In this paper the results are given of an experimental investigation into the flow development and the surface pressures in an unsteady shock boundary layer interaction at Mach numbers of 1.3 and 1.5. The experiment is one in which a naturally grown turbulent boundary layer on the tunnel walls is disturbed by a normal shock wave spanning the test section of the tunnel. The shock wave is oscillated by a periodic pressure disturbance in the tunnel diffuser far downstream of the interaction. Both instantaneous (spatially averaged) and time averaged results for the surface pressures are presented. These results are used to study the flow development and phase changes through the interaction over a frequency range from 34 to 167 Hz. This frequency range was chosen to cover the range of parameters likely to occur in conditions corresponding to stall flutter in the compressor fans of large bypass engines.

The nature of flapping torsion flutter of a helicopter blade in forward flight is discussed. The essential complication in the analysis is the presence of periodic coefficients in the equations of motion; approximate solutions are obtained by use of a perturbation procedure. An unusual behaviour of the flutter equations which occurs when the fundamental frequency of flutter is a half-integer multiple of rotational speed is studied. Two different instability mechanisms can be distinguished and are related to the two energy sources in the system, namely the rotation of the rotor and the forward speed of the helicopter. It is found that forward flight can have a significant stabilising influence on flutter and that, as the tip speed ratio increases, flutter occurs predominantly at half-integer frequencies. The results are confirmed by the use of a numerical method.

A static model of the hose and drogue used to provide data in a realtime flight simulation of air-to-air refuelling is presented. The model predicts (a) the interference effect of the flow around the nose of a combat receiver aircraft on the pre-contact position of the drogue and (b) the hose shape and the loads induced on the receiver probe during contact. The use of the static model in the pre-contact case is validated for a normal closure speed by comparison with the results from a small perturbation dynamic analysis. At higher closure speed the static model is found to be less satisfactory. In a typical receiver approach a movement of the drogue similar to the drogue radius is predicted.

In September 1957 the first small satellite was placed into orbit around the earth by the U.S.S.R. Since that time a very competitive situation has developed between the U.S.S.R. and the U.S.A. Each side has presumably used its military ballistic missiles, together with hurriedly-developed second and third stages. We are now reaching a situation in the West where more carefully prepared and optimised rocket stages are being used, with the basic I.C.B.M. and I.R.B.M. first stages. The result of this is to put larger payloads into orbit for a given launch weight. Recently, cost figures for the past^ present and future satellite launchers have been quoted.

This paper presents the results of some calculations of the incompressible turbulent boundary layer on an infinite yawed wing. A discussion is made of the effects of increasing lift coefficient, and increasing Reynolds number, on the displacement thickness, and on the magnitude and direction of the skin friction. The effects of the state of the boundary layer (laminar or turbulent) along the attachment line are also considered.

A study is made to determine whether the behaviour of the boundary layer can adequately be predicted by a two-dimensional calculation. It is concluded that there is no simple way to do this (as is provided, in the laminar case, by the principle of independence). However, with some modification, a two-dimensional calculation can be made to give an acceptable numerical representation of the chordwise components of the flow.

Experiments were made on a two-dimensional low speed diffuser with passive control. The passive control consisted of porous surfaces made of holes/slots connected by a breather passage. Passive control postpones the stall and reduces buffet levels in the diffuser. This test confirms that passive controls to prevent flow separation are not restricted to transonic speeds.

A simple method is developed for computing the trailing vorticity which arises when a non-uniform stream is turned.

It is shown that, for a sudden and constant deflection of a non-uniform stream, no net trailing vorticity is set up in the exit flow and hence there is no secondary motion.

In the case of an impulse cascade of finite dimensions with constant turning, it is found that the trailing vorticity has three distinct components—the passage vorticity and two components which appear as vortex sheets springing from the trailing edges of the aerofoils. It is shown that for small angles of deflection there is no net circulation associated with the trailing vorticity downstream, of the cascade, and it is inferred that this should still be so for large deflections.

