Experimental results obtained in jets from notched nozzles in a co-flowing free stream are presented. It is confirmed that even at low jet speeds the notches produce trailing vortices and it is found that the effect on the free stream on the vortices is surprisingly small.

A review is made of recent research in the United States relating to runway slipperiness and to slush-drag or fluid-drag effects on aircraft ground performance. The formation of fluid spray and its drag effects on aircraft are traced from spray origin at the intersection of the tyre and runway to impingement on the aircraft over the complete aircraft ground-velocity range. Six manifestations of tyre hydroplaning are presented and discussed: fluid-drag peak, tyre spin-down, loss of aircraft directional stability, loss of aircraft braking traction, suppression of bow wave, and progressive detachment of the tyre footprint from the fluid-covered runway surface as ground speed is increased. A convenient formula for estimating tyre-hydroplaning speed is indicated to be

where Vp=hydroplaning speed, knots

p=tyre-inflation pressure, lb/sq in.

Comparison with available experimental data is shown. The effects of fluid viscosity and density on runway slipperiness are discussed. The effects of some runway-surface textures, tyre-tread patterns, and aircraft landing-gear wheel arrangements on slipperiness are described. The paper also includes indications of areas of interest that need further study.

When accepting the invitation to present the Seventh Louis Blèriot Memorial lecture, the author was deeply conscious not only of the honour which was bestowed upon him by this invitation but, in an equal measure, of the responsibility which attached to it.

For the first time one of the eminent memorial lectures is devoted to the rotating wing; the helicopter has been raised to the level of discussion of the great aeronautical papers.

The natural impulse of the gentler sex when faced with a great occasion is invariably: “how shall one dress?” To the author, a somewhat similar problem became manifest: how can the helicopter, this new-comer, best be “ dressed?” How should it be presented, so as to make not only a good impression but also a full and, above all, a true impression?

Before entering upon the subject of my lecture I beg to express my feelings of gratitude for the honour and appreciation of my efforts, implied in the invitation to speak on metal aeroplane construction before so eminent and learned a body as the Royal Aeronautical Society.

In consideration of the general political situation I deem myself entitled to attribute to this invitation the deeper meaning of a token of amiable disposition given apart and beyond my person, from nation to nation, and to see in it an effort of renewing the ties of a genuine humanity, which is well aware of the unavoidableness of conflict and fight, but does not know hatred, and desires to extinguish the sad traces of a devastating war by hoisting the flag of peaceful competition.

Closed-form equations for the evaluation of forces on a body in compressible, inviscid and viscous flows are presented. The expressions which are used to calculate lift and drag require only knowledge of the velocity field (and its derivatives) on an arbitrarily chosen contour enclosing the body. Various forms of the force equations are presented, both for two-dimensional and three-dimensional flows. The work is an extension of earlier research by Noca) developed for incompressible flows. The equations are implemented on a number of two-dimensional aerofoil cases and the results show excellent agreement with surface and wake integration approaches.

The results are presented of an investigation of a particular type of baffle for the production of symmetrical velocity profiles having high ratios of maximum to mean velocity in ducted incompressible flow. Two similar families of profiles are obtained depending on whether a short (12 diameters) or a long (48 diameters) entry length is used before the baffle. The highest value of the maximum to mean velocity ratio obtained is 1·42 and the pressure loss coefficients associated with the use of the baffle are given together with an indication of the effect of Reynolds number.

A novel automatic grid generation method for complex CFD applications has been developed, driven by the need of the industrial CFD practitioner to reduce the turnaround time from an initial CAD geometry to obtaining a flow simulation. The method uses a subdivision strategy, allowing hanging nodes in a Cartesian unstructured mesh environment and is combined with a non-overlapping, semi-structured surface grid. Unusually, the surface grid is the end rather than the starting point of the present approach. This reversal of gridding strategy allows usability to be enhanced, in that the grid generation is reduced to one automatic step. The capabilities of the method are demonstrated by reference to selected inviscid and viscous flow solutions using the Fluent code Rampant v4.0 as a flow solver.

