A semi-inverse viscous/inviscid coupling technique for the calculation of compressible flows is presented. Euler equations in integral form are solved for the inviscid part and integral compressible boundary-layer equations in direct and inverse form are solved for the viscous one. Interaction is realised using the equivalent sources approach.

The technique has been applied to the calculation of the aerodynamic characteristics of aerofoils both in cruise and take-off or landing configurations. To treat the complex geometries involved in the latter case, a multi-block approach, based on structured ‘H’ grid topology, has been developed and validated for different configurations.

The emergence of electronic computers has brought about a revolution in the analysis of aircraft structures in the past decade which is now generally accepted throughout the Industry. But the major impact of computers on design has, in my opinion, not yet been realised and I hope, in this short paper, to show how powerful a tool they can become in the hands of imaginative design engineers. The direct benefits to the designer lie in the ability of the computers to perform large amounts of routine arithmetic in a short space of time. This capability enables us to obtain an insight into structural behaviour and to consider the influence of the aircraft environment and design alternatives from the earliest stages of design. By the development of general purpose computer programmes for handling routine calculations the structural designer can be freed from much of the laborious calculation associated with complex structures and he is able to devote more time to creative work. An even more important indirect benefit is the basic re-thinking of design processes which is necessary in order to reduce them to routine arithmetic. We begin to realise that many aspects of efficient design such as continuity, cost and weight can be introduced as quantitative parameters instead of relying on intuitive compromise.

The analysis of risks per phases of flight is fundamental for safe helicopter operations, of which night-time offshore oil- and gas-related missions form an important part. The safe execution of such missions also depends on pilots’ recent flying practice and a stable visual approach segment prior to landing. However, the poor quality of the safety data currently available prevents accurate analysis of risk on a per-phase-of-flight basis, establishment of a meaningful flying recency requirement and identification of any preferable visual approach design. To redress these problems, this paper develops a bespoke taxonomy of phases of offshore helicopter flights and uses it as the basis for a questionnaire survey on the phase-specific risk levels experienced by pilots in the night-time, perceived optimal flying recency requirement and preferred visual approach design. With the responses obtained from pilots located in seven countries, extensive statistical hypothesis testing shows that the phases involving visual scan techniques at high speed regimes are problematic, especially the visual segment of instrument approaches. Moreover, the between-night-flights time gaps required for assured flying recency were found considerably shorter than currently standardised across the industry. Finally, no preferred visual approach technique was identified. A number of important implications have been highlighted and should form the basis for future safety interventions.

Features of aircraft which influence flow separations, and hence the onset of buffeting, should be always of interest to an aerodynamicist. The present compilation of such features is based on notes made during a visit to the USAF Aircraft Museum in Dayton, Ohio. Although two low-speed aircraft with high aspect ratio wings are considered, the emphasis is primarily on transonic and supersonic military aircraft, with wings of lower aspect ratio.

Some recommendations are made to stimulate research into improved methods to control flow separation, particularly for future transonic and supersonic aircraft.

A recent paper(1) gives details of an investigation on the low endurance fatigue behaviour of an aluminium-zinc-magnesium alloy (D.T.D. 683). It becomes apparent in subsequent discussion that information on another aluminium alloy to B.S. L65C would satisfy a wider practical interest than the former. Although the equipment used for the earlier work was engaged on another project, it was decided that the latter should be halted for a period to enable a brief programme on low endurance fatigue to be conducted on B.S. L65C material which had previously been kindly provided by the Aluminium Development Association.

This paper describes the application of low-order models to the prediction of the steady performance of axial compressors at sub-idle conditions. An Euler body-force method employing sub-idle performance correlations is developed and presented alongside a mean-line approach employing the same set of correlations. The low-order tools are used to generate the characteristic lines of the compressor in the locked-rotor and zero-torque windmilling conditions. The results are compared against steady-state operating points from three-dimensional (3D) Reynolds-averaged Navier–Stokes (RANS) computational fluid dynamics (CFD) simulations. The accuracy of the low-order tools in reproducing the results from high-fidelity CFD is analysed, and the trade-off with the computational cost of each method is discussed. The low-order tools presented are shown to offer a fast alternative to traditional CFD which can be used to predict the performance in sub-idle conditions of a new compressor design during early development stages.

