Search

This work describes the development and implementation of a system to perform accurate measurements of mass flow rate in the Turbine Test Facility (TTF) at QinetiQ Farnborough. The facility has recently been upgraded so that turbine aerodynamic efficiency measurements can be performed and the implementation of a system for mass flow rate measurement formed part of that upgrade. The measurement system is novel in that accurate measurements can be performed with both combustor representative turbine inlet temperature distortion (hot-streaks) and swirl. The TTF is a short-duration (approximately 0.5 second run time) isentropic light piston turbine facility, which has been used for aerodynamic and heat transfer investigations of – primarily – high-pressure turbine stages, although it has also been configured to operate as a 1½ stage (HP stage with IP or LP vane) turbine. The MT1 turbine is a highly loaded unshrouded design relevant to modern military engine design, or future civil engine design. The turbine is engine scale, and all relevant dimensional parameters for aerodynamics and heat transfer are matched: Re, M, N T01 , Tgas Twall . Hot streaks are simulated in the TTF by the controlled mixing of hot and cold gas streams. The cold stream is introduced though a conventional sonic metering nozzle, from a large reservoir acting in blow-down mode. In a transient facility, the accurate measurement of stage mass flow rate with combustor representative inlet temperature distortion and swirl presents an interesting problem, as the capacity of the ngv row is affected by both of these combustor representative inlet flow-fields. In the TTF, a mass flow rate measurement system has been developed using the exit contraction of the piston tube as a subsonic converging-diverging venturi: upstream and , and throat are measured at a number of circumferential locations around the exit contraction to determine the mass flow rate of the hot stream. The effective area of the venturi was measured using a novel blow-down calibration technique which is described. 0 p T0 p The bias error in the measurement of mass flow rate with and without temperature distortion was 1.37 per cent and 1.13 per cent respectively, of the same order as the accuracy associated with conventional tertiary devices. The precision uncertainty was 0.198 per cent in both cases. Accuracy is unaffected by the introduction of inlet swirl.

A common requirement for turbomachinery testing facilities is the ability to independently control Mach number and Reynolds number. In practice, this means independent control of the inlet total pressure and exit static pressure in a test facility. In this paper, we present a solution to this problem with particular applicability to large-scale annular test facilities. We describe the design and commissioning of a combined annular dump diffuser and deswirl system for back-pressure control in environments with high-whirl transonic flow. The particular application was an annular cascade of nozzle guide vanes from a current civil engine. The purpose was to provide independent control of Mach number and Reynolds number, by controlling the back-pressure in the intermediate annular plenum which forms the dump diffuser. The dump diffuser is necessary to facilitate optical and probe access (without interaction effects) and to reduce the risk of exit static pressure disturbances (due to particular duct design). The system has been installed and validated in the engine component aerothermal (ECAT) facility at the University of Oxford. In this implementation of the system, the high-whirl transonic flow from the nozzle guide vanes passes through a short, parallel annular duct, and is dumped into an annular plenum, before being re-accelerated into a deswirl vane ring. The deswirl ring turns the flow to the axial direction. The flow is then discharged through a variable choke plate into a silencer. Conditioning the flow to have zero whirl at the choke plate reduces the sensitivity of the choke plate effective blockage to the whirl angle. The design, deswirl vane aerodynamic performance, and overall system performance are assessed with detailed experiments and 3D unsteady computational fluid dynamics predictions. The control of high-whirl transonic flow is notoriously challenging, and the deswirl system has application to exhaust conditioning in a number of applications including annular cascade experiments and rocket turbo-pump exhaust systems.

Detailed experimental measurements have been performed to understand the effects of turbine inlet temperature distortion (hot-streaks) on the heat transfer and aerodynamic characteristics of a full-scale unshrouded high pressure turbine stage at flow conditions that are representative of those found in a modern gas turbine engine. To investigate hot-streak migration, the experimental measurements are complemented by three-dimensional steady and unsteady CFD simulations of the turbine stage. This paper presents the time-averaged measurements and computational predictions of rotor blade surface and rotor casing heat transfer. Experimental measurements obtained with and without inlet temperature distortion are compared. Time-mean experimental measurements of rotor casing static pressure are also presented. CFD simulations have been conducted using the Rolls-Royce code HYDRA and are compared with the experimental results. The test turbine was the unshrouded MT1 turbine, installed in the Turbine Test Facility (previously called Isentropic Light Piston Facility) at QinetiQ, Farnborough, UK. This is a short duration transonic facility, which simulates engine-representative M, Re, Tu, N/T, and T g/T w to the turbine inlet. The facility has recently been upgraded to incorporate an advanced second-generation temperature distortion generator, capable of simulating well-defined, aggressive temperature distortion both in the radial and circumferential directions, at the turbine inlet. © 2012 American Society of Mechanical Engineers.

