Experiments in a in (6-35 mm) channel have yielded further information on the precision and convenience of the razor blade technique. It is shown that adhesive tape or carefully located cement can be used to secure a segment of razor blade over a static pressure hole: the resulting calibration for shear stress remains valid if the blade is removed and relocated over the same or a different, similar sized hole. Razor blade segments, calibrated in this manner, have been used to measure wall-shear stress in a turbulent boundary layer with tangential, secondary injection: the results indicate that V. C. Patel’s law of the wall is valid for such flows.

Details of the design, construction and use of an electrical resistance network for solving certain of the problems encountered in the flexure and extension of flat plates are given. Some typical experiments have been carried out with the aid of this network and the method of operation of the analogy is described. The results of these experiments are presented and they are compared with the corresponding theoretical values. The comparison is as good as could be expected in view of the limitations imposed by the very coarse mesh which was used; essentially the same results would have been obtained by the exercise of a relaxation computation, using the same size of net. No practical difficulties would be encountered in building a large network having a very fine mesh.

This paper considers the problem of a reinforced elliptical hole in a plate under the action of a principal stress system of the type found in cylindrical and ellipsoidal pressure vessels. That is, stress systems in which the ratio of the principal stresses is not greater than two to one. It is shown that when the ratio of the major and minor axes of the ellipse can be chosen arbitrarily, practical reinforcements can be designed to give a maximum stress around the hole which is only slightly greater than the maximum stress in a similarly loaded plate with no hole. General expressions are obtained for the stress distribution in the plate around the hole, for the stress acting on a normal cross section of the reinforcement, and for the cross-sectional area of a reinforcement which gives a small stress concentration. These are used to find the variation in the stress distribution around the hole due to reinforcements having different cross-sectional areas when the applied principal stresses are in the ratio of two to one and Poisson's ratio for the material of the plate and reinforcement has practical values.

When a uniform, two-dimensional supersonic flow expands suddenly round a corner in a wall it forms a pattern known as a Prandtl-Meyer expansion or centred simple wave. If the flow is two-dimensional but not initially uniform, or if it is axially-symmetrical, the expansion is still centred, but is not a simple wave. An approximate solution is given in this paper for the isentropic, irrotational, steady two-dimensional or axially-symmetrical flow of a perfect gas in the neighbourhood of the centre of such an expansion. The solution is designed to replace the conventional method of characteristics in such a region.

The main application is to a jet issuing from a nozzle that discharges into a container with a pressure lower than that in the nozzle; in particular, a formula is derived for the initial curvature, at the lip of the nozzle, of the boundary of the jet. The solution also applies to the flow near an edge in a boundary wall, and a formula is derived for the velocity gradient on the wall immediately downstream of the edge.

An appreciation is made of recent experimental and project work of relevance to the design of a very light weight aircraft capable of being taken off and flown by two men. The emphasis is on the aerodynamic problems involved. It is concluded that, although there are many difficulties still needing attention, all the information supports previous assertions that flight by muscular power alone is possible. The merits and de-merits of a particular projected design are studied, and its performance and stability assessed.

This paper presents a multi-disciplinary analysis methodology for a box-wing aircraft configuration optimised for a given mission scenario. This conceptual design methodology and associated toolchain combines well-established vortex lattice analysis and a newly developed structural analysis tool called WingMASS, allowing the design space to be explored from a combined aerodynamics and structural design perspective. For a given mission scenario, the method optimises a box-wing configuration and compares it with an equivalent conventional configuration. This study shows that, for a given mission, a box-wing configuration can lead to a fuel burn reduction of up to 5% by optimising aspect ratio, horizontal and vertical wing separation.

Aircraft conflict resolution is an important part of air traffic control operations. This study presents a mixed integer linear programming model (MILP) using a space discretisation technique to deal with aircraft conflict resolutions in en-route flight operations. The purpose of space discretisation is to concentrate on only the significant points of the airspace. The model integrates the multi entry point approach with an airspeed adjustment technique in the horizontal plane. The model aims to generate conflict-free trajectories while minimising the total changes in entry points and airspeed values. A new heuristic algorithm was developed due to the complexity of the problem. The computational results demonstrated that the proposed approach resolved aircraft conflicts for 450 different traffic scenarios in less than a minute. Considerable fuel savings were achieved with no significant increase in delay or flight time compared to conventional vectoring techniques in a fixed entry point airspace structure.

