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This experimental study documents the development and separation of a hypersonic boundary layer produced naturally on the cold surface of a sharp slender cone. At the base of the conical forebody, the equilibrium turbulent boundary layer was allowed to separate over an axisymmetric rearward facing step to form a compressible base flow. The investigation was conducted in the Imperial College No.2 gun tunnel at a freestream Mach number of 9 and unit Reynolds numbers of 15 and 55 million. The compressible boundary layer study was carried out at both of the available freestream unit Reynolds numbers and the measured data include distributions of wall static pressure and heat transfer rate, together with profiles of pitot pressure through the boundary layer. Using the chordwise distribution of surface heat flux as a means of transition detection, the cone transition Reynolds number was found to be 5.4x10^. This result, together with that obtained from flat plate studies conducted in the same test facility, provided a ratio of cone to flat plate transition Reynolds number of 0.8. Boundary layer integral quantities and shape factors are derived from velocity profiles and in most cases the measured data extended close enough to the wall to detect the peak values of the integrands. The separated flow region formed at the base of the cone was documented only at the higher unit Reynolds number, a condition under which the approaching turbulent boundary layer was found to be close to equilibrium. The data include pitot pressure profiles recorded normal to the surface downstream of reattachment, together with wall static pressure and heat transfer rate distributions measured throughout the base flow region. Reattachment occurred approximately two step heights downstream of separation and a surface flow visualisation study indicated the existence of Taylor-Goertler type vortices, emanating from the reattachment line in the downstream direction. A simple shear layer expansion model is developed and shown to provide a favourable prediction of the measured pitot pressure profiles recorded downstream of the reattachment line. The success of this second order model implies that the dynamics of the corner expansion process, except in the immediate vicinity of the wall, is governed largely by inviscid pressure mechanisms and that the supersonic region of the boundary layer expansion is essentially isentropic.

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The effect of sweep on separation bubbles as occurring in the subsonic flows past thin flat plates with rectangular leading edges has been studied experimentally. The distance between separation and reattachment, at high Reynolds number, was about 5.5 times the plate thickness in the flow region undisturbed by end effects. This distance was independent of sweepback for sweep angles up to and including 45<SUP>o</SUP>. The chordwise distribution of a static-pressure coefficient and a coefficient of the intensity of the static-pressure fluctuations, both measured on the surface of the plate and based upon the free-stream velocity component normal to the leading edge, were independent of the sweep angle up to and including 30<SUP>o</SUP> to a first approximation. The spectra of the static-pressure fluctuations, however, displayed some qualitative changes with increasing sweep angle. The distribution of a coefficient of the chordwise skin-friction component, based upon the free-stream velocity component normal to the leading edge, was independent of sweep up to and including 30<SUP>o</SUP> to a crude first approximation. The chordwise velocity profiles non-dimensionalised by the local external chordwise velocity component, were independent of sweep up to and including 45<SUP>o</SUP> in the separation bubble but downstream of reattachment small but persistent changes occurred with increasing sweep angle. Smoke-flow visualisations in the swept and the unswept flow at low Reynolds number displayed the presence of typical vortex loops in the reattachment region, many of which broke up and were partially entrained into the separation bubble.

