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.

Recent lunar and planetary return vehicle studies have stimulated interest concerning the effect upon stagnation point heating due to electronic heat conduction associated with re-entry at speeds greater than 30,000 ft./sec. In order to provide experimental data for comparison with theory, a series of experiments was undertaken in which shock tube techniques were utilised for the measurement of the stagnation point heating of a blunt body over a stagnation enthalpy (Hst/RT0) range of 650 to 900, corresponding to velocities between 32,000 ft./sec. and 39,000 ft. /sec. respectively. The tests were performed in a 1·5 inch diameter shock tube with 3 foot length driver and 16 foot length driven section.

A flush air data system uses surface pressure measurements to obtain speed and aerodynamic orientation of a flight vehicle. This paper investigates the use of a neural network to calibrate the air data system on a low-observable aircraft forebody with 22 pressure tappings. Wind tunnel data were obtained for between 0 and 25° angle of attack, 0 to 10° sideslip for speeds of 32, 37 and 45ms-1. Experimental data were used to train multilayer perceptron neural networks. A calibration accuracy of 0·322ms-1, 0·811° and 0·552° for speed, alpha and beta respectively was achieved just using the first five tappings on the nose cone. Increasing the number of tappings used as inputs to the neural network reduces the calibration error. A neural network with 22 inputs gives a best accuracy of 0·095ms-1, 0·15° and 0·085° for speed, α and β respectively. Trained networks show poor robustness to single sensor failures. Robustness is improved by preprocessing input data with an autoassociative network or by introducing sensor redundancy.

An integrated artificial immune system-based scheme that can operate over extended areas of the flight envelope is proposed in this paper for the detection and identification of a variety of aircraft sensor, actuator, propulsion, and structural failures/damages. A hierarchical multi-self strategy has been developed in which different self configurations are selected for detection and identification of specific abnormal conditions. Data collected using a motion-based flight simulator were used to define the self for a wide area of the flight envelope and to test and validate the scheme. The aircraft model represents a supersonic fighter, including model-following direct adaptive control laws based on non-linear dynamic inversion and artificial neural network augmentation. The proposed detection scheme achieves low false alarm rates and high detection and identification rates for all the categories of failures considered.

Machining-induced surface integrity has an important effect on reliability and service life of the components used in the aerospace industry where titanium alloy Ti-6Al-4V is widely applied. Characterisation of machining-induced surface integrity and revealing its effect on fatigue life are conducive to structural fatigue life optimisation design. In the present study, surface topography, residual stress, microstructure and micro-hardness were first characterised in peripheral milling of titanium alloy Ti-6Al-4V. Then, low-cycle fatigue performances of machined specimens were investigated on the basis of the tension-tension tests. Finally, the effects of surface integrity factors (stress concentration factor, residual stress and micro-hardness) on fatigue performances were discussed. Results show that stress concentration can reduce the fatigue life while increasing the residual compressive stress, and micro-hardness is beneficial to prolonging the fatigue life, but when the surface material of the specimen is subjected to plastic deformation due to yield, the residual stress on the surface is relaxed, and the effect on the fatigue performance is disappeared. Under the condition of residual stress relaxation, the stress concentration factor is the main factor to determine the low-cycle fatigue life of titanium alloy Ti-6Al-4V. While for the specimens with no residual stress relaxation, micro-hardness was the key factor to affect the fatigue life, followed by residual stress and stress concentration factor, respectively.

A study of the effect of linear damping on flutter speed is described. The energy exchange at flutter is considered in each degree of freedom of a general n-degrees-of-freedom system and an energy compatibility equation is formed. An analytical discussion of binary systems, based on the energy compatibility equation, is then presented.

