Presented in this paper are the preliminary results from a blade-vortex interaction (BVI) test series conducted in the University of Glasgow's 1·61m x 2·13m closed-return, low-speed windtunnel, in which a single vortex interaction with a rigid, loaded rotor blade was studied. This work is an extension of that by Horner and Galbraith, which describes the flow phenomena present during single vortex interaction with an unloaded blade. The paper presents blade surface pressure data recorded by 72 pressure transducers located in the outer regions of the blade during these interactions. Integrated Cn and Cm1/4 data are also presented to provide a detailed history of a BVI event. In addition, a method of isolating the bound circulation and vortex induced effects is presented, from which it is concluded that for a BVI event the vortex-induced effects are (to a first approximation) independent of rotor pitch setting.

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.

In this paper, a surrogate model approach for non-linear aerodynamics is presented in order to reduce the computational effort of coupled aeroelastic analyses. The usability of the approach is demonstrated in static as well as transient aeroelastic analyses of the HIRENASD wing-fuselage configuration. Furthermore, it is shown that the surrogate model approach is able to cover variations of flow conditions at a fixed Mach and Reynolds number.

In recent years, various strategies for the concurrent operation of fixed-and rotary-wing aircraft have been proposed as a means of increasing airport capacity. Some of these strategies will increase the likelihood of encounters with the wakes of aircraft operating nearby. Several studies now exist where numerical simulations have been used to assess the impact of encounters with the wakes of large transport aircraft on the safety of helicopter operations under such conditions. This paper contrasts the predictions of several commonly-used numerical simulation techniques when each is used to model the dynamics of a helicopter rotor during the same idealised wake encounter. In most previous studies the mutually-induced distortion of the wakes of the rotor and the interacting aircraft has been neglected, yielding the so-called ‘frozen vortex’ assumption. This assumption is shown to be valid only when the helicopter encounters the aircraft wake at high forward speed. At the low forward speeds most relevant to near-airfield operations, however, injudicious use of the frozen vortex assumption may lead to significant errors in predicting the severity of a helicopter’s response to a wake encounter.

This paper presents findings from research conducted at the University of Liverpool aimed at optimising the motion cues available from a short-stroke hexapod motion platform. Piloted simulation experiments were conducted to gather both objective and subjective measures of pilot performance and motion fidelity for a typical helicopter low-speed yaw axis capture task. These measures included subjective ratings taken for the first time using a new motion fidelity rating scale. During the experiment four baseline configurations were presented to the pilot; – No motion, Rotational-only motion, Translational-only motion and Rotational + Translational motion, and within these baseline configurations various values of motion filter gains were examined. The key findings from the experiment show that the most acceptable motion fidelity was obtained when congruent cues were presented in both the rotational and translational axes. A possible explanation for this result is postulated to lie within the relative motion perception thresholds of the human semicircular canals. These results contrast with previous research which suggested that yaw axis motion cues were of less importance than translational motion cues.

While it has long been known that added fluid mass may be important in the dynamics of parachutes, due to inadequate or incorrect derivation and/or implementation of the added mass tensor its full significance in the stability of parachutes has yet to be appreciated. The concept of added mass is outlined and some general conditions for its significance are presented. Its implementation in the parachute equations of motion is reviewed, and the equations used in previous treatments are shown to be erroneous. A general method for finding the equivalent external forces and moments due to added mass is given, and the correct, anisotropic forms of the added mass tensor are derived for the six degree-of-freedom motion in an ideal fluid of rigid body shapes with planar-, twofold- and axisymmetry, These derivations may also be useful in dynamic stability studies of other low relative density bodies such as airships, balloons, submarines and torpedoes. Full nonlinear solutions of the equations of motion for the axisymmetric parachute have been obtained, and results indicate that added mass effects are more significant than previously predicted. In particular, the component of added mass along the axis of symmetry has a strong influence on stability. Better data on unsteady forces and moments on parachutes are needed.

The stress-concentration factor is calculated for an infinite plate in tension containing a doubly-symmetrical hole whose boundary consists of parts of three intersecting circles. A suggestion is made for modifying the results to apply to a strip.

