An eddy viscosity relation for the inner region is combined with an empirical wake function for the outer region of a turbulent boundary layer. The calculated distributions for the mean velocity and Reynolds shear stress are compared with both direct numerical simulation and experimental data in a zero pressure gradient boundary layer, mainly at low Reynolds numbers. The comparison is favourable in both the inner and outer regions of the flow.

The tailless aircraft configuration is relatively rare in the field of aircraft design and for many years was regarded as a “freak” design. It is intended to trace the development of tailless and flying wing aircraft and to outline the major problems in control and stability.

First it is necessary to arrive at an arbitrary definition of the term tailless aircraft. Later an attempt will be made to classify the different tailless types in detail, but for the moment it is sufficient to regard a tailless aircraft as “an aircraft having no auxiliary horizontal stabilising plane, deriving lift and controlling moment solely from the main plane.” This definition, as it stands, includes what are often termed flying wing aircraft.

A description is given of the distribution of vorticity in the surface of thin wings with large leading edge sweep. Although the delta wing is chosen as the basic plan form the deductions are general and applicable to other types of wing. The conclusions are illustrated with experimental evidence from a water tunnel.

The natural frequencies of vibration of a cantilever beam allowing for rotary inertia and shear deformation are obtained by the approximate Ritz method.

The workability of the method is dependent upon the approximating functions chosen for the dynamic displacement curves. A series of characteristic functions representing the normal modes of vibration of cantilever beams in simple flexure is used as the approximating functions for both deflections due to flexure and shear deformation. Good agreement is shown between frequencies obtained by the Ritz method and those resulting from an analytical solution.

The effect upon the natural frequencies of allowing for rotary inertia alone is shown and it is seen to increase rapidly with mode number.

For years, the development and testing of aero gas turbines has been hampered by the inability of development and test engineers to perceive and readily measure the positional changes of static and moving components within the engines due to the opaque nature of the engine casing walls. ‘Touch’ or abradable probes have been used but require special engine builds to employ them.

With the introduction of high strength materials and the thinning of engine structures and as the performance requirements have become more demanding, small component movements during engine handling have become more important. This is particularly true in areas where the airflow has the greatest work potential such as the tip and root seals of the high pressure turbine and the root seals of the final compressor stages.

Control of flight is essentially control of lift. When one proposes—as we shall do here—to tamper with one of the most fundamental principles of aircraft control, it is evident that this may have far reaching repercussions and that such a venture cannot therefore be undertaken lightly. The elevator has been the primary longitudinal control since the earliest pioneering days and has served the aircraft well from its modest beginnings to this age of supersonic flight. Why do we want to change and what do we want to take its place; how will the pilot react to an altogether novel way of controlling the aeroplane? These are the questions one is bound to ask when direct lift control is suggested as an alternative and these are the questions which we shall attempt to answer here.

Data mining is an iterative process in which progress is defined by discovery through either automatic or manual methods. A data cleaning procedure is proposed to improve the quality of classification tasks in the knowledge discovery process by taking into account both redundant and conflicting data. The redundancy check is performed on the original dataset and the resultant dataset is preserved. This resultant dataset is then checked for conflicting data and, if any are found, they are corrected and updated on the original aircraft dataset. This updated dataset is then classified using a variety of classifiers such as Bayes, functions, lazy, MISC, rules and decision trees. The performance of the updated datasets on these classifiers is examine, and the result shows a significant improvement in the classification accuracy after redundancy and conflicts are removed. The conflicts after correction are updated in the original dataset, and when the performance of the classifier is evaluated, great improvement is observed. This paper aims to address how data mining techniques can be used to understand complex system accidents in the aviation domain. Decision trees are considered to be the one of the most powerful and popular approaches in knowledge discovery and data mining. The objective is to develop a classification model for aviation risk investigation and reduction using a decision tree induction method that enhances the ability to form decision trees and thereby proves that the classification accuracy of decision trees is greater. Different feature selectors are used in this study in order to reduce the number of initial attributes.