This paper aims at contributing to a better understanding of the effect of Tyler–Sofrin Modes (TSMs) on forced vibration responses by analysing a 4.5-stage research axial compressor rig. The first part starts with a brief review of the involved physical mechanisms and necessary prerequisites for the generation of TSMs in multistage engines. This review is supported by unsteady CFD simulations of a quasi 2D section of the studied engine. It is shown that the amplitude increasing effect due to mistuning can be further amplified by the presence of TSMs. Furthermore, the sensitivity with respect to the structural coupling of the blades and the damping as well as the shape of the expected envelope is analysed.

The second part deals with the Rotor 2 blisk of the research compressor rig. The resonance of a higher blade mode with the engine order of the upstream stator is studied in two different flow conditions realised by different variable stator vane (VSV) schedules which allows to separate the influence of TSMs from the impact of mistuning. A subset of nominal system modes representation of the rotor is used to describe its mistuned vibration behaviour, and unsteady CFD simulations are used to characterise the present strength of the TSMs in the particular operating conditions. Measured maximum amplitude vs blade pattern and frequency response functions are compared against the predictions of the aeromechanical models in order to assess the strength of the TSMs as well as its influence on vibration levels.

The aeroelastic modeling and flutter characteristics of a shear deformable wing/stores configuration under pull-up angular velocity is investigated. An isotropic non-uniform wing, which structural model incorporates flexibility in transverse shear and warping effects, is considered. The aeroelastic governing equations and boundary conditions are determined via Hamilton’s variational principle. In order to exactly consider the span wise location and properties of the attached stores the generalised function theory is used. The partial differential equations are transformed into a set of eigenvalue equations through the extended Galerkin’s approach. Numerical simulation highlighting the effects of the pull-up angular velocity and store parameters and configurations, such as mass ratio and their attachment locations, on the flutter speed are presented. The results of flutter analyses are validated with the published results and good agreement is observed. Furthermore, the procedure for an optimal deployment of stores is obtained for the case of the wing with four stores.

This paper investigates a method for autonomous obstacle avoidance for fixed-wing unmanned aerial vehicles (UAVs), utilising potential fluid flow theory. The obstacle avoidance algorithm needs only compute the instantaneous local potential velocity vector, which is passed to the flight control laws as a direction command. The approach is reactive, and can readily accommodate real-time changes in obstacle information. UAV manoeuvring constraints on turning or pull-up radii, are accounted for by approximating obstacles by bounding rectangles, with wedges added to their front and back to shape the resulting fluid pathlines. It is shown that the resulting potential flow velocity field is completely determined by the obstacle field geometry, allowing one to determine a non-dimensional relationship between obstacle added wedge-length and the corresponding minimum pathline radius of curvature, which can then be readily scaled in on-board implementation. The efficacy of the proposed approach has been demonstrated numerically with an Aerosonde UAV model.

The present paper presents oil flow visualisations and pressure measurements over a hemisphere-cylinder body attached with a forward facing spike at Mach 6 and Reynolds number of 1·38 × 108 at 0° and 5° angle-of-attack. The oil flow pictures depict the separation region in the vicinity of the spike on the hemisphere-cylinder body. The oil flow visualisations will help to locate the reattachment shock wave on the hemisphere-cylinder body and also understand the flow field behavior on the blunt-nosed spike configuration. The pressure measurements over the hemisphere-cylinder body depend on the shape and the length of the spike. The pressure distributions over the blunt-nosed body show significant influence of the angle-of-attack. The maximum pressure coefficient on the hemisphere-cylinder body is a function of the spike length, shape of the aerodisk and angle-of-attack. The windward and leeward sides pressure variations show dependence of the geometrical parameters of the spike and shape of the spike. The hemisphere and the flat-faced aerodisk cause considerable reduction of pressure leading to decrease of aerodynamic drag compared to the conical spike.

