Project reviews are carried out to enable the right decisions to be taken to achieve project objectives. However, these decisions are often taken on incomplete, optimistic, inaccurate, misleading or simply wrong information. Whilst incompleteness is inevitable with the complex and rapidly changing nature of high technology engineering projects, it is important that the project manager has a good appreciation of the extent to which these other factors may affect the progress information upon which judgements are made. This paper explores a few of the traps the author has come across, and suggests possible approaches to avoid some of them.

A systems perspective is recommended, with a focus on risk management, resolving ambiguity, rapid response, trend analysis, earned value principles and taking account of the effects of organisational changes.

Recent experiments have demonstrated that humans can fly, with great exertion of the leg muscles using fixed wings and a propeller. However, understanding of aerodynamics and physiology show that using existing technology it is not possible for humans to fly on Earth using their arm muscles to flap wings like birds. Human flapping flight might, however, be possible in a partial gravity environment such as a chamber in Earth orbit or on the surface of the Moon. This paper considers the feasibility of this and provides preliminary estimates of the main parameters involved.

The performance of axial line singularity methods has been investigated systematically for various solution parameters using carefully chosen test cases. The results indicate that increasing the number of elements and using stretched node distribution improves the solution accuracy until the matrix becomes near-singular. The matrix condition number increases with these parameters as well as with the order of intensity variation and profile thickness. For moderate fineness ratios, the linear methods outperform zero-order methods. The linear doublet method performs best with control points at the x-locations of nodes while the source methods perform best with control points mid-way between nodes. The doublet method has a condition number an order of magnitude lower than the source method and generally provides more accurate results and handles a wider range of bodies. With appropriate solution parameters, the method provides excellent accuracy for bodies without slope discontinuity. The smoothing technique proposed recently by Hemsch has been shown to reduce the condition number of the matrix; however it should be used with caution. It is recommended to use it only when the solution is highly oscillatory with a near-singular matrix. A criterion for the optimum value of the smoothing parameter is proposed.

The circular cross-sectional Royal Aerospace Establishment (RAE) 2129 S-shaped intake diffusing duct series, shown in Fig. 1, which has an offset of the inlet and exit centerline of 0·3 and 0·45 of the axial length of the duct, was designed at the RAE (Bedford) and tested at British Aerospace Filton, at low forward speeds (freestream Mach no. range from 0 – 0·2 but a range of duct inlet Mach numbers up to choking speeds) in the last two decades to fulfil the objectives of collecting systematic experimental data on engine face pressure recovery, total pressure flow distortion and wall static pressures for computational fluid dynamics validation. Some of the measurements have been presented, as for example, by Willmer, Brown & Goldsmith. The measurements from one version of Model 2129 (the 0·3 length offset S-duct geometry) have been used to compare with calculated results from a number of CFD programs, as for example, by Anderson, Horton and by the AGARD Working Group 13. The series of comparison also highlight the need for measurements other than the engine face total pressure and wall static pressure which have already been made.

An experiment is described which was undertaken specifically to provide CFD validation data for the case of transonic flow over nozzle afterbodies. The tests were undertaken with the AGARD standard 10° and 15° axisymmetric boat-tail geometries. Onset Mach numbers in the range 0·80-0·99 and subsonic and under-expanded jet plumes were employed in the tests. Test conditions were selected which provided a range of afterbody flow features from largely attached to shock-induced separated flows. A uniquely detailed set of surface pressure and flowfield data are presented. The flow data were acquired with a two-component laser Doppler anemometer (LDA) and define the mean and fluctuating flow components at about 500 spatial locations for each of these complex transonic flowfields. Additional information was recorded which fully defines the required computational boundary conditions.

Also presented is a detailed study of the necessary attributes of windtunnel CFD validation data. It is demonstrated that relatively high blockage experiments using cost-effective windtunnels can be used to generate CFD validation data if proper account is taken of the model/tunnel interference.

It would be difficult to exaggerate the impact of models on our lives — in recreation, education, investigation, inspiration, and intuition. The models can be real or abstract, static or dynamic, simple or complex, and scaled in size or time. Their use can be either explicit or implicit. Appreciation of the value of models has been heightened by the availability of desktop computing power. Experts on any topic cannot be interviewed on TV without a personal computer in the background, to reinforce their image.

The “post-stall” manoeuvre capabilities of current and projected combat aircraft have prompted a resurgence of interest in the dynamic response to wing motion of the vortex burst at high angles of attack. A large number of experimental studies have been reported, but these have tended to be restricted to a limited range of motion profiles and a single specific wing configuration. Presumably due to the complexity of the phenomenon, no attempt appears to have been made to undertake a systematic study of the burst response for a range of wing planforms and motion profiles. Hence, a survey was made of the published data and a semi-empirical correlation derived using a system identification methodology. For preliminary performance studies, the dynamic response of the vortex burst is modelled adequately by a second order lag transfer function. The physical significance of this system model and the implications for the design of a systematic experimental study are discussed.

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.

The dynamic stall characteristics of a NACA 0015 performing ramp-up motions are considered. Pressure data show the development and convection of structures at the leading and trailing edges, which indicate differing degrees of what are described as leading edge and trailing edge stall mechanisms. When a transition strip is placed at the leading edge, the characteristics of leading edge stall predominate, and the growth of the structure at the trailing edge is suppressed. In addition the stall vortex convection speed is observed to alter significantly, which is interpreted as a change in the vortex origin, size, strength and trajectory.