From the first, let me nail my colours firmly to the mast by quoting Beethoven’s eventual dedication of the “Eroica” Symphony. For this particular Lanchester Lecture we are “to celebrate the memory of a great man”. My meaning here, I hasten to add, has none of that bitter intention of the Republican Beethoven who, so the story goes, having ripped out his earlier dedication to Bonaparte then substituted those words to commemorate the man once revered but who had recently proclaimed himself Emperor of the French. My intention is that we celebrate the memory of a man whose thinking remains as alive, as vital, as the day he set pen to paper. Each day of our lives we see his vision made manifest reality in the vapour trails of the Big Jets high overhead. And to borrow such a dedication seems appropriate since, not only was Lanchester a great admirer of Beethoven’s music, but also Lanchester’s largeness of spirit, his grandness of vision, his capacity for unremitting sheer hard labour put him in the heroic mould.

May I thank the Society for the immense compliment it has paid me tonight by the award of the Society's Gold Medal. Earlier this year I was prevailed upon by Council to prepare the Wright Brothers Lecture. This is an occupational hazard of the Director, RAE, and has led to a situation—unique in my experience—in which the Society has presented the Medal and then within a few minutes asked the recipient to stand up and earn it. I am not at all optimistic about the outcome—but will do my best.

Active Control Technology (ACT) is not a new subject; it has been utilised in engineering circles for over a century in various guises, the first application of any significance being the steam engine governor. The history of the subject is accordingly long and is well referenced, and we shall depart from the time-honoured rigmarole of providing a detailed survey of early developments by directing our audience instead to Ref. 1 where such a survey may be found. In similar mildly unconventional vein, we note that this platform, and many others like it in this country and elsewhere, has been used many times by eminent individuals to expound their own idiosyncratic views as to how ACT should be defined in its constituent functions and how, in turn, these might possibly be applied without much in the way of reference to what has been done or to the many practical limitations. While our own views presented here will also be idiosyncratic, we shall attempt to relate what we say to that which has been achieved in actual applications and, based on these achievements, to what might follow in the foreseeable future.

Previous work by R. T. Jones on the drag minimisation of elliptic wings is extended to the case of the slewed wing with thickness. These results are used to calculate lift/drag ratios of idealised configurations related to a supersonic transport aircraft. The values of lift/drag ratio and optimum slenderness ratio found are comparable with those calculated earlier in studies of delta-like plan forms.

In the usual form of thin-shock-layer theory it is assumed that the flow about a lifting body can be expanded in terms of the inverse density ratio across a basic oblique shock wave lying in the plane of the leading edges of the body. In this paper it is shown that more accurate results can be obtained by moving the basic shock closer to the calculated shock wave below the body. The results obtained show why the original form of thin-shock-layer theory often gave good agreement with experiment in conditions which appeared to be outside the range of validity of the theory.

The purpose of this paper is to describe an approach which performs data fusion on the output of multiple, spatially separate, sensors engaged in the real time tracking of obstacles in a helicopter's environment. The generated information can be used either as a flight director aid or as feedback required by an automatic collision avoidance system. Obstacle track estimation has been commonly carried out using the Kalman filter (KF) for linear estimation, or the extended Kalman filter (EKF) for use on nonlinear problems. However, certain assumptions made in the derivation of the EKF algorithms render it sub-optimal for aerial obstacle track estimation. Additionally, the EKF has problems with initialisation and divergence (stability) for many non-linear processes.

Research at the University of Southampton has highlighted a link between fuzzy networks and associative memory neural networks. This link is important as it allows new learning rules to be developed for training fuzzy rules, and learning convergence to be proved. This paper explores methods of fusion of estimates using neuro-fuzzy models, and addresses some of the weakness of the Kalman filter approximation introduced by the assumptions made in its derivation.

This paper is the second of a series of three. The first dealt with the steady forced harmonic motion of a simple oscillator with one degree of freedom when “ hysteretic damping ” is present; by this it is meant that a damping force exists which is in phase with velocity but whose magnitude is proportional to displacement. The present paper is devoted to the formulation of a general theory of small hysteretically damped vibration in which the existence is postulated of damping forces which act between pairs of points such that they are in phase with the relative velocities and their magnitudes are proportional to relative displacements. The theory is presented in general terms and use is made of Lagrange’s equations with a new type of “ dissipation function ”; this is comparable with that of Rayleigh for viscous damping. It is shown that this theory generally leads to simpler algebra than does that of viscous damping. Moreover, it will be shown in a later paper that the nature of forced and damped oscillation may be thought of in terms of modes more simply than is the case with the more familiar viscous damping.