The present paper describes a new, agent-based computer model that can simulate the evacuation of narrow body transport aeroplanes in the conditions prescribed by the airworthiness regulations for certification. The input data are extracted from a complete plan view of the cabin. The results include full egress details of all occupants, passengers and crew-members, and the most significant evacuation figures and diagrams. The model has been tuned and verified with real data of narrow body certification demonstrations. Numerical simulations of six narrow body aircraft, representative of current designs, show the capabilities of the model and provide relevant information on the relationship between cabin features and emergency evacuation results. Although the computer model has been developed for helping in the certification process it would be useful too in the design of new cabins.

A key challenge in open rotor design is getting the optimum aerodynamics at both the cruise and take-off conditions. This is particularly difficult because the operation and the requirements of an open rotor are very different at cruise compared to takeoff. This paper uses CFD results to explore the impact of various design changes on the cruise and take-off flow-fields. The paper then considers how a given open rotor design is best operated at take-off to minimise noise whilst maintaining high thrust. The main findings are that various design modifications can be applied to control the flow features that lead to lost efficiency at cruise and increased noise emission at take-off. A breakdown of the lost power terms from CFD solutions demonstrates how developments in open rotor design have led to reduced aerodynamic losses. At take-off, the operating point of the open rotor should be set such that the non-dimensional lift is as high as possible, without causing significant flow separation. This can be achieved through suitable amounts of re-pitch and speed up applied to a design. Comparisons with fully three-dimensional CFD show that the amount of re-pitch required can be determined using simplified methods such as two-dimensional CFD and a Blade Element Method.

The purpose of this paper is to design and develop an advanced co-flow fluidic nozzle, based on the Coanda effect concept, for multi-directional thrust vectoring of small jet engines. Recent progress on finding an optimal geometry with a fixed momentum ratio to achieve maximum thrust deflection angle is discussed here. The efficiency of the system is found to be a weakly nonlinear function of the secondary to primary flow momentum as well as three geometric parameters. A useful empirical formulation is derived for thrust vectoring angle, based on a series of tests carried out on different nozzles. An accurate computational fluid dynamics model is also developed and evaluated against the experimental data. Moreover, quasi-Newton optimisation algorithm is employed to find an optimal geometry with a constant relative jet momentum and a constant secondary slot size. In this technique, the optimal wall geometric parameters are calculated in the direction of the steepest gradient with the help of the numerical simulation model in every iteration step. Additionally, an optimised fluidic nozzle is constructed to experimentally verify the numerical results and the empirical equation.

This paper presents an experimental investigation related to controlling the unsteady characteristics of the separated shear layers occurring over highly swept wings, and in particular focuses on application of synthetic jet actuators for modification of unsteady dynamic loading on the wing surface due to the phenomenon referred to as vortex breakdown (vortex burst).

In the post burst flow region the surface pressure measurements reveal the presence of certain characteristic spectral peaks that are thought to represent the presence of a spiralling filament of vorticity inside the expanded vortex that is known to be present in the burst flow over swept wings.

This paper details an investigation into how the use of an array of 18 discrete synthetic jet actuators, distributed along the leading edge of a delta wing with a 60° sweep angle, can be used to alter the spectral content of this unsteadiness and reduce the level of unsteady pressure found in the post-burst region toward the wing trailing edge by up to 40%.

Measurements of the surface pressure spectral distributions over the wing are presented together with PIV measurements of the vortex cross-section, conducted in the successive planes parallel to the wing trailing edge. Additional surface flow visualisation indicates that the effect of the actuators on the leading edge boundary layer is to induce local separation delays close to each actuator orifice, which introduce ‘ripples’ into the shear layer as it separates. The results obtained are used to formulate an interpretative hypothesis attempting to explain the mechanisms responsible for modification of the spectral content and the level of excitation measured on the wing surface.

