The aim of the present work is the investigation about the use of blade trailing-edge flaps for the reduction of vibratory loads arising at the hub of helicopter main rotors in forward flight. The alleviation of these loads is achieved through multicyclic higher harmonic actuation of the blade flaps, which is related to measured vibratory loads amplitude. The feedback control law is obtained by an optimal control process based on the minimisation of a cost function, under the constraint of compatibility with the nonlinear equations governing blade aeroelasticity. In the numerical investigation concerning a four-bladed rotor in level flight conditions, a computationally efficient local controller methodology is applied, with the attention focused on the effectiveness of the control algorithm, along with its robustness with respect to differences (existing in real applications) between the aeroelastic models used for control law synthesis and validation.

The potential positive effects of ground proximity on the aerodynamic performance of a wing or aerofoil have long been established, but at transonic speeds the formation of shock waves between the body and the ground plane would have significant consequences. A numerical study of the aerodynamics of an RAE2822 aerofoil section in ground effect flight was conducted at freestream Mach numbers from 0·5 to 0·9, at a range of ground clearances and angles of incidence. It was found that in general the aerofoil’s lifting capability was still improved with decreasing ground clearance up until the point at which a lower surface shock wave formed (most commonly at the lowest clearances). The critical Mach number for the section was reached considerably earlier in ground effect than it would be in freestream, and the buffet boundary was therefore also reached at an earlier stage. The flowfields observed were relatively sensitive to changes in any given variable, and the lower surface shock had a destabilising effect on the pitching characteristics of the wing, indicating that sudden changes in both altitude and attitude would be experienced during sustained transonic flight close to the ground plane. Since ground proximity hastens the lower surface shock formation, no gain in aerodynamic efficiency can be gained by flying in ground effect once that shock is present.

High speeds on water are being attained in racing sport and in attempts on world speed records in various classes. Success, safety and stability of these craft depends upon the favourable interaction of their aerodynamic and hydrodynamic characteristics under the influence of two media, one about 800 times denser than the other. Speed on the straight course and in turns is important.

As its velocity increases, a craft experiences increasing dynamic pressure in water and to maintain the balance, the ‘wetted’ area of the craft reduces as it rises up (‘planing’). Modern fast craft have ‘tunnel-hulls’ and lifting areas to generate aerodynamic lift and to assist the craft to attain planing attitudes rapidly.

The ‘lifting’ areas may not necessarily be in the correct locations however. An example often seen is that of a power boat riding virtually on its propeller and ‘wallowing’ in an unstable manner. In this case the variation in riding height alters the relative positions of centres of aero- and hydro-lift such that the ‘stability-margins’ are near critical in both the longitudinal and the lateral sense.

The equivalent continuum beam-rod model of an aircraft wing structure with composite laminated skins has been developed based on the concept of energy equivalence. The equivalent structural properties of the continuum beam-rod model are obtained by directly comparing the reduced stiffness and mass matrices for a typical segment of aircraft wing with those for a finite element of continuum beam-rod model. The finite element stiffness and mass matrices are condensed through the well known finite element formulation procedure to be used in calculating the reduced stiffness and mass matrices for the aircraft wing segment in terms of the continuum degrees of freedom introduced in this paper. The present method of continuum modelling may yield every equivalent structural property including all possible couplings between bending, torsional and transverse shear deformations. To evaluate the equivalent continuum beam-rod model developed herein, the free vibration and aeroelastic analyses for a box-beam type aircraft wing structure have been conducted by using the continuum beam-rod model, and the numerical results are compared with the results by using other different models of the aircraft wing structure.

The following is a report of preliminary tests to determine the effect of vortex generators on the drag of a circular cylinder. It was thought that with suitable placement of the vortex generators on the cylinder, the laminar separation of the boundary layer at subcritical Reynolds numbers might be delayed, and a reduction in drag coefficient obtained.

The program VICONOPT is used to find the optimum (least mass) dimensions of a range of stiffened wing panels which are subject to buckling and material strength constraints and are loaded in axial compression with a sinusoidal manufacturing imperfection. Design plots are presented to show the effects that various rib spacings and stiffener types have on optimum design mass. A simplified model of a complete wing box is used to illustrate the design of a full wing panel and plots of optimum values of design variables at various stations along the wing have been obtained. The results were chosen to illustrate the practicality of optimisation with reference to manufacture of a full wing panel and to show the effect of changing the sophistication of modelling and theory used for the range of panels considered. The important aspects of the choice of design variables and design concepts are highlighted and percentage savings in mass, compared with an optimum metal panel design, are given for the various (global) optima found along with some examples of (rejected) local optima.