Some in-service deterioration in any mechanical device, such as an aero-engine, is inevitable. As a result of experiencing a deterioration of efficiency and/or mass flow, an aero-engine will automatically adjust to a different set of operating characteristics; thereby frequently resulting in changes of rpm and/or turbine entry temperature in order to provide the same thrust. Rises in the turbine entry-temperatures and the high-pressure turbine’s rotational speed result in greater rates of creep and fatigue damage being incurred by the hot-end components and thereby higher engine’s life cycle costs. Possessing a better knowledge of the effects of engine deterioration upon the aircraft’s performance, as well as fuel and life usages, helps the users to take wiser management decisions and hence achieve improved engine utilisation. For a military aircraft, using a computer performance simulation, the consequences of low-pressure (LP) compressor’s deterioration upon an aero-engine high-pressure (HP) turbine blade’s life-consumption have been predicted.

Shock-tunnel experiments are carried out to study the strong interaction between an impinging shock wave and boundary layer on a flat plate, accompanied by large separation bubble with a length comparable to the distance of the location of shock impingement from the leading edge of the plate. For nominal freestream Mach numbers ranging from 6 to 8.5, moderate to high total enthalpies of 1.3MJ/kg to 6MJ/kg are simulated in the Indian Institute of Science's hypersonic shock tunnels HST-2 (a conventional Hypersonic Shock Tunnel) and Free Piston Shock Tunnel (FPST) with freestream Reynolds numbers ranging from 4 × 106/m to 0.3 × 106/m. The strong impinging shock is generated by a wedge (or shock generator) at an angle of 30.96° to the freestream. From the time-resolved Schlieren flow visualisations using a high-speed camera and surface pressure measurements on the flat plate using fast response sensors, a statistically steady flow field with a large separation bubble was established within the short test time of the shock tunnels (around 600µs in HST-2 and 300µs in FPST). The role of various parameters on the interaction – Mach number, location of shock impingement and flow total enthalpy – are investigated from the measured separation length and surface pressure distribution. For the nominal Mach number of 8.5, with shock impingement at 100mm from the leading edge, the separation length increased from 60mm to 70mm as the total enthalpy is increased from 1.6MJ/kg to 2.4MJ/kg; whereas it dropped drastically to 30-40mm at 6MJ/kg. This is due to the prominence of real gas effects at higher enthalpies.

A stratospheric airship flies at a working altitude of 20km when it takes off from the ground. During ascent and descent, the wind field and thermal environment are highly complex. The thermal environment affects altitude, whereas wind influences the horizontal position of the airship. At a low altitude, this horizontal position cannot be controlled by thrusts given the low thrust-to-weight ratio, especially under a large wind field. However, it may be controlled indirectly by the pitch angle during ascent and descent with a certain vertical velocity. This study therefore proposes ascending and descending schemes for a stratospheric airship based on the thermal model. In this model, altitude is determined by the net lift/weight, whereas the horizontal position is controlled by the thrust and pitch. The pitch angle is determined by ballonets and an elevator. To allocate pitch control between the ballonets and the elevator under different airspeeds, pseudo-inverse dynamics of varied weight are introduced. In horizontal position control, the method of chain allocation is then applied between a pitch angle and vectored thrust to control the position of a stratospheric airship during ascent/descent.

