A three-dimensional integral formulation for the solution of the full potential equation and the associated numerical algorithm, the field-integral method, are presented. The method is a direct extension of a two-dimensional method and in particular retains the simple grid generation requirements noted in that method. Results are presented for the flow over body shapes and a complex winglet configuration, and are compared with existing transonic methods and experiments with good agreement. The further work necessary to provide a fast, robust method for use in design is outlined.

Automated manufacturing of thermoplastic composites has found increased interest in aerospace applications over the past three decades because of its great potential in low-cost, high rate, repeatable production of high performance composite structures. Experimental validation is a key element in the development of structures made using this emerging technology. In this work, a $750\times640\times240$ mm variable-stiffness unitised integrated-stiffener out-of-autoclave thermoplastic composite wingbox is tested for a combined shear-bending-torsion induced buckling load. The wingbox is manufactured by in-situ consolidation using a laser-assisted automated tape placement technique. It is made and tested as a demonstrator section located at 85% of the wing semi-span of a B-737/A320 sized aircraft. A bespoke in-house test rig and two aluminium dummy wingboxes are also designed and manufactured for testing the wingbox assembly which spans more than 3m. Prior to testing, the wingbox assembly and the test rig were analysed using a high fidelity finite element method to minimise the failure risk due to the applied load case. The experimental test results of the wingbox are also compared with the predictions made by a numerical study performed by nonlinear finite element analysis showing less than 5% difference in load-displacement behaviour and buckling load and full agreement in predicting the buckling mode shape.

Flow separation results in many problems to single expansion ramp nozzle (SERN) and hypersonic vehicle. However, little research has been conducted on the separation patterns and their effects on SERN’s performance. In the present paper, the numerical simulation is adopted to get the intuitive results and help to analyse the separation phenomena in SERN thoroughly. The main separation pattern is the restricted shock separation (RSS) in SERN, and the free shock separation (FSS) only appears in a small range of the nozzle pressure ratio (NPR), which is much different from the axisymmetric rocket nozzle. Further CFD results show that the separation pattern transition makes great effects on the performance of SERN, especially the lift. Moreover, the performance of SERN has an extreme in the separation pattern transition because of the main jet impinging on the expansion ramp. The transitions occur in both the startup and shutdown processes but the critical nozzle pressure ratios of the separation pattern transitions are different, which leads to a hysteresis loop of SERN performance.

Weapon target allocation (WTA) is an effective method to solve the battlefield fire optimisation problem, which plays an important role in intelligent automated decision-making. We researched the multitarget allocation problem to maximise the attack effectiveness when multiple interceptors cooperatively attack multiple ground targets. Firstly, an effective and reasonable fitness function is established, based on the situation between the interceptors and targets, by comprehensively considering the relative range, relative angle, speed, capture probability and radiation source matching performance and thoroughly evaluating them based on the advantage of the attack effectiveness. Secondly, the optimisation performance of the particle swarm optimisation (PSO) algorithm is adaptively improved. We propose an adaptive simulated annealing-particle swarm optimisation (SA-PSO) algorithm by introducing the simulated annealing algorithm into the adaptive PSO algorithm. The proposed algorithm can enhance the convergence speed and overcome the disadvantage of the PSO algorithm easily falling into a local extreme point. Finally, a simulation example is performed in a scenario where ten interceptors cooperate to attack eight ground targets; comparative experiments are conducted between the adaptive SA-PSO algorithm and PSO algorithm. The simulation results indicate that the proposed adaptive SA-PSO algorithm demonstrates great performance in convergence speed and global optimisation capabilities, and a maximised attack effectiveness can be guaranteed.

It is fairly generally known that if, at any fixed speed of a compressor, either axial or centrifugal, the flow is reduced by throttling the outlet, then a point is eventually reached at which a complete breakdown of the air flow occurs and in most cases an actual flow reversal through the compressor takes place. Sometimes this phenomenon is gradual, taking the form of a sort of burbling, but more generally it takes the form of a sudden “ bang ” associated with a violent shake of the whole foundation of the bed on which the compressor is mounted. In most centrifugal compressors the flow reversal which takes place stops rapidly, the performance recovers, and a second “ surge ” occurs if the throttling is not reduced. The frequency of these individual surges varies greatly according to the degree of throttling and many other conditions; it may be only one isolated occurrence, in which case it would be concluded that the working point was just, but only just on the surge point, or it may take the form of a rapid series of thuds indicating that the working point is beyond the surge. In all cases of centrifugal compressors known to the authors reduction of the throttling will restore the compressor to its normal performance.

An historical sketch shows how the basic ideas of the jet flap were evolved. The exploratory work included the prediction that lift would be induced on an aerofoil by a two-dimensional jet deflected from the trailing edge, and that the magnitude of the total lift would be of the order of four times the direct jet lift. The prediction was confirmed experimentally and a first quantitative theory evolved. Meanwhile the hypothesis was put forward that, ideally, the forward thrust on the aerofoil would be independent of the deflection angle of the jet.

This paper is the first of a series of papers that will describe the development and present state of knowledge of the jet flap.

