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.

This paper begins with an outline of the procedure for predicting the infrared signature emissions from the airframe, engine casing, and the plume, and their attenuation by the intervening atmosphere. These emissions are contrasted against the background, to obtain the infrared signature levels. The infrared detector’s — noise equivalent flux density, is proposed as an operational constraint on the flight envelope. The shift of this newly imposed constraint on the flight envelope for several engine-operating conditions, and for turbojet and turbofan engines is studied. The signature levels from the casing and plume, of a turbofan and equivalent turbojet engine, are compared at different operating points on the flight envelope. Result in the form of a polar plot of infrared signature level variation with aspect is also examined for low flying missions. The results are analysed to direct stealth design and operation.

Benchmarks are defined, against which the performance of a given engine may be judged, with general indications for engine selection and aspects of design through which improvements may be sought. These are derived from published compilations of engine performance data obtained from suppliers around the world, which, if regularly refined and updated, will allow progress in small engine design to be systematically monitored and encouraged.

An experimental study was conducted to analyse the pressure distribution along the floor of a cavity, with and without the presence of an upstream tandem cavity, at a constant freestream Mach number of about 0-911. Measurements were made for single cavities and the results compared with those obtained in the presence of an upstream tandem cavity. This comparison was made over a wide range of geometries, covering open to closed classes of cavities with both identical and different dimensions for the two cavities. The effect of the spacing between the two cavities was also studied. The downstream cavity is shown to be significantly affected by the presence of an upstream cavity, with both the overall net static pressure and its gradient being affected, dependent upon the class of cavity geometry and spacing under consideration.

Basic requirements for aero engine parts are to carry loads at high temperatures safely over service periods of 30-40 years. This generally calls for the best and latest materials and for highly sophisticated design and manufacturing procedures.

These requirements are frequently in contrast to the desire to keep costs at affordable levels. This aims at the costs at both the customer’s side as well as the manufacturer’s side. The measures to reduce cost in the design and production phase are various:

• redesign to allow for less complex manufacturing processes

• reassessment of quality characteristics without reduction of overall quality (only those characteristics to be maintained which are really critical)

• acceptance of minor damage in new parts which would be acceptable for repaired parts according to technical standards

• assessment of all steps in the production process with respect to their influence on the final product properties

• more stabilised manufacturing processes which use only a portion of accepted tolerance bands to reduce the subsequent test effort

• change of material.

In the service phase, cost reduction can be achieved by optimum exploitation of part’s life potential. The means are improvement of the lifing concepts where portions of the crack propagation life may be utilised in addition to the traditionally cleared crack initiation life. Accurate records of life consumption can be obtained with modern systems for life usage monitoring.

Aerospace designers are increasingly interested in predicting unsteady flowfields such as those associated with store release, rotating propellers etc. However, the cost of performing fully unsteady calculations is usually prohibitively expensive. In order to address this problem for unsteady flows driven by a moving surface, a novel method is presented which calculates the time derivates as an analytic function of the instantaneous flowfield. This allows an accurate solution of the unsteady flow equations to be calculated using a quasi-unsteady approach. The validity of this approach is demonstrated for a store release and a propeller test case. Possible extensions to this method for more complex unsteady flows are presented.

This paper describes an investigation of methods of controlling transition and separation in the leading edge region of military aircraft wings. For wings with the high leading edge sweep relevant to some military aircraft, if attachment line contamination can be prevented then transition is predominantly caused by crossflow instability close to the leading edge. The use of surface suction or cooling for suppressing these instabilities in order to delay transition, has been investigated in a parametric study. The placing of a short suction panel close to the leading edge has been found to be an effective means of controlling instability. Conversely, the level of cooling required to suppress crossflow instability may be too high for practical aircraft applications. The use of suction for preventing laminar separation for pressure distributions with a leading edge suction peak has also been included in the parametric study. The suction quantity required is strongly dependent on the peak height. The suction quantity that can be achieved in practice will limit the maximum peak height that can be attained without laminar separation. An investigation of leading edge stall and control has also been carried out. The analysis suggests that it is important to be able to identify whether the stall is due to laminar bubble bursting or turbulent re-separation, since different methods of controlling the stall may be required.

In this paper an approach to fault-tolerant flight control system design based on the integration of sensor/actuator fault detection and isolation (FDI) and reconfigurable control is presented. The effects of the sensor and actuator faults in the innovation process of the Kalman filter are investigated, and a decision approach to isolate the sensor and actuator faults is proposed.