Numerical simulations were performed to analyse the possibility of damping abrupt incoming free stream disturbances upon a porous aerofoil using an unsteady Reynolds-averaged Navier-Stokes (URANS) model. To mimic the turbulence disturbance levels that are typically encountered in the atmosphere, two flow configurations were considered. In the first configuration, the unsteadiness of the flow was created with vortices shed from a circular cylinder installed ahead of a WTEA-TE1 aerofoil. The continuous von Kármán shedding vortices contained within the cylinder wake were convected downstream and projected upon the aerofoil. In the second configuration, an instantaneous pair of discrete vortices was created by a rotational snapping of a flat plate, installed upstream of the aerofoil. Solid and porous aerofoil configurations, with porosity settings of 11 and 22%, were applied on 50% of the chord of the aerofoil starting from the leading edge. Both steady and unsteady flow simulations were performed to assess the performance of the porosity under steady and unsteady effects. The steady state flow simulations revealed a noticeable reduction in the aerofoil lift coefficient for the porous aerofoil. For unsteady solutions with a continuous or distinct series of vortices interacting with the aerofoil, the porosity showed insignificant damping of the lift coefficient amplitude. The porosity values investigated in the current exercise had indiscernible effect upon the unsteady lift-load alleviations caused by free stream disturbances.

In an analysis by the matrix force method of a structure containing thin tubes of open cross-section, it is commonly assumed that the torsional stiffness of such members may be taken into account by considering the St. Venant shear stresses only. This effectively introduces the assumption that the cross-section of an open tube is able to warp freely, whereas in practice the warping may be constrained by the surrounding structure.

In this note a method is presented which allows such warping constraint effects to be taken into account in a finite element analysis using the Force Method. Because of the assumption of linear variation of direct stress along the element the full flexibility matrix does not have to be used, which may be an advantage in certain cases. The “exact”, but more complex, flexibility matrix has been derived and will be presented in a later note.

Pressure recovery measurements in a straight pipe with crossflow-generated entry velocity distributions suggested substantial effects of inflow distortion on the pressure recovery downstream of a sudden enlargement in area. This note reports results of actual measurements in the same pipe equipped with entry sections of various reduced diameters providing sudden area enlargement ratios of 1.33, 1.71, 2.04, and 3.52. The results have practical bearing on the performance of settling pipes located at diffuser exits where, in general, velocity distributions are significantly non-uniform. In addition, the results provide information on the momentum content of the crossflow-generated velocity distribution.

In this paper, a non-linear tracking control algorithm is extended. The control objective of this research is to track a desired time-varying attitude of a satellite in the presence of inertia uncertainties and external disturbances, in order to be more suitable for implementation in a real-world application. In this investigation, the actuators are reaction wheels and the actuator dynamics are modelled in addition to the spacecraft dynamics. Thus, the control signal is DC motor voltage which is the most fundamental control variable and can be generated easily by a motor driver in practical cases. To achieve robust tracking of the desired time-varying attitude, a sliding mode controller is designed, and adaptive techniques are developed based on sliding mode control to overcome the inertia uncertainties and to estimate and compensate external disturbances. The kinematic equations of the satellite are expressed using quaternion parameters, and a novel control law will be derived by using a new facilitating approach in controller design, which is based on quaternion algebra, because of quaternion advantages, such as singularity rejection. Using this approach it will be more comfortable to deal with tedious mathematical operations, and on contrary with most of the previous studies, the terms corresponding to derivatives of the desired attitude are not neglected, and tracking capability is retained. The global stability of both methods (Sliding Mode Control (SMC) and adaptive sliding) is investigated using Lyapunov’s stability theorem. In order to validate the control methods, first, Simulink-ADAMS co-simulation of a 3-DOF attitude control is used to verify the algorithm performance and integrity, and finally, the control strategy is implemented on the Amirkabir University of Technology (AUT) 3-DOF attitude simulator for different types of non-linear attitudes. Both co-simulation and implementation results clearly illustrate the designed attitude control algorithm’s excellent performance in the various manoeuvres.

A number of panel methods are now currently in use for solving potential flow about single and multi-element aerofoils. All these methods use a combination of singularities, both source and vorticity, on or inside the aerofoil surface and solve the singularity strengths by satisfying the boundary condition of zero normal flow on the surface of the aerofoil.

In this note the method developed by Basu for solving single aerofoil problem using the concept of internal singularity distribution has been applied to the case of multi-element aerofoil problem.

