The forces due to normal pressure have been measured in the region of the junction between a rear-loaded wing swept forward at 28°, and a flat plate on which a turbulent boundary layer had developed. The Reynolds number was 1·03 × 106, based on the streamwise wing-chord of 500 mm. At low incidences, 0° < α ≤ 6°, the local pressure drag coefficients were found to be negativeaway from the junction, and the interaction with the junction was found to be favourable;that is the pressure drag coefficient became more negative as the junction was approached. At α = −3° the pressure drag coefficient changed from a positive value away from the junction to a negative value close to the plate. These results contrast with those for a swept-back wing-plate interaction, which has been reported as being unfavourable so far as the pressure drag was concerned. At 6° incidence, the interaction remained favourable but there was evidence of root stall beginning. At an incidence of 9°, root stall was much more marked and the positive pressure drag coefficient increased as the junction was approached.

This paper describes the use of oct-trees to represent digitised terrain elevation data (DTED) in terrain reference navigation (TRN) systems. Oct-trees provide a regular method to represent digitised terrain data where the level of detail of the terrain is encoded in the tree structure. The use of oct-trees also provides a basis for a significant reduction in the on-line storage requirement for DT-EDs. A method of encoding using Morton ordering is introduced which allows DTEDs to be accessed as oct-trees and quad-trees where the tree structures are represented as pointerless structures. Algorithms to construct and access terrain oct-trees are presented which form a set of access primitives for tree operations required in TRN applications. The paper concludes with examples to illustrate the efficiency of the methods described in the paper for two DTEDs, in terms of performance of the tree access operations and reductions in storage.

The high enthalpy and high Mach number effects on flat plate boundary layer flow are examined. In particular, the behaviour of the density profile under high enthalpy, real gas conditions is considered. Mach-Zehnder interferograms of the flow over a flat plate are analysed using a two-dimensional Fourier transform procedure in order to obtain the density field. The comparison between theory and experiment is only fair. It must be pointed out, however, that the real gas effects on the density profiles were found to be small. Estimation of the velocity boundary layer thickness as obtained from a measurement of the thermal boundary layer thickness is seen to compare well with a theoretical prediction.

The paper is focuses on the study of the profile optimisation of a unidimensional structure in idealised critical flutter conditions. The problem has been addressed already by other authors, but here an original technique is used to search for some variables at the left hand end of a simply supported vibrating beam, the knowledge of which is necessary for the numerical integration of the governing equations. Additionally, the frequency is considered as a control parameter. A composite orthotropic symmetric panel with four layers is considered and the optimised profiles of all the structural components determined.

Steady, inviscid, incompressible, two dimensional flows containing vortex patches bounded by vortex sheets are used to model finite separated regions on certain closed bodies as an approach to modelling separation bubbles. First of all the problem of a vortex patch on a circle is considered. Calculations were done for symmetric flows covering a wide range of values of the separation position, the strength of the vortex sheet and the circulation at infinity and for non-symmetric flows at various angles of incidence. Smooth separation was detected in some cases. This model is used as an intermediate step for the calculation of the flow around an ellipse and a Joukowski aerofoil with the use of the Joukowski transformation.

A method of determination of the wing lift-to-drag ratio of a propelled hang-glider in steady, level flight is described. The method requires a relatively small number of flight parameters to be measured. These measurements have been reduced successfully to the tilt angle of the propelling “trike”, i.e. a closed, wheeled, cabin hanging under the wing, and to the force on the flying controls only. The recording of these parameters and actual flight speed and altitude is sufficient to determine, after simple processing, the experimental speed and drag polars of the wing. The method is useful, in particular, for flexible wing research and development and allows quick aerodynamic tests.

The space industry of the late 20th century is not simply a subset of the aerospace industry: it is a producer, an employer and a wealth creator in its own right. This paper reviews the current state of the art in space technology by analysing the development of an industry which produces both a wide variety of spacecraft and the launch vehicles which deliver them to orbit. In doing so, it considers the application of spacecraft for communications, navigation. Earth observation and science, as well as their application to manned space flight.

