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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 ([Formula: see text]) with nitrogen, air, argon, krypton, and helium injection gases and mass flow rates ranging from 0.01 to [Formula: see text]. 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 ([Formula: see text]) 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.

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 ±10%. 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.

Shock tube flows can be used to investigate nonequilibrium thermochemistry and radiative processes found in hypersonic flows. The flow in a shock tube contains many flow nonuniformities, in particular boundary-layer effects. For the purpose of studying the properties of the test slug, the flow behind the shock is usually considered analogous to the stagnation line flow over a blunt body. In reality, radial and longitudinal velocity distributions in the two flows are different, and so a coordinate transformation is needed to match time-of-flight profiles. This work develops a method to approximately transform the postshock distance for normal and blunt body shocks to a shock tube flow using a Mirels analysis. The effects of shock tube diameter and shock speed variation are highlighted along with the importance of postshock temperature and density rises. This is then applied to blunt body simulations of NASA EAST data for air and Titan tests. This showed that minimal difference is encountered for fast reacting flows. However, large differences can be seen in the slower reacting Titan tests, which to date have been misinterpreted when compared to blunt body simulations.

Abstract 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

The influence of injection gas type on second mode instabilities is researched on a 7 degree half-angle cone at Mach 7. The wind tunnel model of 594.5 mm length with a sharp nosetip is fitted with a porous aluminium patch that spans 60 degree in azimuth and 25 mm in axial length. Four different gases are injected into the boundary layer flow, namely nitrogen, carbon dioxide, helium and argon. Three different tunnel test conditions with different unit Reynolds numbers provide a variety of amplification levels of the second mode instability at the injection location. Frequency data is obtained using PCB sensors and the density boundary layer thickness is inferred from high speed z-type schlieren images. Early analysis of initial data shows a drop in second mode frequency and second mode power behind the injector. If transition is not caused the effect weakens further downstream. Higher blowing ratios caused a stronger decrease in frequency and also reduced the power in the frequency band more. Larger blowing ratios are required at lower unit Reynolds numbers to achieve the same effect. The observed boundary layer thickness, as inferred from schlieren images, followed comparable dynamics. A larger blowing ratio of the same coolant gas lead to more thickening, while Helium had a significantly stronger total effect than carbon dioxide. Preliminary results suggest that helium injection at a moderate rate causes the second mode peak to be disproportionally damped and its bandpower to be significantly reduced. A strongly simplified analytical model suggest that the mechanical increase in boundary layer thickness alone may not be sufficient to explain all observed outcomes.

This paper focuses on the development process of a floating element shear stress measurement device for testing in a short duration hypersonic wind tunnel. First experiments have been carried out in the Oxford High Density Tunnel at a nominal Mach number of 5 with unit Reynolds numbers 44 - 62×106 m−1 . Testing has successfully provided proof-of-concept demonstration of the measurement of a smooth and three rough surfaces (𝒌 + s ranging from 2.7 to 826) in turbulent hypersonic flows. The measured shear stress for the smooth surface shows encouraging agreement with the theoretically predicted levels, using heat transfer measurements from identical test conditions. Overall better agreement with the predictions was observed for the transient calibration approach.

Arc-jets and impulse facilities have been extensively used to investigate the flow physics of planetary re-entry, but they each address different aspects of re-entry flow-fields. This work presents a conceptual design of a dual-facility setup, utilising an arc-jet as a model preheating device in the impulse test facility. This novel preheating apparatus will replicate realistic ablation coupling effects in hypervelocity flows. This paper outlines the major challenges and key considerations for the system integration of arc-jet in the T6 Stalker Tunnel. A small-scale arc-jet facility, OPG1, is currently being developed at the University of Oxford to test the plasma generator, model movement system, and sub-components for the expansion tube testing. Analyses were performed to examine the thermal and structural loads on the model movement system exposed to both plasma and hypervelocity flow. The results showcased that the model movement system will effectively serve dual-facility operations.