The insufficient depth of modelling to capture the flow physics within primary combustion zone is the prime reason behind limited accuracy of semi-empirical correlations. Flame volume concept establishes a better connection between LBO performance and flame parameters, which improves the modelling depth and hence the prediction accuracy. Nonetheless, estimation of flame parameters is a challenging task. In addition, the iterative loop to approach convergence for a single geometry demands several numerical simulation runs. In this study, the association of LBO performance has been extended to flow structures, they are uniquely associated with the geometric features and can efficiently relate global LBO performance with primary zone geometry. The lean blowout phenomenon was presented as a contest between igniting and extinction forces within Reverse Flow Zone. These forces were quantified by four performance parameters including area, minimum axial velocity, average temperature, and average velocity. Selected parameters provide valuable information regarding the size of recirculation bubble, the intensity of flow reversal and the amount of entrained hot gases. For the purpose of validation, 11 combustor geometries were selected. The RANS simulation was carried out to estimate performance parameters, and predicted performance was compared against experimental data. The excellent agreement highlights the efficiency and promising future for the proposed methodology. Moreover, the association of prediction process with flow structure, instead of geometric features/dimension, makes it universal prediction methodology for wide range of combustor configurations.

In the present study, fuzzy uncertainty and reliability analysis of aeroelastic aircraft wings are investigated. The uncertain air speed and structural parameters are represented by fuzzy triangular membership functions. These uncertainties are propagated through the wing model using a fuzzy interval approach, and the uncertain flutter speed is obtained as a fuzzy variable. Further, the reliability of the wing flutter is based on the interference area in the pyramid shape defined by the fuzzy flutter speed and air speed. The ratio between the safe region volume and the total volume of the pyramid gives the reliability value. Two different examples are considered—a typical wing section, and a clean wing—and the results are given for various wind speed conditions. The results show that the approach considered is a low-cost but suitable method to estimate the reliability of the wing flutter speed in the presence of uncertainties.

Dry wind-tunnel (DWT) flutter test systems model the unsteady distributed aerodynamic force using various electromagnetic exciters. They can be used to test the aeroelastic and aeroservoelastic stability of smart aircraft or high-speed flight vehicles. A new parameterised modelling method at the full system level based on the generalised force equivalence for DWT flutter systems is proposed herein. The full system model includes the structural dynamic model, electromechanical coupling model and fast aerodynamic computation model. An optimisation search method is applied to determine the best locations for measurement and excitation by introducing Fisher’s information matrix. The feasibility and accuracy of the proposed system-level numerical DWT modelling method have been validated for a plate aeroelastic model with four exciters/transducers. The effects of key parameters including the number of exciters, the control time delay, the noise interference and the electrical parameters of the electromagnetic exciter model have also been investigated. The numerical and experimental results indicate that the proposed modelling method achieves good accuracy (with deviations of less than 1.5% from simulations and 4.5% from experimental test results for the flutter speed) and robust performance even in uncertain environments with a 10% noise level.

This note considers two cases of bending beyond the limit of proportionality which have not received a great measure of attention in the past. These are Case (i). A beam subjected to a pure bending moment acting in a plane other than one of symmetry. Case (ii). A beam subjected to a bending moment and an axial load.

A mathematical form is used for the stress–strain curve in order that the results shall be applicable to a large number of materials and to avoid the tedious arithmetical summations which would be necessary if actual stress-strain curves were used. All the results are presented in terms of form factors since this notation is convenient for design purposes and is consistent with current British practice. Although the method adopted in the analysis is applicable to a number of different types of beam cross section, only rectangular cross sections are considered in detail at the present time.

Curves are presented showing the variation of the form factor with respect to the angle of the plane of loading, for Case (i), and with respect to the ratio of bending moment to axial load for Case (ii). These curves are based upon a standardised form of the stress–strain curve which may be taken as representative of a range of aircraft materials. Consideration is given to the automatic satisfaction of present military proof loading requirements when these form factors are used for the estimation of maximum permissible bending moments and axial loads under ultimate loading conditions.

The recommendations of F. P. Cozzone, for dealing with these two problems are examined and comparison is made between the theoretical work of this note and a limited number of test results available for Case (i).

Exactly what constitutes a design criterion is a matter on which there may well be differences of opinion. For practical purposes design criteria may be taken to be empirical numerical rules to aid the designer in forming his initial conception of a design. The formulation of such rules for aircraft fatigue is still a matter of some difficulty. Not only is the subject of aircraft structural fatigue still in its infancy, but its study has to be linked with systematic experiment in a somewhat unusual way.