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A series of experiments were undertaken at the Defence Evaluation Research Agency (DERA) Farnborough, within the Aero Physics group into the phenomenon called Jet Interaction (JI). Jet Interaction (JI) is produced by the interaction of a jet with the external flow around a vehicle. This study focused on investigating the effects of a divert thruster employed to provide a vehicle with a rapid divert capability on the external flow-field and on the induced forces and moments exerted on the vehicle by the jet. The research was based on studying the effects on a hypersonic interceptor sometimes referred to as a KKV (Kinetic Kill Vehicle). There are many important parameters in JI. One of these is the jet Amplification Factor (AF). This is caused by the deflection of the free-stream around and over the jet and results in a pressure increase on the vehicle surface which adds to the divert thrust force. The experiments were carried out in the intermittent hypersonic gun tunnel, at a free-steam Mach number of 12.1. This produced a Reynolds number (based on diameter) Red of 300,000 with a one-tenth scale vehicle; these conditions correspond to a full-scale vehicle flying at an altitude of 41 km. To simulate the divertthruster, nitrogen was supplied to the model through a purpose-made force balance. Measured forces ineluded normal, axial and side as well as pitch and yawing moments. The experimental results have been compared with that of a full size vehicle featuring a 2kN divert thruster. The results have then been matched to the effective altitude as a function of the thrust coefficient (Cr). This allows the experimental data to be interpreted for a full-scale vehicle in order to answer design questions important to system engineers. The results from this investigation show that the effectiveness of a divert jet is influenced by the vehicle's altitude, achieving a negligible increase in AF with a Cr > 2.5 at an altitude > 50 km. The seeker will suffer from jet induced problems at low Cr levels for a. = 10°. An increase in Cr causes the separation region in front of the jet to extend to the nose of the tested configuration for M1 = 12.1 while complete separation is achieved at Cr > 1.2. Injection Mach number (MJ) has a small influence on AF. However it does not influence the separation region. Penetration height (h) of the jet is increased for higher Mach number injection. Both AF and the separation region are influenced by nozzle geometry. A series of different nozzle geometries were tested. These had the effect of reducing the measured amplification factor to a maximum of 1, except for the dual circular orifice combination, which doubled the measured AF achieved for a single circular orifice. The influence of nozzle geometry reduced the Cr levels required to produce a negligible increase in AF and the corresponding altitude. The angle of attack (œ) has a strong influence on AF at low Cr levels, however it becomes negligible when complete separation is achieved.

A comprehensive experimental study of fountain flow has been completed. The study was concerned with the characteristics and effects of the flow domain below an idealised small-scale model of the underside of the fuselage of a vertical take-off or landing aircraft with a pair of matched jets. The fountain flow was generated by the normal impact of the jets with a ground board and it in turn impacted with a rectangular blocking tray representing the undersurface of a fuselage. A range of test programmes were carried out. After having examined the fountain flow pressure field and conducted simple flow visualisation tests, a series of tests were undertaken on a range of blocking trays to determine the magnitude and variation of the normal forces on the trays due to the effect of impinging fountain flows. A few tray force tests were carried out with a heated compressed air supply to the jets' nozzles. Following these tests, programmes of pressure distribution tests were carried out on a 'long' tray and on a 'short' tray. Temperature distribution tests were conducted on a long tray, these tests were exploratory only and could form the basis for future experimental studies. The test facility was appraised for jets flowing at high subsonic velocities and found to be suitable but no further tests were undertaken. Finally use was made of computational fluid dynamics (CFD) theoretically to solve the governing Navier-Stokes fluid flow equations for domains which simulated the experimental studies just completed and analysed. The analyses of the data obtained from the test programmes revealed that fountain flow caused towards 20% of jet lift augmentation at maximum with good correlation between direct tray force measurements and integrated tray pressure measurements. A useful plot of the variation of tray force against tray height was obtained from moving ground board tests but the velocity of vertical movement of the board towards the nozzles and tray was found not to make a significant difference for the given range of velocities. The plot displayed a distinct peak when the gap between the tray and the ground board was about 2% jet nozzle diameters. At smaller gaps the tray force fell steeply to a suckdown value approaching 20% of jet lift reduction. A reasoned explanation of this phenomenon is presented. The tray force tests using hot jets indicated that slightly lower tray forces occurred than for jets at ambient temperature, everything else being equal. For the general configuration of the model it was concluded that hot gas tests were not necessary as tests at room temperature gave more useful results for design purposes. The CFD computations were particularly effective in obtaining a qualitative impression of the flow patterns in the blocked fountain flow domain. A useful insight was also gained into entrainment into the domain. The tray force computaions were less satisfactory, however, resulting in underestimations of about 30% although the general pattern of pressure distributions across and along the tray were quite similar experimentally determined patterns. Overall it was concluded that the use of small scale model tests had been shown to be effective for certain design purposes for full-scale aircraft and that a dual approach using CFD and model tests could be more rewarding than either the one or the other technique in isolation.