High Density Tunnel Research Articles

Comparison of force measurement techniques in a short duration hypersonic facility

Force measurement experiments have been conducted within the University of Oxford’s High Density Tunnel with a 7∘\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$7^{\\circ }$$\\end{document} half-angle cone. The purpose of the study was to provide a direct comparison between two independent force techniques in the same facility at the same free-stream conditions to provide a quantitative and qualitative discussion of the advantages and disadvantages of both techniques. The first force measurement technique used a conventional 4-axis sting-mounted force balance which was calibrated both statically and through the stress wave deconvolution method, whilst the second technique used the less established static free-flight methodology. Experiments were conducted at a Mach 5 test condition which provided sufficient dynamic pressure to generate aerodynamic forces suitable for the measurement range of the force balance. Results for lift, drag and pitching moment coefficients were obtained over a range of angles of attack and compared with predictions from a hypersonic panel method code. Agreement between the independent force techniques and numerical data sets was good over the range of angles of attack. Maximum uncertainties were shown to be 38.46 ± 0.56 N and 22.52 ± 0.44 N for free-flight in lift and drag, respectively, and 38.74 ± 1.59 N and 22.13 ± 1.27 N for the dynamically calibrated force balance which demonstrates the superiority of free-flight.

Investigation of fluidic thrust vectoring for scramjets

Fluidic thrust vectoring (FTV) offers a novel approach to aerodynamic control, circumventing some of the issues associated with mechanical systems. One method is shock vector control which involves injecting a fluid into the exhaust nozzle of an engine to redirect the gases and thus, produce a control force. An experimental model which incorporated FTV was designed and tested at Mach 6 in the Oxford high density tunnel (HDT). The model was a simplified two-dimensional scramjet geometry with two different configurations to compare an internal and external exhaust nozzle. The FTV injection system consisted of a slot at the rear edge of the exhaust nozzle fed from an internal plenum. In the experimental campaign, a range of gas injection pressures and free stream stagnation pressures were tested to assess the effectiveness of both configurations. Two new measurement methods were successfully implemented in the HDT: pressure sensitive paint and a 6-axis load cell. The FTV system has been shown to be effective with observable increases in lift and pitching moment. A linear relation between the injection pressure ratio and the control forces could be observed for both configurations.Graphical abstract

Experimental Studies of Hypersonic Shock Impingement on a Transpiration-Cooled Flat Plate

Hypersonic vehicle design requires mitigation of the high heat fluxes present in regions of shock-wave/boundary-layer interactions. A candidate technology that may be applied locally to these regions is transpiration cooling. In this work, experiments were conducted in the University of Oxford’s high-density tunnel at Mach 6.1 in both laminar and turbulent undisturbed boundary-layer regimes where a 10 deg shock generator impinged a strong oblique shock wave onto a transpiration-cooled microporous injector. For the laminar boundary layer, due to the strength of the incident shock, a transitional shock-wave/boundary-layer interaction region was formed with peak heating over 50 times greater than the nominal laminar level. Both nitrogen and helium were used as coolants. Relatively low levels of helium injection of for the transitional and for the turbulent scenarios were sufficient to reduce the heat transfer downstream of shock interaction to approximately 50% of the value without cooling. In fact, helium is highly effective with a similar cooling performance achieved as eight times the equivalent mass flux of nitrogen. The experimental data are correlated, and both the turbulent and transitional shock-impingement scenarios display a similar trend of reduced surface heat transfer with higher blowing parameters. Empirical fits are proposed that may be used for initial systems design.

Transpiration-Cooling Heat Transfer Experiments in Laminar and Turbulent Hypersonic Flows

The design of a transpiration-cooled system requires detailed local heat transfer information on and in the vicinity of the porous injector; however, limited spatially resolved experimental studies exist, particularly in hypersonic flows. In this work, experiments were conducted in the University of Oxford’s high-density tunnel at Mach 6.1 in both laminar and turbulent regimes. Spatially resolved two-dimensional surface heat transfer measurements were acquired by imaging directly on and downstream of two microporous transpiration-cooled injectors (METAPOR® CE170 and zirconia) using high-speed infrared thermography. Whereas injection in the laminar regime results in a steady, monotonic reduction in heat transfer from the start of the injector, a flatter profile is present for the turbulent cases where turbulent mixing inhibits surface heat transfer reduction. It was found that a modification to existing relations from film theory successfully correlates the streamwise heat transfer distribution on the injector for different blowing rates of nitrogen and helium. A key result is that helium performs much better than reported in previous experiments. Finally, the downstream thermal effectiveness is characterized for turbulent flows. A collapse of the thermal effectiveness is achieved and a modified analytical correlation proposed.