In the criteria described in this paper, the structure has first to be subjected to a standardised laboratory test. This test is itself a gross simplification since it has to represent in a rough way a diversity of conditions occurring in flight.

There is considerable lack of information regarding shear stresses in tapered beams. In this brief report these stresses are examined by a simple extension of the conventional engineering method for uniform beams, and where possible the results are compared with those of classical theory.

Complete agreement is obtained with the few existing classical solutions. The use of the engineering method is illustrated by its application to various types of beam, and curves are drawn illustrating the general nature of the influence of taper on shear stresses.

This paper reviews work on the flexural centre of elastic cantilever beams and contains a number of hitherto unpublished results, including a formula giving the position of the flexural centre in terms of the Prandtl torsional stress function.

The appearance of the note by Jacobs has prompted the preparation of a review of this subject which is all the more desirable as several of the investigations made shortly before the 1939-45 War were never published and others, although published, seem to be in danger of being overlooked. The present paper contains a number of hitherto unpublished results. The question of nomenclature is worthy of mention since general agreement is lacking. Some of the names used are flexural centre, centre of flexure, elastic centrum and centre of shear but a complete search of the literature would probably lead to the discovery of yet other names. The terms “centre of shear” and “elastic centrum” seem vague and fail to indicate that the point referred to has any special relation to flexure. We shall here use the name “flexural centre” exclusively.

The Dynamic Relaxation solution of the Laplace equation introduces dynamic terms into the basic equation. When this is written as an explicit finite difference formulation it can be solved by an iterative process which only requires a simple substitution routine. The method is easy to programme and requires small storage in the computer. By studying problems involving wind tunnel interference in steady flow, the potentialities of the method are demonstrated.

The dynamic inflow model is a powerful tool for predicting the induced velocity distribution over a rotor disc. On account of its closed form and simplicity, the model is especially practical for studying flight mechanics or for designing control systems for helicopters. Scant attention has, however, been paid so far in utilising the dynamic inflow model to analyse an autorotating rotor, which is different from a powered rotor in the geometric relation between the direction of the inflow and the rotor disc. Autorotation is an abnormal condition for helicopters, but for gyroplanes it is the normal mode of operation. Therefore the theoretical discussion on an autorotating rotor is of importance not only to improve the understanding of present gyroplanes, but also in the development of new gyroplanes and to analyse the windmill-brake state of helicopters. Dynamic inflow modelling is reviewed from first principles, and this identifies a modification to the mass flow parameter. A qualitative assessment of this change indicates that it is likely to have a negligible impact on the trim state of rotorcraft in autorotation, but a significant effect on the dynamic inflow modes in certain flight conditions. This is confirmed by numerical simulation, although considerable differences only become apparent for steep descents with low forward speed. It is concluded that while modification of the mass flow parameter is perhaps mathematically accurate, for practical purposes it is required only in a limited area of the flight envelope of autorotating rotorcraft.

Among various aviation meteorological conditions, heavy rain is an important one that may seriously affect aircraft flight safety. Over the past decades, appreciable efforts have been made to study the impacts of heavy rain on aircraft flight performance. Although there has been a consistent conclusion that heavy rain can cause great static aerodynamic performance degradation, such as lift decrease and drag increase, little has been known on the effects of heavy rain on aircraft dynamic flight performance. This article explores the static and dynamic aerodynamic performance of an approximated model of the DLR-F12 transport aircraft in simulated heavy rain environment. A novel synthesised approach is proposed to study the stability dynamic derivatives in a heavy rain condition. The results suggest that heavy rain not only causes more fuel consumption to compensate the lost lift performance but also induces great dynamic flight performance degradations, especially the short-period mode performance, thus seriously threatens aircraft flight safety.