The reconfigurable control algorithm based on the Extended Kalman Filter (EKF) is presented. The control reconfiguration procedure is executed by considering the identified control distribution matrix. In the simulations, the longitudinal dynamics of an aircraft control system are considered, and control reconfiguration is examined. In the paper, the principal block diagram of a fault tolerant aircraft control system is offered.

Hybrid laminar flow control (HLFC) is an active drag reduction technique that permits extended laminar flow on an aircraft surface and thus offers the potential for significant fuel savings. This is at the expense of an increase in system weight and specific fuel consumption. An overview of HLFC system failure types and consequences is presented as an introduction to this study, which investigated the impact of a potential loss of laminar flow due to flight in cirrus clouds. At typical cruising altitudes, the ice crystals are of a sufficient size and may result in sufficient particle flux to cause a temporary transition of the boundary layer. A computer performance model of a twin engine aircraft in the class of the Boeing 757 has been used to study the impact of alternative fuel planning scenarios on the fuel consumed by a HLFC aircraft, taking into account a model of probable cloud encounter. Based on the models, the study showed that if the fuel planning assumed 25% time-in-cloud (TIC) during the cruise, then in the extreme case of 55% TIC during the cruise, the contingency fuel (taken as 3% of the trip fuel), would be sufficient for the aircraft to complete the mission (including the alternate leg and hold).

The goal of life extending control (LEC) is to enhance structural durability of complex mechanical systems, such as aircraft, spacecraft, and energy conversion devices, without incurring any significant loss of performance. This paper presents a concept of robust life-extending controller design for reusable rocket engines, similar to the Space Shuttle Main Engine (SSME), via damage mitigation in both fuel and oxidiser turbines while achieving the required performance for transient responses of the main combustion chamber pressure and the oxidant/fuel mixture ratio. The design procedure makes use of a combination of linear robust control synthesis and nonlinear optimisation techniques. Results of simulation experiments on the model of a reusable rocket engine are presented to this effect.

High performance fighter aircraft have problems with handling qualities, at high angles of incidence. Pilots are limiting combat effectiveness by cautiously avoiding hazardous regions of the flight envelope. This paper presents and extends existing work on the F-16 aircraft's handling characteristics during the deep stall condition. Inflight experience shows that there are two types of control problems when flying at high angles-of-attack, the ‘pitch departures’ and the ‘deep stall trim’. This paper investigates the critical deep stall condition, assesses the effectiveness of the proposed method of recovery and suggests an augmented and reliable method of returning to normal flight.

A computational study has been performed to evaluate the predictive capabilities of some existing eddy-viscosity (both linear, LEVM, and non-linear, NLEVM) and Reynolds stress transport turbulence models (RSTM) by reference to a transonic shock-induced separated flow over a 10% axisymmetric bump. The calculations have been carried out during the course of a collaborative research programme including both UK universities and industry. The findings of the project demonstrate that improved results can be obtained for such flows by using more advanced turbulence models. For linear eddy-viscosity models, only the SST approach gave good predictions of shock location, recirculation size and pressure recovery, although this was accompanied by deficiencies in the prediction of post-shock velocity profile shape. Non-linear eddy-viscosity models, particularly at the cubic level, provided a more consistent level of agreement with experiments over the range of shock location, wall pressure and velocity profile parameters. Some improvement was also seen in the prediction of turbulence quantities, although only a move to an RSTM closure model reproduced the measured peak stress levels accurately. It was notable that the use of low-Re variants of the models (instead of wall functions) produced no significant improvement in predictions. There are, however, some shortcomings in all models, particularly in the development of flow after reattachment, which was always predicted to be too slow.

A method is proposed by which a direct numerical simulation of the compressible Navier-Stokes equations may be embedded within a more general aeronautical CFD code. The method may be applied to any code which solves the Euler equations or the Favre-averaged Navier-Stokes equations. A formal decomposition of the flowfield is used to derive modified equations for use with direct numerical simulation solvers. Some preliminary applications for model flows with transitional separation bubbles are given.

The present work is concerned with the aerodynamics of the turbulent boundary-layer flow over yawed rectangular cavities with the focus on the steady and unsteady pressures generated by the interaction. Cavities with a planform aspect ratio of 4-85 and streamwise length to depth ratios from 1 to 3 were studied experimentally in a low-speed wind tunnel.

