivalent to flying at an altitude of 10,000m.

Low velocity impact testing has shown that the energy to perforate the sandwich panels is dependent on the properties of the core. It has been shown that increasing the density of the crosslinked PVC foams by a factor of two yielded a 600% increase in the perforation resistance of the sandwich structures. At higher densities, the crosslinked foam sandwich structures offered a superior perforation resistance to the linear PVC structures. The numerical analysis accurately predicted the perforation energies of the sandwich panels, as well as the prevailing failure mechanisms following impact. Finally, it has been shown that sandwich panels impacted at high altitude offer a similar perforation resistance to those tested at sea level.

Experimental investigations were carried out to study the acoustic radiation from a rectangular wall mounted cavity in a confined supersonic flow. The free-stream Mach number was maintained at 1·5 and the cavity length-to-depth ratio was varied from 0·43 to 5·0. Acoustic measurements made on the top wall show jumps in the dominant frequency as the cavity behaviour changes from shallow-to-square-to-deep cavity. Numerical simulations of this unsteady two-dimensional flow using the commercially available software FLUENT have also been carried out. Unsteady pressure data at the same location in the flow field as the pressure transducers in the experiments was collected. FFT analysis of the unsteady pressure data was performed to obtain the dominant acoustic frequencies. The values for these dominant frequencies predicted by the numerical calculations agree well with experimental data. The numerical study also predicts the frequency jump observed in experiments.

When an aero-engine is operating on the ground the formation of a potent inlet ground vortex can arise which has the ability to ingest foreign object debris. The ingestion of foreign objects can cause notable damage to engine components as well as overall performance degradation. The assessment of foreign object ingestion has been conducted using a combination of computational fluid dynamics, analytical modelling and an Euler-Lagrange uncoupled discrete phase particle tracking method. The flow fields for a full-scale aero-engine have been simulated for a range of ground clearances and intake velocity ratios under crosswind conditions. The sensitivity of the debris ingestion thresholds and characteristics to particle size and material density has been evaluated. The characteristics of the ingestion location within the intake are also considered for a range of operating conditions and particle parameters.

The aeroelastic stability of a uniform, untwisted hingeless ‘smart’ helicopter rotor blade in hover has been analysed. The concept of a ‘smart’ blade is achieved by implementing a piezoelectric stack at an appropriate location along a host blade such that upon actuation it enters the load path becoming an integral part of the host structure. Thus, the stiffness characteristics of the rotor are altered causing modal damping augmentation of the blade. The perturbation equations of motion for the ‘smart’ blade that describe the unsteady blade motion about the equilibrium operating condition are obtained using Galerkin’s method. These differential equations with periodic time coefficients are analysed for stability utilising the Floquet method. Six different regimes of actuation are investigated, and a parametric study is carried out by considering six different design cases. It is shown that, compared to a ‘host’ blade the stability characteristics of the ‘smart’ blade are not affected adversely. In fact, a judicious design and actuation of the ‘smart’ spring has the potential of improving the stability boundaries of individual blades.