Browsing by Author "Stollery, J. L."
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Item Open Access A case study on the aerodynamic heating of a hypersonic vehicle(Royal Aeronautical Society, 2012-09-30T00:00:00Z) Mifsud, Michael; Estruch-Samper, David; MacManus, David G.; Chaplin, Ross; Stollery, J. L.A Parabolised Navier-Stokes (PNS) flow solver is used to predict the aerodynamic heating on the surface of a hypersonic vehicle. This case study highlights some of the main heat flux sensitivies to various conditions for a full-scale vehicle and illustrates the use of different complimentary methods in assessing the heat load for a realistic application. Different flight phases of the vehicle are considered, with freestream conditions from Mach 4 to Mach 8 across a range of altitudes. Both laminar and turbulent flows are studied, together with the effect of the isothermal wall temperature, boundary-layer transition location and body incidence. The effect of the Spalart-Allmaras and Baldwin-Lomax turbulent models on the heat transfer distributions is assessed. A rigorous assessment of the computations is conducted through both iterative and grid convergence studies and a supporting experimental investigation is performed on a 1/20th scale model of the vehicle's forebody for the validation of the numerical results. Good agreement is found between the PNS predictions, measurements and empirical methods for the vehicle forebody. The present PNS approach is shown to provide useful predictions of the heat transfer over the axisymmetric vehicle body. A highly complex flow field is predicted in the fin-body-fin region at the rear of the vehicle characterised by strong interference effects which limit the predictions over this region to a predominately qualitative level.Item Open Access The effects of bluntness and sweep on glancing shock wave turbulent boundary layer interaction(Cranfield University, 1986-09) Fomison, N. R.; Stollery, J. L.An experimental investigation has been performed into the effects of leading edge bluntness and sweep on the mean flow characteristics of a glancing shock wave turbulent boundary layer interaction. A series of blunt un swept shock generators (ranging in leading edge diameter from O to 1.0 ins) and a series of sharp swept generators (covering angles of sweep from 0 to 75°) were tested at incidences up to 30° at a Mach number of 2.4 and a free stream Reynolds number of 2.6 x 106 m 1. The results obtained, using a combination of oil flow visualisation, static pressure measurements, schlieren photography and vapour screen visualisation, indicate that existing flow field models can be extended to include the more general configurations _ tested. Leading edge diameter was found to be the major parameter controlling the scale of the interaction produced by the blunted models, with model incidence playing a secondary role, even at large distances from the generator. Existing methods for predicting the scale of the un swept sharp generator flow field are shown to provide a reasonable estimate of the variation of upstream influence with Reynolds number.Item Open Access The effects of junction modifications on sharp-fin-induced glancing shock wave/turbulent boundary layer interaction(1994-06) Koide, S.; Stollery, J. L.The effects of junction modifications on the glancing shock wave/turbulent boundary layer interaction generated by a sharp fin placed on the wall of a supersonic wind tunnel were examined experimentally at a Mach number of 2.46 and a Reynolds number of 2.59xl06/m. The interactions between a turbulent boundary layer on the wall and shock-wave systems caused by a fin with a fillet and several fins with different strakes were examined individually in order to find an effective modification technique. The flow features obtained by oil flow visualization, surface pressure measurements and liquid crystal thermography were compared with the data from an unmodified fin to evaluate the effects of each modification. The comparisons indicated that a "srake-type" modification had a weakening effect on the interaction-induced separation. To understand the flowfield structures, three-dimensional surveys using laser-light-sheet flow visualization were employed with schlieren photography and oil flow visualization. In addition to these experimental observations, an Euler CFD solver was used to help understand the inviscid flow structures which play important roles in the interaction behaviour. Based on the data experimentally and numerically obtained, a method was proposed for predicting the junction shapes needed to prevent separation.Item Open Access The experimental and theoretical aerodynamic characteristics of