The results indicated three main types of cavity flows. The shear layer bridges the cavity for small angles between mean flow direction and minor cavity axis. The flow field remains almost two-dimensional with little change in drag coefficient. Strong instabilities, associated with a rise in drag coefficient, are found when the cavity is yawed to greater angles. An aerodynamic feedback mechanism depending on interactions between the separated shear layer and the cavity fluid is suggested as the mechanism responsible for the generated oscillations. The influence of the cavity depth is, hereby, found to be fundamental as it determines the degree to which interactions between the separated shear layer and the cavity base occur. As a result both the magnitude and the frequency of the instabilities are a function of the cavity depth. When rotating to higher angles a greater portion of the shear layer reattaches to the cavity base which leads to a loss of flow organisation and a significant increase in drag.

In this paper an algorithm for the trajectory tracking of a three-dimensional trajectory, assigned as a function of time, is presented. The proposed control system is suitable for application on unmanned aerial vehicles (UAVs) or for aircraft that require accurate path tracking, as in the case of rotorcraft in nap-of-the-earth (NOE) flight conditions. The control system logic features (i) an external loop based on a simple guidance scheme and a two-time-scale inverse simulation algorithm, and (ii) an inner loop, based on a linear-quadratic (LQ) full-state-feedback controller. In this way the control action is split into two contributions, i.e. a feedforward command, in order to follow the trajectory generated by the guidance scheme, and a feedback increment, for compensating external disturbances and model uncertainties. A rotorcraft model is used to demonstrate the algorithm capability in a NOE–like flight task. System robustness is analysed and control system performance are discussed in terms of the error between vehicle state and desired trajectory at a given time. Simulation of a representative manoeuvre shows that the feedforward estimate of the control action is accurate and only minor compensation is required from the LQ tracker. The algorithm is suitable for a number of applications, as (i) no simplifying assumptions are postulated for the model, (ii) there are no restrictions on the flight condition, and (iii) the computational time should allow for real–time implementation.

The objective of the work described in this paper was to identify differences in the flow fields immediately downstream of air-jet and vane vortex generators. The flows were assumed to be incompressible and fully turbulent and were solved using the finite volume, Navier-Stokes code CFX 4 (CFDS, AEA Technology, Harwell) on a non-orthogonal, body-fitted grid using the k-ε turbulence model and standard wall functions. The behaviour of the longitudinal vortices produced by the vanes and air jets is presented in terms of stream-wise and cross-stream velocity profiles, circulation and peak vorticity decay, peak vorticity paths in the cross-stream and streamwise direction, cross-stream vorticity profiles, and cross-stream distribution of shear stress. The predicted results show that the vanes and air jets considered produce vortices with significantly different circulation strengths, but that the enhancement of skin friction was of similar magnitude in both cases.

A computational investigation of the subsonic and transonic turbulent open flow over cavities was conducted. Simulations of these oscillatory flows were generated through time-accurate solutions of the Reynolds-averaged form of the Navier-Stokes equations. The effect of turbulence was included through the k–ω model. The results presented include calculations of the acoustic pressure distributions along the cavity floor, which compare well with experiment. The results are then used to describe the behaviour of the flow.

This paper evaluates a time marching simulation method for flutter which is based on a solution of the Euler equations and a linear modal structural model. Jameson’s pseudo time method is used for the time stepping, allowing sequencing errors to be avoided without incurring additional computational cost. Transfinite interpolation of displacements is used for grid regeneration and a constant volume transformation for inter-grid interpolation. The flow pseudo steady state is calculated using an unfactored implicit method which features a Krylov subspace solution of an approximately linearised system. The spatial discretisation is made using Osher’s approximate Riemann solver with MUSCL interpolation. The method is evaluated against available results for the AGARD 445.6 wing. This wing, which is made of laminated mahogany, was tested at NASA Langley in the 1960s and has been the standard test case for simulation methods ever since. The structural model in the current work was built in NASTRAN using homogeneous plate elements. The comparisons show good agreement for the prediction of flutter boundaries. The solution method allows larger time steps to be taken than other methods.

This paper describes power requirements for micro air vehicles, flying in the Reynolds number regime of -lO*. Three flight modes have been researched: fixed wing, rotary wing and flapping wing. For each mode, the literature in the public domain has been reviewed to obtain appropriate lift and drag coefficient data at these low Reynolds numbers. Energy and power requirements for the three flight modes have been calculated and an optimisation procedure has been utilised to evaluate the most efficient flight mode and configuration for a variety of specified missions. The effect of wind-speed on the optimal solution has been examined. It has been discovered that when there is no hover requirement, fixed wing flight is always most energy efficient for the micro air vehicle. However, if there is a hover requirement, the suitability of flapping or rotary wing flight is dependent on the mission profile and ambient windspeed.