This paper presents the development and testing through simulation of an integrated scheme for aircraft sub-system failure detection and identification (FDI) based on the artificial immune system (AIS) paradigm augmented with artificial neural networks. The features that define the self within the AIS paradigm include neural estimates of the angular accelerations produced by the abnormal conditions. The simulation environment integrates the NASA Generic Transport Model interfaced with FlightGear. A hierarchical multi-self strategy was investigated for developing FDI schemes capable of handling malfunctions of a variety of aircraft sub-systems. The performance of the FDI scheme has been evaluated in terms of false alarms and successful detection and identification over a wide flight envelope and for several actuator and aerodynamic surface failures. For all cases considered, the performance was very good, confirming the potential of the AIS paradigm augmented with the proposed neural network-based approach for feature definition to offer a comprehensive solution to the aircraft sub-system FDI problem.

Any time a new aircraft is introduced into service, or an old aircraft undergoes substantial modifications or needs to be certified to carry and employ new stores, the store separation engineer is faced with a decision about how much effort will be required to provide an airworthiness certification for the aircraft and stores. Generally, there are three approaches that have been used: wind-tunnel testing, computational fluid dynamics (CFD) analyses and flight testing. During the past twenty years there have been considerable advances in all three areas. In particular, there has been a considerable improvement in the speed and validity of CFD results for store separation. The Holy Grail of CFD has long been the reduction/replacement of wind-tunnel testing. This would mean in store separation the ability to go from a CFD calculation to flight testing at the end point. The paper will describe how this was achieved for the F/A-18C/Litening pod program.

Sudden release of accreted ice in fuel systems could pose a serious challenge in aircraft operation. The resultant snowshower may reach the filter and fuel-oil heat exchanger, causing a restriction in fuel flow to the engine. It is fundamental to have an appreciation of the character and the interface shear strength of the accreted ice in aircraft fuel systems. This helps to recognise factors for the sudden release of the accreted ice and the intensity of the consequential snowshower. An experimental study was carried out to quantify the character and the interface shear strength of accreted ice on subcooled surfaces submerged in jet fuel. Ice was accreted on naked aluminium, painted aluminium and carbon fibre composite surfaces at various subcooled temperatures. The accreted ice was akin to fresh snow and exhibited soft and fluffy attributes. The character may be expressed quantitatively in terms of the porosity and was found to be c. 0·95. The ice weakly adhered to the substrate surfaces, and the interface shear strength was found to be c. 0·36Pa and c. 2·19Pa at the top surface and at the vertical surface of a specimen block, respectively. It was not possible to detect any variation in the porosity and the interface shear strength for different types of surface finishes and differences in water affnity in fuels due to the crude approach in the estimation of these parameters.

In this paper, fire and evacuation computer simulations are used to reconstruct the 1985 Manchester Airport B737 fire that resulted in the loss of 55 lives. First the actual fire and evacuation are reconstructed. Secondly, the impact of exit opening times and external wind on the fire and evacuation are investigated. Finally, the potential benefit offered by modern materials is evaluated. The results suggest that the number of fatalities could have been reduced by 87% had the forward right exit not malfunctioned and by 36% had the right over-wing exit been opened without delay. Furthermore, given the fuel pool size and location, a critical wind speed of 1.5m/s is necessary to cause the fire plume to lean onto the fuselage eventually resulting in fuselage burn-through. Finally, it is suggested that the use of modern cabin materials could have made a significant difference to the fire development and survivability.

Analysis of the performance of a 1/4.71 model-scale and full-scale Sikorsky S-76 main rotor in hover is presented using the multi-block computational fluid dynamics (CFD) solver of Glasgow University. For the model-scale blade, three different tip shapes were compared for a range of collective pitch and tip Mach numbers. It was found that the anhedral tip provided the highest Figure of Merit. Rigid and elastic full-scale S-76 rotor blades were investigated using a loosely coupled CFD/Computational Structural Dynamics (CSD) method. Results showed that aeroelastic effects were more significant for high thrust cases. Finally, an acoustic study was performed in the tip-path-plane of both rotors, showing good agreement in the thickness and loading noise with the theory. For the anhedral tip of the model-scale blade, a reduction of 5% of the noise level was predicted. The overall good agreement with the theory and experimental data demonstrated the capability of the present CFD method to predict rotor flows accurately.