aerofoil sections suitable for remotely piloted vehicles(Cranfield University, 1984-03) Render, P. M.; Stollery, J. L.Using the design requirements of Remotely Piloted Vehicles (RPV's), selected for wind tunnel testing over the Reynolds number range 3 x 105 to 1 x 106. The first aerofoil, NACA 643-418, showed a degradation of performance in terms of lift-to-drag ratio as the Reynolds number was reduced. There was also a laminar separation bubble of notable extent on both the upper and lower surfaces at most incidences throughout the Reynolds number range. The second aerofoil, Göttingen 797, had good performance in terms of lift-to-drag ratio and maximum lift coefficient, even at the lowest Reynolds number. This was attributed to the flat bottom of the aerofoil, which allowed the formation of extensive laminar flow on the lower surface without the formation of a laminar separation bubble. The third aerofoil, Wortmann FX63-137, generally exhibited the best aerodynamic performance in terms of maximum values of both lift-to-drag ratio and lift coefficient, throughout the Reynolds number range considered. Four alternative lower surface geometries for this aerofoil were also tested. The modifications reduced the maximum values of both the lift coefficient and lift-to- drag ratio of the original aerofoil throughout the Reynolds number range, but generally improved the lift-to-drag ratios at low values of lift coefficient. The notable exception was the modification which resulted in a flat bottomed section. This had maximum values of lift-to-drag ratio which were within a few percent of those of the original aerofoil throughout the Reynolds number range. Wind tunnel results were used to evaluate low-speed aerofoil analysis computer programs written by Eppler and Somers (13) and Van Ingen (18). The results were disappointing. However, using the same wind tunnel results it was noted that computer programs using semi-inverse viscous methods show great promise.Item Open Access Experimental study of slender vehicles at hypersonic speeds(Cranfield University, 1996-04) Singh, Amarjit; Stollery, J. L.An experimental investigation of the hypersonic flow over (i) a wing-body configuration, (ii) a hemi-spherically blunted cone-cylinder body and (iii) a one-half- power-law body has been conducted for M,, = 8.2 and Re,, = 9.35x104 per cm. The tests were performed at model incidences, a=0,5 and 10° for flap deflection angles, (3 = 0,5,15, and 25° for the wing-body. The incidence ranged from -3 to 10° for the cone- cylinder and -5 to 15° for the power-law body. (i) The schlieren pictures showing top and side views of the model indicate that the body nose shock does not intersect the wing throughout the range of a under investigation. Detailed pressure measurements on the lower surface of the wing and flap along with the liquid crystal pictures suggest that the body nose shock does not strike the flap surfaces either. The wing leading edge shock is found to be attached at a=0 and 5° but detached at a= 10°. The liquid crystal pictures and surface pressure measurements indicated attached flow on the lower surface of the wing and flap for 13 =0 and 5° at all values of a under test. However at a= 0°, as the flap angle is increased to 15° the flow separates ahead of the hinge line. As incidence is increased the boundary layer becomes transitional giving rise to complex separation patterns around the flap hinge line. The spherically blunted body nose causes strong entropy layer effects over the wing and the trailing edge flap. A Navier-Stokes solution indicated a thick entropy layer of approximately constant thickness all around the cylindrical section of the body at zero incidence. However, at an incidence of 10° the layer tapers and becomes thinner under the body. The surface pressure over the wing and the plateau pressure for separated flow was found to increase from the root to the tip. This is partly because of the decrease in local Reynolds number across the span, however in the present case, entropy layer effects also affected separation. The entropy layer effects were found to reduce the peak pressures obtainable on the flap. The peak pressures, over the portion of the flap unaffected by entropy layer effects, could be estimated assuming quasi two dimensional flow. (ii) Force measurements were made for the blunted cone-cylinder alone as well as with the delta wing, with trailing-edge flap, attached to it. The lift, drag, and pitching moment characteristics for the cone-cylinder agree reasonably well with the modified Newtonian theory and the N-S results. The addition of a wing to the cone-cylinder body increases the lift as weil as the drag coefficient but