The main purpose of this paper is to illustrate the use of the stability functions for long plates in combined compression and shear defined in a previous paper by the authors. The panel is supported on its two edges and on a series of N intermediate simple supports. Each of the two edges is either clamped or simply-supported and numerical results are given for each of the three possible combinations, with varying numbers of supports. In all cases it is found that a simple parabolic interaction formula between shear and compression is extremely accurate. Also simple approximate formulae are given for calculating the wavelength of the buckle and the lead of the nodal lines for any combination of compression, shear, edge conditions and number of supports.

The selective bleed turbofan is a two shaft, three compressor, variable cycle aircraft engine. Its bypass ratio (BPR) can be modulated to suit the operating flight conditions. At subsonic flight speeds it operates as a medium bypass turbofan (low pressure mode or LPM). It becomes a low bypass turbofan (high pressure mode or HPM) when flying faster and is capable of supersonic cruise without reheat (RH).

The aim of this paper is to describe the selection of the main engine design parameters in the HP mode and to outline the challenges to be met to enable the powerplant to operate in the other mode. The components have to perform satisfactorily over a wide range of varying conditions. The resulting demands are reflected in many ways, for example in the operating lines of the low pressure compressors and in the overall efficiency changes with increased bypass ratio. In the case of the present investigation the two bypass ratios selected are 0·3 for the high pressure mode and 0·7 for the low pressure mode. The latter was selected after an examination of a choice of bypass ratios ranging from 0·4 to 1·0. The change in variable stators required to achieve the LP mode cycle is indicated.

The effect of changing bypass ratio on engine performance and handling is shown. The main conclusion is that the design of this engine seems to be feasible within current technological capabilities and further investigation is encouraged because it appears to yield significant benefits.

This paper studies hovering capability of flapping two-dimensional tandem wing sections inspired by a real dragonfly wing configuration and kinematics. Based on unsteady numerical simulations, the dragonfly corrugated wings have been benchmarked against a flat wing in terms of the aerodynamic forces and flow structures generated during a flapping cycle. The timing of rotation at each stroke is studied by pitch rotation at three different rates, i.e., 80%, 60% and 40% of a flapping period. The results suggest that the longer time pitch rotation with the period of 80% of the overall flapping period is closer to the force calculations obtained of a balanced flight, that is, the mean vertical force supports the dragonfly weight of 0.754 g with a small difference of 0.92% and the mean horizontal force indicates negligible thrust. However, the corrugated wing performs aerodynamically differently from the flat plate with differences in and in by ±2.06% for the corrugated shape. The vorticity flow field for both wings have been recorded at some instants of flapping motions which give more explanation of such dissimilarity.

The necessary conditions of Clebsch and Weierstrass and of the multiplier rule in the calculus of variations, which arise from the study of the first variation of a function, are summarised. A further necessary condition associated with the second variation is stated. The latter condition is applied to two problems: (i) the determination of the thrust-time programme which maximises the altitude of a sounding rocket, (ii) the determination of the thrust direction programme for a rocket with a known propellant expenditure programme which yields a maximum range. In both problems it is found that the additional necessary condition is satisfied.

The general theory and formulation of Part 1 can be applied to a whole family of element shapes. However, quadrilaterals and triangles which are compatible with each other are found to be most useful in idealising general structures. The lowest order polynomial shapes of this kind which are generally curved are the ‘biquadratic’ quadrilateral and the ‘quadratic’ triangle. These have nodes (or nodal pairs for the sandwich elements) situated at the corners and the mid-sides of the ‘parent’ elements. The quadrilateral therefore has eight nodal pairs.

The development during the last two decades of analytical techniques for the solution of problems relating to the optimisation of rocket trajectories is outlined and the present position in this field of research is summarised. It is shown that the determination of optimal trajectories in a general gravitational field can be expressed as a Mayer problem from the calculus of variations. The known solution to such a problem is stated and applied, first to the special case of the launching of an artificial satellite into a circular orbit with minimum expenditure of propellant and, secondly, to the general astronautical problem of the economical transfer of a rocket between two terminals in a gravitational field. The special cases when the field is uniform and when it obeys an inverse square law of attraction to a point are then considered, and the paper concludes with some remarks concerning areas in which further investigations are necessary.

