Dye visualisation of both inclined synthetic jets and skewed synthetic jets was undertaken in a cross-flow experiment and the results were compared with those of normal synthetic jets. The process of vortex roll-up near the orifice exit and how the structure develops and interacts with the cross-flow as it propagates downstream was investigated so as to obtain an understanding of the effect of orifice orientation on the behaviour of synthetic jets. The effects of varying Reynolds number, velocity ratio and Strouhal number due to changes in diaphragm displacements and freestream velocities on the characteristics of synthetic jets were also examined. It is observed that in comparison to the normal jets vortical structures produced by both inclined and skewed jets tend to stay closer to the near wall region where maximum flow control effect is required. In both cases, at a relatively low Reynolds number and velocity ratio the active structures produced by the synthetic jet appear to be hairpin vortices which turn into vortex rings that migrate away from the wall as the Reynolds number and velocity ratio increase. These hairpin vortices are persistent in the near wall region hence are believed to be desirable structures for delaying flow separation.

This paper deals with an unsteady-state three-dimensional heat conduction problem in a prismatic body which has a complicated boundary shape. Numerical techniques such as a relaxation method must be used to solve a problem of this type; this method, however, requires unusually large amounts of labour. The method used in this paper consists of the application of conformal mapping and numerical solution of a partial differential equation and is applied first to a prismatic body of square cross section for which the exact solution is well known. As a second application, the approximate solution for a problem involving a star-shaped cross section is presented.

A method is given for the calculation of incompressible inviscid flow through non-axisymmetric contracting ducts with rectangular cross-sections. The method is based on a finite difference approximation to Laplace's equation and solved by the method of successive over relaxation. In an attempt to provide practical criteria for the design of such contractions, the flows through a series of contraction shapes were calculated, each shape being based on a pair of matched elliptic arcs. This permitted choice of such parameters as length, local (ie two-dimensional) contraction ratio, position and magnitude of the maximum slope. It was found that reducing the length of the contraction also reduced its effective length, although increasing the effects of overshoot and undershoot. This could be compensated for by designing a contraction with a steep maximum slope which with associated low curvatures at entrance and exit reduced the values of overshoot and undershoot. The axial positions of maximum slope on wall and roof should be the same.

The predictions of the numerical method were tested against experiment and, in general, satisfactory comparisons were obtained.

The special application environment puts forward the higher requirement of reliability of parts made from titanium alloy Ti–6Al–4V, which is closely related to the machining-induced residual stress. For the fact of the non-linear distribution of residual stress beneath the machined surface, distribution of peripheral milling-induced residual stress and its effect on fatigue performance of titanium alloy Ti–6Al–4V are still confusing. In the present study, residual stress profile induced by peripheral milling of Ti–6Al–4V is first studied. And then, energy criteria are proposed to characterise the whole state of the residual stress field. Finally, the effects of residual stress profile and surface energy on tensile–tensile fatigue performance of titanium alloy Ti–6Al–4V are discussed. The conclusions were drawn that the variation trend of surface residual stress (σr,Sur), maximum compressive residual stress (σC,ax), location (hr0) and response depth (hry) of residual stress profile with cutting parameters showed a similar pattern for both measure directions those parallel (σ1) and perpendicular (σ3) to the cutting direction. Cutting speed and feed rate have a main effect on surface residual stress, and the depth of cut has little effect on all the four key factors of residual stress profile. With the increase of cutting speed and feed rate, machining-induced surface energy tends to become larger. But increasing the depth of cut caused the strain energy stored in unit time to decrease. Furthermore, the effect of depth of cut on surface energy was weakened when the value of cutting depth becomes larger. Both the surface compressive residual stress and the maximum compressive residual stress are beneficial for prolonging the fatigue life, while large value of machining-induced surface energy leads to a decrease of fatigue life. Analysis of variance result shows that maximum residual compressive stress has a greater impact on fatigue life than other residual stress factors.

Incompressible turbulent flow in smooth pipe bends of circular-arc curvature and constant cross section is discussed. The effects of bend angle, radius ratio, duct cross-sectional shape and Reynolds number are considered. Particular attention is paid to the influence of the tangents on the flow and losses in a bend. In addition consideration is given to the definition and measurement of the bend loss.

A correlation of experimental pressure loss data is made. It is shown that large variations in experimental results, which cannot be ascribed to the effects of geometrical parameters or Reynolds number, are accounted for by the effects of the downstream tangent and position at which the pressure measurements were made. The inadequacy of existing experimental information is an obstacle to the presentation of comprehensive data on the effect of geometrical parameters on bend losses.

Experimental work is described showing that a major cause of changes in the calibrations of hot-wire anemometers consists in the precipitation of dust on the upstream surface of the wire during use. A simple means of cleaning the air in closed-circuit wind tunnels is described. The rate of change of calibration due to dust was reduced in the Manchester University Fluid Motion Laboratory by at least a factor of 15, by adequate cleaning of the air and by minimising the amount of interchange between the tunnel air and the atmosphere. It is concluded that where accurate results are required from hot wires used in a wind tunnel, the tunnel should be designed to be dust-free.

