This paper presents a survey of published works on ducted fans for aeronautical applications. Early and recent experiments on full- or model-scale ducted fans are reviewed. Theoretical studies, lower-order simulations and high-fidelity CFD simulations are also summarised. Test matrices of several experimental and numerical studies are compiled and discussed. The paper closes with a summary of challenges for future ducted fan research.

A choked nozzle with an appropriate wall contour has adischarge coefficient, CD, so close to unity that a theoretical calculation of (I—CD) would allow the nozzle to be used as an absolute meter for air flow. The high discharge coefficient results basically from the fact that ∂(ρv)∂p=0 at M=1.

Simplified calculations yield formulae for the boundary layer displacement thickness and for the flow reduction resulting from the variation in static pressure across the throat. The optimum profile for the wall at the throat of an absolute meter is suggested to be a circular arc of radius of curvature equal to about twice the throat diameter. For such a meter the theoretical discharge coefficient is found to be within ¼ per cent of 0·995 over a wide range of Reynolds numbers.

The uncertainty in the discharge coefficient for a steady flow at Reynolds numbers of 106 and over appears to be less than ±0·15 per cent, both when the boundary layer is known to be entirely turbulent and when it is known to be entirely laminar. When the state of the boundary layer is not known the corresponding figure appears to be ±0·25 per cent. Experimental information might therefore be helpful on transition—under the appropriate conditions of flow unsteadiness and rig vibration. Available experimental results with known boundary layers tend to confirm the theoretical discharge coefficients down to a Reynolds number of 0·4x106.

A pressure ratio of about 1·1/1 or less would probably be sufficient to establish fully supersonic flow if the nozzle were followed by a suitable diffuser.

This paper deals with the effects of a superimposed radial outflow upon the motion of the fluid between a rotating and a stationary plane disc. This is an idealisation of the situation found in advanced turbo-machinery in which cooling of the rotor faces is necessary. The boundary layer approximations to the equations of fluid motion are integrated by an extension of the numerical technique proposed in Ref. 11. Predictions of drag torque and radial pressure distributions are compared with experimental results from a 30 in (762 mm) diameter system, using a range of speeds up to 4000 rev/min, while varying the axial spacing and rate of imposed radial flow. Qualitative agreement was satisfactory throughout the range, but quantitative discrepancies in the estimates of drag and in the pressure distributions at certain flow conditions suggest that the simple mixing length hypothesis used in this analysis for the turbulent shear terms in the equations of motion is not universally adequate.

The results of extensive measurements which are presented here show the presence of secondary flow in the laminar corner boundary layer which is opposite in rotation sense to that found in the turbulent corner flow. This secondary flow is associated with progressive changes of the form of the velocity profiles. These changes become marked prior to transition and may be linked with the transition process. The laminar corner flow separates readily under the influence of an adverse streamwise pressure gradient.

The elastic instability of a complete spherical shell under uniform external pressure is studied experimentally and theoretically. The premature snapping of a thin elastic shell, made of polyvinyl chloride, is seen to be classical in nature. The experimental maximum pressure and pre-snapping bending deformation are correlated with the theoretical behaviour of an initially imperfect shell. The large deflection behaviour of a perfect shell is assessed experimentally, and the stable post-buckling states are observed to be rotationally symmetric. A fairly precise theoretical analysis of these states is performed, the use of a high-speed digital computer allowing a considerable advance over previous treatments. The experimental and theoretical post-buckling curves are in good agreement, yielding the first detailed correlation of post-snapping equilibrium states in the field of shell instability.

The general picture of research in active flow control for aircraft applications has been continuously changing over the last 20 years. Researchers can now obtain design sensitivities by using numerical flow simulations, and new optical experimental methods can be used that measure flow field data non-intrusively in planes and volumes. These methodological advances enabled significant knowledge increase. The present paper reviews recent progress in active flow control by steady blowing. It appears that two strategies of blowing deserve particular attention. The first uses tangential blowing of thin wall jets to overcome the adverse pressure gradients from locally very large flow turning rates. This approach exploits the potentials of the Coanda effect. The second strategy employs oblique blowing of air jets designed to generate longitudinal vortices in the boundary layer. The longitudinal vortices provide convective redistribution of momentum in the boundary layer, and they also enhance turbulent momentum transport. The sensitivities of these two approaches as observed in fundamental flow investigations and in applications to high-lift aerofoils are described and suited efficiency parameters of blowing are analysed.

