This paper presents a new approach to force fighting equalisation in a redundant active-active-active rudder actuation system that is used for the primary flight control system of a turboprop regional aircraft. The related coupled problem of force fighting scenario, and the hydraulic architecture of electronic-hydrostatic actuator (EHA) are analysed, the mathematical model of the EHA system is built. The virtual test bench is designed to evaluate the performance of the force fighting equalisation strategy. The proposed methodology is tested on an iron bird test rig. The physical experiment shows that the fighting force is minimised under all flight conditions, meets the low cost requirement and can be a very reliable system. The proposed methodology can be applied to other types of aircraft’ flight actuation systems.

The aviation industry plays a critical role in modern society, supporting global, regional and local economies and facilitating long distance travel for business and leisure. However the environmental impact of the industry is significant and this threatens the sustainability of future growth and the consequential benefits which could arise from it. Technological, operational and infrastructure improvements designed to reduce emissions and noise from aircraft are continually being eroded by the rapid growth of the industry, with the result that the environmental impact of air travel is increasing. Consequently, the sustainable development of the aviation industry will require that in future, technological innovation and fleet replacement continues at a rapid pace and that the highest priority is given to achieving a ‘step change’ in aircraft technology.

The quadratic constitutive relation was proposed as an extension of minimal complexity to linear eddy-viscosity models in order to improve mean flow predictions by better estimating turbulent stress distributions. However, the successes of this modification have been relatively modest and are limited to improved calculations of flow along streamwise corners, which are influenced by weak secondary vortices. This paper revisits the quadratic constitutive relation in an attempt to explain its capabilities and deficiencies. The success in streamwise corner flows cannot be entirely explained by significant improvements in turbulent stress estimates in general, but is instead due to better prediction of the particular turbulent stress combinations which appear in the mean streamwise vorticity equation. As a consequence of this investigation, a new formulation of turbulent stress modification is proposed, which appears to better predict the turbulent stress distributions for a variety of flows: channel flow, equilibrium boundary layers, pipe flow, separated boundary layers and square duct flow.

Even the shortest flight through unknown, cluttered environments requires reliable local path planning algorithms to avoid unforeseen obstacles. The algorithm must evaluate alternative flight paths and identify the best path if an obstacle blocks its way. Commonly, weighted sums are used here. This work shows that weighted Chebyshev distances and factorial achievement scalarising functions are suitable alternatives to weighted sums if combined with the 3DVFH* local path planning algorithm. Both methods considerably reduce the failure probability of simulated flights in various environments. The standard 3DVFH* uses a weighted sum and has a failure probability of 50% in the test environments. A factorial achievement scalarising function, which minimises the worst combination of two out of four objective functions, reaches a failure probability of 26%; A weighted Chebyshev distance, which optimises the worst objective, has a failure probability of 30%. These results show promise for further enhancements and to support broader applicability.

Artillery rockets are sensitive to disturbances (total impulse variation, wind, thrust misalignment, etc.). As the range of ground-to-ground rockets increases, the accuracy of free flight rockets decreases. Requirements for the increase of the range and minimisation of the impact point dispersions can be solved by adding guidance and control systems to rockets. Based on the differences between the measured flight parameters and the calculated parameters for the nominal trajectory, the flight path angle correction algorithm is obtained by adding the correction to the nominal value. The flight path steering guidance system with lateral acceleration autopilot in the inner loop, is used for the guidance of hypothetical artillery rockets. The guidance algorithm given in this paper eliminates the time correction due to rocket velocity variations. Efficiency of the proposed algorithm is illustrated by numerical simulation. There is no practical influence of the external disturbances on impact point dispersion. It is also shown that the measurement errors of the inertial measurement unit are the dominant factor affecting impact point dispersion of artillery rockets modified by adding a guidance system based on the flight path steering method.

The supersonic flow past a combination of a thin wing and a slender body of revolution is discussed by means of the linearised equation of motion. The exact equation is first established so that the linearised solution can be fed back and the order of the error terms calculated. The theory holds under quite general conditions which should be realised in practice.

The wing-body combination considered consists of a wing symmetrically situated on a pointed body of revolution and satisfying the following fairly general conditions. The wing leading edge is supersonic at the root, and the body is approximately cylindrical downstream of the leading edge. The body radius is of an order larger than the wing thickness, but is small compared with the chord or span of the wing.

It is found that if the wing and body are at the same incidence, and the aspect ratio of the wing is greater than 2 (M2-1)-½, where M is the main stream Mach number, the lift is equivalent to that of the complete wing when isolated. If the wing only is at incidence then the lift is equivalent to that of the part of the wing lying outside the body.