Heat Transfer Measurements of a Transpiration-Cooled Stagnation Point in Transient Hypersonic Flow

This paper presents a novel experimental technique where infrared thermography is employed to directly measure the surface heat transfer of a transpiration-cooled porous material in transient hypersonic flow. Experiments were conducted in the Oxford High Density Tunnel on a flat-faced hemispherical probe at a single Mach 7 freestream condition () with nitrogen, air, argon, krypton, and helium injection gases and mass flow rates ranging from 0.01 to . Surface heat transfer measurements were extracted by imaging directly on the porous material using a FLIR A6751 high-speed long-wave infrared camera. Porous alumina was chosen due to its favorable thermal properties for infrared analysis and its very small pore sizes () enabling a uniform outflow. It was found that the surface Stanton number reduction matched to within 10% of both computational fluid dynamics results and correlations.

Numerical Simulation of Transpiration Cooling in a Laminar Hypersonic Boundary Layer

Two-dimensional simulations of transpiration cooling in a laminar, hypersonic boundary layer were performed using the thermochemical implicit nonequilibrium algorithm (TINA): a Navier–Stokes solver. Coolant concentration and heat flux results are compared to data obtained from laminar transpiration cooling experiments conducted in the Oxford High Density Tunnel employing a flat-plate geometry at Mach 7. TINA successfully predicts the mixing rate at the wall as a function of the streamwise direction for all blowing ratios. The simulations are more successful in predicting the mixing downstream of the injector as compared to the mixing on the injector: especially at low blowing ratios. A collapse of the thermal effectiveness values calculated from TINA simulation data is achieved, which agrees with laminar correlations within an absolute value of . It is shown that, when the concentration effectiveness is close to one at the injector, the temperature gradient becomes negative at locations immediately downstream of the injector, resulting in a negative heat flux. The acceleration of the coolant in the streamwise direction downstream promotes dissipation of energy, which results in a reduction in the temperature of the coolant, and thereby induces a negative temperature gradient close to the injector.

Laminar transpiration cooling experiments in hypersonic flow

The mixing between the coolant and the boundary-layer gas downstream of an injector—for transpiration/film cooling—has been extensively studied for turbulent flows; however, only a handful of studies concerning laminar mixing exist, particularly in hypersonic flows. In this paper, the concentration of the coolant gas at the wall and the heat flux reduction downstream of a transpiring injector in a hypersonic laminar flow are experimentally measured and examined. Experiments are performed in the Oxford High Density Tunnel at Mach 7. A flat-plate model is coated with pressure-sensitive paint (PSP) to spatially resolve the film and obtain a film effectiveness based on coolant concentration. Thin-film arrays are installed to measure the heat flux reduction. Six different cases are studied featuring nitrogen and helium as the coolant gas, where the blowing ratio is varied from 0.0406% to 0.295%. The unit Reynolds number of the flow is 12.9times 10^6;mathrm {m^{-1}}. A coolant concentration of up to 95% is achieved immediately (2 mm) downstream of the injector. The film concentration drops in a monotonic fashion farther downstream; however, a constant film coverage of 5–20 mm immediately downstream of the injector is observed in cases with a higher blowing ratio. A film coverage above 15% over three injector lengths is present even for the lowest blowing ratio. Heat flux reduction is achieved in all cases. The concentration effectiveness obtained from PSP is compared with the thermal film effectiveness calculated from the heat flux reduction. The latter is found to be higher than the former for all data points. Finally, a collapse of the thermal effectiveness is achieved and a modified analytical correlation is proposed.Graphical abstract

Free-Flight Aerodynamic Testing of the Skylon Space Plane

The determination of aerodynamic coefficients for complex high-speed vehicles still requires experimental measurement. This paper details the development of the free-flight measurement technique within the University of Oxford High Density Tunnel. In particular, a novel image processing technique is developed for calculating the position of the model from high-speed video. This study focuses on the measurement of high Reynolds number experimental aerodynamic data for a subscale model of the Skylon space plane of Reaction Engines. Testing was undertaken at a Mach 7 test condition replicating flight at an altitude of 63.5 km. Results for lift, drag, and pitching moment coefficients were obtained over a range of angles of attack. Lift coefficient was nearly linear over the range of angles of attack tested, and drag coefficient was parabolic in shape, but sensitive to model yaw. The vehicle was also shown to be statically unstable, a common characteristic of canard configuration vehicles.