there is an overall increase in the lift/drag ratio. The deflection of a flap from 0° to 25° increases the lift and drag coefficients at all the incidences tested. However, the lift/drag ratio is reduced showing the affects of separation over the wing. The experimental results on the wing-body are compared with the theoretical estimates based upon two-dimensional shock-expansion theory. (iii) The lift, and drag characteristics of a one-half-power-law body are compared with other existing results. The addition of strakes to the power-law body are found to improve its aerodynamic efficiency without any significant change in its pitching moment characteristics.Item Open Access Hypersonic control effectiveness(Cranfield University, 1995-01) Kumar, D.; Stollery, J. L.; Mapes, J.The present study analyses the effects of a number of geometric parameters on the performance of a trailing edge control flap on a hypersonic body. The tests were conducted in a gun tunnel at Mach 8.2 and Mach 12.3. The study revealed that flap deflection promoted separation lengthscales and boundary layer transition. The latter significantly increased the local aerothermal loads on the flap. For well separated flows, flap heat transfer rates were successfully predicted by reference temperature theory. The promotion of transition caused a progressive reduction in the lengthscales of separated flows. In a free-flight environment, vehicle incidence varies considerably. Incidence was found to promote transition on both flat plates and control flaps. The latter resulted in a considerable increase in flap heat transfer. A modified version of reference temperature theory successfully predicted the aerothermal loads on the flap. For laminar and transitional interactions, the separated flow lengthscale was found to have a complex variation with incidence. A number of relevant flow parameters were identified. The intense heat loads on a vehicle in hypersonic flight dictates the blunting of the leading edge. This strengthens the leading edge shock structure and generates an entropy layer. Bluntness was found to significantly decrease the separation interaction scales on the flap. This was due to a reduction in the pressure recovered on the flap. The latter adverse affects control effectiveness. The aerothermal loads on the control flap was successfully predicted by reference temperature theory. An investigation into the efficiency of an under-expanded transverse jet controls was conducted on an axi-symmetric slender blunt cone. Force measurements found that the interaction augmented the jet reaction force by 70% at zero incidence. This increased to 110% at low incidence. The experiments found that the scale of the interaction region was determined by Poj/pes. Using this parameter, a closed loop algorithm for the shape of the separation front was developed. The latter can be used to predict jet reaction control effectiveness.Item Open Access Hypersonic interference heating in the vicinity of surface protuberances(Springer Science Business Media, 2010-09-30T00:00:00Z) Estruch-Samper, David; MacManus, David G.; Stollery, J. L.; Lawson, Nicholas J.; Garry, Kevin P.The understanding of the behaviour of the flow around surface protuberances in hypersonic vehicles is developed and an engineering approach to predict the location and magnitude of the highest heat transfer rates in their vicinity is presented. To this end, an experimental investigation was performed in a hypersonic facility at freestream Mach numbers of 8.2 and 12.3 and Reynolds numbers ranging from Re (a)/m = 3.35 x 10(6) to Re (a)/m = 9.35 x 10(6). The effects of protuberance geometry, boundary layer state, freestream Reynolds number and freestream Mach numbers were assessed based on thin-film heat transfer measurements. Further understanding of the flowfield was obtained through oil-dot visualizations and high-speed schlieren videos. The local interference interaction was shown to be strongly 3-D and to be dominated by the incipient separation angle induced by the protuberance. In interactions in which the incoming boundary layer remains unseparated upstream of the protuberance, the highest heating occurs adjacent to the device. In interactions in which the incoming boundary layer is fully separated ahead of the protuberance, the highest heating generally occurs on the surface just upstream of it except for low-deflection protuberances under low Reynolds freestream flow conditions in which case the heat flux to the side is greater.Item Open Access Measurement of shock wave unsteadiness using a high-speed schlieren system and digital image processing(American Institute of Physics, 2008-12) Estruch-Samper, David; Lawson, Nicholas