A ‘two dimensional’ air intake comprising a wedge followed by an isentropic compression has been tested in the Cranfield Gun Tunnel at Mach 8·2. These tests were performed to investigate qualitatively the intake flow starting process. The effects of cowl position, Reynolds number, boundary-layer trip and introduction of a small restriction in the intake duct were investigated. Schlieren pictures of the flow on the compression surface and around the intake entrance were taken. Results showed that the intake would operate over the Reynolds number range tested.

Tests with a laminar boundary layer demonstrated the principal influence of the Reynolds number on the boundary-layer growth and consequently on the flow structure in the intake entrance. In contrast boundary layer tripping produced little variation in flow pattern over the Reynolds number range tested. The cowl lip position appeared to have a strong effect on the intake performance. The only parameter which prevented the intake from starting was the introduction of a restriction in the intake duct.

The experimental data obtained were in good qualitative agreement with the CFD predictions. Finally, these experimental results indicated a good intake flow starting process over multiple changes of parameters.

In this paper, a simple static compressible flow model for circular synthetic jet actuators is described. It is used to undertake a systematic computational investigation of the effect of changing actuator geometrical and operating parameters on the magnitude of peak jet velocity at the orifice exit of an actuator whose diaphragm displacement and frequency are allowed to vary independently. It is found that, depending on the flow conditions inside the orifice duct, the actuator may operate in two distinct regimes, i.e. the Helmholtz resonance regime and the viscous flow regime. In the Helmholtz resonance regime, the resultant synthetic jet is generated by the mass physically displaced by the oscillating diaphragm coupled with the Helmholtz resonance in the actuator. In the viscous flow regime, the Helmholtz resonance is completely damped by viscous effect such that the jet is produced by the diaphragm oscillation alone. The relationship between actuator geometrical and operating parameters at the optimum condition which yields the maximum peak jet velocity at a given diaphragm displacement is also established for these two regimes. Finally, a preliminary procedure for designing synthetic jet actuators for flow separation control on an aircraft wing is proposed.

The problem of transferring a space vehicle between two points in a given gravitational field such that the minimum amount of fuel is used has been called the fundamental navigational problem of astronautics. In such a problem it may be required to find the optimum thrust magnitude and thrust direction which yields a minimum fuel trajectory. Furthermore, certain end conditions may be specified which the optimal trajectory must satisfy. In a large number of published papers the velocity of the vehicle is supposed known both at the beginning of the transfer and at the end whereas the time taken to complete the manoeuvre may or may not be given. Also, other performance criteria have been chosen besides minimum fuel. For example, minimum time of transit or maximum orbital altitude at perigee.

Papers mentioned in this review deal mainly with flight in two dimensions apart from those sections on general theory. Furthermore, all space vehicles considered are assumed to have a fixed exhaust velocity unless otherwise stated.

A high order boundary element formulation is presented and applied to the solution of potential, incompressible flows around non-lifting and lifting configurations. The high order numerical algorithm is based on a bicubic interpolation of both geometry and quantities over each element of discretisation of the boundary. Numerical validation of the formulation is performed by studying the aerodynamic solution around fuselages and wings, and making comparisons with existing numerical results.

Passive control experiments were made on an 18% thick biconvex aerofoil model in Mach number range giving shock oscillations and at a Reynolds number of 0.7 x 106 with fixed boundary layer transition. The control methods studied included a passive control of shock wave boundary layer interaction by a porous surface and a cavity, and a buffet breather made of holes connecting upper and lower surfaces of the aerofoil. The experiments show that the shock oscillations can be virtually eliminated by these methods. A physical explanation is given for the success of these passive controls.