An efficient method is developed for calculating the non-equilibrium properties of the test gas in a shock tube in the shock frame of reference. The one dimensional method is based on the parabolised Navier-Stokes equations, resulting in a form similar to a stagnation line problem but with appropriate consideration of the mass loss to the boundary layer present in a shock tube. Gas properties are determined using Park’s two temperature model. Transport properties are evaluated using second order Chapman-Enskog theory. The centreline solution is coupled to an artificial radial pressure profile which mimics the effect of a boundary layer. The method was tested on a variety of air cases ranging from 5.5 km/s to 9.6km/s, and demonstrated improved modelling of the non-equilibrium regions compared to a Rankine-Hugoniot solver. A 6.1 km/s, 13.3 Pa test case relevant for Titan entry demonstrates the necessity of appropriately modelling mass loss to the boundary layer in a shock tube. The effect of shock structure and mass loss to the boundary layer in a non-equilibrium flow within a shock tunnel is shown to substantially impact the test gas properties and non-equilibrium radiance profiles in the UV/Vis and Vis/IR regions.

This paper presents a methodology for adapting the Infrared Thermography (IRT) diagnostic technique for use on large scale rough surfaces. Situations where more established methods of surface heat transfer measurement, such as thin film gauges (TFG) and calorimeter gauges, are typically employed are met with a considerable challenge when applied to rough walls. In order for the gauges to provide meaningful data, their surface must possess the same topography as the surrounding roughness, which adds to the complexity and expense of gauge construction. Thus, a benefit of IRT over gauge methods is IRT’s ability to provide data in a non-intrusive manner that is easily adapted to any roughness geometry, reducing the necessity for a multitude of different gauges. Furthermore, IRT produces spatially continuous heat transfer data over a given surface area, providing insight into the heat flux augmentation that rough walls experience. This methodology also solves the general problem that arises from the change in directional emissivity that accompanies the varying surface height of rough walls. As this is a consequence of 3D effects, the 2D image produced by IRT does not inherently account for this and so an emissivity map must be evaluated and given as an input. The methodology in this paper is then applied to an idealised two-dimensional roughness pattern and a three-dimensional pattern that represents an ablated surface of HEEET material developed by NASA, showcasing the Stanton number enhancement maps and emissivity maps produced.

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.

An efficient method is developed for calculating the non-equilibrium properties of the test gas in a shock tube. The method is based on the parabolised Navier-Stokes equations, resulting in a form similar to a stagnation line problem but with appropriate consideration of the mass loss to the boundary layer present in a shock tube. Gas properties are determined using Park’s two temperature model. Transport properties are evaluated using second order Chapman-Enskog theory. The centreline solution is coupled to an artificial radial pressure profile which mimics the effect of a boundary layer. The effect of shock structure for a non-equilibrium flow within a shock tunnel is shown to substantially impact the test gas properties and non-equilibrium radiance profiles in the UV/Vis and Vis/IR regions.

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 on a flat plate model in the Oxford High Density Tunnel at Mach 6.1 in both laminar and turbulent regimes. Spatially resolved 2D surface heat transfer measurements were acquired by imaging directly on and downstream of two micro-porous transpiration cooled injectors (METAPOR CE170 and Zirconia) using high-speed infra-thermography. Whilst 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 stream-wise heat transfer distribution on the injector for different blowing rates of Nitrogen and Helium injection. A key result is that Helium performs much better than reported in previous experiments for a turbulent boundary layer.

Two new, longer duration, modes of operation for Ludwieg Tunnels have been implemented in the Oxford High Density Tunnel (HDT): Extended Ludwieg Mode (ELM) and Plenum Augmented Ludwieg Mode (PALM). In standard Ludwieg Tunnel operation, the duration of the steady test time is limited by the return of the rarefaction wave to the facility nozzle throat. Operation in ELM extends the rarefaction wave such that the tail is generated as the head returns and results in a steady decrease in supply pressure. PALM takes advantage of the dual-throat arrangement of the HDT, which features a plenum region between the facility barrel and the nozzle throat, and tailors plug valve opening to produce a steady supply pressure to the facility nozzle at the expense of total pressure capability. This paper describes the theory of operation of these new modes, their implementation in the existing Oxford High Density Tunnel and presents transient quasi-1D numerical simulations confirming their principles of operation. It then presents experimental results from initial testing of the HDT in the new modes. These results demonstrate a successful implementation of ELM, producing a test flow that is over fifteen times longer, but of comparable steadiness to a Mach 7 Ludwieg Mode plateau in the Oxford HDT. PALM was implemented with partial success, resulting in a test flow that is both ten times the duration of, and significantly steadier (for supply pressure and unit Reynolds number), than a Ludwieg Mode plateau. Thus, the implementation of ELM and PALM significantly expand the capability of the facility to investigate long duration flow phenomena.