J.; MacManus, David G.; Garry, Kevin P.; Stollery, J. L.A new method to measure shock wave unsteadiness is presented. Time-resolved visualizations of the flow field under investigation are obtained using a high-speed schlieren optical system and the motion of the shock wave is determined by means of digital image processing. Information on the shock’s unsteadiness is subsequently derived with Fourier analysis. A sample study on shock unsteadiness in a shock-wave/turbulent boundary-layer interaction with separation is included. The method presented enables a measure of shock unsteadiness at locations in the imaged flow field not accessible by intrusive methods.Item Open Access Passive Techniques for Controlling the Flow in Supersonic Wing - Body Junctions(1994-03) Blank, S. C.; Stollery, J. L.Junction flows are common to all flight speeds and they are associated with undesirable features such as increases in drag, limitations on performance and at supersonic speeds high heat transfer rates. Junction flows are associated with performance losses in turbomachinery (around 30% of the total pressure losses in an axial turbine) and they can lead to premature detection of military underwater vessels. Junction fairings are widely used at subsonic speeds and correct fairing of the C-141 wing alone, could have resulted in potential fuel savings of approximately US $ 40,000 per year per aircraft which can be roughly translated into a $ 600,000 saving during the lifetime of each airframe. Typically, for a modern transport type of aircraft the wing juncture accounts for between 1 and 2 % of cruise drag and therefore, careful design of the wing junction is necessary. At supersonic speeds, by far the most disadvantageous feature associated with juncture design is high heat transfer rates due to shock / shock interaction. These heat transfer rates are sufficient to cause severe structural damage leading to component burn-off. Typical leading edge temperatures during re-entry for an aerospaceplane, like HOTOL, are around 2000 K, exclusive of increases in temperature caused by the interaction. Although drag reduction may not be so relatively critical at these flight speeds, the potential loss of control components, like stabilising fins, needs to be carefully examined and some configuration re-design may be necessary as a consequence. This research project is aimed at developing a technique through which the disadvantageous features associated with supersonic junction designs can be alleviated. It was found that through re-design of the wing / body junction the maximum mean static pressure local to the fin leading edge could be halved, the strength of the junction vortices could be lowered and the amount of separated flow reduced. The applications of the technology span all vehicles operating within the supersonic flight regime and therefore, the markets to which the technology is applicable are military aircraft, defence systems, aerospaceplanes and commercial supersonic transports. As the technology is, in principle, applicable to the design of commercial supersonic transport aircraft (HSCT) and the market for these vehicles is forecast to be worth up to $ 200 billion (FY 1987) an examination of the issues behind marketing this type of vehicle is presented in the non¬technical section. The presently available data produced by the major manufacturers were found to be lacking in the following areas (a) evaluations of market elasticity, (b) distribution techniques, (c) the availability of acquisition finance and financing techniques, (d) political sensitivity analysis (d) product life cycle analysis and (f) the relationship marketing of the venture.Item Open Access The performance of 60 degrees delta wings : the effects of leading edge radius on vortex flaps and the wing(Cranfield Institute of Technology; College of Aeronautics, 1990) Hu, B. K.; Stollery, J. L.Item Open Access Projectile aerodynamics: Measurement and computation(Cranfield University, 1997-10) Kontis, Konstantinos; Stollery, J. L.An experimental study has been performed at M∞=8.2 and Re∞/cm=93000 to examine: 1. The effect of strakes on the aerodynamic characteristics and performance on slender elliptic cone missile configurations. Some information regarding the shock layer was obtained from schlieren pictures. Detailed flow properties in the shock layer were obtained, for some elliptic cone configurations with and without strakes, using a threedimensional, high resolution, iterative, finite volume parabolized Navier-Stokes solver. Surface flow visualisation, using an oil-dot technique, and pressure measurements were made on one of the models to determine the effect of