The use of finite element analysis in the modelling of an unloaded countersunk hole revealed that there is a consistent relationship between the maximum stress concentration factor and the design parameters of the hole — countersunk depth, edge distance, t/d ratio and countersunk angle. It has been shown that countersunk depth and edge distance have a strong influence on the Kt value while the t/d ratio has a lesser effect. Variation of countersunk angles between 90° and 110° was found to have a negligible influence on the maximum Kt values.

It has also been shown that normalising the stress concentration factors of a countersunk hole by dividing the Ktvalues at various countersunk depths by the Ktvalues for a plain hole of similar configuration, enables the stress concentration factors in a countersunk hole to be determined from values in a plain hole. This is very useful since extensive work had been done on stress concentration factors for a plain hole and data are therefore more commonly available.

An extensive experimental investigation into the nature of the wall jets produced by single and twin normal jet impingement has been undertaken. Wall jet velocity profiles have been recorded up to 70 jet diameters from the impingement point, at pressures representative of current VStol technology. The tests used fixed convergent nozzles, with nozzle height and spacing and jet pressure being varied. Single jet impingement displays a consistent effect of nozzle height on wall jet development. For twin jet cases a powerful reinforcement exists along the wall jet interaction plane. Remote from the interaction plane the wall jets are weaker than those produced by a single jet impingement.

Changing airline scenarios and aggravated congestion have led to an increasing demand for new regional aircraft which will cover ranges of more than 1500 nm and accommodate 80 to 130 passengers. These far exceed the capability of today's commuter types. In the required thrust range of 15-5-20 klb only a very limited number of older engines with high SFC is available. Recently the manufacturers have begun offering new turbofans based on modern core technology and internal mixing (long duct) with lower SFC.

The underwing turbofan installation aspects are well understood from earlier wide and narrowbody aircraft but installing a long-duct nacelle on such a small aircraft means a significant increase in interference problems which may lead to close coupling and new methods for minimisation of the mutual interference.

After a survey of installation problems on previous jet transports up to the Airbus family the state of the art of experimental and theoretical methods for interference analysis is discussed. By means of an inverse design method a local lofting change on a given wing design confined to 20% span area was derived in order to sustain the isobar pattern of the clean wing with the pod/pylon on.

The design modification was checked by means of a 3D-Euler code and verified on a large half-model in the NLR-HST wind tunnel. From the analysis of the resulting conclusions concerning the feasibility of such a lofting change, further improvements are made.

Navigation is an art practised by all who travel but it developed as a science when the means of travel took travellers beyond sight of landmarks. Thus it was seamen who initially developed the science of navigation and the requirements of the marine fraternity drove its development until the invention of the aeroplane.

The capabilities of early aircraft limited their navigation requirements but as they could fly in conditions where visual contact with the ground was lost, the need for sophisticated navigation methods grew. At first marine techniques and equipment were adopted but these were found to have severe limitations in the new environment.

A calculation procedure for modelling the interaction between shock waves and attached or separated turbulent boundary layers is introduced. The numerical framework, applicable to general curved grids, combines cell-vertex storage, a Lax-Wendroff time-marching scheme and multigrid convergence acceleration. The main numerical ingredients of the procedure are documented in some detail, with particular emphasis placed on the inclusion of viscous and turbulence transport within the cell-vertex framework, which was originally formulated for inviscid flow. An investigation of the predictive performance of alternative transport models of turbulence has been the primary objective of the present work. Particular attention has been focused on a comparison between variants of low Reynolds number k-ε models and an algebraic variant of a Reynolds-stress transport closure in strong interaction situations, including shock-induced separation. The turbulence models are introduced, and important numerical issues affecting their stable implementation are discussed. The calculation procedure is then applied to two confined transonic flows over bumps — one incipiently and the other extensively separated (Delery Cases A and C) — and to the transonic flow around the RAE 2822 aerofoil at two angles of incidence (Cases 9 and 10). The investigation demonstrates that the eddy-viscosity models tend to seriously underestimate the strength of interaction, particularly when separation is extensive. The performance of the Reynolds-stress model is not entirely consistent across the range of conditions examined. In the case of bump flows, the model displays strong sensitivity to the shock, predicting excessive boundary layer displacement in Case A, a broadly correct separation pattern in Case C and insufficient rate of post shock recovery in both cases. The aerofoil flows are either attached or incipiently separated, and the benefits arising from Reynolds-stress modelling are modest. Neither the k-ε model nor the Reynolds-stress closure is able to return a satisfactory representation of the most challenging RAE 2822 Case 10; at least not with the recommended windtunnel corrections for freestream Mach number and angle of incidence.