In the theoretical and experimental work on the buckling of sandwich panels with laminated faces reported in Refs. 1 and 2, panels which had certain coupling terms present in the constitutive equations for the faces were solved in an approximate manner. The previous theoretical solution for angle-ply lay-ups assumed that there were sufficient layers present to treat the faces as orthotropic, with the Bij coupling terms in the constitutive equations put equal to zero. Further simplifying assumptions were made when comparing theoretical and experimental results in Ref. 2 where tests were carried out on three different types of carbon fibre faces; orthotropic, balanced angle-ply and unbalanced cross-ply. It was only in the first case that an exact theoretical analysis was carried out, and the second case was again treated as an orthotropic panel neglecting coupling between direct and shear deformations and bending and twisting curvatures. The unbalanced cross-ply lay-up, which consisted of two fibre reinforced plastic orthotropic layers, one at 0° to the plate axis and one at 90°, gave rise to large B11 and B22 coupling terms which were taken into account approximately using “reduced stiffness” terms with the orthotropic equations. However, the disparity between theory and experiment was greatest in this case, particularly as far as wrinkling was concerned, where the experimental failure load was 78% higher than the theoretical wrinkling load. This type of cross-ply layup is important since it is used when making carbon fibre sandwich floor panels; it being chosen from the point of view of manufacturing convenience rather than from considerations of strength and stiffness. Furthermore, the bending stresses in the faces, which result from the normal loads, predominate in one direction so that one face is subjected to essentially a uniaxial compressive stress which might cause the faces to wrinkle. In order to study this problem in greater depth, it was decided to carry out further theoretical and experimental work on the simpler problem of the wrinkling of sandwich columns with carbon fibre cross-ply faces.

A novel vortex generator design positioned upstream of a normal shock followed by a subsequent diffuser was investigated using large eddy simulations. In particular, “ramped-vane” flow control devices with three different heights relative to the incoming boundary layer thickness (0.34δ, 0.52δ and 0·75δ) were placed in a supersonic boundary layer with a freestream Mach number of 1.3 and a Reynolds number of 2400 based on the momentum thickness. This is the first numerical study to investigate the size effect of the ramped-vane for flow control device in terms of shape factor, flow separation and flow unsteadiness. The results showed that these devices generated strong streamwise vortices that entrained high-momentum fluid to the near-wall region and increased turbulent mixing. The devices also decreased shock-induced flow separation, which resulted in a higher downstream skin friction in the diffuser. In general, the largest ramped-vane (0.75δ) produced the largest reductions in flow separation, shape factor and overall unsteadiness. These results and a careful review of the literature study also determined the quantitative correlation of optimum VG height with Mach number, whereby h/δ~1 is often optimum for incompressible flows while higher Mach numbers lead to small optimum heights, tending towards h/δ~0.45 at M=2.5.

Very Large Transport Aircraft (VLTA) pose considerable challenges to designers, operators and certification authorities. Questions concerning seating arrangement, nature and design of recreational space, the number, design and location of internal staircases, the number of cabin crew required and the nature of the cabin crew emergency procedures are just some of the issues that need to be addressed. Other more radical concepts such as blended wing body (BWB) design, involving one or two decks with possibly four or more aisles offer even greater challenges. Can the largest exits currently available cope with passenger flow arising from four or five aisles? Do we need to consider new concepts in exit design? Should the main aisles be made wider to accommodate more passengers? In this paper we demonstrate how computer based evacuation models can be used to investigate these issues through examination of staircase evacuation procedures for VLTA and aisle/exit configuration for BWB cabin layouts.

It is generally recognised that fuel can be saved on long haul routes if non-stop flights are broken down into a number of short stages. As fuel costs rise, these savings may eventually become economically attractive to airlines. ‘Staging’ may also play a role in the reduction of aviation’s impact upon climate change. To allow an accurate estimate of the potential benefits, an analytic model has been developed from first principles, with closure provided by data from existing aircraft. The analysis captures the relevant parameters and their influence is quantified. It has been found that significant savings are only available for distances greater than about 5,500km and a stop at the halfway point is either the best, or close to the best, strategy. For the longest routes, optimal staging with existing long range aircraft can give savings of up to 12%. However, if aircraft optimised for the stage lengths are used, this may be increased to up to 28% and, if the floor area per passenger is reduced to the short haul level, a limiting value of about 45% may be possible. The aircraft design passenger number and, provided it is constant, the load factor have little, or no, effect upon the percentage savings. Maximum use of staging on all current routes longer than 5,500km could reduce the total fuel consumed by between 3% and 20%, with a consequent reduction in global fleet annual fuel burn of between 1% and 6%. However, the savings for a given route are reduced as aircraft become more efficient.