This paper presents an experimental and computational study of the effect of severe inlet temperature distortion (hot streaks) on the efficiency of the MT1 HP turbine, which is a highly-loaded unshrouded transonic design. The experiments were performed in the Oxford Turbine Research Facility (OTRF) (formerly the TTF at QinetiQ Farnborough): an engine scale, short duration, rotating transonic facility, in which M, Re, Tgas/Twall and N/T01 are matched to engine conditions. The research formed part of the EU Turbine Aero-Thermal External Flows (TATEF II) program. An advanced second generation temperature distortion simulator was developed for this investigation, which allows both radial and circumferential temperature profiles to be simulated. A pronounced profile was used for this study. The system was novel in that it was designed to be compatible with an efficiency measurement system which was also developed for this study. To achieve low uncertainty (bias and precision errors of approximately 1.5% and 0.2% respectively, to 95% confidence), the mass flow rate of the hot and cold streams used to simulate temperature distortion were independently metered upstream of the turbine nozzle using traceable measurement techniques. Turbine power was measured directly with an accurate torque transducer. The efficiency of the test turbine was evaluated experimentally for a uniform inlet temperature condition, and with pronounced temperature distortion. Mechanisms for observed changes in the turbine exit flow field and efficiency are discussed. The data are compared in terms of flow structure to full stage computational fluid dynamics (CFD) performed using the Rolls Royce Hydra code. © 2013 American Society of Mechanical Engineers.

Modern lean burn combustors now employ aggressive swirlers to enhance fuel-air mixing and improve flame stability. The flow at combustor exit can therefore have high residual swirl. A good deal of research concerning the flow within the combustor is available in open literature. The impact of swirl on the aerodynamic and heat transfer characteristics of an HP turbine stage is not well understood, however. A combustor swirl simulator has been designed and commissioned in the Oxford Turbine Research Facility (OTRF), previously located at QinetiQ, Farnborough UK. The swirl simulator is capable of generating an engine-representative combustor exit swirl pattern. At the turbine inlet plane, yaw and pitch angles of over ±40 deg have been simulated. The turbine research facility used for the study is an engine scale, short duration, rotating transonic turbine, in which the nondimensional parameters for aerodynamics and heat transfer are matched to engine conditions. The research turbine was the unshrouded MT1 design. By design, the center of the vortex from the swirl simulator can be clocked to any circumferential position with respect to HP vane, and the vortex-to-vane count ratio is 1:2. For the current investigation, the clocking position was such that the vortex center was aligned with the vane leading edge (every second vane). Both the aligned vane and the adjacent vane were characterized. This paper presents measurements of HP vane surface and end wall heat transfer for the two vane positions. The results are compared with measurements conducted without swirl. The vane surface pressure distributions are also presented. The experimental measurements are compared with full-stage three-dimensional unsteady numerical predictions obtained using the Rolls Royce in-house code Hydra. The aerodynamic and heat transfer characterization presented in this paper is the first of its kind, and it is hoped to give some insight into the significant changes in the vane flow and heat transfer that occur in the current generation of low NOx combustors. The findings not only have implications for the vane aerodynamic design, but also for the cooling system design. © 2013 American Society of Mechanical Engineers.

