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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.

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.

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.

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.

A robust, fast response calorimeter based heat transfer gauge called the Diamond Heat Transfer Gauge (DHTG) has been developed for use in transient hypersonic ground test facilities. Desirable response time and robustness are achieved by using a 75 µm to 300 µm thick layer of synthetic diamond as the calorimeter and a platinum thin film resistance thermometer to measure the temperature rise. Different versions of the gauge have been shown experimentally to have a rise time of between 7 µs and 28 µs. The gauges have measured peak heat fluxes up to 300 MW/m2 in flows with total enthalpies of 50 MJ/kg over multiple shots in the X2 expansion tunnel at the University of Queensland. Experimentally measured heat fluxes have been shown to compare well with surface junction thermocouples and thin film heat transfer gauges.

Expansion tubes are important facilities for the study of high enthalpy hypersonic flows which avoid the non-equilibrium chemical and thermal effects associated with the flow stagnation intrinsic to reflected shock tunnels. X3 is one of the largest freepiston super-orbital expansion tube in the world with an overall length of approximately 69 m and is capable of generating reentry speed flows equivalent to those experienced during a hyperbolic re-entry trajectory. It was originally built with a twostage free-piston driver to achieve the high compression ratio of a large diameter compression tube without the high construction costs of designing the large diameter tube to be strong enough to resist peak driver pressure loads. However, this arrangement proved difficult in operation. This paper describes the upgrades to X3, in respect to its physical layout. The facility has been recommissioned to incorporate a single-piston driver, a steady expansion nozzle and a new test section. Major changes have been made to the free-piston driver with a re-designed piston and launcher and a new end cap tube which is 200 mm thick to contain driver pressures up to 80 MPa. The re-designed piston introduces an area change at the primary diaphragm, ensuring that the maximum increase in total pressure and temperature can be gained as the driver gas undergoes unsteady expansion from sonic to supersonic conditions. The compression process steadily increases up to Mach 1 at the throat then gains of up to an order of magnitude in total temperature and pressure can be realised as the unsteady expansion process takes over. The area change will also increase test times; with a throat at the primary diaphragm, the piston mechanics can be more readily tuned to minimise reflection of waves off the piston which would otherwise reduce the test time. A new Mach 10 steady expansion nozzle has been developed which has increased the core flow and the test time for appropriate conditions. The dump tank has been replaced with a larger tank and test section giving a larger volume with greater potential for instrumentation.

Modelling of the operation of the Oxford High Density Tunnel (HDT) was performed to establish the level and duration of test gas heating required in the facility and the thermal structural implications of this heating. 1-D gas dynamic simulations of the transient flow conditions in the facility were performed for a range of starting pressures and tunnel configurations demonstrating potential test times of 80ms at M6 and unit Re = 1.45e8m-1. Gas-state analysis was used to establish the liquefaction limits for the oxygen and nitrogen components of the air test gas at the outlet of the nozzle to quantify the heater requirements (power and duration) for the facility. The HDT is required to contain high-temperature and high-pressure test gas safely for extended periods to ensure uniform heating and transient FEM simulations showed that the structural and thermal-structural loads on the facility during its heat up and operation were well within the static limits of the facility. FEM was also used to perform transient simulations of the heat loss from the test gas during the experimental runs. CFD simulations of the buoyant mixing were performed to establish the duration of preheating of the test gas required in the facility to ensure gas temperature uniformity prior to operation of the tunnel. This was found to be of the order of minutes.

In this thesis, the free-flight force measurement technique is commissioned, validated and implemented with increasing complexity in geometries within the University of Oxford High Density Tunnel, a short-duration, heated Ludwieg Tube. The motivation of the work was to improve the current capability of force measurement techniques through the free-flight methodology, allowing for the measurement of high quality, flight representative forces with lower uncertainties than established force measurement techniques. Rather than rigidly mount a sub-scale model through a sting and measure forces through a load cell, the free-flight technique allows the model to move in six degrees of freedom during the test time. This allows for forces and moments to be measured that are not impacted by the mounting structure but require non-intrusive methods to measure the model's kinematics. In this work, kinematics are measured through a 3 degree of freedom image tracking algorithm with sub-pixel accuracy and through accelerometers mounted internally to the model on an on-board data acquisition system. This thesis introduces the static free-flight technique, whereby the aerodynamic pitching moment is minimised by matching the location of centre of gravity to centre of pressure, reducing the static margin of the model. This technique limits the dynamic influence on the measured aerodynamic coefficients and allows for only the static derivatives to be measured. This technique was conducted with a 7 degree half-angle cone, a simple geometry in which the measured coefficients can be validated against a numerical panel method code, giving confidence in the methodology. The tests were conducted at a Mach 7 condition representative of a hypersonic vehicle’s trajectory at 35 km altitude. Non-intrusive image tracking and on-board inertial measurement units were used to determine accelerations to calculate the aerodynamic forces acting on the cone. 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. The experimental and numerical data sets agreed well over the range of angles of attack with the experimental uncertainties remaining below 3.1 \% for all coefficients. Following validation of the static free-flight technique, tests were conducted to directly compare free-flight against force balance techniques. The force balance was calibrated using two methods; a static calibration and by dynamically calibrating the balance by generating a global impulse response function of the model and using the stress wave deconvolution methodology. The experimental model was a blunted 7 degree half-angle cone and 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 +/- 0.56 N and + /- 0.44 N for free-flight for lift and drag respectively and + /- 1.59 N and + /- 1.27 N for the dynamically calibrated force balance. Finally, after showing that the free-flight methodology in the High Density Tunnel resulted in lower uncertainties than other force measurement techniques, it was applied to a 7 degree half-angle cone with fins, at a flight-representative Mach 7 condition. The fins resulted in a centre of pressure that was a function of angle of attack. Longitudinal static force and moment coefficients were measured at the Mach 7 condition with the experimental data agreeing well with numerical predictions. Forces were obtained using both image tracking at 76,000 fps as well as the direct measurement of accelerations using an on-board inertial measurement unit. Uncertainties remained below + /- 0.0075 for lift coefficient and + /- 0.0051 for drag coefficient. The fins were also angled to allow the model to roll in free-flight so that roll coefficient could also be determined.

