This paper describes a simple method for smooth curve fitting or function interpolation. Where possible, the method uses continuous piece-wise conies in preference to cubics so as to prevent the occurrence of extraneous inflection points. The method is ideal for single-valued functions with sharply changing gradients such as occur in pressure distributions, containing shock waves. A version that is valid for multivalued functions is also described.

In the past ten years or so a number of changes have taken place in aircraft which have progressively reduced the potential tyre life and tend therefore to increase the operational cost. We now see that the latest aircraft have developed to such a stage that there is a certain hazard in stopping them on wet or slush covered runways. To ease this situation, further changes are recommended to the tyres, likely still further to reduce tyre life.

BAE Systems and the Engineering and Physical Sciences Research Council (EPSRC) have recently formed a partnership to invest in strategic research in aerospace and defence. The framework which has been developed as part of this new alliance places a strong emphasis on collaboration. This contrasts with the conventional approach to industry led research which is normally based on the establishment of ‘centres of excellence’ in specific subject areas. By using a collaborative approach, the funding partners aim to benefit from inter-disciplinary collaboration which will take place during the project, giving rise to a more effective use of the invested funds.

The objective of the research programme is to develop technologies which would support the design of low cost (both to acquire and operate) flapless unmanned aerial vehicles (UAVs). This work includes fundamental aerodynamic research to provide control forces without the use of conventional flaps, coupled with developments in the areas of control systems, manufacturing engineering, structural engineering, the electromagnetic behaviour of these structures and design optimisation. The output from the research has been extended to include the design, manufacture, assembly and flight of a demonstrator vehicle, which will provide the research teams with a platform on which to evaluate the performance of their technology in a realistic flight environment. A total of fourteen research groups at ten universities are involved in the five year programme, which has a total value of £6·5M (€9·75M).

The problem of a delta wing in a high speed stream has been approached using linearised potential flow theory for low supersonic Mach number, and using thin shock layer theory for hypersonic Mach number. Recently the author has given a new theory of supersonic and hypersonic flows with attached shock wave over the compression surface of an unyawed delta wing. It has the advantage of being unified for both supersonic and hypersonic flows, and it gives almost identical results compared with exact numerical solutions. The purpose of this note is to extend the theory of Ref. 4 to include the effects of yaw.

The piloted aircraft provides the supreme example of the highly complex machine performing a precise function in which the control loop is still completed through a human operator. Although both physically and mentally his task has been greatly assisted by the introduction of automatic pilots, auto-stabilisers, power controls, flight directors and the like, he still possesses the ultimate advantage over the machine of being able to use judgment in the face of arbitrarily changing circumstances. Thus, until it is possible in advance to define the required tasks exactly, and to measure the subsequent performance of the system in precise and unambiguous terms, the pilot will be saved from the encroachment of automation.

The performance of the human operator is described by the aptness and speed of his reaction to the pertinent stimuli, in effect his transfer function as an integral element of the overall control loop.

In a recent publication the problem of denning the thrust of a jet engine, or the internal drag of a ducted body, is considered in some detail. This Report was compiled by a special Panel of the Aeronautical Research Council and, during its discussions, concern was expressed at the position regarding the definitions of certain quantities having the dimensions of pressure; in some cases the most widely-used definitions were not rigorous, in others a variety of names were in use. In addition, some quantities of considerable value in fluid dynamics were either unnamed or, had been given misleading names. This state of affairs was felt to be undesirable. As a result, the A.R.C. set up a small Panel to consider the matter more fully. The Panel contained representatives from both of the two important groups interested in obtaining uniformity in this field, those concerned with external aerodynamics and those primarily interested in internal flow problems.

The procedure described applies to aerofoils or cascades with circulation controlled by tangential blowing jets and provides a complete numerical solution for aerodynamic performance in incompressible flow conditions.

The pressure distribution over the blade surface is predicted by a potential flow model in which the region of separated flow is represented by an appropriated source distribution. Boundary layer development is calculated by a finite-difference solution of the parabolic boundary layer momentum equation: the development of the blowing jet and its mixing with the boundary layer over the curved trailing edge is predicted by the same procedure applied to an angular momentum equation, using an intermittency representation of the eddy viscosity distribution.

Predictions are compared with experimentally measured lift coefficients for an isolated aerofoil, with turning angles for a cascade tested by other workers and with experimental turning angles for a cascade tested by the authors. The agreement between theory and experiment is good.

