Browsing by Author "Lilley, G. M."
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Item Open Access An approximate solution of the turbulent boundary layer equations in incompressible and compressible flow(College of Aeronautics, 1960-07) Lilley, G. M.If over the ‘outer’ region/ of the boundary layer, where the mean velocity varies but little from its value outside the shear layer, a virtual eddy viscosity is defined, which is constant over the outer region but varies in the direction of the mainstream, a solution of the turbulent boundary layer equations can be found which satisfies the appropriate boundary conditions. The solution leads to a compatibility condition for the virtual eddy viscosity in terms of the wall shear stress, the boundary layer momentum thickness and the mainstream velocity, at least for the case of a constant external velocity. [...cont.]Item Open Access Experimental investigation of the interference of a body on a low aspect ratio wing of rectangular planform at a Mach number of 2(College of Aeronautics, 1955-06) Busing, J. R.; Marson, G. B.; Lilley, G. M.Results are given of pressure measurements on awing of gross aspect ratio approximately 2/3 mounted on a cylindrical body with an ogival nose, made at a Mach number of 2.00 in the 9in, x 9in, supersonic -wind tunnel at the College of Aeronautics. The wing section was a single wedge having a 6° total nose angle. The ranges of body incidence and roll in these tests were from 0° to 30°, and 0°, 30°, 600 and 90° respectively. The normal force, lift, drag and side force coefficients, and rolling and pitching moments were obtained from the pressure measurements.Item Open Access Experiments on an induction type high speed wind tunnel driven by low pressure steam(College of Aeronautics, Cranfield, 1949-03) Lilley, G. M.; Holder, D. W.The performance of an induction type high speed wind tunnel driven by low pressure steam (up to 120lb. per sq.in. absolute) has been investigated up to a Mach number of about 1.7. It was found that by suitable design a range of Mach numbers could be attained over a wide range of supply pressures and steam quantities. Comparison with previous experiments in which compressed air was used to drive the tunnel show that the required steam and air pressure and quantities are comparable. These results imply that in many cases existing boiler plants can readily be adapted to drive high speed tunnels of useful dimensions. Continues …Item Open Access Ground level disturbance from large aircraft flying at supersonic speeds(College of Aeronautics, 1960-05) Lilley, G. M.; Spillman, J. J.The Whitham-Walkden theory for the estimation of the strength of shock waves at ground level from aircraft flying at supersonic speeds is applied to the case of a typical projected supersonic civil transport aeroplane. If a figure of 2 lb/sq.ft. (including a factor of 2 for ground reflection) is taken as an upper limit for the acceptable strength of the bow wave from such an aircraft it is shown that restrictions on the climb and flight plan will be involved. The advantage of the employment of larger engines with or without afterburning is discussed, with reference also to the penalties involved owing to the increase in weight of the aircraft and its direct operating costs. Finally it is suggested that an aircraft of given volume could be designed, by suitable choice of thickness and lift distribution, to minimise the strength of the shock waves in the far field.Item Open Access An investigation of the flexure-torsion flutter characteristics of aerofoils in cascade.(College of Aeronautics, Cranfield, 1952-05) Lilley, G. M.Part 1 of this report describes the results obtained from a series of tests on the flexure-torsion flutter characteristics of cascades of similar aerofoils having symmetrical sections. Continues … Part 2 of this report is a review of the theoretical studies on oscillating aerofoils in cascade. Continues …Item Open Access An investigation of the flow in an axial compressor inlet casing and its effect on the performance of the first stage of the axial compressor(College of Aeronautics, 19) Lilley, G. M.Summary The present report is a continuation of an interim report dated 8th August, 1951. The latter- report dealt with the flow in a model casing whereas the present reports deals with a similar investigation on the ‘Reavell’ compressor inlet casing which is to be part of the 9in x 9in supersonic wind tunnel installation, In the basic condition the flow had a circumferential instability* caused by the splitting of the inlet flow around the central shaft fairing. In addition the main flow through the casing was under-turned by approximately 200, Consequently the flow at the exit annulas was unsteady and changes of 300 in the exit flow direction were found. It was also found that at least four guide vanes were stalled in any one flow configuration. The instability was prevented by a flat dividing plate placed behind the central shaft fairing, but smooth flow at exit, in the vicinity of the plate, was only possible when curved fairings were added to both sides of the plate. A gauze screen, of 30 wires per inch, placed at entry, partly reduced the unsteadiness at exit but it had little effect on the flow direction and did not unstall the guide vanes. A perforated plate, having 15. inch diameter holes and 50 per cent blockage, similarly placed, had little or no effect. The guide vanes were unstalled by adding curved extensions to their leading edges. The large changes in the flow direction across the annulus at exit were, however, only reduced by modifying the internal contour of the casing. With these modifications flow angle variations of between 5° and 10° were obtained at exit. The velocity at exit was not uniform over the exit passages between the guide vanes, but the variations, which were not unduly large, were caused by secondary flows and the wakes from the