For the purpose of the design and certification of inflight icing protection systems for transport and general aviation aircraft, the eventual re-definition/expansion of the icing environment of FAR 25/JAR 25, Appendix C is under consideration. Such a re-definition will be aided by gathering as much inflight icing event data as reasonably possible, from widely-different geographic locations. The results of a 12-month pilot programme of icing event data gathering are presented. Using non-instrumented turboprop aircraft flying upon mid-altitude routine air transport operations, the programme has gathered observational data from across the British Isles and central France. By observing a number of metrics, notably windscreen lower-corner ice impingement limits, against an opposing corner vortex-flow, supported by wing leading edge impingement limits, the observed icing events have been classified as ‘small’, ‘medium’ or ‘large’ droplet. Using the guidance of droplet trajectory modelling, MVD values for the three droplet size bins have been conjectured to be 15, 40 and 80mm. Hence, the ‘large’ droplet category would be in exceedance of FAR/JAR 25, Appendix C.

Data sets of 117 winter-season and 55 summer-season icing events have been statistically analysed. As defined above, the data sets include 11 winter and five summer large droplet icing encounters. Icing events included ‘sandpaper’ icing from short-duration ‘large’ droplets, and a singular ridge formation icing event in ‘large’ droplet. The frequency of ‘large’ droplet icing events amounted to 1 in 20 flight hours in winter and 1 in 35 flight hours in summer. These figures reflect ‘large’ droplet icing encounter probabilities perhaps substantially greater than previously considered. The ‘large’ droplet events were quite localised, mean scale-size being about 6nm.

This paper is concerned mainly with that application of counter-rotating propellers referred to as “ dual-rotation ” propellers. This term is here defined to apply to a pair of oppositely-rotating propellers operating in close tandem, mounted on concentric shafts. Only the application to aeroplanes is discussed herein, for the application of these principles to helicopter design is considered another subject.

This paper presents a summary of the development and background of dual-rotation propellers and the benefits that can be derived from their use. Complete references are furnished for designers, no attempt being made to reproduce the somewhat involved theory and lengthy reference data.

A short history of the three applications of counter-rotating propellers is presented. The applications include multi-motored aeroplanes with oppositely-rotating propellers, aeroplanes with engines mounted in tandem (back to back) so that the propellers revolve in opposite directions, but primarily, aeroplanes with dual-rotation propeller installations.

Most present-day aeroplanes with maximum velocities of 350 m.p.h. and over have single-rotation propellers. At this velocity the single propeller is operating past the peak of its efficiency. The loss in efficiency increases rapidly with increasing forward velocity. Dual-rotation propellers offer an excellent means of maintaining peak propeller efficiencies up to aeroplane velocities of about 450 m.p.h. by absorbing the rotational energy losses in the propeller slipstream. At velocities above 450 m.p.h. up to the limiting velocity there will be only a small reduction from the peak in dual-rotation propeller efficiencies.

Sustained level flight aeroplane velocities appear to be limited to about 550 m.p.h. when propellers are used as the means of propulsion. The efficiency of a conventional single-rotation propeller will be so low, however, that this limiting velocity can be approached with such a propeller only at a prohibitive cost in engine horse-power. Any given velocity above 300 m.p.h. can be obtained with less horse-power if dual-rotation propellers are used to replace a single-rotation propeller with the same total number of blades. The percentage gain in performance will increase with increasing forward velocity when dual-rotation propellers are used.

It is possible that single-rotation propellers with more than four blades will never be commonly used. Instead, dual-rotation propellers with a total of four, six or perhaps even eight blades, will be used in the next decade on many high performance aeroplanes, especially those operating at high altitudes. The types of aeroplanes will include small single-seater racers and fighters and high performance, long range bombers and transports, particularly those with supercharged cabins.

The discussion brings forth the fact that there are few experimental data available in this country to allow a quantitative engineering evaluation to be made of the worth of dual-rotation propellers. Sufficient substantiated theory exists, however, to permit the general statements made above. An outline of the experimental tests that are required to check the theory is presented.

Following the discussion of tests made by the U.S. Air Corps on a pursuit type aeroplane equipped with a dual-rotation propeller system, comments are made on the relative advantages and disadvantages of dual and single-rotation propellers as they affect aeroplane design. Hub design problems associated with dual-rotation propellers are discussed.

Because of the increasing importance of dual-rotation propellers, additional pertinent references other than those used in this paper are presented in the Appendix.