Swirling flows are now widely being used in modern gas turbine combustors to improve the combustion characteristics, flame stability, and reduce emissions. Residual swirl at the combustor exit will affect the performance of the downstream high-pressure (HP) turbine. In order to perform a detailed investigation of the effect of swirl on a full-scale HP turbine stage, a combustor swirl simulator has been designed and commissioned in the Oxford Turbine Research Facility (OTRF), previously located at QinetiQ, Farnborough UK, as the Turbine Test Facility (TTF). The swirl simulator is capable of generating engine-representative combustor exit swirl distributions at the turbine inlet, with yaw and pitch angles of up to 40 deg. The turbine test facility is an engine scale, short duration, rotating transonic turbine facility, which simulates the engine representative M, Re, Tu, nondimensional speed, and gas-to-wall temperature ratio at the turbine inlet. The test turbine is a highly loaded unshrouded design (the MT1 turbine). This paper presents time-averaged experimental heat transfer measurements performed on the rotor casing surface, and on the rotor blade surface at 10%, 50%, and 90% span. Time-averaged rotor casing static pressure measurements are also presented. Experimental measurements with and without inlet swirl are compared. The measurements are discussed with the aid of three-dimensional steady and unsteady CFD simulations of the turbine stage. Numerical simulations were conducted using the Rolls-Royce in-house code HYDRA, with and without inlet swirl. © 2012 American Society of Mechanical Engineers.

This paper presents an investigation of the aerothermal performance of a modern unshrouded high-pressure (HP) aero-engine turbine subject to nonuniform inlet temperature profile. The turbine used for this study was the MT1 turbine installed in the QinetiQ turbine test facility based in Farnborough (UK). The MT1 turbine is a full scale transonic HP turbine, and is operated in the test facility at the correct nondimensional conditions for aerodynamics and heat transfer. Datum experiments of aerothermal performance were conducted with uniform inlet conditions. Experiments with nonuniform inlet temperature were conducted with a temperature profile that had a nonuniformity in the radial direction defined by (T max-T min)/T-=0.355, and a nonuniformity in the circumferential direction defined by (T max-T min)/T-=0.14. This corresponds to an extreme point in the engine cycle, in an engine where the nonuniformity is dominated by the radial distribution. Accurate experimental area surveys of the turbine inlet and exit flows were conducted, and detailed heat transfer measurements were obtained on the blade surfaces and end-walls. These results are analyzed with the unsteady numerical data obtained using the in-house HybFlow code developed at the University of Firenze. Two particular aspects are highlighted in the discussion: prediction confidence for state of the art computational fluid dynamics (CFD) and impact of real conditions on stator-rotor thermal loading. The efficiency value obtained with the numerical analysis is compared with the experimental data and a 0.8% difference is found and discussed. A study of the flow field influence on the blade thermal load has also been detailed. It is shown that the hot streak migration mainly affects the rotor pressure side from 20% to 70% of the span, where the Nusselt number increases by a factor of 60% with respect to the uniform case. Furthermore, in this work, it has been found that a nonuniform temperature distribution is beneficial for the rotor tip, contrary to the results found in open literature. Although the hot streak is affected by the pressure gradient across the tip gap, the radial profile (which dominates the temperature profile being considered) is not fully mixed out in passing through the HP stage, and contributes significantly to cooling the turbine casing. A design approach not taking into account these effects will underestimate the rotor life near the tip and the thermal load at midspan. The temperature profile that has been used in both experiments and CFD is the first simulation of an extreme cycle point (more than twice the magnitude of distortion of all previous experimental studies): It represents an engine-take-off condition combined with the full combustor cooling. This research was part of the EU funded Turbine AeroThermal External Flows 2 program. © 2012 American Society of Mechanical Engineers.