This thesis presents results from the first experiments conducted in the Oxford T6 Stalker Tunnel in Reflected Shock Tunnel mode and observations on the boundary layer instabilities established over a cone model at Mach 7 and Mach 8 conditions characterised using high speed schlieren image processing techniques. The Oxford T6 Stalker Tunnel was developed and commissioned in a new mode of operation, Reflected Shock Tunnel mode which was crucial to the establishment of boundary layer instabilities over a model. The process by which this was done is outlined, from an initial analytical model to a quasi one-dimensional facility simulation. Piston dynamics were modelled and characterised by blank-off experiments, followed by validation of the diaphragm properties with tests done in shock tube mode. Finally, reflected shock tunnel experiments at half and then full pressure were conducted. This was used iteratively to improve the test conditions, bringing them to within 1% of the required nozzle supply enthalpy and 2.6% of the required nozzle supply pressure. A nitrogen driver condition was developed to improve testing time and these conditions were used with both a Mach 7 and Mach 8 nozzle. Traditional sensors were not reliable at the high frequencies of instabilities expected (∼1 MHz) and would be damaged in the harsh flow environment. Thus, optical diagnostic techniques were explored as a means of extracting both quantitative and qualitative information from the flow. These techniques were developed based on high speed schlieren imaging and first applied to the Oxford High Density Tunnel, which is a cold flow facility for which pressure transducers could be used as a means of validating the developed techniques. Development of schlieren image processing techniques both for frequency and wavepacket convection speed detection as well as feature detection was undertaken. These methods were applied to experiments on a 7deg half-angle cone with a nose tip radius of 1.25 mm at Mach 7. Second mode instabilities were detected at a frequencies of 208-216 kHz with pressure transducers which allowed for validation of the image processing techniques. Nonlinear interactions were explored with the use of selected image processing techniques showing significant second harmonics in the flow. An investigation was conducted using three different high speed cameras and the effect of varying camera parameters on the quality of the data extracted. Methods of extracting the frequencies of second mode instabilities using camera frame rates at much lower and much higher sampling rates than the expected frequency of the instabilities were developed and applied with varying levels of success. Useful results were gained from these experiments, both as a review of the application of current schlieren processing methodologies, novel methodologies, and a comparison of different camera parameters which would inform future testing. Finally, both the commissioning work on the facility and the developed image processing techniques culminated in establishing and observing excellent first images of boundary layer instabilities in T6, namely high frequency second mode instabilities and non-modal instabilities at Mach 7 and Mach 8. Frame rates of 1-2 million frames per second allowed for capture of these instabilities. The extracted frequencies of these instabilities ranged from 0.8-1.4 MHz, at significantly higher frequencies than those observed in HDT. The methodologies allowed for the extraction of frequencies within 1-7% from the theoretical prediction. A blunt nose tip of 3.4 mm was used, allowing for the establishment of non-modal instabilities which could be visualised. These were fully characterised in terms of their convection speed and structure angle. These instabilities were swept along at just under the edge velocity (0.93-0.94Ue). Interesting results were obtained by tracking and identifying the structure angle variation in time, which clearly showed the instability collapsing from a maximum angle of 30deg to a more linear feature at approximately 8deg to the surface. The results show significant progress in the development and study of hypersonic boundary layers in the T6 high enthalpy facility and allows for the extraction of accurate boundary layer information in harsh flow conditions without the use of traditional pressure sensors. This enables the UK to have the capability of high enthalpy testing for the study of hypersonic vehicle models and future boundary layer instability research at flight representative conditions.