The laminar flow over a slender delta wing at incidence has been extensively studied both experimentally and theoretically using vortex sheet methods. These vortex sheet methods have generally been successful apart from the prediction of the secondary boundary-layer separation induced by the primary vortex. This paper revisits the problem using computational fluid dynamics (CFD) and focusses on the effects of the secondary flow separation. The modelling approach is briefly summarised, and the results are compared with flow measurements and results from vortex sheet methods. The computations show very good agreement with measurements for the surface pressures and total head contours. The results help to understand the complex structure of the leading edge vortex flow, and the associated secondary separation of the boundary layer. They indicate that inviscid mechanisms dominate the larger scale features, and highlight a possible mechanism for the development of an instability in the leading edge vortex sheet.

During the past few years, it has become recognised that one of the major objections to the introduction of helicopters fitted with pressure jets at the blade tips is the high level of noise set up by the jets The source of this noise lies m the jet stream itself When air travelling at high speed mixes with the surrounding air, turbulence is formed in the region of high velocity shear and noise is emitted into the stationary or near stationary medium surrounding the jet The large velocity gradient or shear actually amplifies the sound (Ref 1) which increases as a very high power of the jet velocity, normally about the eighth power Needless to say, if turbulence exists initially in the jet stream, either from poor internal shaping of the engine or from incomplete burning of the fuel, the noise will be aggravated and a different noise law achieved With well designed jet umts, however, the eighth power law is obeyed surprisingly well

It has long been recognised that landing and taking—off are the most hazardous periods of the average flight. In this respect an aeroplane is similar to a boat being launched or beached at high speed; the change from one medium of support to another is bound to involve possible damage especially if speeds are high.

So far as risk of mechanical damage, due to uneven aerodrome surface (or swell in the case of a seaplane), is concerned, the speed of landing should be the minimum practicable. Very low landing speeds have a possible disadvantage in that one may not have sufficient aerodynamic control to compete with gusts, etc., but there are few aeroplanes of to-day which have landing speeds so low that this objection can be raised.

The trend of design in recent years has been to increase aerodynamic efficiency by streamlining and to increase the wing loading.

The physics behind the transitions of natural Samaras, or the bio-inspired counterparts, to steady autorotation has been unclear. Theoretical and experimental investigations explore the inertial and aerodynamic characteristics required to guarantee stable transitions of an artificial Samara-like decelerator from chaotic tumbling motions to azimuthal autorotation. A non-dimensional inertial criterion is proposed, which is in accord with experiments.

In recent years, there has been significant momentum in applying deep learning (DL) to machine health monitoring (MHM). It has been widely claimed that DL methodologies are superior to more traditional techniques in this area. This paper aims to investigate this claim by analysing a real-world dataset of helicopter sensor faults provided by Airbus. Specifically, we will address the problem of machine sensor health unsupervised classification. In a 2019 worldwide competition hosted by Airbus, Fujitsu Systems Europe (FSE) won first prize by achieving an F1-score of 93% using a DL model based on generative adversarial networks (GAN). In another comprehensive study, various modified and existing image encoding methods were compared for the convolutional auto-encoder (CAE) model. The best classification result was achieved using the scalogram as the image encoding method, with an F1-score of 91%. In this paper, we use these two studies as benchmarks to compare with basic statistical analysis methods and the one-class supporting vector machine (SVM). Our comparative study demonstrates that while DL-based techniques have great potential, they are not always superior to traditional methods. We therefore recommend that all future published studies of applying DL methods to MHM include appropriately selected traditional reference methods, wherever possible.

A contra-rotating fan offers several aerodynamic advantages that make it a potential candidate for future aircraft engine configurations. Stall in a contra-rotating axial fan is interesting since instabilities could arise from either or both of the rotors. In this experimental study, a contra-rotating axial fan is analysed under clean or distorted inflow conditions to understand its performance and stall inception characteristics. The steady and unsteady measurements identified the relative contribution of each rotor towards the performance of the stage. The tip of rotor-1 is identified to be the most critical region of the contra-rotating fan. The contribution of rotor-2 to the overall loading of the stage is observed to be relatively less than rotor-1. The penalty due to distortion in the stage pressure rise is mostly felt by rotor-1, while rotor-2 also shows a reduction in performance for distorted inflows. Rotor-2 stalls at a high flow coefficient marking the initiation of partial stall of the stage, and the stall of the whole stage occurs once rotor-1 stalls. A fluid phenomenon that is attached to the blade surface marks the stall of rotor-1, and this fluid phenomenon initially rotates at a speed close to the speed of rotation of the blade. As the stage moves towards the fully developed stall, this fluid phenomenon sheds from the blade surface. The fluid phenomenon thus propagates at a speed much lower than the rotational speed of the blade during fully developed stall.