The presence of the body has a more significant effect on the drag. If, for example, the body is an infinite cylinder of radius a, and the wing is rectangular with aspect ratio greater than 2(M2-1)-½, then the drag of the wing is decreased by a factor (1-2a/b), where 2b is the span of the wing.

When these conditions do not hold the results are not quite so simple but are by no means complicated.

A numerical method for the solution of three-dimensional supersonic flows and the associated computer programme have been devised. The present algorithm is based on a reference-plane characteristic network. The physical region between the shock wave and the body surface is replaced by a computational mesh facilitating the incorporation of different boundary conditions and arbitrary body shapes. For reasons of stability and computing time a Hartree-type grid was used where the characteristic segments are projected upstream. The base points do not coincide with nodal points and interpolation by means of Tchebycheff polynomials is required to obtain the flow properties. Numerical differentiation is carried out by means of spline functions. The programme has been used and verified under a wide variety of flows past circular and elliptic cones, tangent ogives and a delta wing body. After a series of numerical experiments it was found that accurate solutions could be found using a 7 × 7 grid, and that the essence of the solution is preserved using a very coarse mesh of 4 × 4. Furthermore, the good agreement obtained with experimental results indicates that it could be applied successfully to problems of practical interest.

The design of ducted fans, and the estimation of their performance for conditions other than those of the design, are usually based on the equations of the so-called strip-theory. Existing methods of solution invariably employ some form of iteration which is more or less rapidly convergent according to the experience of the computer and the detail of the method.

In this note, a solution is given which, although it is strictly an iteration, is explicit in the sense that in practice no iteration is necessary. It is shown how both design for a given performance and performance for a given fan can be calculated in one direct step.

Conditions for maximum efficiency are considered in an Appendix. Here a suggestion of R. Baldur's is used to give a simple approximate solution of the strip-theory equations which can be developed to give conditions for maximum efficiency, for the special case of a constant pressure rise through the fan.

This paper discusses some of the special problems of wind-tunnel testing which arise with high-lift jet-blowing models, supplementing earlier comments. Further detailed remarks are made about recent developments on tunnel-wall interference, test rigs and methods of minimising constraints from air-feeds to models, and on general test and model design techniques.

The President of the Royal Aeronautical Society (Sir Arnold Hall) invited Professor J A J Bennett (Chairman, Lecture Committee, of the Helicopter Association of Great Britain) to introduce the Lecturer

Professor J A J Bennett said that this was a very important occasion, both for the Royal Aeronautical Society and for the Helicopter Association of Great Britain Tonight, they were to hear about the development of an aircraft which was certainly not an aeroplane, and it was certainly not a helicopter according to present-day definitions To accept the F A I nomenclature, it was in an intermediate class, the so-called “Convertiplane” category As, however, there was neither a Convertiplane Association of Great Britain nor yet a Convertiplane Section of the Royal Aeronautical Society, it was appropriate that the two societies should get together and discuss the project

An approximate method is presented for the calculation of heat transfer rates to cooled turbine blades. The method is based on a combination and extension of methods which have been developed in recent years for the calculation of the skin friction and heat transfer coefficients on wings in high speed flight. The use of the method is demonstrated by application to a specific cascade for which an experimental determination of overall heat transfer coefficient is known. Very close agreement with the experimental results is found over the range of Reynolds number tested. The calculated distribution of local heat transfer coefficient indicates that local pressure gradients have a marked effect on the heat transfer. A first-order estimate of the effect of blade cooling on the rate of mass flow through a blade passage shows that an increase of the order of one per cent in the mass flow rate may be obtained by a reasonable degree of blade cooling.

The theory developed in Part I is applied to a number of problems of aeronautical interest, the most important of which is to the setting of “ streamline “ walls about a symmetrical aerofoil placed in the centre of a channel. It is shown how the position of the streamline wall can be deduced from the (experimentally determined) position of a constant pressure wall. This theory is applicable to symmetrical aerofoils of any given shape, and makes allowance for the presence of the aerofoil's wake. To illustrate the theory, and to test it by an extreme example, the flow is calculated about a circular cylinder, with a diameter about half the tunnel height, for both straight and constant pressure walls. The solid blockage is calculated in each case and compared with the standard first order theory. For this extreme example the standard theory fails badly for straight walls, but is reasonably accurate for constant pressure walls.