Analysis and testing have been used to develop a flapping wing which can efficiently produce lift and thrust for ornithopter flight. With the assumption that the only imposed motion is dihedral flapping at its root, the wing's spanwise distribution of pitching and bending is due to its aeroelastic response. Numerical studies revealed criteria for efficient designs, which then guided the development and construction of an example wing tested in a windtunnel. The experimental results for average thrust and lift compared very favourably with the theoretical predictions, confirming that such aeroelastically tailored designs provide a means for achieving mechanical flapping-wing flight.

A design-oriented model for the unsteady aerodynamics of a flapping wing has been developed using a modified strip theory approach. Within this constraint, vortex-wake effects are accounted for as well as partial leading edge suction and post stall behaviour. Also, the contributions of sectional mean angle of attack, camber, and friction drag are added, which allows this model to be used for the calculation of the average lift, as well as the thrust, power required, and propulsive efficiency of a flapping wing in equilibrium flight. An example of such calculations is presented in the performance prediction of a mechanical flying pterosaur replica.

Lanchester was undoubtedly a great British engineer. He died just 47 years ago, at which time the name Lanchester could be seen on the front of some motor cars, but many knew little about the great contributions that Frederick William Lanchester had made to both mechanical and aeronautical engineering.

Whether Lanchester was interested in automobile aerodynamics or not, the subject combines two of his great loves. He designed and built petrol-engined cars, and his 1898 machine won the RAC’s Gold Medal (Fig. 1). As might have been expected, the drag coefficient of this vehicle is not recorded.

In this paper the results are given of an experimental investigation into the flow development and the surface pressures in an unsteady shock boundary layer interaction at Mach numbers of 1.3 and 1.5. The experiment is one in which a naturally grown turbulent boundary layer on the tunnel walls is disturbed by a normal shock wave spanning the test section of the tunnel. The shock wave is oscillated by a periodic pressure disturbance in the tunnel diffuser far downstream of the interaction. Both instantaneous (spatially averaged) and time averaged results for the surface pressures are presented. These results are used to study the flow development and phase changes through the interaction over a frequency range from 34 to 167 Hz. This frequency range was chosen to cover the range of parameters likely to occur in conditions corresponding to stall flutter in the compressor fans of large bypass engines.

There can be no doubt that Sir Sydney Camm has left an enduring legacy, both within the aerospace industry as a whole and in the RAF in particular. Indeed, the fact that the RAF has been flying one or more aircraft designed either in part or wholly by Sir Sydney Camm for the majority of the 75 years of our existence provides ample testimony to the extent of that legacy (Fig. 1). A military theme for this paper would therefore seem to be entirely appropriate.

The challenge, of course, stems from the reduced threat of global war and the rather greater likelihood of further conflicts such as that seen in the Gulf (Operation GRANBY). These developments have required the RAF to undertake a radical reappraisal of its size, composition and operational posture.

This paper describes the incompressible flow through contracting ducts as used in windtunnel circuits. The value of the velocity gradients that are critically adverse to the boundary layer development are specified and the contraction profile is derived for two-dimensional flow. Solutions to the flow through three-dimensional contractions of circular, square, and square to rectangle shape are obtained. These lead to simple rules for scaling the profile of two-dimensional flow to provide three-dimensional shapes having short lengths with the required velocity distributions. The development of both laminar and turbulent boundary layers is calculated to provide criteria for the avoidance of separation.

The feasibility of mechanical flapping-wing flight has been studied by analyses and experiments. The key results from this work include the development of an efficient wing with unique features for twisting and lift balance, as well as a lightweight and reliable drive mechanism. These were incorporated into a radio-controlled, engine-powered, flapping-wing aeroplane (ornithopter), whose flight tests have been the proof-of-concept focus of this research. In September 1991, this aircraft achieved successful sustained flights, demonstrating the practicability of this particular solution for mechanical flapping-wing flight.

The development of jet propulsion and the aircraft gas turbine in Britain has two main lines. One began in 1926 when Dr A.A. Griffith, then of the Royal Aircraft Establishment (RAE), proposed the use of a gas turbine as an aircraft power plant. This line of development led to the Metropolitan-Vickers F2 jet engine and its derivatives. Our story began two years later in 1928 when Frank Whittle was a cadet at Cranwell. Each term, the cadets were required to write a thesis and in his fourth and final term Whittle chose as his subject “Future Developments in Aircraft Design”. He concluded that for very high speeds combined with long range, it would be necessary to fly at very great heights where low air density would greatly reduce resistance in proportion to speed. He was thinking in terms of speeds of 500 mph at heights where the, air density was less than one quarter of its sea level value.

The compressible flow Navier-Stokes equations, in conservation form, have been solved using a modified MacCormack implicit predictor-corrector scheme. For turbulent flow the solution is modified by incorporating an eddy viscosity, to solve the Reynolds-averaged Navier-Stokes equations. The algorithm is accurate to second order in time and space and is unconditionally stable. It requires the solution of block bidiagonal matrices and the programme can run on commonly available minicomputers. It also has the advantage that a steady state solution is approached, if it exists, independent of the time step employed.

