This paper reports the findings of a flutter investigation on a low-pressure turbine blade using a 3D, non-linear, integrated aeroelasticity method. The approach has two important features: (i) the calculations are performed in a time-accurate and integrated fashion, whereby the structural and fluid domains are linked via an exchange of boundary conditions at each time step, and (ii) the analysis is performed on the entire bladed-disk assembly, thus removing the need to assume a critical vibration mode shape. Although such calculations are both CPU and in-core memory intensive, they do not require prior knowledge of the flutter mode and hence allow a better understanding of the aeroelasticity phenomena involved.

The flow is modelled inviscidly but the steady-state viscous effects are accounted for using a distributed loss model. The structural model was obtained from a standard finite element (FE) representation and a large number of assembly modes were included in the calculations. The study focused on three part-speed conditions at which a number of unstable modes were known to exist from the available experimental data. The whole assembly was modelled using about 664,000 mesh points and predictions were made of aeroelastic modal time histories. From these time histories it was possible to identify the forward and backward travelling waves and to deduce the unstable modes of vibration. The theoretical predictions were found to be in very good agreement with the experimental findings.

The Paper describes a metal and composite recurrent cost model of a large civil aircraft wing structure for a multidisciplinary design, analysis and optimisation (MDO) environment. The work was part of a recent European MDO project (BE95-2056) which investigated methods for the integration of structures, aerodynamics, dynamics and manufacturing cost at the preliminary design stage.

The paper discusses the cost modelling approach, which is based on parametric and process cost model methods, and the integration of the cost models into an MDO process. Results for the cost models are shown.

A framework has been successfully developed which allows the incorporation of manufacturing cost models into an MDO environment. It allows a designer to evaluate cost changes with respect to specific design changes such as rib pitch, stringer pitch, wing area and wing sweep.

An experimental study has been carried out to explore the effect of propeller-induced slipstream on the vortex flow field on a fixed-wing Micro Air Vehicle (MAV). Experiments were conducted at a freestream velocity of 10 m/s, corresponding to a Reynolds number based on a root chord of about 1.6 × 105. Surface flow topology on the surface of the MAV wing at propeller-off and propeller-on conditions was captured using surface oil flow visualisation at four angles of incidence. The mean off-body flow over the MAV was documented in the four spanwise planes at different chord position using Stereoscopic Particle Image Velocimetry (SPIV) technique at angle-of-attack of 24° for both conditions. The oil flow visualisation showed minimal differences in flow patterns for propeller-off and propeller-on conditions at 10° and 15° incidence. The small asymmetry between port and starboard side observed at 20° during the propeller-off condition became significantly pronounced at 24°. The fuselage stub which is necessary for housing the motor of the propeller was seen to have a significant effect on the flow symmetry at large incidences that can occur when the MAV encounters sudden vertical gusts. Switching on the propeller restored the symmetry at both incidences. SPIV measurements were carried out at the incidence of 24° which exhibited the highest asymmetry. The off-body data shows the re-establishment of symmetry during propeller-on condition owing to the increase in the magnitude of spanwise and vertical velocities as a result of the propeller slipstream. The findings emphasise the importance of considering the propeller flow and design of the motor housing while evaluating the aerodynamics of low-aspect-ratio MAVs.

This experimental investigation concentrates on the aerodynamic behaviour of a generic fan-in-wing configuration. The effects of the fan(s) on the flow circulation in a short take-off and landing or a transition flight condition without ground effect are evaluated. A wind-tunnel model has been constructed and tested to quantify the aerodynamic effects. Force measurements, surface pressure measurements, stereo-particle image velocimetry and wool tufts flow visualisation are performed. Different fan-in-wing configurations with the fans rotating in the wing plane, one fan either at the rear or front part of the wing and two fans are compared to the closed wing without fans set as reference. A fan placed near the trailing edge improves significantly the lift coefficient due to a jet flap effect on the wing lower side combined with enhanced suction on the wing upper side. The jet exiting the nozzle rolls up in a counter rotating pair of vortices affecting significantly the wing behaviour.

This experimental investigation constitutes also a useful database for further CFD comparison.

Commercially-driven air traffic management (ATM) innovations typically aim to increase air space capacity and/or reduce delays. Here, their potential application for environmental mitigation is discussed. Both carbon dioxide (CO2) and non-CO2 climate impacts are considered, as are noise and air quality issues. We outline the technological, scientific and political barriers to an integrated approach to applying ATM technologies to environmental mitigation. These issues highlight the need to improve comparison and prioritisation of the emissions and effects of aviation.