It is well known that the performance of the practical gas turbine cycle, in which compression and expansion are non-isentropic, is critically dependent upon the maximum temperature of the working fluid. In engines in which shaft-power is produced the thermal efficiency and the specific power output rise steadily as the turbine inlet temperature is increased. In jet engines, in which the gas turbine has so far found its greatest success, similar advantages of high temperature operation accrue, more particularly as aircraft speeds increase to utilise the higher resultant jet velocities. Even in high by-pass ratio engines, designed specifically to reduce jet efflux velocities for application to lower speed aircraft, overall engine performance responds very favourably to increased turbine inlet temperatures, in which, moreover, these more severe operating conditions apply continuously during flight, and not only at maximum power as with more conventional cycles.

A numerical technique for the stability analysis of linear mechanical dynamic systems with periodically varying parameters is proposed. The technique is based on representation of the solution vector in terms of Chebyshev polynomials defined over the principal period. Two formulations have been presented. The first formulation is suitable for systems described by state space equations, while the second can be applied directly to a set of second order equations with periodically varying mass, damping and stiffness matrices. As an illustrative example, the flap-lag stability of a multi-bladed rotor is examined. The numerical accuracy and efficiency of the proposed technique is compared with standard numerical codes based on Runge-Kutta, Adams-Moulton and Gear algorithms. The results indicate that the suggested approach is by far the most efficient one, particularly for systems with larger dimensions.

The force on a circular cylinder inclined to a uniform flow may be resolved into three components, one perpendicular to the plane defined by the axis of the cylinder and the upstream velocity vector (described in this report as the out-of-plane force) one within this plane and perpendicular to the axis (the in-plane normal force) and one within this plane and along the axis (the in-plane axial force).

It has become common practice, since Allen and Perkins, to model the flow over a circular cylinder inclined to a uniform flow by resolving the incident flow into two perpendicular component flows. One flow is taken to be parallel to the axis of the cylinder (axial flow) and the other normal to it (cross flow). The cross flow is not taken to be uniform over the whole length of the circular cylinder. Instead it is considered to be started impulsively at the nose and to develop over the length of the body in the same way that the impulsively-started flow over a cylinder develops with time.

The equations governing the flow in a boundary layer are a simplified form of the full Navier-Stokes equations; the simplification being based on the argument that within the boundary layer the velocity gradients normal to the solid surface are much larger than those parallel to the surface so long as the boundary-layer thickness is much smaller than (a) a representative streamwise length of the surface and (b) the principal radii of curvature of the surface.

Recent advances in the development of ground cushion vehicles have indicated the need for a more efficient means of providing lift at the higher speeds envisaged in the future. A wing operating in ground effect could provide this lift and might be far more attractive economically than the “Hovercraft”. This note investigates some of the stability aspects of such a “Ground Effect Wing” Vehicle, using linear analysis, by developing the relevant equations of motion and examining the characteristic equations using quasi-steady derivatives, at fixed heights, obtained from wind-tunnel tests. In addition comparison is made between an analogue simulation of a GEW and the flight paths obtained from a free-flight model. This investigation showed that the longitudinal mode of oscillation for both single and tandem-wings with endplates was unstable and artificial stability was necessary. The lateral mode of oscillation for a single wing was stable in both roll and yaw, the latter being subject to fin and endplate configurations. The effects of the force, and moment, rate of change of height derivatives were not considered, as values for these derivatives were not available. Some proposals for future work include the development of suitable aerofoil sections for operation near the ground and the investigation of slender bodies in ground effect.

This paper presents the results of experimental and computational investigations on the effect of slot cavities on a supersonic projectile. Experimental work was carried out to show the effects of the slots on the drag at Mach numbers M = 1·36, 1·65, 1·83. The computational analysis was done at M = 1·36. A single configuration of the slot pattern was used with two different slot widths (0·5mm and 2·0mm). Flow features were investigated in the slotted area and at the base. The analysis presented includes the pressure distribution, the supersonic cavity flow and the effect of the slots on the overall aerodynamic drag. Unlike the case of slots at transonic speeds, the suction and blowing mechanism is not found at supersonic Mach numbers. Streamwise cavity slots cause a small base drag reduction. The reduction in total drag is modest when the width is 0·5mm. However, the experiments showed that with the wider slots (2mm) the drag actually increases at Mach numbers from 1·36 to 1·83.