Seven composite laminates, three carbon, one Kevlar and three hybrid as supplied by an aircraft manufacturer were assessed for damage tolerance. Rectangular coupons were subjected to low velocity impact in the range 0-4 Joules by a swinging pendulum. The coupons were ultrasonically C-scanned to measure delamination area and visually inspected for damage. Residual strength tests in tension, compression, flexure and shear were then carried out on the damaged coupons. Visual damage, delamination damage and strength reduction thresholds were determined from the data and compared. Of the three thresholds, delamination damage and strength reduction tended to coincide. For the various lay-ups these were in the range of 1·5 to 2·0 J for the longer (tensile) specimens or in the case of the smaller (shear) specimens between 0·2-0·3 J. In all cases onset of visible damage occurred at slightly higher values of impact energy. The significance of specimen stiffness was noted.

Boundary-layer noise produced by rough surfaces is a potential contributor to airframe noise. In this paper, an attempt is made to assess the surface roughness noise for a conceptual Silent Aircraft design SAX-40 using a prediction model described in previous theoretical work (Liu and Dowling(12)). Estimates of three idealised test cases show that surface roughness could produce a relatively high noise level and enhance the trailing-edge noise somewhat. Roughness height and roughness density are two significant parameters which affect surface roughness noise, with roughness height having the more dominant effect. The distribution of roughness noise sources decreases in the streamwise direction on the SAX-40 surface. Two candidate rough surfaces are selected for SAX-40 to keep surface roughness noise at a negligible level and to meet an aggressive noise target.

Some observers believe that there will be such growth in the travelling public and such constraint on airport and airspace usage that, by early in the 2000s. very large passenger carrying aircraft will be required. By ‘very large’ is meant anything from 1,000 to 2,000 seaters. Whilst the problems which come to mind when musing on this possibility are mostly those of ‘passenger flow’ in the widest sense and rather less related to the vehicle itself, there are some issues related to the vehicle which suggest themselves. One is the choice of fuselage cross section.

Because cabins have to be pressurised (10 psi is a typical figure for cruise altitudes up to say 40,000 ft), it has been the norm to utilise circular cross sections (or double-bubble type) since the pressure forces are contained by hoop loads only, without frame bending, which affords a light and fatigue-insensitive structure (Fig. 1).

It is well-known that, at large distances from the flight path, the waveform of the sonic bang from an aircraft, without lift at any rate, has the characteristic N-wave shape, as illustrated in Fig. 1. When we depict a waveform such as this we are stating what would occur at large distances in the absence of viscosity and other diffusive effects, such as turbulence. We know, of course, that in reality, because of these effects, the bow and stern shocks will be “softened”, i.e. the pressure jumps associated with them will not be instantaneous, but will have a finite, but small, “rise time”. However, this is not the matter under discussion. The essential feature of the N-wave illustrated in Fig. 1 is that the pressure jumps associated with the bow and stern shocks are equal, and that the positive and negative phases are of equal duration. The arguments for the waveform to take up this shape at large distances are put in a classical paper by Whitham.

In common with other pneumatic and electronic cabin pressure control installations the two new pneumatic systems discussed in this paper have been designed to control cabin pressure automatically by throttling the out-flow of the cabin conditioning air. For those unfamiliar with the subject Fig. 1 shows diagrammatically a typical modern military aircraft installation. Both the systems discussed should however be of particular technical interest warranting general description, as for the first time with this type of equipment low cost and robust integrated fluidic circuits have been extensively used in their design. Patents covering these have now been granted in the United Kingdom, the United States and other countries.

A method based on the shedding and convecting of discrete vortices has been developed to calculate the transient loading on an aerofoil-spoiler configuration when the spoiler is suddenly deployed to a particular angle. An innovation in the method is the introduction of Kutta conditions at the separation points.

Although the variation of CL with time is highly irregular, qualitative trends can be deduced. For a range of aerofoil angle of incidence and spoiler angle deployment adverse lift effects (ie, when the lift initially increases) have been obtained, similar adverse lift effects have been measured in experiments.

It is my privilege to present to the Royal Aeronautical Society a brief summary of the work done and the results achieved in developing the autogiro in the last few years. Many fundamental improvements have been the product of that work, but of still greater importance is the considerable progress in the knowledge of the aerodynamics, dynamics and engineering requirements of the new aircraft. The improvements in the formula have been made possible or practical only because of the painfully acquired technique, which, as is the case in all new arts, is a compound of theoretical developments and empiricism.