The thermal effects of plume impingement on the uselage of a hypersonic vehicle were investigated using a plume generator mounted to a flat plate model in the Oxford High Density Tunnel at Mach 7. The model was instrumented with thin film heat transfer gauges, pressure transducers, Pressure Sensitive Paint and thermocouples. Schlieren photography was used to visualize the flowfield. Exit Pressure Ratios of up to 22 were tested, with CO2 as the cold injected gas. Three nozzles were used to explore the effect of changing the expansion profile on heat transfer: a 5 mm conical nozzle, a 6 mm conical nozzle and a 6 mm Rao nozzle. It has shown that initially Stanton number decreases by up to 90% with cold gas injection, but flow separation causes the Stanton number to increase up to 125% of the no injection condition. This significant decrease in heat transfer indicates that there will be direct contact between the rocket engine plume and the fuselage, allowing for significant heat transfer to the fuselage.

A novel mass capture device for testing high speed inlets in short duration facilities has been developed. The device improves upon the accuracy and sensitivity of current methodologies by separating the mass flow measurement from the back-pressure inducing capability. It also has sufficient temporal response for reliable application to short duration test facilities. This paper describes the principles of the operation and the design process of the mass capture device. Calibration was performed using a steady flow bench top test using an orifice mass flow meter. The device’s suitability for use in transient hypersonic flow facilities was validated in the Oxford High Density Tunnel by connection to a simple Pitot inlet that achieves full mass capture.

The design of hypersonic vehicles typically incorporates the use of simple, geometric shapes with smooth surfaces. There are many reasons why aircraft cannot have perfectly smooth surfaces with small, unavoidable imperfections frequently being present. These can appear in many forms, such as gaps between tiles and rivets joining different panels. Surface features like this are mainly detrimental to aircraft performance for two reasons: they can cause premature boundary layer transition, leading to higher integrated heat loads and unexpected aerodynamic loads; and they cause highly localised areas of heat flux augmentation – up to many times the undisturbed level. This work is part of a project which aims to characterise some of these effects using a combination of experimental and numerical techniques with the overall aim being to produce useful engineering level correlations for use in vehicle design. This paper presents initial experiments performed in the Oxford High Density Tunnel (HDT). Heat flux and pressure data have been acquired for a range of step and cavity geometries, and capability of a sophisticated experimental model has been shown.

In the study of atmospheric entry, one of the most challenging environments to model is the lower phase of the transition regime between free molecular and continuum behaviour. This regime, known as slip flow, occurs at Knudsen numbers between 0.001 and 0.1, i.e. when the mean free path of the gas around an object starts to become significant compared to said object’s physical dimensions. The Oxford Low Density Wind Tunnel is a continuous flow facility specifically designed to model this regime (at Mach numbers in the range 5 6 Ma 6 10) which has recently been re-commissioned with a range of new subsystems including upgraded instrumentation. This paper presents experimental flow characterisation measurements for two conditions, specifically Mach 5 flows at Kn = {0.0066˙ , 0.02}, performed using Pitot pressure and total temperature surveys. These results confirm the suitability of the wind tunnel for simulating hypersonic slip flow problems such as those associated with atmospheric entry calculations.

The T6 Stalker tunnel is a multi-mode high enthalpy pulse facility for testing of aerothermodynamics for high speed flight. It operates with a free-piston driver and can be coupled to several different components downstream to become a shock tube, reflected shock tunnel or an expansion tube. This allows for a wide range of testing from subscale model testing to exploration of fundamental high-speed flow processes. Its development was initiated in 2014 with its first commissioning beginning in 2017. Eight years from its conception, has seen its successful commissioning in all modes of operation. Research has been conducted in several fields including shock layer thermochemistry and radiation, convective heating and boundary layer transition. Additional hardware has been developed to expand its capability and operability. This paper will highlight the performance of the facility in each of its modes now that it has been operated against the overall performance map for the facility. Some of the research performed in the facility, including radiation testing, satellite demise and sub-scale planetary probe testing, is also presented.