Numerical results are presented for the two-dimensional shock-boundary layer interaction in both laminar and turbulent flow and the results are compared with experimental values and the numerical results of other workers. The method is calibrated against the analytical solution for the boundary layer on a flat plate and for the time dependent Couette flow problem.

This report describes a series of steady and time dependent pressure measurements on a half-model of a rectangular wing of aspect ratio 4. The wing was mounted on a halfbody attached to a sidewall of the RAE 8 ft x 8 ft windtunnel and the tests were made at Mach numbers of 0·60, 0·70 and 0·80 with fixed transition and a Reynolds number of about 2·4 x 106 at the highest Mach number. The wing was oscillated about its mid-chord axis at four frequencies giving frequency parameters up to 0·27 at M = 0·80. The static angle of incidence was varied from -10° to + 10°; here interest is centred on angles of incidence for which the flow on the upper surface of the wing is separated.

Selected mean and oscillatory measurements are offered as a challenge to computational fluid dynamicists for this simple threedimensional configuration. The measurements are of particular interest because the transonic flows include both incipient and well developed separations.

A new rotor model for helicopter flight mechanics simulation is presented. The rotor dynamics are described using multiblade coordinates, and the aerodynamic loads include nonlinear effects such as stall and compressibility. These loads are described as harmonic series. The number of harmonics in the series determines the model accuracy. Thus by changing the number of harmonics from one to a large number, it is possible to obtain models that range between a quasi tip path plane approach and an accurate blade element model. The user of the model can very easily change the model accuracy and consequently its efficiency. The new rotor model is investigated and its application for trim and manoeuvre calculations is presented and discussed.

Transonic shockwave/turbulent boundary layer interactions on a porous surface above a closed plenum chamber have been studied experimentally in the choked flow of a windtunnel test-section. The equivalent freestream Mach number is 0.76 and results were obtained for three shock strengths. Without the porous surface the Mach numbers ahead of the shock were 1.13, 1.18 and 1.26. The respective shock Mach numbers with the porous surface were 1.10, 1.11 and 1.19. Laser holographic interferometry results are used to measure the density flowfield and examine the nature of the interaction. These results show that the interaction on the porous surface is modified by a thin shear layer adjacent to the surface and the weakening of the Shockwave is attributed to this. The interaction was also studied by solving the two-dimensional Reynolds-averaged Navier-Stokes equations together with the two-layer algebraic eddy-viscosity model of Baldwin-Lomax modified with appropriate corrections for surface transpiration. The computed results show excellent agreement with the experimental data. The examination of these numerical results shows that the surface transpiration occurs at a low subsonic velocity and suggests that the effect of the transpiration through the porous surface on the interaction may be optimised.

Aircraft plying between urban and outlying communities usually enjoy a runway with noise restrictions at the urban end and perhaps little to land on at the other end, but with no objection to noisy operation. The existence of this situation in northern Canada persuaded the National Research Council to embark — some thirty years ago — on an investigation of powered lift possibilities, with strong support from the trunk and regional airlines, the charter and helicopter operators, and various manufacturers.

From a number of generic lift-propulsion arrangements examined in greater or lesser detail, a selection are described here in increasing order of what the certifying authorities call “damage tolerance” or reduced dependence on “critical parts”. Freedom from demands for hovering endurance and quietness, combined with emphasis on small aircraft, offered encouragement to attack what is otherwise a very difficult task. Of the alternatives, only one, which incorporated transverse axis fans and the light weight gas turbines now available, emerged as a potentially satisfactory choice for small rugged transport aircraft.

The project was discontinued because of pressure of other work, but is recounted here in case others may discern wider market possibilities. A small Vtol aircraft should, if successful, be a useful precursor to sophisticated airliners capable of taking off and landing without a run.

This study details the design of an automatic control system for a sailplane in towed flight, the principal objective of which is control the relative position of the tow-aircraft and glider, since a large vertical distance is expected to drive unstable motions of the whole system. The synthesis of the control system is based on an approximated model for the constrained sailplane. The controlled motion of the sailplane is simulated by a general dynamic model of the three bodies, namely the two aircraft and the rope. Significant numerical results are presented and discussed in order to determine the effectiveness of the stability augmentation.

The planar vortex lattice method is extended to the case of wings in small-amplitude oscillations in ideal flow. Vortex rings are used to represent the bound vorticity on the wing and the trailing vorticity in the wake. A modified procedure for solving the unknown vortex strengths gives greater computing efficiency. Comparisons with results reported in the literature show good agreement for the various test cases considered. Convergence with increasing lattice density is uniform.