The hot-wire anemometer is one of the few instruments which can be used to make velocity measurements in turbulent and unsteady flows. However, the probe supporting the wire inevitably interferes with the local flow and it has been found that the effect of this interference on the reading of the anemometer varies considerably as the orientation of the probe to the flow direction is changed (the wire itself being maintained in the same direction). This leads to errors in any measurements taken where the instantaneous local flow direction differs significantly at any time from the direction for which the anemometer was calibrated. Such errors are quite separate from, and in addition to, errors due to finite wire length, incidence of the wire to the local stream direction, etc.

The effects of aspect ratio on elliptical sonic jet decay at different levels of under-expansion has been studied experimentally. Elliptical sonic jets from orifices of same area with aspect ratios (AR) 2, 4 and 6 at nozzle pressure ratios 2 to 5 in steps of 1 have also been studied. A circular jet from an orifice with an area equal to that of elliptical orifice was also studied for comparison. Jet centreline pressure decay, spread and waves present in the jet core were analysed. The results show that the mixing of the elliptical jet is superior to the circular jet, at all the nozzle pressure ratios of the present study. Also, the aspect ratio of the elliptical orifice has a strong influence on the jet mixing. Elliptical jets of aspect ratio 4 and 6 experience a significantly higher mixing than the aspect ratio 2 jet, till the under-expansion level corresponding to Mach disk formation. For higher under-expansion levels, the mixing of AR 4 and 6 jets become inferior to that of the AR 2 jet.

The non-equilibrium fluid mechanics and thermodynamics of two-phase vapour-droplet and gas-particle flow are considered. The formation of the droplets as well as their subsequent interaction with the vapour are discussed. Five topics have been given particular attention: (i) CFD application to unsteady condensation waves, (ii) CFD application to shock waves moving through a vapour-droplet mixture, (iii) a new theory of nucleation of water droplets in steam turbines based on Monte Carlo simulation (steam turbines are responsible for 80% of global electricity production and the presence of moisture significantly reduces the turbine efficiency costing £50m per annum in the UK alone), (iv) a unified theory for the interpretation of total pressure and total temperature in two-phase flows and, (v) a unified theory of particle transport in a turbulent flowfield.

Unsteady high Mach number flows past oscillating cones and oscillating wedges are studied according to Newtonian impact theory, with and without centrifugal force corrections. It is found that, contrary to existing theories, the centrifugal force contribution in unsteady flow is significant and not negligible; for example, it contributes one-half to the dynamic stability derivatives of an oscillating wedge and one-third to that of an oscillating cone. It is also shown that in both cases the dynamic stability derivatives calculated by Newtonian impact theory with centrifugal force correction agree exactly with those obtained by gas-dynamic theory in the double limit γ → 1 and M∞ → ∞, independently. These conclusions are expected to hold for more general three-dimensional unsteady Newtonian flow when the Newtonian shock layer is attached.

A procedure for defining an empirical formula for the mass estimation of the fore and aft wings field Uof a medium range box wing aircraft is described. The procedure is based upon the work of Howe for estimating the wing mass of conventional cantilever wing aircraft. The paper outlines the procedure used to relate conventional cantilever wings to box wing aircraft wings. Using a vortex lattice tool, finite element methods and regression analysis, the modification performed on the coefficient in Howe’s method to enable its use on a medium range box wing aircraft is outlined. The results show that the fore and aft wings would use the same correction coefficient and that the aft wing would therefore be lighter than the fore wing on a medium range box wing aircraft.