This paper presents boundary layer instability measurements obtained in The University of Queensland’s T4 Stalker Tunnel at Mach 7. The model geometry is a 7◦ half angle cone with a nose bluntness of 1.55 mm radius, pitched at an angle of attack of zero degrees. Results are presented at a constant enthalpy (H0 = 2.0–2.4 MJkg−1 ), a sweep of unit Reynolds numbers (Reunit,∞ = 5.0–12.0 × 106 /m), and three wallto-total temperature ratio distributions (Tw/T0 = 0.12 – 0.19). The cone nosepiece is instrumented with a DC powered heater cartridge, allowing it to reach temperatures up to 433 K at the tip. Continuous wavelet analysis and power spectral density results suggest that a disturbance may exist at a frequency of ≈ 750 kHz. The frequency and amplitude of the suspected disturbance did not appear to change for the limited Tw/T0 distributions tested. The results suggest that only the early stage of growth was detected, with a maximum amplitude of 8.1% reached.

Free-flight experiments have been conducted within the University of Oxford’s High Density Tunnel (HDT) with a subscale model of Reaction Engines’ Skylon spaceplane. The purpose of this study was two-fold; firstly, to provide the first high Reynolds number experimental aerodynamic data of the Skylon vehicle for comparison with prior public-domain numerical studies; secondly, to implement and commission the necessary infrastructure to enable free-flight experiments to be conducted within the HDT. Experiments were conducted at a Mach 7 test condition replicating flight at an altitude of 63.5 km. Results for lift, drag and pitching moment coefficient were obtained over a range of angles of attack. Lift coefficient is generally linear with high angles of attack and drag coefficient is approximately quadratic in shape but was sensitive to model yaw. The vehicle was also shown to be statically unstable, a common characteristic of canard configuration vehicles.

This paper presents a numerical model which assesses the coupled effect when transpiration cooling is applied to the external skin of a thermal protection system and the substructure of a high speed flight vehicle. The PIRATE transpiration cooling code has been extended and validated to account for quasi-two-dimensional lateral conduction effects allowing for analysis of more complex geometries. This enables very fast calculations of the two dimensional transient temperature response of a transpiration cooled thermal protection system, ideally suitable for large scale systems studies. To solve for the coupled transpiration cooled material and two-dimensional substructure, PIRATE has been coupled with the commercial finite element package COMSOL. This enables modelling of the longer term thermal effects of the integrated heat load over a flight trajectory. Transpiration cooling has been applied to a simplified blunt body model with an aluminium substructure. Results for the substructure temperature history for the Space Shuttle trajectory are obtained showing that transpiration cooling can lead to a 40% reduction in peak substructure temperature as well as a reduction in thermal stresses near the stagnation point.

To provide data for the design of re-entry vehicles and validation of computational models, ground testing facilities such as the Oxford T6 Stalker Tunnel are crucial. The current work focuses on the commissioning of the tunnel in reflected shock tunnel mode. For a given test condition, the fill pressures and temperatures in the reservoir, compression tube and shock tube are calculated to ensure a tailored condition and tuned piston motion. A 90kg Steel piston has been refurbished for this mode of operation to maximise testing time. In order to verify the loss factors used in the numerical modelling of the tunnel with the refurbished piston, blank-off tests were undertaken and the losses taken into account to adjust the fill conditions accordingly. Shock tube tests were undertaken to determine the necessary score depth of the diaphragms to achieve the required rupture pressure. This was performed for a nominally Mach 7 nozzle, with the test conditions of 2.4 MJ/kg total enthalpy and unit Reynolds number of 10 million/m. A Pitot survey was undertaken to fully characterise the test flow. Reflected shock tunnel mode was commissioned successfully, with the final condition being within 4.3% and 1% of the nominal nozzle supply pressure and enthalpy respectively. The core flow diameter was over 160 mm with a test time of approximately 0.5 ms.