strakes. Lift, drag and pitching moment characteristics for the elliptic cones with and without strakes were obtained using a three component strain-gauge balance. No gross external flow differences were detected from the schlieren pictures for models tested due to the addition of strakes. Oil-dot visualisation demonstrates that the strakes alter the surface flow characteristics and tended to inhibit the cross-flow. The addition of strakes caused a reduction of pressure on the leeward side and an increase of pressure on the windward side. The strakes produced a significant increase in the lift and drag coefficients, in the incidence range of 0° to 200. The right elliptic cone without strakes with its major axis horizontal exhibits higher lift coefficients than the cone with its major axis vertical. The numerical study predicted the complex flowfield surrounding the right elliptic cone with its major axis horizontal, gave a better understanding of the complicated nature of the flow and good indications of the shock shape and vortex core positions. An estimation model of the aerodynamic forces and moments for the right elliptic cone with and without strakes was developed based on the standard Newtonian theory. The model successfully predicted the experimental trends in the aerodynamic coefficients. 2. The aerodynamic effectiveness of a cylinder flare body at zero incidence under laminar and turbulent boundary layer conditions. Two nose geometries, namely a 10° half-angle sharp cone and a hemisphere, were used. The surface flow over the cylinderflare body was studied using oil-dot and liquid crystal techniques. Some information regarding the shock layer was obtained from schlieren pictures. The effects of entropy layer and boundary layer state on flare effectiveness were deduced from pressure measurements over the cylinder and the flare. The most important difference between the laminar and turbulent boundary layer interaction is that a much smaller angle is necessary to cause laminar separation than that necessary for turbulent separation. The determination of incipient separation is very sensitive to the detection method employed. The existence of a small scale separation bubble can explain the differences in the determination of incipient separation angles if different experimental methods are used. The addition of a hemisphere nose reduces the surface pressure and heat transfer levels on the flare. This is due to loss of reservoir pressure across the bow shock wave. The reduction of flare pressure also reduces the separated flow lengths for the laminar case, whereas for the turbulent case the separated flow lengths are increased. This may be due to the boundary layer along the cylinder body not being fully developed. The effect of Mach shear on the flare pressures distribution has been calculated theoretically. The model predicted the experimental results satisfactorily.Item Open Access Ram-jet combustion based on shock/flame interaction(Cranfield University, 1983-12) Edwards, J. A.; Stollery, J. L.An experimental investigation into the effects of shock/wake and shock/flame interaction on the base pressure of axisymmetric bodies at Mach 2 has been carried out. This investigation has determined the effects of various forms of shock generator (axisymmetric cowls, twodimensional wedges and 'delta' wings) on the base pressure. Shock waves generated by over-expanding the airflow in an open-jet wind tunnel have been used to determine the effect of shock strength on the base pressure of an axisymmetric fuel injector. Both peripheral bleed and axial bleed of hydrogen fuel have been examined and the effect of shock compression on the resulting flame has been determined. In the axial bleed case nitrogen and hydrogen bleed without combustion has also been examined. The effect of varying the airflow stagnation temperature has also beeninvestigated. It is demonstrated herein that there is a distinct shock/wake interaction position that maximises the base pressure, that with interaction at this optimal position the static pressure rise across the shock wave can be communicated in full to the base of the centrebody, and that favourable aerodynamic interference between the wake and a cowl of 50 convergent-divergent internal section can give rise to a net drag reduction. The shock/wake and shock/flame experiments demonstrate that a significant base thrust can be generated, however, the fuel efficiency decreases with increasing shock strength. It is shown that the fuel specific impulse is a function of shock strength, interaction position and bleed mode (peripheral or axial). The onset of boundary layer separation due to the adverse pressure gradient encountered