Separated and vortical flows are dominant over various unmanned air vehicles (UAVs). In this article, issues and challenges of vortical flows for future UAVs are reviewed. These include shear layer instabilities, vortex breakdown and wing stall, vortex interactions, nonslender vortices, multiple vortices, and manoeuvring wing vortices. There are also issues relating to vortical flows in certain flow/structure interactions, as well as in aerodynamics/propulsion interactions. Separated and vortical flows are even more dominant at low Reynolds number flows. The main features of vortical flows, unsteady aerodynamics, and propulsion related vortical flow isssues relevant to mini- and micro air vehicles, are discussed.

Based on earlier theories of mechanical, electrical and magnetic relaxation phenomena, the paper discusses a simple mathematical model representing qualitatively the basic characteristics of “internal” damping in solid materials. The model is assessed with respect to variation with frequency of stiffness and loss factor at simple harmonic oscillations, and is applied to oscillatory systems with discrete parameters and to continua.

A simple method is given for the calculation of the heat transfer from a laminar flow surface. Computation is by a quadrature. The method is essentially a simplification and extension of the Eckert(1) method, and is applicable both to two–dimensional and to axisymmetric flows.

Optical pressure measurement using fluorescent techniques is being widely researched for use in windtunnels. The technique offers global coverage of model surface pressures. A paint, the fluorescence of which is dependent on air pressure, is applied to the surface of a windtunnel model and the pressure distribution is obtained from the intensities measured. There are three principle techniques used to produce the surface pressure distributions: radiometric imaging, lifetime measurement and fluorescent lifetime imaging.

Radiometric imaging produces global images. A constant illumination is used to excite the paint. The fluorescent intensity distribution is dependent on the surface pressure. Error analysis reveals that the technique has the highest precision at low absolute pressures. However, wind-off reference images are required to compensate for illumination and paint non-uniformity. Experimental investigation found that although this is able to compensate for rigid structures, model movement introduces registration and geometric errors. These errors can be corrected by processing at the expense of accuracy.

Lifetime measurement is a technique that uses pulsed illumination. The fluorescent lifetime is dependent on the pressure. A wind-off reference image is not required and error analysis predicts that the technique holds an advantage at high absolute pressures.

Fluorescent lifetime imaging produces global images using a modulated light source to excite the paint. The phase and modulation depth of the fluorescent signal are proportional to the pressure. A wind-off reference image is not required and error analysis predicts that the technique has the same advantages as lifetime imaging. Preliminary results from fluorescent lifetime imaging indicate that this technique can be implemented using robust, solid-state technology.

The UK OST Technology Foresight for defence and aerospace identified multi-sensor data fusion as a future critical enabling technology for the UK, requiring a coordinated research agenda. This review paper provides an overview of past research and applications of data fusion. The process of data fusion and sensor integration is formally introduced together with a variety of implementation architectures, that recognise data fusion as a critical element in overall systems integration. The various benefits and attributes of data fusion are discussed together with a brief review of potentially fruitful areas of future research.

The purpose of the work was to investigate the effect of the finite width of plate on the elastic-stress distribution due to a central crack in a flat plate loaded in tension.

The range of investigation: (i) Two-dimensional photoelastic tests were carried out on flat-plate specimens. The photoelastic specimens were geometrically similar to the 10 in. wide fatigue specimens used by Frost and Dugdale, with the central crack simulated by a slot bounded by holes of small radii, (ii) A theoretical solution of the problem was derived and compared with the present and other photoelastic results.

It is shown that the effect of the crack-length/plate-width ratio on the elastic-stress concentration at the head of the crack can be expressed by the following formulae:

Considerable effort has been extended over the past years in adapting the matrix displacement method to the specific problems of thin shells under membrane and bending action and developing suitable elements of varying sophistication. Some of the difficulties arising in the process of idealisation were reviewed in ref. 1. For example, simple considerations show that a representation of a shell by polyhedron surfaces may lead to serious errors, especially in the presence of pronounced bending and so-called boundary-layer effects. For this and other reasons it appears imperative to allow for the curvature of the shell. Much ingenuity has been shown in evolving elements for shells of specific geometry.