This paper presents the performance of pressure sensitive paint (PSP) for the direct investigation of a transpiration cooled surface. This technique allows the quantification of the pressure distribution on a porous surface with blowing. Additionally, it can be used to detect where the coolant is on the surface, thus measuring the transport of molecular oxygen to a transpiration cooled surface. For highly turbulent flows, it can also be used to evaluate the film effectiveness. A porous aluminium sample was anodised and dip coated in a PSP luminophore solution. It was fitted into a flat plate model and exposed to a Mach 5 cross-flow in the Oxford High Density Tunnel. Tests were conducted with no coolant injection, air injection and with nitrogen injection at increasing blowing ratios. The film effectiveness of the transpiration cooled surface was obtained for several conditions at Re = 15.5 - 31.5 10^6/m and F = 0.001 - 0.002. The film effectiveness increases as the Reynolds number decreases and the blowing ratio increases, which is in good qualitative agreement with the literature. Furthermore, it shows the same features as a velocity map of the outflow.

As part of the EU LAPCAT II programme, the University of Oxford tested and analysed a 1:100 scale Scramjet intake for starting capabilities. Tests were performed at a Mach 7 condition in the Oxford Gun Tunnel, with a unit Reynolds number of 20x106/m. The model was instrumented with 61 measurements of wall static pressure throughout the intake and isolator, 24 thin film heat transfer gauges on the end of the intake and start of the isolator and video Schlieren of the cowl closure region of the inlet. A comparison of the intake wall pressures showed that the LAPCAT MR2 intake exhibited higher intake pressures than the tests completed by DLR using the same geometry and flight condition. Analysis of the Schlieren and the pressure data has revealed that the most likely cause of the higher inlet pressures is a small separation that originates from the corner at the mid-plane of the intake. With a reduced exit area 25.4%, which is analogous to an increased back pressure from the combustor, a normal shock sits stable in the isolator. At extreme AoA (i.e. above +4°), the external part of intake shows a very low pressure as it is being processed by a Prandtl-Myer expansion, resulting in low pressure in the isolator with minimal external compression achieved. Testing showed that the intake was capable of self-starting across the range +4 degrees to -12 degrees AoA. This was observed through rapid start-unstart or buzz on the intake once a sufficient level of blockage was applied. Body wall pressure, body heat flux and crotch area Schlieren for various times throughout the inlet start/unstart/start process will be presented and discussed in the final paper. The presence of intake buzz confirms that this geometry is self-starting.

This paper presents measurements of equilibrium radiation obtained in the NASA Ames Research Center's EAST facility and the University of Queensland's X2 facility. These experiments were aimed at measuring the level of radiation encountered during conditions relevant to Orion lunar return into Earth's atmosphere. The facilities have targeted the same nominal test conditions of 10 km/s and 26.6 Pa (0.2 Torr). In addition, variations on the nominal shock speed have also been the focus of recent testing in the EAST facility. A comprehensive comparison between the EAST data and NEQAIR is presented in this paper with preliminary X2 comparisons where appropriate. Since the two facilities have different dimensions, and the tests have different shock speeds, NEQAIR simulations are used as a point of reference for the EAST and X2 comparison. Results obtained by independently reducing the data from both facilities are compared. The present analysis endeavors to provide a better understanding of the uncertainty in the measurements, as well as provide an initial comparison between EAST and X2. Furthermore, the present analysis explores various radiative mechanisms to determine if they are due to physical processes relevant to flight, or are just facility dependent phenomena. These phenomena include effects such as the magnitude of the background continuum. © 2010 by the American Institute of Aeronautics and Astronautics, Inc.

A Diamond Heat Transfer Gauge has been used to measure the stagnation point heat flux at three flow conditions in an arc heated plasma wind tunnel. The gauge was coated with a thin layer of copper to provide a known surface catalycity and reflectivity. The temperature dependent impulse response of the probe was determined from a finite element model of the probe, and from the result of a laser-based calibration experiment using the Non-Integer System Identification method. The two sets of impulse responses were used to calculate heat flux from a measured temperature rise in the wind tunnel. Differences between the impulse responses from the two methods, particularly at higher initial temperatures, led to differences in measured heat fluxes of up to 55 %. A measurement taken using a steady state calorimeter gauge did not conclusively agree with heat flux measurements calculated using impulse responses from either method.