I must start by thanking the Royal Aeronautical Society for the invitation to present this 27th Lanchester Memorial Lecture. It is an honour and a privilege to follow in the footsteps of the distinguished scientists and engineers who have given the first 26 lectures in this series. These lectures have included many outstanding reviews of a wide range of different topics and I am very conscious that they have set a standard that I, for my part, will find difficult to match. I hope, however, that by choosing a topic that has only been mentioned in passing in just a few of the previous lectures, I will be able to make my own distinct, individual contribution to this tribute to the memory of a man who was not only a great scientist and engineer and talented musician but who, by his writings as long ago as 1907, still carries a message for us today in 1987.

Some velocity profile and skin friction measurements for turbulent boundary layers at subsonic and supersonic speeds are reviewed to derive the magnitude of the wake component and the effects of its variation with Reynolds number on predictions of skin friction. As Reynolds number increases the wake component increases rapidly, reaches a maximum and then decreases. The measurements considered suggest that this maximum may be higher at supersonic speeds than at subsonic speeds. This wake component variation has a significant influence on predictions of skin friction in the Reynolds number range of interest on wind tunnel models.

Our engineering habits form slowly, and once formed are slow to change. This is perhaps less true in aeronautics than in other branches of engineering. Nevertheless some aeronautical customs are deeply rooted, so much so that it is easy to confuse the superstructure of habit with the basic foundations. In this article I have reviewed some of the customary methods of specifying strength conditions and have tried to disentangle the underlying principles. Current changes in these methods emphasise the need of a proper appreciation of the principles, and a study of these principles suggests even more radical changes of methods in future.

A review is given of developments in the field of aeroelasticity during the past ten years. The effect of steadily increasing Mach number has been two-fold: on the one hand the aerodynamic derivatives have changed, and in some cases brought new problems, and on the other hand the design for higher Mach numbers has led to thinner aerofoils and more slender fuselages for which the required stiffness is more difficult to provide. Both these aspects are discussed, and various methods of attack on the problems are considered. The relative merits of stiffness, damping and massbalance for the prevention of control surface flutter are discussed. A brief mention is made of the recent problems of damage from jet efflux and of the possible aeroelastic effects of kinetic heating.

In this paper a generalised presentation for symmetrically stiffened orthotropic cylinders under compression is developed, based on a linear orthotropic stability theory of cylinders. Similarities and differences in the minimum weight behaviour of stiffened cylindrical shells and flat transversely-stiffened wide columns are investigated in some detail to provide a satisfactory physical picture. The concluding results provide a comparative evaluation of various forms of stiffening systems for cylindrical shells under compression.

A meeting of the Society was held in the Great Hall of the Institution of Civil Engineers, Great George Street, Westminster, S.W.1, on Thursday, 31st October, 1946, at which a paper on “Aeroelastic Problems at High Speed,” by Professor A. R. Collar, M.A., D.Sc, F.R.Ae.S., was presented and discussed. In the Chair, Dr. H. Roxbee Cox (Vice-President).

The prediction of the forced vibration response of rotating cantilever blading is of considerable importance in the design of turbine and compressor blading where the rotational speeds are relatively high. Non-linearities in the blades arise from the effects of Coriolis accelerations due to the rotation of the blades mounted on the periphery of the disc.

Lo and Renbarger derived the differential equations of motion of a cantilever beam mounted on a rotating disc at a given stagger angle and show, for a linear case, that the frequencies of a bar vibrating transversely to a plane inclined at an angle to the plane of rotation can be found by a simple transformation of the frequencies of vibration perpendicular to the plane of rotation. Boyce, Di Prima and Handelman applied the Rayleigh-Ritz and Southwell methods to the case of a turbine blade vibrating perpendicular to the plane of rotation and determined the upper and lower bounds of the natural frequencies.

The 887th Lecture to be given before the Royal Aeronautical Society was held at the Institution of Mechanical Engineers, Storey's Gate, London, S.W.I, on 29th October 1953, and was presided over by Sir William Farren, C.B., M.B.E., M.A., F.R.S., F.R.Ae.S., President of the Society. Sir William said that this was the first main lecture of the session. In one paragraph of his paper Mr. Mercer said that those who were specialists in fields of a highly specialised kind should get together with their friends in other branches of aeronautics, otherwise they might all go on in ignorance of what was being done by others; he hoped therefore, that there were a number of specialists present who would join in the discussion.

Introducing the Lecturer, Sir William said that he had known Mr. Mercer for many years; Mr. Mercer had joined the Blind Landing Experimental Unit at Martlesham Heath when it was formed and it was there that he had done the work described in the paper. The President said he had always felt that one of the things most needed for the real development of aviation was a means of landing in bad weather. He hoped that Mr. Mercer's paper was a contribution towards that end; it was concerned with one phase only of blind landing—approach—but he hoped there would be other papers on the subject. He now called on Mr. Mercer to give his paper.