The elementary engineering theory of longitudinal waves in beams allows for no dispersion, whereas Love's corrected version of it predicts that very short waves are not propagated. A new correction is added which overcomes these defects. To do this, a simple theorem is proved concerning certain types of closed region. A method of improving Love's formula and the present one is illustrated.

This paper begins with an outline of the procedure for predicting the infrared signature emissions from the airframe, engine casing, and the plume, and their attenuation by the intervening atmosphere. These emissions are contrasted against the background, to obtain the infrared signature levels. The infrared detector’s — noise equivalent flux density, is proposed as an operational constraint on the flight envelope. The shift of this newly imposed constraint on the flight envelope for several engine-operating conditions, and for turbojet and turbofan engines is studied. The signature levels from the casing and plume, of a turbofan and equivalent turbojet engine, are compared at different operating points on the flight envelope. Result in the form of a polar plot of infrared signature level variation with aspect is also examined for low flying missions. The results are analysed to direct stealth design and operation.

The design methodology and performance of Loughborough University’s new 1·9m × 1·3m, indraft wind tunnel is discussed in the following paper. To overcome severe spatial and financial constraints, a novel configuration was employed, with the inlet and exit placed adjacent to each other and opened to atmosphere. Using a fine filter mesh, honeycomb, two turbulence reduction screens and a contraction ratio of 7·3, flow uniformity in the working area of the jet at 40ms-1 is shown to be within 0·3% deviation from the mean velocity, with turbulence intensity in the region of 0·15%. Working section boundary layer characteristics are shown to be consistent with that of a turbulent boundary layer growing along a flat plate, which originates at the point of inflection of the contraction. A maximum velocity of 46ms-1 was achieved from a 140kW motor, compared to a prediction of 44ms-1, giving an energy ratio of 1·42. Comparison between theoretical and measured performance metrics indicate differences between the way modules perform when part of a wind tunnel system compared to data gathered from test rigs.

Dye flow visualisation of circular synthetic jets was carried out in laminar boundary layers developing over a flat plate at a range of actuator operating conditions and freestream velocities of 0·05 and 0·1ms–1. The purpose of this work was to study the interaction of synthetic jets with the boundary layer and the nature of vortical structures produced as a result of this interaction. The effects of Reynolds number (Re), velocity ratio (VR) and Strouhal number (St) on the behaviour of synthetic jets were studied. At low Re and VR, the vortical structures produced by synthetic jets appear as highly stretched hairpin vortices attached to the wall. At intermediate Re and VR, these structures roll up into vortex rings which experience a considerable amount of tilting and stretching as they enter the boundary layer. These vortex rings will eventually propagate outside the boundary layer hence the influence of the synthetic jets on the near wall flow will be confined in the near field of the jet exit. At high Re and VR, the vortex rings appear to experience a certain amount of tilting but no obvious stretching. They penetrate the edge of the boundary layer quickly, producing very limited impact on the near wall flow. Hence it is believed that the hairpin vortices produced at low Re and VR are likely to be the desirable structures for effective flow separation control. In this paper, a vortex model was also described to explain the mechanism of vortex tilting.

A survey of information on laminar separation bubbles and their effect on the stalling characteristics of aerofoils is presented. It is shown that there are two principal kinds of bubbles and that their existence can be predicted. It is believed that their behaviour may be related to the kind of boundary layer transition process present.

It is intended that a second part of this paper will describe the characteristics of aerofoils at Reynolds numbers typical of model aircraft and show how the observed phenomena are related to the behaviour of laminar separation bubbles.

This technical article discusses design and integration associated with distributed propulsion as a means of providing motive power with significantly reduced emissions and external noise for future aircraft concepts. The technical work reflects activities performed within a European Commission funded Framework 7 project entitled Distributed Propulsion and Ultra-high By-Pass Rotor Study at Aircraft Level, or, DisPURSAL. In this instance, the approach of distributed propulsion includes a Distributed Multiple-Fans Concept driven by a limited number of engine cores as well as one unique solution that integrates the fuselage with a single propulsor (dubbed Propulsive-Fuselage Concept) – both targeting entry-in-service year 2035+. Compared to a state-of-the-art, year 2000 reference aircraft, designs with tighter coupling between airframe aerodynamics and motive power system performance for medium-to-long-range operations indicated potentially a 40-45% reduction in CO2-emissions. An evolutionary, year 2035, conventional morphology gas-turbine aircraft was predicted to be –33% in CO2-emissions.