Surface heat flux measurements in hypersonic pulse-flow facilities, such as Reflected Shock Tunnels and Expansion Tunnels, provide vital information for the design of hypersonic and hypervelocity vehicles and for the understanding of these complex flow environments. However, test-times in these facilities are on the order of microseconds to milliseconds, and by nature of their operation, the flow contains particles of metal diaphragms, resulting in damage to instrumentation. Additionally, at high flow enthalpies, ionised species and free electrons are present in the gas. Given these significant challenges, this thesis describes the design and manufacture of a new surface heat flux gauge, called the Diamond Heat Transfer Gauge (DHTG), and presents new data reduction methods, the results of calibration experiments, and two experimental campaigns in pulse-flow facilities. The DHTG is shown to overcome limitations of current instrumentation by increasing robustness, and improving response time, electrical isolation, and accuracy. Four types of DHTG have been manufactured, consisting of a 2.5 mm diameter disk of synthetic diamond, either 75 µm, 150 µm, 220 µm or 325 µm thick, mounted into a cylindrical steel housing, with a platinum thin-film resistance temperature detector sputtered on the rear surface of the diamond to measure its temperature rise. An analysis of the gauge as a calorimeter showed that the rise-time of the four types of DHTG are 2.4 µs, 11.6 µs, 36.3 µs, and 54.4 µs respectively, and that the measurement error due to the data reduction method grows with time, limiting measurements to 500 µs for a 1 % drop in accuracy. While the rise-time is sufficiently fast for use in pulse-flow facilities, in order to extend the useful test-time and enable accurate measurements in a wider range of facilities, a non-linear implementation of the Sequential Function Specification Method (SFSM) was developed as an alternative data reduction method. This method deconvolves the impulse response of the gauge from a measured temperature rise to estimate the surface heat flux, and using a numerical study is shown to be accurate to within 0.5 % over a typical 2.5 ms Reflected Shock Tunnel test-time. Additionally, for high signal-to-noise ratio measurements, the SFSM method can improve the rise-time of the DHTG. The typical experimental measurement uncertainty of the DHTG was calculated to be 7.5 % and is primarily driven by uncertainty in the specific heat capacity of diamond. This is comparable to or better than values of uncertainty for existing techniques, such as thin-film gauges and coaxial thermocouples. An experiment to measure the impulse response of the DHTG to reduce measurement uncertainty further was performed using a laser to provide a short-duration pulse of heat flux. Issues stemming from spatial non-uniformity of the laser intensity and measurement uncertainty meant that the measured response was unsuitable for direct use in the SFSM data reduction method. No significant difference was found between the specific heat capacity of diamond estimated from this experiment and values given in the literature. Experimental measurements in the X2 expansion tunnel on blunt- and sharpnosed wedge models showed that the DHTG is a viable method for measuring heat fluxes in an expansion tunnel, but is more susceptible to damage from diaphragm fragments than surface junction thermocouples. The five flow conditions used had total enthalpies ranging from 25 MJ·kg−1 to 70 MJ·kg−1 , resulting in DHTG heat flux measurements of up to 129 MW·m−2 . Agreement better than 12 % was seen between the heat fluxes measured by the two types of gauge for the majority of the conditions. Across the set of tests, DHTGs were mounted on surfaces at angles of attack from 0° to 90°. No DHTGs were damaged at 0°, whereas those mounted at higher angles of attack survived between 1 and 17 shots, with half still operational at the end of the experimental campaign, but with significant signs of surface damage. In a second experimental campaign, a DHTG, a coaxial thermocouple, and a thinfilm heat transfer gauge were used to make stagnation point heat flux measurements in the T6 Reflected Shock Tunnel at a flow total enthalpy of 2.5 MJ·kg−1 . The DHTG and the thermocouple were undamaged after three shots, while the thin-film gauge was damaged beyond use after one shot. The DHTG measurements agree with the Fay and Riddell correlation, but are 20 % lower than the thermocouple measurement. The difference is believed to be a result of error in the effective thermal effusivity used in the data reduction of the thermocouples: the junctions were formed by abrasion using sandpaper, whereas for the experiments in X2 the junctions were created by making two larger-scale scratches. For stagnation point measurements in a Reflected Shock Tunnel, the DHTG is shown to be the preferred measurement technique, demonstrating significant improvements in robustness versus thin-film gauges, and in measurement uncertainty versus surface junction thermocouples.

This thesis details the establishment of a new, transient, high-enthalpy aerothermodynamic ground test facility at the University of Oxford: the T6 Stalker Tunnel. T6 is capable of operating as an expansion tube, reflected shock tunnel or in one of two high-enthalpy shock tube modes. This new facility was commissioned in each of these configurations, generating test conditions from 2.7 to 115.0 MJ kg−1. These correspond to trajectory points for a Mach 7 atmospheric flight condition and proposed Mars Sample Return missions. Overall, high-quality test flows were observed. This new facility now represents a unique resource within Europe for high-enthalpy aerothermodynamic ground testing. The core aim of this work was to investigate ‘radiative heating’ effects. During planetary entry, the high-temperature gas around a return vehicle emits strong electro- magnetic radiation, which can contribute significantly to surface heat flux. However, this heating mechanism is poorly understood, with high associated uncertainties.This thesis contributed to the resolution of this problem through development of a new mode of T6 — the Aluminium Shock Tube, or AST — which was designed specifically to investigate this phenomenon. The new mode was successfully commissioned; a thorough experimental assessment showed good comparison to computational predictions of performance and test time. Notably, the ability of the AST to achieve conditions of relevance for radiative heating in the atmospheres of Earth, Mars and Titan was demonstrated. Utilising the new AST mode, shock-layer radiation experiments were undertaken from 7 km s-1 to 10 km s−1 in air. Using a newly developed dual-channel emission spectroscopy system, simultaneous measurements of spatially and spectrally resolved, absolute spectral radiance were made between 200 nm and 850 nm. These were performed at selected benchmark conditions from the literature to validate the new experimental results. In all cases, large extents of post-shock equilibrium were generated. In general, comparisons were favourable to the benchmark data from other facilities. Although some discrepancies were seen in comparison to numerical codes, these were consistent across the set of experimental data considered from all facilities. Reasons for these differences were proposed. The T6 AST is now one of the very few facilities in the world which has demonstrated the capability to produce validated shock-layer radiation data. Lastly, a combined experimental-numerical study was undertaken to investigate the influence of shock deceleration on the measured post-shock radiance and electron density in shock tube experiments. This effect was conclusively shown to directly affect the measurement. The analysis was then extended to previous results from the NASA Electric Arc Shock Tube, thereby explaining historic discrepancies in that data.