By considering some of the problems that need to be solved in order to develop the field beyond its current state, the paper then looks forward to the developments in these areas expected in the 21st century. The interrelated development of international co-operation — as opposed to political competition — and the rise of space commercialisation is considered throughout.

A non-linear methodology based on high resolution Navier-Stokes solutions has been proposed to predict incipient separation boundaries. Comparison has been made between the prediction and an empirical criterion based on vast experimental data.

This paper describes a simple method to accurately simulate the flow field around a wing in surface effect ship (Wises) model that involves the use of a contraction and a slot to suck the boundary layer from a flat plate. Theoretical aspects of the design are discussed with relation to the required level of suction and the slot width.

The effectiveness is ascertained from measurements of the boundary layer height on the plate through measurement of the mean velocity profile and a reduction in height of 50% was generally observed when suction was used.

Comparison of the forces on a low aspect-ratio wing, obtained from the proposed and the image method, verified the proposed method for wing heights above 2·5% of the mean chord. At lower heights, separation of the boundary layer on the ground plate is thought to have altered the pressure distribution on the wing.

Recent advances in two lifting surface methods are reported. The present subsonic constant pressure method improves upon the program robustness of the doublet lattice method. The present unified supersonic-hypersonic method extends the range of applicability of lifting surface theory to the Newtonian limit. Based on a uniformly-valid series formulation, the method accurately approximates the non-linear thickness effect, whereas this effect is neglected by the linear theory and overestimated by piston theory. In fact, the present results generally agree well with the trends of several Euler solutions. Among other findings, cases computed by the unified method confirm that the effect of thickness is to reduce the supersonic flutter speed.

The paper describes the aeroelastic analysis for the design, the ground verification tests and the flight test programme, for the Ranger 2000 training aircraft, developed jointly by Dasa and Rockwell International between 1991 and 1994, as a competitor for the next generation US Air Force and Navy Joint Primary Advanced Training System (JPats). Special efforts with respect to the aeroelastic stability were required for the T-tail configuration, for the design of the manual flight control system, and for the establishment of sufficient mass balance for the control surfaces. Several ground vibration tests were performed for the complete aircraft and for individual components for all major design improvements during the flight test evaluation. To minimise the required time for these tests, highly modular test equipment was required. For an efficient flutter flight test programme a reliable excitation system was chosen. This system consists of slotted rotating cylinders, mounted on small vanes, which can be attached to any aerodynamic surface. This equipment creates defined unsteady aerodynamic forces to excite the eigenmodes of the structure.

Quasi-on-line frequency and damping data evaluation between consecutive flight test points was made possible by installing the required hardware and software directly at the flight test quick-look control room, for the direct use of telemetry data. With this approach only a small number of dedicated flutter flights was required.

During recent years a new vortex model of the unsteady aerodynamics of a hovering rotor or a rotor in axial flight, named Temura, has been developed in the Technion. This model has succeeded in describing various phenomena of rotor dynamics. In the present paper Temura is used in order to develop a model which describes the blades’ responses to harmonic variations of their pitch angles. This model includes unsteady influences of trailing and shed vortices, together with unsteady geometric effects which represent a special capability of Temura. The model is general and can be applied to any number of blades and arbitrary differences between the pitch angle variations of different blades. The application of the model to analyse cases of collective harmonic pitch variations and differential harmonic pitch variations is presented. Numerical results of the new model are compared with experimental results from the literature and good agreement is shown.

Steady state aerofoil ground effect is studied both numerically and analytically. Discrete vortex and linear vortex panel methods are applied to a parabolic arc and symmetric Joukowski aerofoil, respectively. A single vortex model for the flow over the parabolic arc aerofoil is developed. The single vortex model and other analytical solutions, valid either near or far from the ground, are compared with the numerical results. The numerical results are used to delineate the influences of angle of attack, camber and thickness. For small values of camber and angle of attack, normalised lift is enhanced near the ground and reduced far from it. For a fixed distance above the ground, normalised lift decreases with increasing angle of attack and camber. Thickness reduces lift at all heights above the ground.