Aircraft yaw control at high angles of attack by tangential forebody blowing has been investigated experimentally. Tests were performed in the University of Bath 21 m x 1.5 m low speed wind tunnel using a 6% scale generic combat aircraft model fitted with blowing slots in the nose cone. Six component strain gauge balance force and moment data were measured for angles of attack up to 90° for various slot geometries and locations. The effect of slot azimuthal location is demonstrated and a slot stall phenomenon described. A geometry dependent forebody/wing flowfield coupling has been identified which can lead to unexpected yawing and rolling moments. The primary source of yawing moment is shown to be the enhanced area of attached flow on the blown side of the forebody rather than direct vortex influence.

The optimum slot extent and location depend on the angle of attack range over which control is required. For regions where steady vortex asymmetry is present, slots near the apex of the forebody produce severe control reversals at low blowing rates. These reversals can be minimised by placing the slots away from the apex. For control in regions where the flow is dominated by periodic vortex shedding, long slots offer efficient control to 90° angle of attack. The most suitable compromise for wide range control would appear to be a short slot placed away from the apex of the forebody.

A technique is described whereby experiments may be conducted on models which are geometrically similar to only part of the full scale configuration with the windtunnel walls being contoured to simulate the flow over that part. In this way, higher spatial resolutions can be achieved in a windtunnel of moderate size. The method is demonstrated by studying the effects of tip geometry on the flow over a wing of aspect ratio 15 using a half wing model of chord 500 mm projecting 1031 mm into a test section of width 1200 mm.

Reliability and safety criteria have, in one form or another, made a significant contribution to engineering design since the earliest times. They gained added significance in Victorian times as the scale of civil and mechanical engineering projects increased. Codes of practice came into use, based both on practical experience and theoretical knowledge. Redundancy and diversity, though not given those names, were introduced as a means of achieving improved reliability and safety. Reliability and safety technology in the modern sense only developed, however, when predictive techniques, based on statistical information, were introduced. This technology has now taken its place in a wide range of engineering activities and is adopting an increasingly international role. It makes use of numerical reliability and safety goals and pays attention to problems associated with human reliability and ergonomics. This paper reviews the particularly important contribution made by civil aviation to the development of present day reliability and safety technology.

This study provides the theoretical foundation for a well-established, but.heretofore empirical, criterion for incipient separation in moderately hypersonic shock/boundary layer interactions. It is based on an examination of the leading high Reynolds number approximation to triple deck theory, combined with its reformulation in terms of the reference temperature concept. The analysis further provides extensions giving the effects of both low supersonic Mach numbers and non-adiabatic wall temperatures on incipient separation.

The three-dimensional vortex/boundary layer interaction of a type which may occur on a high-lift aerofoil has been studied. The experimental configuration simulates the trailing vortex system generated by two differentially-deflected slats which interact with an otherwise two-dimensional boundary layer developed on the wing surface under a nominal zero pressure gradient. The mean and turbulent flowfields are measured by a triple hot-wire system. The measurements show that the trailing vortex system includes the vortex sheets shed from the slats and the single vortex formed at the discontinuity between them. The single vortex moves sideways and interacts with the boundary layer as it develops downstream. During the interaction with the boundary layer, the low momentum, high turbulent-kinetic energy flow carrying negative longitudinal vorticity is entrained from the boundary layer and rolled into the vortex at the line of lateral convergence on the test surface. Likewise, at the line of lateral divergence, the high momentum, low turbulent kinetic energy flow carried by the vortex impinges on the boundary layer, suppressing the turbulent energy level and the growth of the boundary layer.

The evolution of low Earth orbit satellites under the action of atmospheric drag and J2 oblateness perturbations has been widely investigated. Simple analyses describing the altitude decay of a circular orbit through energy arguments have also been widely available. This paper extends these simple analyses by providing compact analytic expressions for true anomaly and ascending node angle along with a more accurate expression for altitude decay. The resulting solutions then allow a fast and efficient method of investigating the orbital evolution of decaying, precessing circular orbits. Applications include the rapid evaluation of station-keeping budgets for both single platforms and the orbital evolution of satellite constellations. Passive station-keeping by continuous low thrust acceleration is investigated and shown to be naturally unstable.

In response to the need for ground testing facilities for super orbital re-entry research, a small scale facility has been set up at the University of Queensland to demonstrate the Superorbital Expansion Tube concept. This unique device is a free piston driven, triple diaphragm, impulse shock facility which uses the enthalpy multiplication mechanism of the unsteady expansion process and the addition of a secondary shock driver to further heat the driver gas. The pilot facility has been operated to produce quasi-steady test flows in air with shock velocities in excess of 13 km/s and with a usable test flow duration of the order of 15 μs. An experimental condition produced in the facility with total enthalpy of 108 MJ/kg and a total pressure of 335 MPa is reported. A simple analytical flow model which accounts for non-ideal rupture of the light tertiary diaphragm and the resulting entropy increase in the test gas is discussed. It is shown that equilibrium calculations more accurately model the unsteady expansion process than calculations assuming frozen chemistry. This is because the high enthalpy flows produced in the facility can only be achieved if the chemical energy stored in the test flow during shock heating of the test gas is partially returned to the flow during the process of unsteady expansion. Measurements of heat transfer rates to a flat plate demonstrate the usability of the test flow for aerothermodynamic testing and comparison of these rates with empirical calculations confirms the usable accuracy of the flow model.