Particulate deposition has become increasingly concerning for jet engine manufacturers, it can lead to a variety of problems, from stress corrosion cracking and the blocking of cooling holes in the turbine, to reduced stall margins for the engine. This thesis describes the development and validation of quantitative design tools that will enable manufacturers to make engines particulate tolerant. These tools are needed to allow: engines to fly in high concentrations of dust for a limited time; manufacturers to operate a cost of ownership model by predicting damage; and most importantly the safe certification of engines against these risks. The key capabilities required are the prediction of particulate trajectories through the engine, and their likelihood of adhering to engine components on contact. This thesis focusses on the robust prediction of turbulent particle transport in the context of the highly complex flows encountered in gas turbines. The tools developed are applicable to engine design because of their compatibility with existing design techniques, commercial flow solvers, and their computational efficiency. A stochastic model of turbophoresis (a Continuous Random Walk, CRW) is developed and a methodology for its application detailed. A model from literature provides the basis of the current transport model. While accurate modelling of turbophoresis has previously been shown to change the likelihood of small particles sticking by several orders of magnitude, compared to the exclusion or inadequate inclusion of turbophoretic effects, these results could only be reliably replicated in simple pipe flows as the models required turbulent flow characteristics determined from Direct Numerical Simulation or experiment. These characteristics are impossible to replicate for complex geometries. Thus, the thesis identifies robust methods to calculate instantaneous turbulent flow fluctuations, seen by the particle in the Lagrangian frame of reference, from Reynolds Averaged Navier Stokes computational fluid dynamic solutions. The model is shown to be equally accurate in simple flows when compared to previous implementations, and requires neither external data nor any a priori assumptions regarding the structure of the turbulence. The choice of the turbulence model, formulation of the particulate tracking time step, a local cell co-ordinate system to describe the turbulence, and modification of the near-wall Lagrangian timescale are all shown to be critical to the generation of accurate spatially resolved particulate deposition. The method is implemented to interface with a commercial flow solver. To ensure reliable deposition distribution maps are produced, while maintaining computational efficiency, convergence criteria are developed as well as a methodology to minimise the number of particulate tracking events required in a computational simulation. Here, convergence of particulate tracking based on the well known integral deposition fraction is shown to be insufficient as a measure of convergence of the spatial distribution. Well established metrics for describing variation in multidimensional distributions are compared, and a suitable technique to generate a converged distribution in a computationally efficient way is established. The revised CRW model has been validated both against existing data in the literature, and further experiments. A difficulty here is the very limited spatially resolved deposition data for which sufficient boundary conditions are available. For more complex, gas turbine relevant flows, no data of sufficient resolution were available in the open literature. A new rig was developed to generate spatially resolved distributions of deposition in accelerating flows, at representative temperatures and flow speeds. The range of tests conducted investigated the effects of geometry curvature, flow speed, and metal temperature. Finally, the revised CRW model is run according to the methodology developed, which ensures convergence, and these results are compared to the experimental data. As bounce-stick models are still unreliable, to reduce the number of variables considered, both all-stick and all- bounce conditions were simulated and compared. The chosen geometries allowed the effects of curvature and flow speed to be further investigated, with experimental trends well replicated. The mechanisms responsible for these trends were inferred from these simulations. It was found that modelling of turbophoresis was critical to capturing a number of features of deposition that were seen in experiments. The transport model is seen to provide a robust basis for accurate modelling, however, more quantitative comparison requires a similar improvement in the modelling of bounce- stick behaviour.