The aim of this paper is to contribute to the growing subject of inlet/engine compatibility a discussion of some of the more interesting results of research and development testing of rig compressors and engines carried out at the Bristol Engine Division of Rolls-Royce (1971) Ltd, and to discuss these against a now extensive background literature.

The effects of an inlet system on an axial flow gas turbine are principally on performance, stability and mechanical integrity. In this paper attention is focused on the effects of total pressure distortion on the compressor system's performance and stability.

A computational study has been performed to investigate the effectiveness of film cooling in hypersonic laminar flows. Both the primary and the coolant flow are air. First, three different primary flow conditions are used for validation. A uniform boundary condition at the slot exit is found to give unrealistic predictions of the heat transfer rate, whilst a boundary condition involving extended coolant inlet gives improvement. Then five coolant injection rates and three slot heights are examined. It is confirmed that increasing the coolant injection rate can increase the film cooling effectiveness for laminar cases. But slot height does not play an important role under the flow conditions in this study. The computational results are compared with the experimental results and good agreement is achieved.

Simple expressions are derived for predicting the spread of turbulent jets issuing into parallel moving airstreams. Difficulties with applying these expressions to practical problems are discussed and comparisons with experimental results are given.

An analysis is presented of the buckling of a rectangular plate under combined biaxial compression, bending and shear. The sides of the plate to which bending is applied are simply-supported, the other two sides being supported by arbitrary edge members. The solution is expressed in a form suited to automatic computation and some sample results are given.

The Nonweiler method of designing delta wings (caret wings) to have plane shocks on the lower surface is extended to the design of upper surfaces. Cambered and shock-free surfaces are also deduced, and methods of design for complete configurations suggested.

An important point is that all surfaces are of single curvature.

The investigation of axisymmetric states of stress by the matrix displacement method may, in principle, be carried out as an ordinary exercise of three-dimensional analysis, using TETRA 4, TETRA 10 or other suitable elements. However, this approach is clearly not the most efficient by not taking advantage ab initio of the inherent simplifications arising in an axisymmetric state; this leads to an unnecessary inflation of the number of unknowns and complexity of the mesh. A number of techniques to deal with this limiting condition of a three-dimensional state have been developed at the ISD. The theory for one specific element, TRIAX 6 which evolves most naturally from the tapered TRIM 6 element analysed in the preceding note 8 is presented.

The present study demonstrates that cyclically varying the damping coefficient of controllable lag and flap dampers can reduce the 4/rev vibratory hub loads of a four-bladed hingeless rotor helicopter in high speed forward flight. Gradient-based optimization is used to determine the optimal multi-cyclic damping variation inputs that minimise a composite vibration index comprising of all six components of vibratory hub loads. Optimal 2/rev and 3/rev variations in the lag damping coefficient virtually eliminate the vibratory hub drag force and yawing moments, and produce small reductions in the vibratory hub side force. The optimal lag damping variations, interestingly, produce increases in the 3/rev and 5/rev components of the blade root drag shear, that cancel the contributions of the blade root radial shear to the vibratory in-plane hub forces. Despite some increases in higher harmonics of blade response, damper loads, and blade and flexbeam root loads, the lower harmonics and the peak-to-peak values show little change, implying that blade and damper fatigue life would not be adversely affected. When optimal 2/rev and 3/rev variations in flap damping coefficient are introduced in conjunction with the optimal lag damping variations, 30% reductions in the hub vertical vibrations are obtained, in addition to the previous reductions in the vibratory in-plane forces and yawing moment. The cyclic flap damping variations reduce the higher harmonics of the blade root vertical shear. Reductions in hub vibration levels are obtained over a range of forward flight speeds.

International political and socio-economic developments have led the armed forces of many countries to become more aware of how their increasingly-stringent financial budgets are spent. One major expenditure for military authorities is upon aero-engines, because in-service deterioration in any mechanical device, such as an aircraft's gas-turbine engine, is inevitable. Each deterioration has an adverse effect on the performance and shortens the reliable operational life of the engine, thereby resulting in higher life-cycle costs. For a military aircraft's mission-profiles, the consequences of an aero-engine's deterioration upon the aircraft's operational-effectiveness as well as its fuel consumption and life have been predicted in this project using validated computer-simulations. These help in making wiser management-decisions, so leading to the achievement of improved engine utilisation, lower overall life-cycle costs and optimal mission effectiveness for squadrons of aircraft.