With the aims of overcoming the limitations of the existing basic flow model derived from an axisymmetric generating body and extending the aerodynamic design method of the airframe/inlet integrated waverider vehicle, this study develops an upgraded basic flow model derived from an axisymmetric shock wave. It then upgrades the design method for airframe/inlet integration of an air-breathing hypersonic waverider vehicle, which is termed the ‘full-waverider vehicle’ in this study. In this paper, first, the design principle and method for the upgraded full-waverider vehicle derived from an axisymmetric basic shock wave are described in detail. Second, an upgraded basic flow model that accounts for both internal and external flows is derived from an axisymmetric basic shock wave by use of both the streamline tracing method and the method of characteristics (MOC). Third, the upgraded full-waverider vehicle is developed from the upgraded basic flow model by the streamline tracing method. Fourth, the design theories and methodologies of both the upgraded basic flow model and the upgraded full-waverider vehicle are validated by a numerical computation method. Finally, the aerodynamic performances and viscous effects of both the upgraded basic flow model and the upgraded full-waverider vehicle are analysed by numerical computation. The obtained results show that the upgraded basic flow model and aerodynamic design method are effective for the design of the airframe/inlet integration of an air-breathing hypersonic waverider vehicle.

In this paper, a theoretical study of the turn manoeuvre of an agricultural aircraft is presented. The manoeuvre with changeable altitude is analyzed, together with the, effect of the load factors on the turn manoeuvre characteristics during the field-treating flights. The mathematical model used describes the procedure for the correct climb and descent turn manoeuvre. For a typical agricultural aircraft, the numerical results and limitations of the climb, horizontal and descending turn manoeuvre are given. The problem of turning flight with changeable altitude is described by the system of differential equations which describe the influence of the normal and tangential load factors on velocity, the path angle in the vertical plane and the rate of turn, as a function of the bank angle during turning flight. The system of differential equations of motion was solved on a personal computer with the Runge-Kutta-Merson numerical method. Some analytical and numerical results of this calculation are presented in this paper.

To solve the on-ground lateral direction control problem of the unswept flying-wing unmanned aerial vehicle (UAV) without rudder, steering system or breaking system, a control approach which uses differential propeller thrust to control the lateral direction is proposed. First, a mathematical model of the unswept flying-wing UAV on-ground moving is established. Second, based on the active disturbance rejection control (ADRC) theory, a yaw angle controller is designed by using the differential propeller thrust as the control output. Finally, a straight line trajectory tracking control law is designed by improving the vector field path following method. Experiment results show that the proposed control laws have a shorter response time, better robustness and better control precision compared with proportional integral derivative (PID) controller. The proposed controller has small computational complexity, simple parameter setting process, and uses practical measurable physical quantities, providing a reference solution for further engineering applications.

Modern high-speed aircraft, especially military, are very often equipped with single or compound delta wings. When such aircraft operate at high angles-of-attack, the major portion of the lift is sustained by streamwise vortices generated at the leading edges of the wing. This vortex-dominated flow field can breakdown, leading not only to loss of lift but also to adverse interactions with other airframe components such as the fin or horizontal tail. The wind tunnel and water studies described herein attempt to clarify the fluid mechanics of interaction between the strake and wing vortices of a generic 76°/40° double-delta wing leading to vortex breakdown. Some studies of passive control using fences at the apex and kink region are also described. Various diagnostic methods-laser sheet flow visualisation, fluorescent dyes, and pressure sensitive paints have been used.

Ion thrusters are well suited for near-Earth and deep space missions because of their exceptionally high specific impulse. Those already flying are also proving to be very reliable. The flow of neutral and charged particles within the different components of these devices is very complex and although they have been studied for several years, little of the detail of their internal operation is properly understood. Numerical simulations of these flows can potentially yield an improved understanding of the physical phenomena involved and in turn this should assist in the optimising of new designs and in achieving higher specific impulses. The external plume, which can potentially cause damage to other spacecraft components such as solar panels, has been the object of some studies but the flow within the internal chambers where many complex phenomena are suspected to occur has been seriously neglected. For this reason the present fully kinetic neutral and charged particle simulations, which take account of the ionisation processes and the applied magnetic field, have been conducted. The studies have been made on one specific electrostatic ion thruster - the T5 model produced by the Space Department at the Defence and Evaluation Research Agency (DERA, Farnborough, UK). Two regions within the thruster have been concentrated on, namely the exit area of the hollow cathode and the main chamber.

The coupling of independent structural dynamic and inviscid aerodynamic models, in the time domain, is considered. The accuracy and CPU requirements of the two common approaches, namely ‘weak’ and ‘strong’ coupling procedures, are investigated. It is found that the strong coupling scheme is more accurate than the weak coupling approach, and only for large real time-steps is the strong coupling scheme more expensive. The computational method developed is used to perform transonic aeroelastic and aeroservoelastic calculations in the time domain, and used to compute stability (flutter) boundaries of 2D wing sections. A control law is implemented within the aeroelastic solver to investigate active means of flutter suppression via control surface motion. Comparisons of open and closed loop calculations show that the control law can successfully suppress the flutter and results in a significant increase in the allowable speed index in the transonic regime. The effect of structural nonlinearity, in the form of hinge axis backlash is also investigated. The effect is found to be destabilising, but the control law is shown to still alleviate the destabilising effect.