The discrete element method is modified so as to simulate the movement of passengers evacuating an aircraft. The results display reasonable agreement with timed trials. However, by focusing on the 1985 Manchester air disaster where only some of the exits could be opened and only then a considerable time after landing, it is argued that the certification procedures are inadequate.

This paper briefly reviews the progress made with the Imperial College hypersonic gun tunnel since August 1958 when it came into operation. The tunnel and instrumentation are described together with the results of various measurements made both in the barrel and in the working section. More detailed reports will be published later.

A new method for wing shape optimisation is presented, in which the wing shape is optimised not only for the best aerodynamic properties but also for minimum structural weight. An advanced structural sizing method for wing structural weight estimation is developed. The method is based on elementary, physics based analyses of simplified structure to determine the amount of required material to resist the applied loads. The sizing loads are determined based on the airworthiness regulations using advanced aerodynamic analysis tools. In this new weight estimation method, the effect of wing outer (aerodynamic) shape on the wing-box structural weight is taken into account. A mathematical equation is derived, which relates the wing-box structural weight to the wing outer shape. The weight estimation method is validated for several existing passenger aircraft. A series of optimisations is performed to optimise the wing shape for a general objective function, which includes both aerodynamic and structural properties of the wing. The aircraft fuel weight is used as the objective function. Different structures are tested both metal and composite ones. The latter includes different composite layups. In order to find the effect of the internal structure on the optimum wing outer shape, the design lift coefficient is assumed to be constant for all the test cases. The results of the optimisations showed that the optimum wing shape is depending on the structural concept (metal alloys versus different composite layups) for different internal structures.

The flow and noise created by sawtooth trailing-edge serrations has been studied experimentally at a low Reynolds number. Experiments have been performed on a flat-plate model with an elliptical leading edge and an asymmetrically bevelled trailing edge at Reynolds numbers of Rec = 1 × 105–1.3 × 105, based on chord. Wide serrations with a wavelength (λs) to amplitude (2h) ratio of λs/h = 0.6 were found to reduce the overall sound pressure level by up to 11dB. In contrast, narrower serrations with λs/h = 0.2 produce tonal noise and increase the overall noise level by up to 4dB. Intense vortices across the span of the trailing edge with narrow serrations are shown to be the source of tonal noise. Wide serrations reduce turbulent velocity fluctuations at low frequencies which explains the lower radiated noise. The narrow serrations that produce low Reynolds number tonal noise were shown previously to be effective at higher Reynolds numbers (Rec > 2 × 105), demonstrating that care is needed to fully understand the flow field over serrations for all intended operating conditions.

Recently attention has been given to the effects of the stream turbulence on the characteristics of the turbulent boundary layers and the turbulent shear flows. For the boundary layers considerable information on the effects of the stream turbulence is available but for the shear flows it is limited. This paper describes an experimental investigation for the effects of stream turbulence on the development of an asymmetric jet and a plane mixing layer. It is shown that the characteristics of these flows are considerably affected by the stream turbulence but for the measurements in the plane mixing layers at large downstream distances the influence of the stream turbulence is negligible provided the intensities are less than 0.6%. For the plane mixing layer the results show that the flow characteristics are not affected significantly by the initial boundary layers, the acoustic disturbances and the boundary conditions imposed by the end plates.

The results of measurements in the flow along a rectangular corner are presented in the form of velocity profiles. The profiles form two sets: one for flow in a slightly favourable streamwise pressure gradient and one for flow when the pressure gradient is ‘practically zero’. The appearance of the profiles is quite different from that of previously reported experimental work and it is suggested here that the primary cause of this is the different type of corner leading edges used in each case. The flow appears to be only marginally stable and even very slightly changed entry conditions caused by alterations in the leading edge geometry may have a marked effect on the profiles further downstream. The new results are consistent with the notion of corner layer similarity (an implication in every attempted theoretical solution) but there remains a significant difference between the experimental results and the more reliable theoretical solutions available.

The local buckling of structural sections composed of isotropic plate elements is well documented. It is the purpose of this paper to extend this work to include sections composed of orthotropic plates. It has been shown that the buckling behaviour of single orthotropic plates with certain boundary conditions can be related to that of an equivalent isotropic plate in terms of a single parameter involving the principal plate rigidities. It is proposed to investigate whether this principle may also be applied to the structural buckling problem.

It has been shown that simple approximations to the deflected form of buckled plates can be used to give fast and reasonably accurate results for the plate buckling problem using an energy method. If these simple deflection functions are used in an energy analysis for structural sections then large errors may be introduced.