This thesis investigates the application of transpiration cooling to mitigate material oxidation of hypersonic vehicles. This was motivated by the early onset of surface oxidation of many high-temperature aerospace materials and the resulting limitation of the flight envelope. For example, Ultra-High-Temperature Ceramics have melting points exceeding 3500 K, thus could balance the incoming heat flux through re-radiation alone, but start oxidising at 1000 K, which limits the passive cooling capability severely. The aim of this thesis is to increase this oxidation limit by shielding the surface from freestream oxygen through boundary layer mass injection. A novel pressure sensitive paint diagnostic is developed in this work to measure the concentration of molecular oxygen on a transpiration cooled surface. This is a fundamental tool to quantify the ability of the coolant film to act as a barrier against freestream species that would lead to oxidation. The paint consists of [Ru(dpp)3]2+ luminophores, which are dissolved in a dichloromethane solution and then sprayed onto porous alumina. This method is validated experimentally on a transpiration cooled flat plate in a hypersonic cross-flow in the Oxford High Density Tunnel. Tests were conducted with no coolant injection, nitrogen injection and air injection at increasing blowing ratios. Oxygen partial pressure maps on the transpiration cooled surface were obtained for several conditions at unit Reynolds numbers between 2.58−5.0e7/m and blowing ratios between 0.016%−0.078%. It is found that the oxygen pressure decreases as the unit Reynolds number decreases and the blowing ratio increases. The direct measurements provided by this technique will aid the development of empirical correlations for the oxygen concentration on a porous surface with mass injection. Before applying this diagnostic to a transpiration cooled stagnation point, a theoretical prediction of the species concentration at the wall is derived. This is achieved by combining the self-similar boundary layer equations with thin film theory. The resulting semi-analytical correlation explicitly expresses the concentration of freestream species on a hypersonic stagnation point as a function of freestream conditions, surface geometry and coolant properties. The concentration depends on the boundary layer edge pressure, temperature and velocity gradient, as well as the wall temperature and injected mass flux. The molar mass and diffusion coefficient are further needed for scaling if the injected gas differs from the freestream gas. The method is compared against the numerically obtained self-similar solutions of a wide range of flow conditions and showed an accuracy of ±4%. The correlation was compared to experiments on a transpiration cooled stagnation probe model tested in the Oxford High Density Tunnel. A porous Al2O3 material, developed in collaboration with Imperial College London, features a similar pore size and outflow homogeneity as ZrB2, with the additional capability of bonding PSP. Experiments were conducted at Mach 6.9 at three different Pitot pressures: 10 kPa, 20 kPa and 30 kPa. Nitrogen, Argon and Krypton are used as injection gases at mass flow rates ranging from 0.01 - 0.55 kg/m2s, in order to displace up to 99% of the freestream gas at the surface. The experimental data shows that transpiration cooling is more effective in displacing freestream gas than predicted by analytical models and numerical solutions. The microheterogeneous surface with recessed pores means there is an additional pressure gradient within the first layer of pores, which increases the displacement effectiveness. A test campaign in the plasma wind tunnel at the Institute of Space Systems in Stuttgart provided the first qualitative data of the oxidation behaviour of transpiration cooled ZrB2 in a continuous, steady state high-enthalpy environment. A 42% porous ZrB2 sample is exposed to a cold-wall, fully catalytic stagnation point heat flux of 3.95 MW/m2 at 3.22 kPa Pitot pressure. While the uncooled sample fully oxidised at a surface temperature of 2150 K, 20.25 g/m2s of helium and 620.11 g/m2s of nitrogen injection mitigated oxidation of the transpiration cooled samples. Combining theoretical, numerical and experimental findings, the oxidation response of the stagnation point in a representative hypersonic flight scenario is predicted. An existing oxidation model for ZrB2 is combined with low order correlations for the heat transfer and oxygen concentration at the surface, to yield the surface recession and oxide scale thickness. A 500 s steady state trajectory at 44 km altitude and 3.6 km/s velocity is found to lead to a 2.2 mm recession of the 3 mm nose radius. A constant mass injection at a blowing parameter of 0.6 reduces the recession to just 0.21 mm. The displacement of freestream oxygen by transpiration cooling has a significant effect on oxidation. Not accounting for the displacement of oxygen at the surface would increase the recession by up to 197%. The recession along the transient trajectory of an envisioned hypersonic vehicle with a 3 mm nose radius is found to exceed 0.94 mm with no mass injection. It is shown that nitrogen and helium injection at a blowing parameter of 0.6 can reduce the recession to 0.13 mm and 0.075 mm, respectively.

Numerous turbofan power loss events have occurred in high altitude locations in the presence of ice crystals. Ice crystals enter the engine core, partially melt in the compressor and then accrete onto components that are initially warmer than freezing. The ice accretion causes blockage of the flow path, and if it later sheds can damage downstream components and potentially cause flame-out alongside other issues. Numerical models are currently required for the design and certification of next-generation engines to fly within ice crystal icing conditions. They have successfully predicted ice growth on simplistic geometries, but are often fine-tuned to a single condition or geometry due to the high level of empiricism. In this thesis a review of the current literature is presented, areas of missing research are indicated, some of which form the basis of the PhD. First, a new method for spatially resolving the three-dimensional ice accretion profile was developed and tested in an ice crystal icing wind tunnel. Historical measurement techniques were either two dimensional or could only be applied after the accretion had formed. This new technique provides transient, three-dimensional, non-intrusive measurement of the accretion during the ice exposure. Such measurements are deemed essential for the development of complex three-dimensional ice accretion models. The second set of work was the development of the ice crystal icing numerical model, ICICLE. The most notable development was the inclusion of the ability to update the flowfield based on the numerically predicted accretion profile. This change in geometry alters the potential field, leading to changes in the predicted particle trajectories and resultant accretion. The new capabilities of the code were tested using two test pieces: a single compressor stator, and a triple NACA airfoil linear cascade. The results showed that certain conditions require the update of the flowfield for accurate estimation of both the shape and volume of the accretion.