Influence of instability modes on cooling performance in hypersonic boundary layer with slot injection

A combined numerical-experimental investigation is presented with focus on the effects of boundary-layer instabilities and transition on the wall cooling performance in a Mach 5 low-enthalpy flow over a flat plate, with coolant injection achieved through a row of slots. The numerical study has been performed through direct numerical simulation (DNS) of the compressible Navier-Stokes equations, and is supported by results from linear stability analysis (LST) for the considered boundary layer. The experiments have been conducted in the High Density Tunnel (HDT) of the Oxford Thermofluids Institute, and include several blowing ratio conditions of injected air for the same freestream conditions. Surface heat transfer and pressure measurements, film effectiveness measurements, and Schlieren images are presented. The analysis links the wall cooling performance to the growth of imposed unstable boundary layer modes. Results indicate that 2D and 3D unstable modes, pertaining to the class of first instability modes, exist in the laminar boundary layer, and that imposition of these modes at different amplitudes leads to different states of the boundary layer, which we refer to as a perturbed state and a transitional state for medium and high amplitude respectively. As confirmed by comparison with experimental data, the perturbed and transitional states of the boundary layer significantly affect the wall cooling performance, providing an increase of the wall heat flux that results in a reduction of the beneficial effects of cooling.

Flowering and fruit set in cherimoya (Annona cherimola Mill.) as affected by the tree-training system

SummaryManagement of flowering and fruit set are of prime importance in order to obtain high yields in cherimoya (Annona cherimola Mill.) due to floral dichogamy. This phenomenon is largely dependent on environmental conditions such as temperature, relative humidity, and light intensity, which usually make hand-pollination of flowers necessary in order to achieve acceptable levels of fruit set and yield. Nevertheless, these factors can be modified by means of the growing system. In this work, we studied the effects of two tree-training systems on the environment inside the canopy, on flowering, on pollination, and on fruit set in mature ‘Concha Lisa’ cherimoya trees. The open-centre vase training system (6 m 4 m) and the high-density slender-spindle (tunnel) training system (4 m 1 m) were used. The high-density tunnel training system increased the relative humidity, as well as reducing the temperature and the level of photosynthetically active radiation, inside the canopy, which resulted in an earlier onset of the male floral stage and in an extended maintenance of the highly-receptive female stages. This resulted in an overlap of the sexuallycompatible floral stages, thus increasing the incidence of natural cross-pollination and maintaining the viable, receptive sexual stages and pollen, over time, to improve hand-pollination.

Design and Test of a Modular Waverider Hypersonic Intake

*† A study of some modular stream-traced hypersonic intakes is presented. These intakes capture a part of a known flowfield by replacing the streamlines that pass through the perimeter of a sector capture shape with solid boundary layer-corrected surfaces. Modules are combined to form an engine with circular external profile and with each intake module feeding a separate combustor. The intakes studied are derived from internally compressive, axisymmetric flowfields with either constant slope or constant pressure boundary conditions. The number of modules varies between three and six. Stream thrust averaging is used to determine equivalent one-dimensional flow for input to a scramjet engine cycle analysis. For a given contraction ratio, the full mass capture engine performance of the modular waverider intakes is comparable with the performance of isentropic spike intake designs when combustor skin friction is neglected. However the low wetted area of the modular combustor design has a significant advantage when combustor skin friction is considered. A four-module waverider intake model was built using stereo lithography and tested at Mach 7 in the University of Oxford gun tunnel, and was shown to be self-starting in QinetiQ’s High Density Tunnel at Mach 7. The stream thrust averaged properties were determined experimentally by measurement of axial force and mass capture. Schlieren photographs and surface pressure measurements are presented. Skin friction in the model isolator section was found to affect intake performance by increasing the amount of compression and reducing efficiency.

Design and Test of a Hypersonic Isentropic-Spike Intake with Aligned Cowl

A study of reexpansion scramjet intakes is presented. Reexpansion intakes employ mixed external-internal compression with the internal contraction fixed by starting requirements and a cowl aligned with the external flow to minimize intake contribution to vehicle drag. As a result of the starting limitation, the captured flow is partially reexpanded, resulting in nonuniform flow at the intake exit. Stream thrust averaging was used to determine equivalent one-dimensional flow for input to a scramjet engine cycle analysis. Engine performance increases with compression, and thus optimum turning-angle appears to be at the limit for detachment of the internal cowl shock. An axisymmetric reexpansion intake was built using stereo lithography and tested at Mach 7.1 in the University of Oxford gun tunnel. The stream-thrust-averaged properties were determined experimentally from measurements of axial force and mass capture. Schlieren and surface-pressure measurements were taken. Skin friction in the model isolator section was found to significantly affect intake performance, increasing the amount of compression and reducing efficiency. Starting characteristics were determined experimentally in QinetiQ’s High Density Tunnel.