This paper presents transpiration cooling experiments conducted in the Oxford High Density Tunnel. The flow structure and mixing phenomena between the injected coolant gas and the hypersonic free stream are investigated. A flat plate model is used which is equipped with a gas injection system, four surface mounted heat flux gauges and three surface pressure sensors located downstream of the injection point. Mass injection is realized through a porous ceramic material, and alternatively through a series of slots. Spatially and temporally resolved film effectiveness downstream of the injection point is measured with fast response pressure sensitive paint. The flowfield is visualised by a Z-type Schlieren imaging system. Results show a significant influence of injection on the boundary layer thickness through the additional coolant gas. High blowing ratios can lead to early transition which increases coolant-external flow mixing and leads to a lower film effectiveness. Transpiration cooling requires approximately two orders of magnitude lower blowing ratios to achieve a similar film coverage as in the case of the injection through slots. The film effectiveness increases with an increasing blowing ratio for locations close to the injection point. However, a lower blowing ratio results in a more stable film for locations further downstream.

The Non-Integer System Identification (NISI) method is an useful tool for the determination of surface heat flux from in-depth temperature data even in transpiration cooled environments. Recently it was demonstrated that in these transpiration cooled environments the surface heat flux can be determined using a measurement of the plenum pressure changes in the gas reservoir at the rear side of the porous structure for the Non-Integer System Identification (NISIp). This paper presents fundamental experiments required for the identification of susceptibilities with respect to the thermophysical and fluid properties. Test results with two different porous materials are compared: zirconium diboride and carbon/carbon. The experimental setup for both systems and the respective resulting impulse responses over varying coolant mass flows are presented and discussed in this paper. It was found that the pressure impulse responses of the ZrB2 system is more sensitive to coolant mass flow rate changes than the C/C system with respect to both response time and amplitude pressure. However, the absolute response time of the C/C system was significantly shorter for all tested flow rates.

Two methods for calculating surface heat transfer rates from sub-surface temperature measurements are detailed for a robust fast-response gauge, designed for use in high enthalpy hypersonic ground test facilities. One method treats the gauge as a calorimeter, and the second is a pseudo non-linear version of the Sequential Function Specification Method algorithm, which calculates surface heat transfer by deconvolving the unit impulse response of the gauge out of a measured temperature. An experiment to find the impulse response of the gauge using a short duration pulsed laser to apply a heat flux to the surface of the gauge is performed. The response at three initial temperatures is recorded to characterise variation in thermal properties as a function of temperature. The use of the experimentally determined unit impulse response in the calculation of surface heat transfer can significantly reduce measurement uncertainty by removing the need to assume physical geometry and material properties of the gauge.

Initial experimentation using X2 to simulate a Mach 13 scramjet flight condition indicated the requirement to develop a tuned free-piston driver. This involves running the piston at much higher velocities around the time of diaphragm rupture in order to sustain high driver pressures for significantly increased duration. Central to the tuned driver is a lightweight piston, which enables the piston to be accelerated to high speeds, and subsequently decelerated, over the same fixed length of compression tube. This paper discusses the background to the tuned driver requirement for X2, and describes the design, stress analysis, and final configuration of a new lightweight aluminium piston. Experimental results for X2 using the new tuned driver are presented, indicating the successful implementation of the concept.

Computational simulations of the AIR-1 test condition in the University of Illinois’ Hypervelocity Expansion Tube were conducted to verify facility operation and to obtain free stream properties that are otherwise difficult to measure. Two types of simulation were undertaken. The first was a one-dimensional simulation of the entire facility and the second was a hybrid simulation, combining a one-dimensional simulation of the shock tube section with a two-dimensional simulation of the acceleration tube. The one-dimensional simulation matched the experimental data well, however the two-dimensional simulation did not initially match the experimental measurements of shock speed and test gas pitot pressure. Further investigation showed the shock speed discrepancy was consistent with air contamination into the acceleration tube and subsequent two-dimensional simulations assuming 10% air contamination showed reasonable agreement with experimental data. Using data taken from the two-dimensional simulation of the expansion tube as a transient inflow condition, modelling was undertaken of a free shear layer experiment being conducted in the facility. Results from equilibrium, finite rate, and perfect gas models were compared. The finite rate simulation provides the best agreement with experimental Schlieren images, with the simulation capturing the major flow structures seen in experiments.