This thesis investigates the application of transpiration cooling to reduce surface heat transfer on hypersonic vehicles. Such vehicles experience high peak heat fluxes and time-integrated heat loads throughout the ascent, re-entry, and cruise phases. The most significant heat fluxes occur at the stagnation point of sharp leading edges, regions of laminar-turbulent transition, or shock-wave boundary layer interactions. Whilst previous and current hypersonic vehicles employ ablation for many of these regions, such thermal protection systems are neither reusable nor suitable for all scenarios. In this work, the performance of transpiration cooling is assessed for a generic flight vehicle. Subsequently, the application of transpiration cooling to laminar and turbulent flows and shock-wave boundary layer interaction regions is assessed experimentally. For the first time, the additional benefit of transpiration cooling to the underlying substructure is evaluated numerically. The low fidelity PIRATE code for evaluating transpiration-cooled systems is extended to account for quasi-two-dimensional lateral heat conduction effects, enabling very fast calculations of the two-dimensional transient temperature response of a transpiration-cooled thermal protection system. To solve for the transpiration-cooled outer wall and a two-dimensional solid substructure, PIRATE has been coupled with the commercial finite element package COMSOL, enabling the modelling of the longer-duration thermal effects of the integrated heat load over a flight trajectory. Transpiration cooling using helium coolant has been applied to a wing leading edge with an aluminium substructure. Carbon–carbon ceramic composite and the Ultra-High-Temperature Ceramic (UHTC) zirconium diboride (ZrB2) were chosen as candidate materials. The substructure temperature history for the Space Shuttle re-entry trajectory was obtained, showing that transpiration cooling can lead to a 35% reduction in peak substructure temperature and a 65% decrease in thermal gradients. A novel heat transfer measurement method is developed to enable measurements of surface heat transfer in transpiration-cooled porous injectors in short-duration hypersonic facilities. Traditional surface heat flux methods such as discrete thin film gauges or thermocouples are not suitable for modern micro-porous materials with pore sizes of the order of 10 microns. High-speed infrared thermography was employed to measure the surface temperature of the porous material. A bespoke calibration system was developed, and the method was verified experimentally in the Oxford High-Density Tunnel on a flat-faced hemispherical probe, with a porous Alumina injector. Nitrogen, air, argon, krypton, and helium injection gases were used with mass flow rates ranging from 0.01-0.235 kgs-1m-2. It was found that the Stanton number reduction matched to within 10% of both CFD results and correlations. The direct measurements provided by this technique enable the acquisition of spatially resolved heat transfer on a porous surface with mass injection. After the development of the heat transfer measurement technique, this was applied to two different scenarios. The first concerns the need for detailed local heat transfer information on and in the vicinity of the porous injector to design a transpiration-cooled system. In this work, experiments were conducted 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 infrared thermography. 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 at both laminar and turbulent regimes. A key result is that helium performs much better than reported in previous experiments. Finally, the downstream thermal effectiveness is characterised for turbulent flows and an analytical correlation is proposed. In the final component of this thesis, transpiration cooling is applied to mitigate the peak heat fluxes caused by shock-wave boundary layer interactions. Experiments were again conducted in the Oxford High-Density Tunnel at Mach 6.1 in both laminar and turbulent regimes where a 10 degree shock generator created a strong oblique shock wave which impinged onto a transpiration-cooled micro-porous injector. For the laminar condition, due to the strength of the incident shock, a laminar-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 and relatively low injection rates were sufficient to greatly reduce the heat transfer downstream of shock interaction for both regimes. The experimental data are correlated, and both fully turbulent and laminar-transitional cases display a similar trend. Empirical fits are proposed which may be used for initial systems design.

Shock tube experiments are a key method of reproducing the flow conditions encountered by hypersonic flight vehicles and are the primary source of known thermochemical reaction rates. However, because of the complex nature of hypersonic flows, and also due to the many subtle complexities of the shock facilities themselves, full characterization of the flows they produce is elusive. This work develops a numerical method for the simulation of shock tube experiments using a novel technique of shock tracking to improve the resolution and decrease the computational cost of such simulations. An axisymmetric, Navier-Stokes shock tube simulation code entitled FROSST keeps the shock and contact discontinuity each stationary in their own zone of refinement that slides along a coarse background mesh of the entire facility. Using this solver, the role of shock history in the formation of nonuniformities in shock tube test gases is investigated in the context of ideal gas simulations. Shock speed variations from -29% to +33% are found to influence temperatures in the test slug by as much as 55% in an Argon test gas with a tube-end 2100 m/s. These results give rise to the observation that nearly all nonuniformities in the test gas originate with variations in shock speed along the tube. Based upon this observation, a novel method is devised to use experimentally measured shock speed to predict test flow properties for each experiment. This method tracks the influence of waves associated with shock speed variation upon isentropic slices of shock-processed gas. By tracking a number of such slices, the entire test slug is reconstructed to a high degree of accuracy. The analytical methodology, entitled LASTA, is extensively validated against simulated ideal Argon data and also experiments performed in the Oxford T6 Stalker shock facility with both Argon and thermochemically reacting air. LASTA is found to capture the shot-to-shot nonuniformities to within 1% within minutes on a single core. Real-world application of LASTA is already leading to improved understanding of radiative emissions from shock tube experiments.

In this thesis, a porous Ultra-High-Temperature-Ceramic (UHTC) made of zirconium diboride (ZrB2) is qualified for the purpose for transpiration cooling for the first time. Subsequently, the mixing mechanism between the coolant and a laminar, hypersonic boundary-layer gas at the wall downstream of a transpiration-cooled injector is investigated. This has led to understanding the mixing process at the wall in a laminar boundary layer. Porous UHTCs are a candidate group of materials for transpiration cooling of hypersonic vehicles due to their exceptionally high melting point, typically above 3000 K. Their high operating temperature permits a higher amount of radiative cooling than that achievable with conventional materials, which reduces the required coolant mass flow rate to cool the surface. This thesis experimentally examines the internal and external flow behaviour of porous UHTC made of zirconium diboride (ZrB2) for the purpose of transpiration cooling. A dedicated ISO standard permeability test rig was built. The outflow velocity distribution was acquired employing miniature hot-wire anemometry. The data obtained for the pressure loss across the porous samples agree with the Darcy-Forchheimer model for flow in porous media; respective Darcy and Forchheimer permeability coefficients are calculated and reported. Cleaning the surface of the samples using sandpaper or an ultrasonic bath raised the permeability coefficient by up to 19%. The outflow velocity maps exhibit a good flow uniformity with an average standard deviation of 25.1% with respect to the mean value. Individual jets are absent, and the velocity varies within the same order of magnitude. 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 thesis, 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 12.9 × 10ˆ6 1/m. A coolant concentration of up to 95% is achieved immediately downstream of the injector. The film concentration drops in a monotonic fashion farther downstream; however, a constant film coverage of 5 mm to 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; an onset of boundary-layer transition is not promoted. 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. Subsequently, a collapse of the thermal effectiveness is achieved and a modified analytical correlation is proposed. A two-dimensional simulation study of transpiration cooling in a laminar, hypersonic boundary-layer using the Thermochemical Implicit Non-Equilibrium Algorithm (TINA) – a Navier-Stokes solver was undertaken. Coolant concentration and heat flux results are compared to data obtained from the experiments. TINA successfully predicts the mixing rate at the wall as a function of the stream-wise direction for all blowing ratios. The simulations are more successful in predicting the mixing downstream of the injector compared to the mixing on the injector, especially at low blowing ratios. A collapse of the thermal effectiveness values calculated from 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 1 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 stream-wise 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. Finally, an analytical model based on one-dimensional diffusion is proposed to predict the mixing between the coolant gas and boundary-layer gas at the wall downstream of a transpiring injector in a laminar flow. The model is validated against the experimentally obtained coolant concentration data. It successfully predicts the mixing at the wall downstream within 17% of the experimental data. It is shown that this mixing mechanism at the wall in laminar flows is fully described by the process of diffusion. The coolant coverage at a given downstream location is promoted when the stream-wise velocity decreases, the blowing ratio increases, or the diffusion coefficient drops. Subsequently, a mass budget calculation is performed for a transpiration-cooled hypersonic vehicle employing the analytical model. The model predicts a 3.6 times less coolant mass requirement when Helium is used as the coolant gas as opposed to Nitrogen for the chosen trajectory. However, Helium requires twice the storage volume compared to Nitrogen.