The paper uses the conventional nine-noded Lagrangian element for studying transient linear response analysis of composite stiffened shells. An improved version of the stiffener modelling has been used in which the stiffener can be placed anywhere inside the element. For the first time, concentric or eccentric stiffeners have been used for composite stiffened shells for solving the transient dynamic response of these structures. These are not available in existing commercial packages. Different types of time dependent loading such as short duration air-blast loading, suddenly applied uniformly distributed step loading and sinusoidally harmonic loading have been considered in this paper. The results of stiffened composite cylindrical shells and doubly curved shells with different boundary conditions and various laminate orientations have been presented for eccentric stiffeners. Parametric studies considering different variables have also been carried out.

Asymmetric vibration of multilayered conical shell with core layers of viscoelastic material are investigated in this paper. The analysis presented herein considers bending, extension in plane shear and transverse shear deformations in each of the layers, and also includes rotary, longitudinal translatory and transverse inertias. Appropriate trigonometric series are used as solution functions in the Galerkin method to reduce the governing equations to a set of matrix equations. The corresponding principle of linear viscoelasticity for harmonic motion is used for evaluating the damping effectiveness of shells with elastic and viscoelastic layers. A computer program has been developed for determining the resonance frequencies and associated system loss factors for various modes of families of asymmetric vibration of a general multilayered conical shell consisting of an arbitrary number of specially orthotropic material layers. Variation of resonance frequencies and the associated system loss factors with total thickness parameter and circumferential modal number for three, five and seven layered conical shells, with three sets of classical end conditions: simply supported at both ends, clamped-clamped and free-free, which can be of some use to designers, are been reported.

In planning interplanetary missions which involve an aerobraking manoeuvre, it is necessary to make accurate predictions of the aerodynamic drag acting on a vehicle during its descent. Of interest to the authors is the Nasa initiative for exploration of Mars and its atmosphere. The Mars Pathfinder is a probe that is expected to enter the Martian atmosphere at a relative velocity of approximately 7.6 kms-1;. The forebody of this vehicle is based on a 70° blunted cone and is typical of aerobraking designs.

In this note, a comparison is made between experimental and numerical techniques for predicting drag in hypervelocity flow. Three different models were examined in this study: a 30° sharp cone; an Apollo heat shield; and a Viking heat shield. A relatively simple analytical result for the drag on a cone provides a convenient reference for both the experimental and numerical results. The two heat shields are typical of those used for interplanetary exploration, such as the Mars Pathfinder. Our aim is to give an example of how computational fluid dynamics can be used in conjunction with experiments to obtain information about the hypervelocity flow about re-entry vehicles.

The paper reviews the developments in avionics over the last seventy years, focusing on the impact of computer technology during the last ten years. The development of aircraft navigation systems is described, including present day satellite navigation and terrain reference navigation. The techniques used in aircraft displays are outlined including CRTs and LCDs and covering head up displays, helmet mounted displays and synthetic displays. The systems architecture which forms the basis of modem avionics is described, covering the major aircraft databuses in use including recent developments in integrated modular avionics. Developments in real-time software, which has become a major cost in many programmes, are covered, leading to the techniques used to ensure aircraft reliability, particularly fault tolerant architectures and also safety systems for health monitoring and recent research in ‘pilot associates'. The paper also summarises developments in flight control systems for both civil and military aircraft. The paper concludes with a brief review of future developments in avionics, speculating on the role of pilot in future aircraft.

A knowledge-based computer system has been developed to automatically scan and interpret graphical displays of vibration data taken from strain gauges mounted on the blades of gas turbines performing acceleration tests. The various characteristic lines which appear on these displays are recognised and classified and the lines representing modes of vibration are examined to determine which mode is being excited and the severity of the excitation. All lines are compared, by a “blackboard” technique, against those predicted by the system working from the engine specification, laboratory tests and finite element analyses. Any unexpected features are reported for manual inspection.

Several series of tests have now been made on the system. In the first test on 130 diagrams, every one of over 3000 features visible to the human eye was detected and then classified.

The paper examines the influence on optimal structural designs of changing the finite element selection, the structural design criteria and the design concepts for an unmanned air-vehicle wing. Thus, single and twin-spar designs are considered with a number of composite design configurations. The results indicate that some of the design criteria have little influence on the optimal designs whilst finite element selection can be significant.