A method of analysis is presented for the shear stiffness of a flat-sided corrugated web, attached at a series of discrete points to flanges either on one side or on both sides of the web. Simple expressions are obtained for the shear stiffness of both types of web, and it is shown that there may be a considerable loss of stiffness due to the method of attachment. A comparison is made with experimental results for corrugated webs with both types of attachment.

Measurements of drag have been made, in a shock tunnel, on a simple integrated vehicle-engine combination for hypersonic cruise with hydrogen scramjet propulsion. The test flow Mach number was 6·4, and the velocity was 2·45 kms-1. Zero drag, which is the necessary condition for cruise, was achieved as the equivalence ratio approached one. It was found that an analysis using established aerodynamic concepts was adequate for predicting drag in the case of no combustion. When combustion occurred results of direct connect experiments provided a qualitative guide to the measured levels of drag, and indicated that thrust nozzle combustion was taking place. An heuristic analysis is used to point to the important effect this may have on propulsive lift.

Unsteady pressure data from seven aerofoil models performing constant pitch rate “ramp-down” motions from the fully separated state to low incidence have been analysed. It is observed that the time delay between attachment at the leading edge and the occurrence of minimum normal force is, to a first order, independent of reduced pitch rate. As a result, at sufficiently high pitch rate, negative normal force may be generated at large positive incidence. In contrast to this, the re-establishmeht of attached flow to the upper surface is strongly dependent upon pitch rate. The paper also contains a description of smoke flow visualisation tests. These were at a much lower Reynolds number than the pressure data tests, and consequently are of qualitative interest. Results of flow visualisation show that during a ramp-down test the near wake is much larger than in the corresponding static case, and that attachment is significantly delayed.

This investigation originated when some difficulties were experienced in establishing good flow conditions in a circular pipe to study the effects of longitudinal cooling on slanted hot-wires.

In the past many investigators have presented results on turbulent flow through pipes and it is now generally assumed that this flow field is known in sufficient detail. For example, Schlichting has given a through description of the mean velocity distribution for turbulent flow through pipes based on the experimental results of Nikuradse. On the other hand Townsend describes the structure of the flow based on the results of Laufer. Although the investigation of Nikuradse and Laufer are classical works there appears to be some uncertainty in whether or not their turbulent flows down pipes were really fully developed.

Although the Cebeci-Smith method of calculating turbulent boundary layers is widely used and generally gives acceptably accurate results, highly inaccurate skin-friction values are obtained for relaxing flows and equilibrium layers in strong adverse pressure gradient. In the present paper, these anomalies are removed by suitable modifications to the basic eddy-viscosity model.

Combined Vertical and short take-off and landing, or ‘V/STOL’ capability has been of great demand and interest in the field of aeronautics since the creation of the aircraft. V/STOL capability is a targeted capability for many projected or prototype future aircraft. Past V/STOL aircraft are reviewed and analysed with regard to their performance parameters. This research has found two embedded categories in this class of aircraft based on their propulsion systems, i.e. jet and non-jet propulsion, and highlights the significant performance differences between them. In light of historical experience the performance of a relatively new class of aircraft, the flying cars, has been evaluated.

The passive control of a shock wave-boundary-layer interaction involves placing a porous surface beneath the interaction, allowing high pressure air from the flow downstream of the shock wave to recirculate through a plenum chamber into the low pressure flow upstream of the wave.

The simple case of a normal shock wave at a Mach number of 1·4 interacting with the turbulent boundary layer on a flat wall is investigated both experimentally and numerically. The experimental investigation made use of holographic interferometry, while the computational section of the investigation made use of a Navier-Stokes code to derive pressure gradients, boundary-layer properties and total pressure losses in the interaction region. It is found that the structure of shock wave-boundary-layer interactions with passive control consists of a leading, oblique shock wave followed by a lambda foot. The oblique wave originates from the upstream end of the porous region, and its strength is determined by the magnitude of the local blowing velocities. The shape of the lambda foot depends on the position of the main shock relative to the control region, resembling an uncontrolled foot when the main shock wave is towards the downstream end of the porosity, but becoming increasingly large as the shock moves upstream and eventually merging with the leading, oblique shock to form a single, large, lambda structure.

Improved forms of passive control are suggested based on the findings of this investigation, including the use of passive control systems which incorporate streamwise variations in the level of porosity.