This paper presents an approach to the determination of structural member sizes to provide minimum weight under a number of load conditions consistent with specified limitations on stress and displacement. The procedure amalgamates the disciplines of matrix displacement analysis and operations research into a scheme suitable for automatic digital computation.

A measurement of the pressure distributions on D-section prisms over wide ranges of depth to height ratio and incidence in a smooth and in a turbulent flow was made with particular reference to the onset of transverse galloping. Good correlation between the onset of galloping and the critical depth, established already for rectangular and the related prisms with sharp downstream corners, applied also to D-section prisms both in a smooth and in a turbulent flow. The development of a reattachment type pressure distribution on the side face of the prism as the critical depth is reached is responsible for this correlation. There is very little effect of turbulence on the pressure distribution and the Strouhal number of a D-section prism with a depth to height ratio of 0.5.

The effect of streamwise slots and grooves on a normal shock wave-turbulent boundary-layer interaction has been investigated experimentally at a Mach number of 1.3. The surface pressure distribution for the controlled interaction in the presence of slots featured a distinct plateau. This was due to a change in shock structure from a typical unseparated normal shock wave-boundary-layer interaction to a large bifurcated lambda type shock pattern. Velocity measurements downstream of the slots revealed a strong spanwise variation of boundary-layer properties, whereas the modified shock structure was found to be relatively two-dimensional. Cross flow measurements indicate that slots introduce streamwise vortices into the flow. When applied to an aerofoil, streamwise slots have the potential to reduce wave drag while incurring only small viscous penalties. In the presence of grooves the interaction was initially found to be significantly different. A bifurcated shock structure was observed but the trailing leg appeared stronger and featured a second lambda foot. Oil flow visualisation also revealed differences in the interactions, with the region of suction and blowing being limited to a smaller extent of the grooved control surface. The amount of crossflow present was reduced compared to the slotted control surface. By varying the internal geometry of the grooves it was found that the interaction could be modified to be similar to that in the presence of slots indicating that a more practical control device can be designed.

In an axisymmetric turbulent boundary layer along a circular cylinder at constant pressure, measurements have been made of mean velocity profile and turbulence characteristics: longitudinal velocity fluctuations, Reynolds shear stress, transverse correlation and spectrum. It has been found that the qualitative behaviour of an axisymmetric turbulent boundary layer is similar to that of a two-dimensional boundary layer in the wall region, where as in the outer region the effects of transverse curvature are observed.

A first attempt is made to use recently developed, non-conventional Artificial Neural Network (ANN) models with Multilayer Perceptron (MLP), Radial Basis Function (RBF) and Adaptive Neuro-Fuzzy Interference System (ANFIS) architectures to predict the fuel flow rate of a commercial aircraft using real data obtained from Flight Data Records (FDRs) of the cruise, climb and descent phases. The training of the architectures with a single hidden layer is performed by utilising the Delta-Bar-Delta (DBD), Conjugate Gradient (CG) and Quickprop (QP) algorithms. The optimum network topologies are sought by varying the number of processing elements in the hidden layer of the networks using a trial-and-error method. An evaluation of the approximate fuel intake values against the ideal fuel intake data from the FDRs indicates a good fit for all three ANN models. Thus, more accurate fuel intake estimations can be obtained by applying the RBF-ANN model during the climb and descent flight stages, whereas the MLP-ANN model is more effective for the cruise phase. The best accuracy obtained in terms of the linear correlation coefficient is 0.99988, 0.91946 and 0.95252 for the climb, cruise and descent phase, respectively.

This paper presents a comparison of linear and non-linear methods for the analysis of aeroelastic behaviour and flutter boundary prediction. The methods in question include NASTRAN and ZAERO, based on linear aerodynamics, and the non-linear coupled CFD-CSD methods RANSMB and PMB, developed at the Universities of Bristol and Glasgow respectively. The test cases used for this comparison are the MDO and AGARD 445.6 weakened wing. In general, it was found that the non-linear methods demonstrate excellent agreement with respect to pressure distributions, deflections, dynamic behaviour, and flutter boundary locations for both cases. This is in contrast to previous studies involving similar methods, where notable differences across the MDO span were found, and is taken to imply good performance of the CFD-CSD interpolation schemes employed here. While the linear methods produce similar flutter boundaries to the coupled codes for the aerodynamically simple AGARD 445.6 wing, results for the transonic ‘rooftop’ MDO wing design did not agree as well.

In 1973, automatic flight controls is one of the most exacting fields of technology, requiring a balanced combination of art, science and human understanding, and the efforts of large project management teams backed by adequate resources to bring any new concept to fruition.

Today's systems derive almost entirely from the technology of the past 25 years, and hence it is little known that there was some activity in automatic flight control for fixed wing aircraft as long ago as 1873, some 30 years before the world-changing events at Kittyhawk and when mechanical flight was still confined to balloons and a few rudimentary gliders.