Andrew J. Neely

The Superorbital Expansion Tube concept, experiment and analysis

In response to the need for ground testing facilities for super orbital re-entry research, a small scale facility has been set up at the University of Queensland to demonstrate the Superorbital Expansion Tube concept. This unique device is a free piston driven, triple diaphragm, impulse shock facility which uses the enthalpy multiplication mechanism of the unsteady expansion process and the addition of a secondary shock driver to further heat the driver ps. The pilot facility has been operated to produce quasi-steady test flows in air with shock velocities in excess of 13 km/s and with a usable test flow duration of the order of 15 mus. An experimental condition produced in the facility with total enthalpy of 108 MJ/kg and a total pressure of 335 MPa is reported. A simple analytical flow model which accounts for non-ideal rupture of the light tertiary diaphragm and the resulting entropy increase in the test gas is discussed. It is shown that equilibrium calculations more accurately model the unsteady expansion process than calculations assuming frozen chemistry. This is because the high enthalpy flows produced in the facility can only be achieved if the chemical energy stored in the test flow during shock heating of the test gas is partially returned to the flow during the process of unsteady expansion. Measurements of heat transfer rates to a flat plate demonstrate the usability of the test flow for aerothertnodynamic testing and comparison of these rates with empirical calculations confirms the usable accuracy of the flow model.

Turbulent heat transfer measurements using liquid crystals

Abstract The transient method of measuring heat transfer coefficients that uses liquid crystals, since its beginnings in the early 1980s, has become one of the best ways of determining full surface distributions of heat transfer coefficient. The turbomachinery research group at Oxford has concentrated on the application of the method to numerous mechanical engineering thermal problems specific to the jet engine. The paper summarises some of the recent developments in the technique including the implementation of an elegant way of producing a change in the fluid temperature. Recent, high-density heat transfer coefficient measurements are discussed together with the advantages such resolution offers in terms of flow field interpretation. A means of integrating the measurements into finite element software for subsequent data analysis is presented. The paper should be of interest to engineers interested in using the most modern heat transfer measurement techniques in their research and development programmes.

OH PLIF Imaging of Supersonic Combustion using Cavity Injection

Planar laser-induced fluorescence of the hydroxyl radical is used to investigate a cavitybased fuel injection system for scramjet operation at a flight Mach number of 11.5. Hydrogen and ethylene fuels are compared over a range of fuel-lean global equivalence ratios. Both fuels show evidence of combustion at each fuel injection pressure. Hydrogen burns diffusively in the shear layer and further downstream, with penetration distance increasing linearly with injection pressure. The combustion in the cavity shear layer sheds regular vortices. Ethylene behaves similarly to hydrogen at the lower equivalence ratios but shows evidence of intense, localized combustion at the highest equivalence ratio. In all cases combustion occurs in the shear layer above the cavity rather than in the recirculating cavity flow.

CFD analysis of an ethylene-fueled intake-injection shock-induced-combustion scramjet configuration

A computational analysis of ethylene combustion in scramjet engine applications has been performed. The capability of simplified chemical reaction mechanisms to accurately calculate ignition delay and heat release has been assessed. The 10 step Jachimowski ethylene combustion model was found to over predict ignition delay by almost an order of magnitude. Using this mechanism, CFD calculations have been made of a threedimensional geometry representing a potential configuration for the Australian developed HyShot scramjet. The free stream conditions are those of shock tunnel conditions with equivalent flight speed of Mach 9.4 at 35 km altitude. Within the 3D flow field the Jachimowski reaction mechanism calculated extremely limited combustion with one percent of injected fuel consumed. An increase of the scale of the scramjet by one order of magnitude enhanced the extent of combustion, but not sufficiently to produce any significant pressure rise since it was found that the fuel injection was not delivering the ethylene to the region where it was most likely to burn

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.

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...

Calibration of Thermal Paints for Hypersonic Flight Test

*† A number of techniques for the transient calibration of permanent-change thermal paints intended for use on short-duration hypersonic flight tests are described and tested. These include the convective heating of painted test samples using an oxy-acetylene torch and the resistive heating of test samples. Both of these methods are shown to be able to induce both the magnitudes and trajectory of temperature rise likely to be experienced on a flight vehicle. The use of a laser to radiatively heat the samples is also discussed but not tested. Methods to analyse the calibration data and apply it to interpret hypersonic flight test measurements are also discussed.

Aerodynamics of a Supersonic Projectile in Proximity to a Solid Surface

Flow around a Mach 2.4 NATO 5.56 mm projectile in close proximity to a ground plane was investigated using computational fluid dynamics for a direct numerical reproduction of live-range experiments. The numerical approach was validated against both the live-range tests and subsequent wind-tunnel experiments. A nonspinning half-model and a full, spinning projectile were examined to clarify the influence of rotation. Multiple ground clearances were tested to obtain clear trends in changes to the aerodynamic coefficients, and the three-dimensional propagation and reflection of the shock waves were considered in detail. The behavior of the flow in the near wake was also studied as ground clearance was reduced. Ground proximity was found to significantly increase the drag force acting on the projectile, as well as generate a force normal to the ground and an increased side force, when ground clearance was less than one diameter. For clearances between approximately 0.4 and 1 diameter, the pitching moment produced was nose-down. For lower clearances, a more distinct nose-up trend was produced. The generated side force was orders of magnitude lower than the normal and drag forces.

Aerothermodynamics Behind a Blunt Body at Superorbital Speeds

Results of an experimental, numerical, and analytical study of the hypersonic laminar near-wake flow of a 45° half-angle blunted cone are presented. A superorbital expansion tube facility, which can generate different planetary atmospheric compositions, was used for the experiments. In the present instance, two test gases [air and a mixture of 96%CO 2 -4%N 2 (simulating the Martian atmosphere)] were used. Superorbital speeds above 8 km/s were generated at specific enthalpies of 44.5 MJ/kg (air) and 32.9 MJ / kg (96 % CO 2 -4%N 2 ). Measurements consisted of surface pressure and heat flux, and both showed good agreement with theoretical and numerical predictions. It was also noted that, based on numerical data, chemical reactions seemed to be stronger in the case of the test gas simulating the Martian atmospheric composition when compared with those of air.

The onset of compressibility effects for aerofoils in ground effect

The influence of flow compressibility on a highly-cambered inverted aerofoil in ground effect is presented, based on two-dimensional computational studies. This type of problem has relevance to open-wheel racing cars, where local regions of high-speed subsonic flow form under favourable pressure gradients, even though the maximum freestream Mach number is typically considerably less than Mach 0.3. An important consideration for CFD users in this field is addressed in this paper: the freestream Mach number at which flow compressibility significantly affects aerodynamic performance. More broadly, for aerodynamicists, the consequences of this are also considered. Comparisons between incompressible and compressible CFD simulations are used to identify important changes to the flow characteristics caused by density changes, highlighting the inappropriateness of incompressible simulations of ground effect flows for freestream Mach numbers as low as 0.15.