This thesis report presents the results from low-order model, CFD and experimental analysis to be used for parametric studies of adiabatic film and overall cooling effectiveness for fully cooled systems (internal and film) under wide ranges of mainstream-to-coolant temperature ratio variation. The purpose is to improve understanding of the scaling process from typical rig conditions to engine conditions. The interest is in the variation in overall effectiveness when the controlling non-dimensional groups change in a natural co-dependent way with changes in temperature ratio. This is distinguished from the situation in which individual non-dimensional groups are varied in isolation. The design, development and commissioning data from a new high temperature (600 K) test facility is presented together with detailed uncertainty analysis. This test facility operates at engine-realistic Reynolds number, Mach number and mainstream-to-coolant pressure ration across a mainstream-to-coolant temperature ratio from 0.50 to 2.30. For experiments, a cooled flat-plate configuration was designed incorporating a reverse-pass cooling design and 45° V-shaped broken rib turbulators to enhance the coupling of the internal and external systems. From CFD studies, adiabatic film effectiveness was found to increase with increasing mainstream- to-coolant temperature ratio. The percentage change in adiabatic film effectiveness (from reference case with the temperature ratio of 2.00) were from -21.2 % to -1.38 % for temperature ratio from 1.10 to 1.90. Here for matched Reynolds number, Mach number and mainstream-to-coolant pressure ratio, the increase in adiabatic film effectiveness was found to be due to the increase in both specific heat capacity flux ratio (+70 % of change) and blowing ratio (+40 % of change). The overall cooling effectiveness measurements from experiments conducted in the mainstream-to-coolant temperature ratio range between 1.09 to 1.62 were presented. Overall cooling effectiveness was found to increase with increasing temperature ratio, with predicted percentage changes from the reference condition (temperature ratio of 1.60) in external wall overall cooling effectiveness of -5.92 % to -1.38 % for temperature ratio from 1.10 to 1.50 respectively. On the internal wall the percentage change in overall cooling effectiveness was -3.73 % to -0.78 % for temperature ratio from 1.10 to 1.50 respectively. To explain the trends in overall cooling effectiveness, a bespoke low-order conjugate aerothermal model (calibrated against measurements) was used. From the low-order model, the increase in overall cooling effectiveness was found to be driven largely by the increase in external film effectiveness (+68 % of change) and the ratio of mainstream-to-coolant heat transfer coefficients (+46 % of change). Secondary influence was a Biot number effect (-18 % of change). The influence of internal coolant warming factor coolant factor was found to be very small (+4 % of change).

In this thesis, novel cooling systems for gas turbine blade platforms were developed and assessed experimentally. A new highly modular facility was designed, manufactured, and commissioned to support investigations of a variety of different platform cooling systems. The facility is a high pressure, high temperature, transonic, linear cascade. A high degree of engine fidelity is achieved by: operating at scaled engine conditions; using test blades with full platform and fir tree root geometry; and including engine-representative internal hub seal structures and coolant feeds. The inclusion of hub seals, realistic coolant feeds, and full test blades allowed engine-feasible systems to be tested and representative leakage flows to be reproduced. Two novel platform cooling systems were developed, using low-order numerical models, which utilise internal cooling features embedded directly into the platform surface. These systems are: a convective cooling passage design which draws coolant from the front seal, passes it through the platform, and exhausts into the rear seal; and an impingement jet array design which draws coolant from the shank cavity and similarly exhausts into the rear seal. Each system was integrated into cascade geometry rotor blades, manufactured in titanium via additive manufacturing, and used to experimentally determine metal effectiveness distributions across mainstream platform surfaces. Metal effectiveness distributions are determined from IR measurements of surface temperature. The cooling performance of each concept was quantified and compared to a baseline case with no cooling features, cooled solely by leakage flows. The film cooling performance of front seal purge flow was investigated experimentally, to determine adiabatic film effectiveness, at a range of purge flow angles and mass ratios. The aerodynamic impact was assessed using a downstream area traverse system to determine total pressure loss coefficient distributions. A practical design to alter the purge flow angle in engine architectures was proposed and the viability of the system demonstrated using numerical modelling. A new experimental data processing technique was developed to correct adiabatic film effectiveness distributions determined from experiments with a non-uniform inlet temperature profile, such as that caused by a time-varying inlet thermal boundary layer.