The drive to increase the thermal efficiency and specific power output of the gas turbine results in increasing turbine temperatures which necessitate active cooling of the turbine blades. High performance cooling systems are required which achieve the necessary blade life requirements whilst minimising coolant consumption. This thesis explores one potential turbine cooling system advancement termed the double-wall, effusion cooled system which is envisaged to find application in areas where high cooling performance is required. Single wall effusion cooling performance is initially investigated given its importance in the double-wall system. An experimental investigation is performed and used to validate a computational model. A two-dimensional superposition method for predicting effusion cooling performance is presented and validated based upon the experimental and computational data. The superposition method was used as part of a computationally light, decoupled conjugate method which was developed to predict the cooling performance of double-wall geometries. This decoupled method was used to assess several developed double-wall geometries. Those geometries which exhibited preferable cooling characteristic were manufactured for testing in a novel experimental facility, designed and commissioned as part of the thesis work. The facility incorporates several features to improve the quality of data obtained and matches several relevant non-dimensional parameters. The facility was used to obtain both overall effectiveness and film effectiveness results for five flat plate, double-wall geometries. This data was then used to infer the convective cooling performance of each geometry at varying coolant flow rates, with preferable cooling features identified. The results indicate the high cooling performance achieved in double-wall, effusion cooled geometries. Two computational methods were used to replicate the experimental setup utilising boundary conditions obtained from the facility. The first was a fully conjugate solver, and the second was a modified form of the developed decoupled method. The results from both sets of simulations compared favourably with the experimental data. The decoupled conjugate method was particularly encouraging given the vast reduction in simulation time required when compared to the fully conjugate solver. Thermomechanical analyses were performed to identify how various geometric features influence the stress field developed under thermal load, which has direct implication on blade lifespan. The work was successful in indicating certain double-wall configurations which could reduce the magnitude of the stresses.

High altitude ice particles can accrete inside the core compression system of turbofan engines in cruise and descent. This can lead to severe in-flight events including blade damage, surge and flameout. This thesis describes the development and validation of a new comprehensive computational model to aid prediction of ice crystal icing in turbofan compressors. The Ice Crystal Icing ComputationaL Environment (ICICLE) delivers a step change in modelling of the phenomenon compared to the first generation of models in the open literature. Modelling of this multi-faceted problem is broadly divided into three strands: first, modelling of the ice particles in flight; second their interactions with solid surfaces; and third the thermodynamics of ice accretion. To aid development of models and provide validation data, three different experiments were also undertaken. Treatment of particle size and shape distribution is considered first, and a particle trajectory model based on Lagrangian tracking is presented. A Nusselt number correlation for non-spherical particles is used to develop a phase change model for the particle in flight, incorporating sublimation, evaporation and melting. The model is then validated against measured particle melt data in an ice crystal facility. A model for the change in enthalpy and humidity of the airflow as a result of the particle phase change is proposed. Existing icing codes do not attempt to model these affects, but evidence from engine encounters with ice crystals indicate that they are significant. It was assessed that experimentation was required to develop modelling capability in three areas: particle sticking, erosion and heat transfer. Two experimental campaigns were performed at the ice crystal wind tunnels of the National Research Council of Canada (NRC) using simple geometries (an inclined flat plate and a cone). Data was presented for the first time on heat transfer from a warm substrate under ice crystal conditions, and a method to predict the change in particle melt during surface impacts was proposed. New semi-empirical models were developed for sticking and erosion, with a substantially wider range of applicability than achieved in previous studies. A new thermodynamic ice crystal accretion model was developed. A literature model for supercooled water icing was adapted to ice crystal and mixed phase conditions, and to substrates either above or below freezing. In the former case, an entirely novel three-layer accretion model was developed, which is a substantial advancement in modelling ice crystal growth on initially warm engine surfaces. Finally, the complete model is validated against experimental accretions on the case of a compressor stator test article, also tested at the NRC. Agreement is seen generally to be good, with the transient behaviour of growth rates well predicted, typically within 20% of experimental measurements. It is shown that a substantial improvement in prediction accuracy may be attained by updating the fluid domain at discrete time points. This accounts for the influence of the growing accretion on the flowfield. The successful application of a quantitative code to a more complex, engine-realistic geometry is a significant step forward for the literature, as existing ice crystal codes have only been validated against simpler geometries.