An alternative analytical method is applied to interference problems of an aerofoil very close to the ground, such as a free surface, a solid wall, and a slipstream. The method of matched asymptotic expansions on the small ground clearance is used in the governing integral equation and this analysis is different from Widnall and Barrows’s analysis. They used the matched asymptotic method in the differential equation. To show that the present method is reasonable, the problem of a ground-effect aerofoil in uniform fluid flow is considered and it is found possible to obtain the same results as those of Widnall and Barrows for the case of small ground clearance. By comparing with some earlier studies, it is found that the asymptotic method used with small ground clearance is valid for a clearance less than about 0.3 of the chord length. Next, the interference of ground-effect aerofoils in the channel wall, in a free jet stream, and in a slipstream are examined, and it is found that the present method can be easily applied to these problems.

The flow past an unyawed, flat, delta wing with subsonic edges, carrying a slender half-cone mounted centrally underneath, is found by using the method of supersonic conical flows. The wing plane is taken to be at zero incidence, but it is shown that wing incidence, camber, twist and thickness effects may be incorporated by superimposing on this solution the field of the isolated wing with these properties. The results are shown to agree with previous work on configurations with sonic or supersonic leading edges for the common case of sonic edges. The leading edge loading may be made zero to secure attached flow at any lift coefficient by using camber and incidence, but it is shown that there is a certain positive lift coefficient for which a negative incidence alone will suffice.

The paper describes a key risk area threatening the widespread deployment of unmanned air vehicles (UAVs), that of attaining high levels of autonomy. Autonomy is loosely defined in the context of UAVs and the meaning of ‘level of autonomy’ discussed. The paper argues that the achievement of high levels of autonomy is not merely a function of increasing machine intelligence but also of maintaining the human operator’s engagement with the decision making process and retaining human authority. An assumption is that a human being in the loop will be a requirement for safety, flight clearance and legal reasons on early systems. Therefore, developers of highly autonomous systems are presented with a paradox. It will be argued that the human must be placed at the centre of the design process and consequently human factors, the human machine interface and the system architecture become critical to achieving high levels of autonomy. This quality impacts on the entire knowledge acquisition and design cycle and broadens what is meant by that term placing it as a discipline firmly in the systems design community. The paper concludes by outlining the key barriers to the successful development of highly autonomous UAVs.

This paper is concerned with the drag, at zero lift, of wing-body combinations which are symmetrical about the wing plane. Only two assumptions are made: that surface slopes are sufficiently small for the application of linearised theory, and that it is sufficient to satisfy the body boundary condition on the surface of a circular cylinder.

The configuration is represented entirely by singularities along the body axis, and three drag formulae are derived. The first involves the Laplace transforms of functions representing the strengths of the sources and multisources; the difficult problem of inverting these transforms is thereby avoided. The second involves these strength functions themselves, and is of the form of a series of double integrals of the famous von Kármán type. The third involves functions describing the geometry of a very hypothetical quasi-cylinder which has the same drag as the Whole configuration in question. The convergence of these formulae is established.

These results are then used to make an order-of-magnitude analysis for wing-slender-body combinations: a large number of drag terms are seen to be negligible when the ratio of body radius to wing chord is small, and the dominant terms for smooth bodies are those appearing in Ward's drag formula.

Finally, the geometry of the “ equivalent quasi-cylinder ” is investigated.

As the result of over a decade of intensive Anglo-French effort on both the airframe and the powerplant, the Concorde supersonic transport has completed its development and has convincingly demonstrated its ability to cross the Atlantic non-stop at Mach 2·0 with a full payload.

This represents a major milestone in the evolution of aviation where increases in speed and reduction of journey times have always attracted traffic. It also represents a major break-through in reducing passenger travel fatigue as has been demonstrated to a large number of people during the recent route-proving trials.

Accurate prediction of the response of low-aspect ratio, flexible aircraft requires correspondingly accurate modeling of the aircraft itself and of the aerodynamic forces, both respectable problems. Assuming that the wing can be modeled as a nonuniform plate, the discretisation process of choice is the finite element method (FEM), which demands a very large number of degrees of freedom for good accuracy. Moreover, accurate modeling of the aerodynamic forces acting on the aircraft suggests the use of computational fluid dynamics (CFD), which requires the use of an extremely large number of variables. On the other hand, feedback control design for the aircraft demands an aircraft model of relatively small order, so that the dimension of the FEM and CFD models must be reduced drastically. Based on physical considerations, reasonably accurate model reductions can be achieved, but a problem remains because the FEM and CFD grids are likely to differ from one another. It is shown in this paper how to achieve desirable model reductions for both the FEM and CFD and how to integrate the aerodynamic forces into the aircraft state equations. A numerical example demonstrates how the theory can be applied to the flight of a flexible aircraft. The analytical/computational approach developed here should permit parametric studies ultimately resulting in a reduction in the time required for aircraft design and flight testing.