when the base pressure is high appears to limit the useful addition of wake combustion. Finally, it is demonstrated that the base pressure, with and without combustion, is only a weak function of airflow stagnation temperature.Item Open Access Some hypersonic intake studies(Royal Aeronautical Society, 2006-03-01T00:00:00Z) Lanson, F.; Stollery, J. L.A 'two dimensional' air intake comprisipg a wedge followed by an isentropic compression has been tested in the Cranfield Gun Tunnel at Mach 8,2. These tests were performed to investigate qualitatively the intake flow starting process. The effects of cowl position, Reynolds number, boundary-layer trip and introduction of a small restriction in the intake duct were investigated. Schlieren pictures of the flow on the compression surface and around the intake entrance were taken. Results showed that the intake would operate over the Reynolds number range tested. Tests with a laminar boundary layer demonstrated the principal influence of the Reynolds number on the boundary-layer growth and consequently on the flow structure in the intake entrance. In contrast boundary layer tripping produced little variation in flow pattern over the Reynolds number range tested. The cowl lip position appeared to have a strong effect on the intake performance. The only parameter which prevented the intake from starting was the introduction of a restriction in the intake duct. The experimental data obtained were in good qualitative agreement with the CFD predictions. Finally, these experimental results indicated a good intake flow starting process over multiple changes of parameters.Item Open Access Studies in transonic flow(Cranfield University, 1983-06) Mohan, S. R.; Stollery, J. L.This thesis is divided into, two distinct parts. Part I describes the design and development of an intermittent cryogenic wind-tunnel, in which the cold conditions are generated by the expansion of high pressure gas. The device uses a light piston moving in a tube and conditions during the running time are maintained constant by 'tuning' the piston motion, i. e. by Imatchingt the volumetric flow rate entering and leaving the tube. The results of the pilot tunnel (running time 0.3 secs. ) show that gas temperatures of about 110K can be obtained with a pressure ratio of 35. Part II describes the flow at, transonic speeds on five aerofoil sections (thickness-chord ratios 6 to 14%). The aerofoil sections were 6& 14% biconvex, an NACA 0012, a supercritical aerofoil CAST 7 and an 11.8% Joukowski profile. The tests were made in an intermittent$ perforated wall wind-tunnel which was developed from an existing supersonic wind-tunnel. A periodic flow due to shock-induced separation occurred over a narrow range of Mach numbers (from 0.82 to 0.90) on the 14% biconvex and 0012 sections at zero incidence, for both laminar and turbulent boundary layers. The frequency parameter(uc/Um) was about 1. Tests were also made with the aerofoils at incidence. Periodic flows occurred on the 14% biconvex, 0012 and 3oukowski profiles., The instability Mach numbers ranged from 0.84 to 0.90 and the frequency parameters from 0.40 to 1.36. A detailed study was made to determine the frequency spectrum of the tunnel noise and its influence on the periodic flow. Also, experiments were made to determine the influence of the aerofail geometry on the periodic flow. The experimental results on the 14% biconvex aerofoil have been compared with the numerical computations of the full Navier-Stokes equations performed at NASA, Ames.Item Open Access A study of the interaction between a glancing shock wave and a turbulent boundary layer(Cranfield University, 1980-08) Kubota , H.; Stollery, J. L.An oblique shock generated by a variable-angle wedge on the side wall of a wind tunnel, has been used to investigate the three-dimensional glancing interaction problem. The shock interacts with the turbulent boundary layer growing along the side wall. Two related test programmes have been completed using a 2.5 x 2.5 inch intermittent tunnel and a 9 x 9 inch continuous-running tunnel. For both the test programmes, the Mach number was approximately 2.5 and the Reynolds number relative to the wall boundary-layer thickness 5 x 10 4 . The experimental results include oil-flow pictures, vapour-screen and smoke photographs,wall pressure distributions, local heat transfers, wall surface temperatures and viscous layer surveys. The experimental results suggest that the interaction reg10n consists of two different viscous layers between which an ordinary separation can take place, (the double viscous layer flow-field model). The three- dimensional separation is found to depend significantly on the