The search for new materials for use in high performance structures, such as aircraft, has resulted in the development of carbon fibre reinforced plastics, CFRP. Parts of aircraft have been manufactured from CFRP, particularly in areas where high acoustic loads are experienced. The acoustic fatigue resistance of fibre reinforced structures is therefore of considerable interest. If the acoustic fatigue problem is to be studied, the response of such structures to acoustic excitation must be predicted and for this, knowledge of material characteristics is essential.

The first stage in the investigation of structural materials is to obtain measured values of the dynamic properties. Data of interest to structural dynamicists are moduli of elasticity and damping information.

The continued trend of increased weight and landing speeds of modern aircraft has strengthened the demand for improved materials for the manufacture of aircraft brakes.

The aircraft brake is similar to the car disc brake since both represent a medium for the conversion of kinetic energy to heat energy by means of friction. There are, however, unique problems which confront the aircraft brake designer since he is required to provide an acceptable mechanism involving components of minimum weight. The present paper discusses some general aspects of aircraft brakes and deals in more detail with some of the materials and material requirements for aircraft brakes; in particular it is hoped to show the manner in which advanced materials can provide some answers to the problems of the most recent development in design.

In this paper an attempt is made to analyse the overall gains that have been obtained in aerodynamic cleanness in the past, to review the overall performance improvement that can be expected in the immediate future from increased cleanness along orthodox lines, and to survey critically certain more advanced concepts aimed at giving still lower drag coefficients and better performance characteristics.

In this paper, a new optimisation method incorporating lamination parameters and a guide-based blending model is proposed. Lamination parameters for a guide laminate and ply number of each panel are employed as design variables for optimisation with a parallel real-coded genetic algorithm incorporating structure behaviour and manufacturing constraints. During the optimisation process, with a form of least squares fitting adopted, another genetic algorithm is used to obtain the guide stacking sequence of the guide laminate from the guide lamination parameters, and then the laminate configurations of each panel are obtained from the guide stacking sequence and number of plies for each panel. The proposed framework is demonstrated via design of an 18-panel horseshoe configuration, where each panel is optimised individually with a buckling constraint. Numerical results indicate that the present algorithm is capable of obtaining fully blended designs.

The paper reviews the issue of cruise Mach number and addresses the benefits of operating subsonic commercial aircraft at speeds below the long-range cruise speed. The case considered is the flight of transport aircraft for flight segments up to 1,000nm. It is shown that the fuel burned is decreased by as much as 1·8% on a nominal 1,000nm stage length for operation around the long-range cruise Mach number, or below. This is achieved at a cost of a marginal delay on each flight segment (less than three minutes). The longer flight time is likely not to affect the daily operation of the aircraft. The fuel saving is compounded, because the gross take-off weight (GTOW) is recalculated to take into account the reduced fuel consumption at each flight segment. The analysis into the environmental benefits includes the reduction in,andemissions, and the heat released in the high atmosphere. Sensitivity analyses are carried out on the take-off weight, on the aerodynamic coefficients, on the transonic drag rise and the weight uncertainty. It is predicted that the optimal operation of the example aircraft over a nominal 1,000nm route can reduce the fuel consumption by as much as 150,000kg per year in comparison with an operation at the long-range Mach number. The aircraft model has a maximum take-off weight of 170,000kg and is powered by two GE CF6-80C2 engines.

One hundred and two years ago, after the Wright brothers had just attempted another unsuccessful flight, they predicted that it would be another 50 years before manned flight was achieved. Only two years later and 100 years ago this year, Orville Wright achieved the first powered flight at Kitty Hawk in North Carolina, illustrating just how difficult it is to predict the future and at the same time launching the pioneering age of aviation.

The aerodynamic shape design of a proprotor for a tiltrotor aircraft is a very complex and demanding task because it has to combine good hovering capabilities with high propeller efficiency. The aim of the present work is to describe a two-level procedure and its results for the aerodynamic shape design of a new rotor blade for a high-performance tiltwing tiltrotor aircraft taking into account the most important flight conditions in which the aircraft can operate. Span-wise distributions of twist, chord and aerofoil were chosen making use of a multi-objective genetic optimiser that worked on three objectives simultaneously. A non-linear sweep angle distribution along the blade was designed to reduce the power losses due to compressibility effects during axial flight at high speed. During the optimisation process, the aerodynamic performance of the blade was evaluated with a classical two-dimensional strip theory solver. The optimised blade was than analysed by means of a compressible Navier-Stokes solver and calculations were validated comparing numerical results with experimental data obtained from wind-tunnel tests of a scaled model of the proprotor.