vanes. This investigation has clearly shown that in the basic condition the unsteady, non-uniform flow at the exit from the inlet casing, which is immediately upstream of the first stage of the compressor, will tend to cause periodic stalling of the rotating blades in that stage. Since the relative velocities are high the danger of running into stn3ling flutter cannot be overlooked, especially as the cascade effect may reduce the critical flutter speed of the isolated blade. Stalling flutter, which usually takes place in the fundamental torsional mode, may be aggravated by a high order resonance occuring at the given rotational speed of the compressor. The suggested' modifications to the design of the inlet casing should prevent stalling of the first stage blades 'of the compressor caused by upstream, unsteady, non-uniform flow.Item Open Access An investigation of the noise field from a small jet and methods for its reduction.(College of Aeronautics, Cranfield., 1952-01) Westley, R.; Lilley, G. M.Sound measurements have been made on the noise from the jet of a one inch diameter convergent nozzle at atmospheric temperature and at speeds above and below choking. The noise level and spectrum have been investigated in both the near and distant fields. The results agree in some measure with the predictions of the Lighthill theory, that the elementary sound radiator is an acoustic quadrupole. The agreement is more marked if attention is confined to the higher frequencies. Simple empirical formulae are derived giving the overall sound intensity and frequency spectrum in terms of the position relative to the jet, the stagnation pressure excess over the atmospheric pressure, and the frequency. The results of tests on various noise reduction devices are discussed. These tests indicate promising lines of investigation. The maximum reduction in total noise level was about 10 db.Item Open Access A note on the decay of aircraft trailing vortices(College of Aeronautics, 1964-03) Lilley, G. M.An elementary theory of aircraft trailing vortex decay is presented based on an assumed law for the variation of the mean eddy viscosity with distance from the wing. This law is based on the experimental data of Rose and Dee (1.963). The analysis gives results, as might be expected, in agreement with their data. The justification for the analysis must however be in doubt until more data are available covering a wide range of variables such as aircraft size, distance, incidence, etc.Item Open Access On some aspects of the noise propagation from supersonic aircraft(College of Aeronautics, Cranfield, 1953-02) Lilley, G. M.; Westley, R.; Yates, A. H.; Busing, J. R.The noise problem associated with an aircraft flying at supersonic speeds is shown to depend primarily on the shock wave pattern formed by the aircraft. The noise intensity received by a ground observer from a supersonic aircraft flying at high as well as low altitudes, is shown to be high although it is of a transient nature. Continues…Item Open Access On surface pressure fluctuations in turbulent boundary layers(College of Aeronautics, 1960-04) Lilley, G. M.Existing work on the pressure fluctuations in turbulent shear flaws is briefly reviewed with special reference to the problem of wall turbulence. An approximate theory for the pressure fluctuations on the wall under both a turbulent boundary layer and a wall jet is given and indicates in the latter case an intensity many times that corresponding to the flow over a flat plate at zero pressure gradient, as typified by measurements on the wall of a wind tunnel. Experiments on a wall jet confirm these predictions and details of the few preliminary data are presented. The results from the wall jet suggest that the intensity of the pressure fluctuations in the regions of adverse pressure gradient, on wings and bodies approaching and beyond separation will be higher than in regions of zero pressure gradient. Appendices are included which deal with the necessary extensions to the analysis to fit the velocity correlation functions as measured by Grant (1958), the effects of time delay and eddy convection.Item Open Access A preliminary investigation of the flow over a particular wing body combination at Mach number 2(College of Aeronautics, 1955-04) Busing, J. R.; Lilley, G. M.Results are presented of visual tests, by surface flow and Schlieren techniques, of the flaw over a particular wing body combination at M = 2.0, made in the 9in x 9in supersonic wind tunnel at the College of Aeronautics. It should be noted that the photographs in this report represent only a small part of all the data collected. The effects of providing roughened bands and air jets for transition fixation on the body are discussed. It is shown that although such methods may be satisfactory at zero or low angles of incidence they do not fix transition uniformly over the complete body circumference at higher incidences. For this reason the main part of the tests have been performed with free transition. The formation of vortex sheets on both wings and body is discussed for a range of incidence and roll angles and some conclusions are drawn as to the shape of the vortex pattern and its induced flow effects.Item Open Access A preliminary report on the design and performance of ducted windmills(College of Aeronautics, Cranfield, 1956-04) Lilley, G. M.; Rainbird, W. J.A preliminary study is made of the theoretical gain in power output obtained with a fully ducted land-land type windmill is compared with the standard unshrouded type windmill. ...Item Open Access Pressure fluctuations in an incompressible turbulent boundary layer(College of Aeronautics, 1960-06) Lilley, G. M.Item Open Access The response time of wind tunnel pressure measuring systems(College of Aeronautics, 1960-11) Lilley, G. M.; Morton, D.The time response of a wind tunnel pressure measuring system, comprising a pressure transducer of fixed volume and