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 ‘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.

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 Repetitively Enhanced Neural Networks (RENN) method is developed and presented for complex and implicit engineering design problems. The enhanced neural networks module constructs an accurate surrogate model and avoids over-fitting during neural networks training from supervised learning data. The optimiser is executed by the enhanced neural networks models to seek a tentative optimum point. It is repetitively added into the supervised learning data set to refine the surfaces until the RENN tolerance is reached. The RENN method demonstrates the effectiveness and feasibility for a 2D highly non-linear numerical example and the structure design of a two-member frame reaching a convergent solution at 10 and 15 iterations at the maximum error of 1% when compared with the exact solution. Then, the RENN method is applied for a long endurance unmanned aerial vehicle (UAV) aerofoil design optimisation. A Class/Shape function transformation (CST) geometry parameterisation method represents an accurate UAV aerofoil with ten geometry design variables. The high-fidelity analysis solver with structured mesh is used for a UAV aerofoil design problem. Using the RENN method, an optimal UAV aerofoil is obtained using 88 high fidelity evaluations at an error of 1·24%. The process reduces the computational time by 81·2% compared with the full high fidelity model. The optimal aerofoil shows a drag reduction of 10·8% in the cruise condition and an improvement in the maximum lift coefficient and stall angle-of-attack when compared with the baseline AG24 aerofoil.

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 theoretical study is made of the problem of a long, thick cylindrical tube subjected to two equal and diametrically opposite normal loads. The stress state is analysed by the three-dimensional theory of elasticity, with the Papkovich and Neuber stress-function approach. Numerical results for stresses and radial displacement are obtained to show the nature of stress concentration in the neighbourhood of the load. Some related experiments using the photoelasticity technique are also accomplished. Favourable correlation is shown between the theory and the test results.

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.

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.

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.

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.

The complex aerodynamics of rectangular underslung helicopter loads can lead to severe stability problems, but are difficult to represent in flight dynamics models. Current models for box aerodynamics are highly unsatisfactory, being entirely empirical and requiring large amounts of experimental data to generate. This paper presents a new modelling approach, which takes account of the bluff-body nature of the flow, where loads are dominated by normal pressure forces. Existing experimental data is recast in body-axes form, with α and β replaced by velocity components perpendicular and parallel to the box faces. Force and moment data for a wide range of boxes then collapse onto a set of simple generic characteristics, with features that can be related directly to the underlying flow physics. Modelling of container aerodynamics is greatly simplified, and allowance for effects of turbulence, Reynolds Number, wind tunnel interference and geometry modifications becomes possible.

The stability and control derivatives are essential parameters in the flight operation of aircraft, and their determination is a routine task using classical parameter estimation methods based on maximum likelihood and least-squares principles. At high angle-of-attack, the unsteady aerodynamics may pose difficulty in aerodynamic structure determination, hence data-driven methods based on artificial neural networks could be an alternative choice for building models to characterise the behaviour of the system based on the measured motion and control variables. This research paper investigates the feasibility of using a recurrent neural model based on an extreme learning machine network in the modelling of the aircraft dynamics in a restricted sense for identification of the aerodynamic parameters. The recurrent extreme learning machine network is combined with the Gauss–Newton method to optimise the unknowns of the postulated aerodynamic model. The efficacy of the proposed estimation algorithm is studied using real flight data from a quasi-steady stall manoeuvre. Furthermore, the estimates are validated against the parameters estimated using the maximum likelihood method. The standard deviations of the estimates demonstrate the effectiveness of the proposed algorithm. Finally, the quantities regenerated using the estimates present good agreement with their corresponding measured values, confirming that a qualitative estimation can be obtained using the proposed estimation algorithm.

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.

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.

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.

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.