S. L. Gai

The interaction of a shock wave with a laminar boundary layer at a compression corner in high-enthalpy flows including real gas effects

The high-enthalpy, hypersonic flow over a compression corner has been examined experimentally and theoretically. Surface static pressure and heat transfer distributions, along with some flow visualization data, were obtained in a free-piston shock tunnel operating at enthalpies ranging from 3 MJ kg -1 to 19 MJ kg -1 , with the Mach number varying from 7.5 to 9.0 and the Reynolds number based on upstream fetch from 2.7 x 10 4 to 2.7 x 10 5 . The flat-plate similarity theory has been extended to include equilibrium real gas effects. While this theory is not applicable to the current experimental conditions, it has been employed here to determine the potential maximum effect of real gas behaviour. For the flat plate, only small differences between perfect gas and equilibrium gas flows are predicted, consistent with experimental observations. For the compression corner, a more rapid rise to the maximum pressure and heat transfer on the ramp face is predicted in the real gas flows, with the pressure lying slightly below, and the heat transfer slightly above, the perfect gas prediction.

Influence of Trailing-Edge Flow Control on Airfoil Performance

Experiments were conducted on a NACA 0012 airfoil equipped with Gurney flap. Two types of Gurney flap were used: one was solid with a height of 5% chord and the other was triangularly serrated with a height of 5% chord. The Reynolds number of the tests, ∼2 x 10 5 based on chord, was typically in the range of interest for microaerial vehicles and rotor blades of helicopters. The results confirmed the advantages of a Gurney flap, in general, and provided additional insight into their effectiveness. A simple analysis that extends the hypothesis of Liebeck on the functioning of the Gurney flap is offered

Mole-Fraction-Sensitive Imaging of Hypermixing Shear Layers

A theoretical model that determines the optimum excitation frequency for obtaining a e uorescence signal with a strong dependence on fuel mole fraction is presented for supersonic fuel ‐air compressible mixing studies. The challenge associated with this is to maintain a high sensitivity to fuel mole fraction with minimal sensitivity to temperature and pressure in a e ow with large temperature variations and pressure gradients. The results of the modelareappliedtothemixingregion behindvariousscramjetfuelinjectorsinashocktunnelto measurefuelmole fraction. Hydrogen fuel at a Mach numberof 1.7 is injected into a mostly N 2 freestream at Mach 4.8. Experimental e uorescence images are presented in streamwise and spanwise planes.

Free piston shock tunnels: developments and capabilities

Abstract The free piston shock tunnel has shown itself to be uniquely capable of adequately simulating real gas effects and has been successfully operating in Australia for more than two decades. As a consequence of the recent interest in hypersonic aerospace planes and general resurgence in hypersonic research, attention has been focussed on this type of facility and free piston shock tunnels have recently been built or being considered in Europe, U.S.A. and elsewhere. It is in the light of these events that this paper discusses the development, performance and capabilities of free piston shock tunnels.

Laminar heat transfer to blunt cones in high-enthalpy hypervelocity flows

Measurements of heat transfer on spherically blunted cones of various bluntness ratios in high-enthalpy hypervelocity air flows obtained using a free-piston-driven shock tunnel are described. The stagnation enthalpies considered were in the range 14—20-MJ/kg (equivalent to flight speeds 4—6-km/s). The corresponding equilibrium stagnation temperatures were in the range 7000-9000 K. Comparison of measured data showed reasonable agreement with the classical theory of Lees 1 for small bluntness ratios while the large bluntness ratio results compared well with the empirical theory of Griffith and Lewis.4

Bluff body wakes with free, fixed, and discontinuous separation at low Reynolds numbers and low aspect ratio

Some comparative experimental results are presented of bluff body wakes with free, fixed, and discontinuous separation. The particular examples considered are a circular cylinder, a plain blunt trailing edge aerofoil and a segemented blunt trailing edge aerofoil. The evolution of the near wake and vortex shedding modes are compared and discussed.

Laminar Near Wake of a Circular Cylinder at Hypersonic Speeds

The laminar near-wake flow behind a circular cylinder at hypersonic speeds has been examined experimentally and analytically. Surface pressure and heat flux measurements were obtained in a free-piston shock tunnel at a nominal Mach number of 10. The freestream unit Reynolds numbers were 3.02 x 10 5 /m and 11.7 x 10 5 /m at total specific enthalpies of 13.35 and 3.94 MJ/kg, respectively. The experimental data of surface pressure and heat flux showed good agreement with theory based on perfect gas. Unlike surface pressure, the surface heat flux depended strongly on the wall-to-total-temperature ratio, and it increased with increase in the ratio. The surface pressure, however, depended on the Reynolds number, although this dependency was found to be weak. The flow separation angles behind the cylinder showed an inverse of the square root of Reynolds number dependence similar to the low-speed laminar flow separation behind a circular cylinder.

Large-Scale Structures and Growth of a Flat Plate Compressible Wake

Experiments conducted on a flat plate wake in a compressible flow at Mach 2 are described. Large-scale organized motions are revealed both in the separating boundary layer upstream and in the wake downstream. The organized motions in the wake contained within them embeded vortex structures, whose periodicity was found to be weak. The broadband spectra contained a distinct frequency with a Strouhal number of 0.3. The study of mean flow characteristics showed that some qualitative similarity exists in the near- and intermediate-wake regions. The relative Mach number distribution showed the compressibility effects to be small

Supersonic Flow over a Shallow Open Rectangular Cavity

An experimental investigation was conducted on a supersonic flow at a freestream Mach number of 2 over a shallow open cavity, including the effects of adding streamwise serrated edges. These flows have relevance to weapons bays and airframe gaps on high-speed aircraft. The measurements consisted of single-shot and time-resolved schlieren visualization, as well as unsteady pressure spectra. The length-to-depth ratio of the cavity was 8. The tests conducted at different Reynolds numbers with the baseline cavity (straight leading and trailing edges) showed that increasing the Reynolds number increases the root-mean-square pressure inside the cavity. The addition of serrations to the cavity leading or trailing edge did not show any significant effect on the separating shear layer nor in controlling the oscillations of the shear layer. There was also no noticeable effect on the overall sound pressure levels inside the cavity. A new expression for calculating shallow-cavity resonant frequencies applicable at su...

Boundary Layers on a Flat Plate at Sub- and Superorbital Speeds

The heat flux and boundary layers over a flat plate at suborbital and superorbital speeds were measured in a freepiston expansion tube at enthalpies of 26.0 and 53.4 MJ/kg, respectively. Estimates of the gas-phase and surface Damkohler numbers were made that indicate that although the boundary layer might be frozen, there is a possibility of surface reactions occurring. The heat-flux results were compared with theoretical predictions of heat flux for both frozen and equilibrium chemistries. The results indicated that the influence of real gas effects, such as recombination and surface catalycity, were minimal for the present flow conditions. Interferograms were obtained using resonance enhancement of the thermal boundary layer over a flat plate. Comparison of the measured boundary-layer thickness with several expressions used to predict the boundary-layer thickness allowed an evaluation of their effectiveness at high enthalpies.