This paper focuses on the trajectory planning for a UAV on a low altitude terrain following/threat avoidance (TF/TA) mission. Using a grid-based approximated discretisation scheme, the continuous constrained optimisation problem into a search problem is transformed over a finite network. A variant of the Minimum Cost Network Flow (MCNF) to this problem is then applied. Based on using the Digital Terrain Elevation Data (DTED) and discrete dynamic equations of motion, the four-dimensional (4D) trajectory (three spatial and one time dimensions) from a starting point to an end point is obtained by minimising a cost function subject to dynamic and mission constraints of the UAV. For each arc in the grid, a cost function is considered as the combination of the arc length, fuel consumption and flight time. The proposed algorithm which considers dynamic and altitude constraints of the UAV explicitly is then used to obtain the feasible trajectory. The resultant trajectory can increase the survivability of the UAV using the threat region avoidance and the terrain masking effect. After obtaining the feasible trajectory, an improved algorithm is proposed to smooth the trajectory. The numeric results are presented to verify the capability of the proposed approach to generate admissible trajectory in minimum possible time in comparison to the previous works.

The hypersonic aeroelastic problem of a generic hypersonic vehicle having a lifting-body type fuselage and canted fins is studied using third order piston theory and Euler aerodynamics. Computational aeroelastic response results are used to obtain frequency and damping characteristics, and compared with those from piston theory solutions for a variety of flight conditions. Aeroelastic behavior is studied for the range of 2·5 < M < 28, at altitudes ranging from 10,000ft to 80,000ft. Because of the significant computational resources required, a study on optimal mesh selection was first carried out for use with Euler aerodynamics. The three dimensional flow effects captured using Euler aerodynamics was found to lead to significantly higher flutter boundaries when compared to those based on nonlinear piston theory. The results presented here illustrate some of the more important three dimensional effects that can be encountered in hypersonic aeroelasticity of complex configurations.

The synergistic combination of two promising engine architectures for future aero engines is presented. The first is the Composite Cycle Engine, which introduces a piston system in the high pressure part of the core engine, to utilise closed volume combustion and high temperature capability due to instationary operation. The second is the Intercooled Recuperated engine that employs recuperators to utilise waste heat from the core engine exhaust and intercooler to improve temperature levels for recuperation and to reduce compression work. Combinations of both architectures are presented and investigated for improvement potential with respect to specific fuel consumption, engine weight and fuel burn against a turbofan. The Composite Cycle alone provides a 15.6% fuel burn reduction against a turbofan. Options for adding intercooler were screened, and a benefit of up to 1.9% fuel burn could be shown for installation in front of a piston system through a significant, efficiency-neutral weight decrease. Waste heat can be utilised by means of classic recuperation to the entire core mass flow before the combustor, or alternatively on the turbine cooling bleed or a piston engine bypass flow that is mixed again with the main flow before the combustor. As further permutation, waste heat can be recovered either after the low pressure turbine – with or without sequential combustion – or between the high pressure and low pressure turbine. Waste heat recovery after the low pressure turbine was found to be not easily feasible or tied to high fuel burn penalties due to unfavourable temperature levels, even when using sequential combustion or intercooling. Feasible temperature levels could be obtained with inter-turbine waste heat recovery but always resulted in at least 0.3% higher fuel burn compared to the non-recuperated baseline under the given assumptions. Consequently, only the application of an intercooler appears to provide a considerable benefit for the examined thermodynamic conditions in the low fidelity analyses of various engine architecture combinations with the specific heat exchanger design. Since the obtained drawbacks of some waste heat utilisation concepts are small, innovative waste heat management concepts coupled with the further extension of the design space and the inclusion of higher fidelity models may achieve a benefit and motivate future investigations.

A novel second-order sliding-mode-based impact angle and autopilot lag guidance law for engaging manoeuvering targets with unknown acceleration is presented in this study. A backstepping technique is applied to the design of the sliding surface. The proposed guidance law is based on a new sliding surface. It exhibits the advantage of ensuring that the sliding surface and its derivative will converge to zero in finite time while guaranteeing that the sliding surface will not cross zero until the ultimate time. The method effectively eliminates the undesired chattering of the sliding surface. To compensate for the uncertainty caused by target manoeuvering, a new observer is developed to estimate target manoeuvering. The convergence of the system is proven through a Lyapunov function and finite time convergence theory. Lastly, mathematical simulations results show that the proposed guidance law can achieve precise interception with a wide range of impact angles, thereby verifying the effectiveness of the guidance law.

