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.

A method for the analysis of flapping-wing flight using lifting-line theory and actuator disc theory is proposed for the prediction of aerodynamic loads, propulsive efficiencies and optimum lift distributions.

The dynamics of the wing tip vortex in the near-field of a NACA 0012 aerofoil has been analysed by means of flow visualisations in a water tunnel. Different axial distances near the wing up to four chords, Reynolds numbers up to 42,000 and three angles-of-attack are studied to characterise the behaviour of the vortex meandering. The spatio-temporal vortex centre positions show distorted elliptical shapes in a (x,y)-plane. The Reynolds number has no significant influence on the axial evolution of the meandering amplitude. In addition, the flow visualisations obtained with a low speed camera are analysed by the singular value or proper orthogonal decomposition. Thus, the most energetic displacement modes are obtained. The frequency associated to these modes is computed by FFT. In all the cases studied, our results show that the most unstable mode corresponds to the azimuthal wavenumber |n| = 1 in the so-called Kelvin helical modes and the frequency is lower or close to 1Hz.

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.

Analysis and testing have been used to develop a flapping wing which can efficiently produce lift and thrust for ornithopter flight. With the assumption that the only imposed motion is dihedral flapping at its root, the wing's spanwise distribution of pitching and bending is due to its aeroelastic response. Numerical studies revealed criteria for efficient designs, which then guided the development and construction of an example wing tested in a windtunnel. The experimental results for average thrust and lift compared very favourably with the theoretical predictions, confirming that such aeroelastically tailored designs provide a means for achieving mechanical flapping-wing flight.

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:

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 tendency in the past has been to assume that when wakes or non-uniform total head profiles are fed into an axial compressor then substantially constant static pressure prevails at the entry, the variations in total head appearing as variations in velocity. This variation in velocity causes variation in incidence on the early stage blade rows and thus can give rise to excitation of blade vibration. This assumption is implicit, for instance, in References 1 and 2, but we think has been a common assumption by most of the people working in this field.

Where the compressor is fed by a duct of substantially parallel walls for a reasonable length ahead, such an assumption appeared justifiable. Such a duct when given an air flow test with its outlet discharging, for instance, to atmosphere instead of to the compressor, then the distribution assumed would normally be obtained and in fact many surveys of such ducts have been represented in this fashion. The object of this note is to show that, in fact, this distribution will not normally occur when the compressor is present and we may normally expect much more nearly a constant velocity into the compressor with attendant static pressure distributions to match with the total head variations ahead of the intake, with of course, the attendant curved flow to support the static pressure gradients.

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.

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.

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.

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.

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.

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.

The buckling of simply-supported sandwich panels, made up from fibre reinforced laminated face plates and honeycomb cores, is considered when they are subjected to uniform end loading. Numerical results are given for both overall buckling and face plate wrinkling for panels having typical carbon fibre faces with the laminations arranged to give effectively orthotropic properties.

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.

One of the early triumphs of the Lanchester-Prandtl wing theory was that it explained the effects of wind tunnel boundaries and provided explicit and simple formulae to correct wind tunnel measurements for these effects. As every aerodynamicist knows, this theory tells us, with the help of “images”, that the boundaries, in the presence of a lifting wing, produce extraneous velocity components in the flow. At least two generations of experimenters have routinely corrected their angles of attack and their measured drag values for the presence of the vertical component.

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.

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.