pressure rise in the direction normal to the swept shock. In this sense the separation is similar to the two-dimensional case.Item Open Access A study of the interaction between a glancing shock wave and a turbulent boundary layer : the effects of leading edge bluntness and sweep(Cranfield University, 1985-11) Hussain, S.; Stollery, J. L.The effects of leading edge bluntness and sweep angle on the three dimensional glancing shock wave - boundary layer interaction have been investigated. A large number of hemi-cylindrically blunted fins with leading edge diameter ranging from 0 to l.Oin, with leading edge sweep angles between 0° and 75° were tested. The incidence angle was varied from 0° to 21°. The shock wave from each configuration interacted with a fully developed turbulent boundary layer growing along the tunnel side wall. The free stream Mach number in the 9in x 9in continuous flow supersonic wind tunnel was 2.4 and the Reynolds number based on boundary layer thickness was 5 x 10^. Experimental investigations included oil smear tests, surface pressure surveys, schlieren pictures of the inviscid shock envelopes and shock structure in the plane of symmetry. The study highlighted the significant effects of bluntness and sweep on the scale and character of the interaction. While bluntness intensified the interaction, sweep alleviated its intensity. The most dramatic effect of sweep angle was observed when the leading edge was swept from 0° to 30°. Sufficiently outboard of the plane of symmetry, the features of blunt and sharp fins became similar. The boundary between the inner "bluntness dominated" and the outer "viscous dominated" regions shifted inboard at the higher incidence and sweep angles. The characteristic surface oil flow patterns showed little change for sweep angles up to A = 60°. Leading edge bluntness increased the scale of the interaction almost linearly while leaving its character unchanged. The multiplicity of the separation and attachment lines on the side wall and the fin surface, suggested a system of vortices in the interaction region. Flow field models have been proposed over the range of sweep angles considered in the present study. The number and strength of the vortices is seen to depend on the leading edge bluntness, sweep and the incidence angle. The important parameters governing the primary separation distance and the peak pressure in the plane of symmetry have been identified. Correlation formulae suggest a strong interdependence of the various parameters concerned.Item Open Access Wind tunnel techniques for reducing commercial vehicle aerodynamic drag(Cranfield University, 1982-09) Garry, K. P.; Stollery, J. L.Increases in the price of petroleum fuels have significantly affected the importance of aerodynamic drag on commercial vehicle operating costs. The considerable savings to be made have resulted in: (i) the appearance of numerous 'add-on' devices intended to reduce the drag-of existing vehicles, and (ii) an acceptance by vehicle manufecturers of the importance of aerodynamics to their new designs. The majority of drag optimisation programmes are carried out using scale models in a wind tunnel, and the effectiveness of resulting modifications is often confined to the individual vehicle concerned. The relatively crude simulation techniques have been acceptable on the basis that potential errors are small compared to the reductions in drag that can be achieved. If the trend in reducing drag is to be maintained a greater understanding of the flow around commercial vehicle configurations will be needed, especially under simuleted,crosswind conditions, together with improvements to the wind tunnel techniques used to simulate the full scale environmen The experimental programme presented in this report is intended to Illustrate ,the influence of wind tunnel simulation technique, on the methods for reducing commercial vehicle aerodynamic-drag. result in wind tunnel tests over a 'range of simulated crosswind conditions and levels of free stream turbulence are present illustrate the significance of variations in Reynolds number, vehicle geometry, and wind tunnel size on the pressure distribution 2. and resulting diagonal forces on the vehicle. I.Jrfece pressure contours are ceteidgued.t0 give insight into the flowfield, around the vehicle in relation to the problems of body fouling and water spray generation although analysis here specifically concerns aerodynamic drag. Techniques for reducing drag relate primarily to modifications Of the forebody flowfield and results from a number of experimental Programmes are collated to illustrate the effectiveness of various,,, techniques on different vehicle geometries.