Ever since the realisation of the matrix Displacement Method, the introduction of a kinematically satisfactory spar element has remained a difficult and, as yet, not fully resolved problem. Two cardinal points have to be considered in establishing a suitable spar model: first, it must reproduce adequately both shearing and bending deformations, and second, it must satisfy everywhere the kinematic compatibility conditions, especially at the junction with the wing cover. Of course, if we were to restrict ourselves to a uniform rectangular element, the corresponding theory could be written down in a few lines. However, from the practical point of view, it is essential to take account, ab initio, of taper, and this necessitates some subtle considerations, especially if we have to allow for asymmetrical taper, as becomes indispensable with some modern flying vehicles.

Many investigators have studied the aerodynamics of axial flow turbomachinery but none has produced a complete map of the three-dimensional flow behind a rotor row. This is of considerable interest to the aero-acoustician. A system is described which uses a constant temperature hot-wire anemometer to analyse the flow behind such a rotor. Although much information may be extracted by using the technique, its interpretation depends to a large extent on its form of presentation. An analysis of the flow behind a research fan is used as a means of discussing various forms of visual presentation.

Free vibration analysis of laminated composite beams, plates and stiffened plates using the finite element method has been presented. To give due importance to the shear deformation in composite materials, a high-order element which considers the quadratic variation of shear strain along the thickness of the element has been developed. The element has been developed using the quadratic isoparametric shape functions. Therefore, it can accommodate curved boundaries. Moreover, the stiffener element can have an arbitrary planform and it need not pass through any nodal point. The stiffener element has been developed in such a fashion to make the mesh division free from the location of the stiffener. This is extremely advantageous for optimisation of the path of the stiffener in a stiffened plated structure, while it is not necessary to modify the mesh division whenever the stiffener path is modified. Two different shape functions — the eight-node Serendipity and the nine-node Lagrangian — have been employed and a comparison between their performances has been presented. Two schemes for the generation of mass matrix, namely the lumped mass scheme and the consistent mass scheme, have been developed. The elements have been implemented in the ASKA element library and they can be used along with any standard ASKA element.

The objective of this paper is to present the principles of a mathematical model of aircraft prescribed motion. Requirements imposed on the aircraft motion are treated as program constraints on the system and both the transient dynamic solution of motion equations and the control ensuring the exact realisation of the prescribed motion are obtained as a result. The approach used is equivalent to the Lagrange multiplier method, generalised for the purpose of this paper. It consists of the solution of the set of differential/algebraic equations of index exceeding three. The presented mathematical model has been applied to the simulation of aircraft prescribed motion in a loop. The flight along an ideal circle and the flight with additionally demanded constant velocity are described. Some results of numerical calculations are demonstrated.

The motion of an aircraft along a straight flight path inclined to the vertical is considered from an arbitrary initial velocity which may be far from steady flight speeds. The velocity and incidence, as a function of distance or time, provide a solution to the ‘inverse phugoid problem’ of finding the control laws that keep an aircraft on a straight flight path by exactly cancelling the phugoid instability (the short period mode is omitted by neglecting the rotational inertia of the aircraft). This is a non-linear problem, namely finding the motion of an aircraft from an initial condition far removed from a stable steady state, or finding the large amplitude deviations from an unstable equilibrium. The stability and instability problems are solved analytically using the following assumptions:

1. (i) the aircraft flies subsonically so that only form and induced drag need be considered;

2. (ii) the thrust consists of a constant term, minus a term specifying a dependence on velocity similar to that of the total drag; and

3. (iii) the aircraft incidence is away from stall so that lift is a linear function of incidence.

The response curves show that:

1. (i) flight at velocities below the lower steady flight speed, which is unstable, is difficult to maintain due to fast divergence towards the stall; and

2. (ii) for an initial velocity above the lower steady flight speed, there is convergence towards the higher steady flight speed.

The convergence, in the latter case (ii), may be slow along a straight flight path, although it can be hastened by suitable manoeuvres.

For types of structure where safety standards cannot be preserved by reliance on inspection, a permissible “safe life” has to be determined by relating the loads experienced by the aircraft in service to the fatigue performance of the structure, with due allowance for scatter in fatigue strength and in the frequency with which loads are encountered. This paper gives a working method for estimating the “safe life”—at which the risk of fatigue failure is negligible—of the wing structure of a transport aircraft, where the flight loads giving rise to fatigue are overwhelmingly due to atmospheric turbulence but where allowance has also to be made for “ground-to-air” loads.