New alternative jet fuels have provided many advantages in the aviation industry, especially in terms of economics and environment. However, fuel–seal compatibility is one of the major issues that restricts alternative fuel advancement into the market. Thus, to help understand and solve the problem, this study examines the swelling effect of prepared and non-prepared O-rings in different fuels and aromatic species. Stress relaxation experiments were carried out to evaluate seal compatibility under compression, which mimics engine operation conditions. Seals were compressed and immersed in a variety of fuels and their blends for about 90h while maintaining a constant temperature 30°C and constant compression force of 25% seal thickness. The two types of elastomers investigated were fluorosilicone and nitrile O-rings, which are predominantly used in the aviation industry. Meanwhile, three different fuels and aromatic species were utilised as the variables in the experiments. The fuels used were Jet-A1, SPK and SHJFCS, while the aromatic species added were propyl benzene, tetralin and p-xylene. The swelling effects were determined from the P/Po value. Results indicate that Jet-A1 has the highest swelling effect, followed by SHJFCS and SPK. It was observed that the higher the percentage of aromatics in fuel, the higher the rate of swelling. Furthermore, prepared seals had a lower swelling rate than did non-prepared seals. Meanwhile, the intensity of the swelling effect in the Jet-A1-SHJFCS blends was in the order of 60/40, 85/15 and 50/50 blend. The work done in this study will aid in the selection of suitable aromatic species in future fuels. The novelty of this research lies in the determination of the appropriate amount of aromatic content as well as the selection of type of aromatic and its mixture fuel. Moreover, the various proportions of fuel blends with aromatic are investigated. The primary aim of this study is to understand the behaviour of prepared and non-prepared seals, and their compatibility with alternative fuels.

The mixing of single or multiple jets is relevant in several engineering applications ranging from flame stabilisation in combustion to high lift devices for V/STOL aircraft. In some of these applications, expeditious manufacturing and ease of installation may dictate the use of sharp-edged slots in preference to nozzles with contoured upstream shaping.

It has been suggested that cruciform nozzle exit shapes may be suitable for improving mixing and thus thrust augmentation in STOL aircraft. The effects of nozzle geometry on the performance of an axisymmetric thrust augmentor have been investigated experimentally and a very slight improvement in the performance of the axisymmetric thrust augmentor was observed when cruciform nozzles were used after the tests had been performed with circular nozzles.

Ever since I first began to study Aeronautics I have been annoyed by the vast gap which has existed between the power actually expended on mechanical flight and the power ultimately necessary for flight in a correctly shaped aeroplane. Every year, during my summer holiday, this annoyance is aggravated by contemplating the effortless flight of the sea birds and the correlated phenomenon of the beauty and grace of their forms.

In constructing an X-type hot-wire probe it has been the policy of a number of experimenters and manufacturers to place the two wires forming the X close to each other to assure that they are both measuring in effectively the same plane. A number of important experiments have been made using such a probe design. Recent experiments at the University of British Columbia and McGill University have shown that an X-wire probe which has two wires almost in the same plane is quite sensitive to movements of the velocity vector out of that plane (defined as pitching motion). This has been attributed to the influence on one wire of the hot wake produced by the other. In this note a Disa X-type probe (type 55A32), with the wires 0·15 mm apart, is tested and found to have a static sensitivity to small angles of pitch, which is very significant for low wire Reynolds numbers (< 5) but becomes small for Reynolds number greater than 10. A modified probe, having the wires one wire length apart, is suggested and when tested was found to have no pitch sensitivity.

The results of experiments on the reattachment of a laminar boundary layer, separating from a rearward facing step in a flat plate aerofoil, are correlated with the properties of the short leading edge bubble which forms on thin aerofoils near the stall.

The experiments, comprising pressure measurements, Pitot explorations, liquid film and smoke studies, indicate that for all Reynolds numbers above the value given by the Owen-KIanfer criterion the reattachment is turbulent behind a stationary air reverse flow vortex bubble. It is also found that the reattachment is laminar for Reynolds numbers below the critical, which further supports Crabtree's interpretation of the Owen-KIanfer criterion in terms of the condition for the growth of turbulent bursts.

The characteristic functions for beam vibration modes are used to derive an approximate solution for the calculation of thermal stresses in rectangular isotropic flat plates subjected to arbitrary temperature distributions in the plane of the plate and constant temperatures through the plate thickness. The thermal stresses are obtained in the form of generalised Fourier expansions in terms of the characteristic functions, and their derivatives, representing normal modes of vibration of a clamped-clamped beam. Since these functions have recently been tabulated, the practical application of this new method to the thermoelastic stress analysis of plates presents no difficulty.

This paper presents a study of the W3-Sokol main rotor equipped with Gurney flaps. The effect of the active Gurney is tested at low and high forward flight speeds to draw conclusions about the potential enhancement of the rotorcraft performance for the whole flight envelope. The effect of the flap on the trimming and handling of a full helicopter is also investigated. Fluid and structure dynamics were coupled in all cases, and the rotor was trimmed at different thrust coefficients. The Gurney proved to be efficient at medium to high advance ratios, where the power requirements of the rotor were decreased by up to 3.3%. However, the 1/rev actuation of the flap might be an issue for the trimming and handling of the helicopter. The current study builds on the idea that any active mechanism operating on a rotor could alter the dynamics and the handling of the helicopter. A closed loop actuation of the Gurney flap was put forward based on a pressure divergence criterion, and it led to further enhancement of the aerodynamic performance. Next, a generic light utility helicopter was built using 2D aerodynamics of the main aerofoil section of the W3 Sokol blade along with a robust controller, and the response of the rotorcraft to control inputs was tested. This analysis proved that the 1/Rev actuation of the Gurney did not alter the handling qualities of the helicopter, and as a result, it can be implemented as a flow control mechanism for aerodynamic enhancement and retreating blade stall alleviation.