Uncertainties in the use of the discrete-vortex method in modelling the time development of the wake of a circular cylinder at very high Reynolds numbers are investigated. It is shown that simply introducing vorticity at generally accepted separation positions at a rate of ½Us2, Us being the velocity at separation, gives wholly unrealistic wake predictions. In the base region pressure fields occur which would promote separation in steady flow and so a first approximation for ‘secondary’ separation is incorporated into the model. This brings pressure distributions and vorticity structures at subcritical and supercritical Reynolds numbers into good agreement with experiment. The convection of the vortices is calculated using the cloud-in-cell technique and comparisons are made with direct summation methods.

In this paper the way in which an inviscid incompressible fluid changes its total pressure is discussed. The basic equations are set out and it is shown that changes in total pressure are essentially associated with time variable flows. Simple examples are worked to illustrate this and finally application is made to single-stage flow machines, using a simple vortex representation.

This paper considers some of the special problems of wind tunnel testing which arise in high-lift work. Discussion is confined to current experience in the R.A.E. and N.P.L. low-speed wind tunnels, and refers mainly to tests on blowing over flaps or jet-flaps. Comments are made on suitable size of model, methods of feeding compressed air into models without affecting balance readings, and general test technique. Methods of model construction are not specifically discussed.

This paper, which describes a photoelastic determination of the elastic stress concentration factors associated with shouldered shafts subjected to pure bending, forms the second part of an investigation initiated by the Structures Committee of the Royal Aeronautical Society. A general description of the manufacture of the models, the photoelastic technique employed to evaluate the s.c.f. and the results of the tests on shouldered shafts subjected to torsion have been given in Part I (Ref. 1).

A theoretical model had been previously developed to predict the behaviour of a semi-rigid helicopter rotor blade during rotor engagement and disengagement under high wind conditions. This has now been extended to cater for an articulated rotor by modelling the droop and anti-flap stop mechanisms by high rate linear springs.. Much increased blade deflections are predicted, as anticipated, along with the possibility of blade tip/tail boom contact at a high enough wind speed. The importance of the higher modes in the description of the phenomenon is demonstrated. A comparison of the theoretical model with an experiment using a full size Puma aircraft is presented.

The non-linear boundary-value problem resulting from an analysis of the stresses and deflections of an annular membrane is transformed into an initial-value problem, replacing the iterative and series methods of Pifko and Goldberg. A general solution valid for the complete range of values of the ratio of inner and outer radii is obtained and tabulated by this non-iterative method.

Evvard’s technique is applied to the problem of a thin finite wing moving at a supersonic speed when the leading edge is subsonic. It is developed in two methods:—

• (i) in which the relationship between the pressure loading and the integrals of the downwash over the wing surface is extended as far as possible, and which has to be computed numerically;

• (ii) in which approximations are made for the upwash velocities in the neighbourhood of the leading edge, resulting in a series of standard integrals for the estimation of the pressure loading.

Method (ii) is applied to the pressure loading on a flat plate triangular wing and cropped delta wing, and the application to more general shapes is discussed.

Some recent theoretical work on slender pointed wings at zero lift is co-ordinated and extended. The wings considered may have any pointed plan form shape, provided that the trailing edge is straight and unswept. The root section profile and cross-section shapes are arbitrary, provided that, on any one wing, the latter are “descriptively similar” (diamond or parabolic biconvex for instance), though not necessarily geometrically similar. The chief aim of the work is to find wings with simple geometry, low wave drag and pressure distributions which are unlikely to be seriously affected by viscous effects. Wave drag and pressure distributions are calculated by slender-wing theory. General formulae, which are both simple and instructive, are given for the wave drag and the overall pressure distribution, with particular emphasis on the root pressure distribution. Results for a number of wings of special interest are presented and discussed.

An exact method has recently been presented for computing the critical buckling loads, or the natural frequencies of vibration, of prismatic assemblies of rigidly inter-connected, thin, flat, rectangular plates. The plates can carry uniform in-plane longitudinal, transverse and shear stresses and may be isotropic or anisotropic, with the elastic coupling stiffness matrix B = 0. This method can be used even if some of the plates have curved cross sections, are non-uniformly longitudinally compressed or, in certain circumstances, have B ≠ 0. However, approximations are then involved when representing the plates. Results presented herein indicate that good answers can be obtained despite these approximations.