a length of capillary tubing, is analysed and the results compared with experiments. It is shown that the approximate analysis of Kendall (1958), in which the friction losses at any given time are assumed equal to the steady state losses, has a wide range of validity, provided the L/R ratio for the capillary tube is large and the inlet and exit losses are included as equivalent lengths of the capillary tube.Item Open Access A review of pressure fluctuations in turbulent boundary layers at subsonic and supersonic speeds(College of Aeronautics, 1963) Lilley, G. M.The equation for the pressure fluctuations in a turbulent boundary layer are derived with special reference to their values at a rigid wall. It is shown that in incompressible flow the pressure field is defined completely once the velocity field is known. The results obtained from the theory are compared with experiment. The work is extended to include the effects of compressibility and it is found that a treatment similar to that given by Phillips is appropriate. It is shown that under conditions of zero heat transfer the theory slightly modified by the effects Mach numbers eddy Mach waves ace pressure distribution throughout obtained in incompressible flow is only of compressibility. However, at higher formed and these modify to some extent the the layer, and in particular at the wall. In addition strong radiation of sound occurs under conditions of flow external to the boundary layer. supersonic Comparisons with experiment show moderate agreement. The theories as mentioned previously are for the case of zero external pressure gradient, but it is shown that even when these conditions are relaxed similar results occur. Some experimental results in support of this conclusion are presented.Item Open Access A simplified theory of skin friction and heat transfer for a compressible laminar boundary layer(College of Aeronautics, 1959-01) Lilley, G. M.The compressible laminar boundary layer equations for a perfect gas in steady flow at arbitrary external Mach number and wall temperature distribution are solved approximately by the combined use of the Stewartson- Illingworth transformation and application of Lighthill's method to yield the shin friction and rate of heat transfer. Appendices are added which give the necessary modifications to the method for the separate cases of very low Prandtl number and for the flow near a separation point. A further appendix describes Spalding's method for improving the accuracy of the wall value of shear stress and rate of heat transfer distributions along a wall having a non-uniform temperature distribution.Item Open Access Some notes on the factors influencing the possible damage to buildings due to the passage of shock waves from aircraft flying at supersonic speeds(College of Aeronautics, 1952) Lilley, G. M.The purpose of this note is to discuss, from the limited information available, the order of magnitude of the strength and the configuration of the shock waves, close to ground level, generated from aircraft flying et supersonic speeds. The level flight case is the only one considered in detail but brief reference is also made to other manoeuvres such as the supersonic dive of limited duration. The alterations in path and strength of a shock wave in passing through the heterogeneous atmosphere are shown to be very important. For normal temperature variation with altitude the attenuation of shock strength, with distance downward from the aircraft, will be decreased. A similar effect will occur when shock waves are propagated against a wind, whose velocity increases with height.. These results should be of interest to civil engineers and others making estimates of the possible damage to buildings due to the above cause. A preliminary, mainly qualitative discussion of this part of the problem is included, which although not complete should at least give some idea of the damage to be expected. In this connection it is found that for aircraft flying at and above 5000 ft. at Each numbers up to 2 the damage to buildings should be limited to a number of ‘freak’ cases of window breakage and similar minor damage. In the case of aircraft diving at supersonic speeds at a high altitude the excess pressures caused by the resulting shock waves and their time duration are so small that the chance of any damage occurring is almost negligible.Item Open Access The use of a potential flow tank for testing axi-symmetric contraction shapes suitable for wind tunnels(College of Aeronautics, Cranfield., 1951-04) Babister, A. W.; Marshall, W. S. D.; Lilley, G. M.; Sills, E. C.; Deards, S. R.The report gives details of tests in the potential flow tank on a series of axi-symmetric contraction shapes. The tests were in connection with the design of the 8ft x 6ft wind tunnel and a water tunnel. Continues...Item Open Access Wall pressure fluctuations under turbulent boundary layers at subsonic and supersonic speeds(College of Aeronautics, 1963-03) Lilley, G. M.The problem of pressure fluctuations at a rigid wall under a turbulent boundary layer has attracted much attention in the past decade. At low Mach numbers the theory is well established from the work of Kraichnan and Lilley, and reasonable agreement is obtained with the experiments of Willmarth, Hodgson and others. At high Mach numbers, measurements exist due to the work of Kistler and Chen but so far no theory is available, apart from that due to Phillips, which is however related to the noise radiated from supersonic turbulent shear flows. The present paper reviews the theory of wall pressure fluctuations in incompressible flow, and shows how the character of the pressure fluctuations changes in passing from the flow to the wall. Attention is drawn to the more important interactions giving rise to the pressure fluctuations, as well as to the region of the boundary layer mainly responsible for the wall pressure fluctuations … [cont.].