This thesis describes validated improvements in the modelling of micron-sized particle deposition within gas turbine engine secondary air systems. The initial aim of the research was to employ appropriate models of instantaneous turbulent flow behaviour to RANS CFD simulations, allowing the trajectory of solid particulates in the flow to be accurately predicted. Following critical assessment of turbophoretic models, the continuous random walk (CRW) model was chosen to predict instantaneous fluid fluctuating velocities. Particle flow, characterised by non-dimensional deposition velocity and particle relaxation time, was observed to match published experimental vertical pipe flow data. This was possible due to redefining the integration time step in terms of Kolmagorov and Lagrangian time scales, reducing the disparity between simulations and experimental data by an order of magnitude. As no high temperature validation data for the CRW model were available, an experimental rig was developed to conduct horizontal pipe flow experiments under engine realistic conditions. Both the experimental rig, and a new particulate concentration measurement technique, based on post test aqueous solution electrical conductivity, were qualified at ambient conditions. These new experimental data compare well to published data at non-dimensional particle relaxation times below 7. Above, a tail off in the deposition rate is observed, potentially caused by a bounce or shear removal mechanism at higher particle kinetic energy. At elevated temperatures and isothermal conditions, similar behaviour is observed to the ambient data. Under engine representative thermophoretic conditions, a negative gas to wall temperature gradient is seen to increase deposition by up to 4.8 times, the reverse decreasing deposition by a factor of up to 560 relative to the isothermal data. Numerical simulations using the CRW model under-predict isothermal deposition, though capturing relative thermophoretic effects well. By applying an anisotropic Lagrangian time scale, and cross trajectory effects of the external gravitational force, good agreement was observed, the first inclusion of the effect within the CRW model. A dynamic mesh morphing method was then developed, enabling the effect of large scale particle deposition to be included in simulations, without continual remeshing of the fluid domain. Simulation of an impingement jet array showed deposition of characteristic mounds up to 30\% of the hole diameter in height. Simulation of a passage with film-cooling hole off-takes generated hole blockage of up to 40\%. These cases confirmed that the use of the CRW generated deposition locations in line with scant available experimental data, but widespread airline fleet experience. Changing rates of deposition were observed with the evolution of the deposits in both cases, highlighting the importance of capturing changing passage geometry through dynamic mesh morphing. The level of deposition observed, was however, greater than expected in a real engine environment and identifies a need to further refine bounce-stick and erosion modelling to complement the improved prediction of impact location identified in this thesis.

Modern gas turbine engines operate at temperatures well above the melting point of the metal components. This has driven manufactures to develop sophisticated cooling methods which minimise the use of coolant to maximise engine efficiency by enabling further increases in operating temperature. This thesis investigates the cooling performance of engine representative mid-chord internal cooling passages for turbine blades. The work forms part of a larger E.C. FP7 project ERICKA (Engine Representative Internal Cooling Knowledge Applications).This thesis provides detailed maps of heat transfer coefficient (HTC) under a number of conditions, new experimental techniques, and has lead to a better understanding of the impact HTC distributions have on the thermal performance of a turbine blade at engine conditions. Transient liquid crystal experiments have been conducted on a large scale model of an engine representative internal cooling passage at three aspect ratios (width:height (chord length:spanwise length), 1:2, 1:3 and 1:4). Spatially resolved maps of Nusselt number have been produced for the full surface of the internal cooling passages. Little information exists in the literature for more engine representative geometries, and it is rare for spatial measurements to be presented over the full surface. The detailed maps provide validation data for CFD within the ERICKA programme. A novel method which produces spatially resolved maps in areas with highly non-one-dimensional heat transfer has been developed and validated. This method couples transient finite element analysis and data from transient liquid crystal experiments. Applied to the ribbed passage geometry, this produced spatially resolved maps of HTC over the rib surface. To the author’s best knowledge this is the first time spatial HTC maps have been presented for an engine representative rib. Industry best practice methods for internal cooling passage design typically apply averaged values of HTC, in part due to lack of spatially resolved data. To determine the significance of this approximation on blade design and life, experimental measurements have been applied to finite element (FEA) models at typical engine conditions. Application of a 3D HTC distribution to a FEA model of a section of ribbed wall demonstrated a significant under prediction (up to 58%) of localised thermal gradients when an average value is applied compared to a spatially resolved profile. This work demonstrated good agreement between distributions taken from experimental data and CFD predictions, indicating that CFD distributions may be more appropriate than bulk values in the design process. A 2D FEA study was undertaken to quantify the impact of HTC distribution approximations and aspect ratio on cooling of a generic turbine section. This study considered multiple adjacent internal cooling passages. It was confirmed that multi-pass arrangements offer greater heat removal for a given mass flow rate. Also a symmetric heat transfer profile with a higher HTC on the ribbed wall is the most desirable distribution. Use of average values significantly impacted the metal temperature, causing an underprediction up to 13◦C and 8◦C in the maximum and average values respectively. Based on the experimental HTC data, the 1:3 aspect ratio passage offered the lowest metal temperatures. Applying HTC distributions from CFD data (calculated with using the centreline temperature) showed, in general, good agreement, with the lowest metal temperatures (by up to 8◦C) in the 1:4 aspect ratio passage. Use of and HTC distribution provided by CFD prediction based on the mixed bulk temperature, produced average and peak metal temperatures 16◦C and 17◦C, respectively, lower in the 1:4 aspect ratio passage than the next best design. This highlights the need for appropriate and consistent method to be used in the analysis. As expected, reducing the passage aspect ratio led to increases in both thermal gradient and total pressure loss.