High pressure (HP) nozzle guide vane (NGV) capacity is one of the most important parameters for setting overall turbine power output, for stage matching, and for understanding turbine performance. In accelerated engine development programmes in particular, accurate and early assessment of NGV capacity is a significant advantage. Whilst the capabilities of computational methods have improved rapidly in recent years, the accuracy of absolute capacity prediction capability is lower than experimental techniques by some margin. Thus, experimental measurement of NGV capacity is still regarded as an essential part of many engine programmes. The semi-transient capacity measurement technique has been developed and refined at the University of Oxford over the last ten years. It allows rapid and accurate measurement of engine component (typically fully-cooled NGVs) capacity at engine-representative conditions of Mach number, Reynolds number, and coolant-to-mainstream pressure ratio. The technique has been demonstrated to offer considerable advantages over traditional (industrial steady-state) techniques in terms of accuracy, time and operating cost. The present work describes in detail improvements made to the semi-transient method to optimise its uncertainty, and presents a detailed uncertainty analysis of the updated facility which is shown to achieve a bias uncertainty of ±0.495% (to 95% confidence) and a precision uncertainty of ±0.025%. The extremely low precision uncertainty in particular allows very small changes in capacity to be resolved. This, combined with rapid interchangeability of test modules, allows studies of the sensitivity of capacity to secondary influences with much greater flexibility than was previously possible. In the present work, this capability is used to examine the effect of changing trailing edge (TE) geometry—particularly suction side (SS) TE overhang length—on the flow capacity. Experimental measurements and complementary numerical predictions of capacity are presented for four geometries bridging the gap between a classical centred ejection design and a SS-overhang design. A monotonic increase in capacity is observed as TE overhang length reduces, but with a greatly reduced gradient when overhang length becomes very short. Strong agreement was achieved between experiment and CFD in terms of both absolute capacity prediction (average error of 0.67%) and capacity changes between TE geometries (average error of 0.20%). Comprehensive examination is given to the fundamental mechanisms responsible for the observed capacity changes. Two models of the vane are proposed which describe the interactions between multiple controlling areas which can be defined from the physical geometry. Whilst these models are incomplete, they represent an important aspect of the changing flow interactions between different TE geometries and help to explain the change in flow behaviour when overhang length becomes very short. Further insight is gained by studying the area and static pressure changes at several controlling areas, although this approach is demonstrated to be somewhat arbitrary. Whilst the capacity trend can be explained by considering particular controlling areas in this way, a more complete system description is provided by holistically analysing the aerodynamic change in the entire controlling region of the vane passage. The SS-overhang TE design has generally been favoured by HP NGV designers in recent years because it has been presumed to be aerodynamically advantageous on account of providing a thinner ultimate TE than the equivalent centred ejection geometry (at the cost of increased cooling requirement for the overhanging length). The prior literature is not in consensus on this point, however, and so the present work re-examines the impact of TE overhang length on NGV aerodynamic loss characteristics. High-fidelity experimental traverse measurements downstream of the HP NGV cascade are presented, and compared against steady and unsteady RANS CFD predictions. This is a highly useful data set for benchmarking typical CFD methods for absolute loss prediction of cooled engine components. URANS CFD under-predicts the rate of mixing out of the non-uniform flow-field, and significantly over-predicts the profile loss coefficient in a particular axial plane by an average of 17.9%. The agreement for mixed-out loss coefficient is much better, differing by just 4.0%. Detailed examination is given to local distributions of loss coefficient. A significant feature of the experimentally-measured flow-field is a radial non-uniformity caused by internal cooling features within the TE slot. This sort of effect would be very difficult to detect without testing engine components. Significant differences in local loss coefficient distributions between different radial sections are also observed due to the 3D shaping (lean, sweep, turning) of the engine component geometry. Analysis of the development of these distributions provides the basis for a model of wake profile decay based on streamwise path length. Detailed analysis of aerodynamic loss measurements are then brought together with comparison of the parametric set of TE designs with four different TE overhang lengths. Unsteady pressure fluctuations in the wake predicted by URANS CFD show a large rise in magnitude, and fall in shedding frequency, when overhang length becomes short. This is attributed to a change in shedding mode from one in which small vortices are shed at relatively high frequency from the overhanging SS TE, to one in which larger vortices are shed at lower frequency from the entire base region. Experiment and CFD are largely in agreement regarding the trend in mixed-out average loss coefficient, corrected for changes in TE thickness, between TE geometries: loss coefficient gradually rises before plateauing at short overhang length. However, the experimental data show an additional large increase in loss coefficient at intermediate overhang length which is not predicted by URANS. This discrepancy is attributed to the shedding of pressure waves associated with vortex shedding in the transonic range, which is not captured in the CFD simulations (a known limitation of the URANS method). Analysis of the base region pressure distributions predicted by URANS shows that the PS base region pressure falls substantially with reducing overhang length, whilst the SS base pressure is almost insensitive to overhang length. The differing behaviours of the PS and SS base regions are explained by interactions with changes in the entire vane aerodynamic at their different locations. There does not seem to be a simple relationship between changes in base region pressure and loss coefficient, suggesting that the overall loss trend cannot be understood without studying the aerodynamic change in the entire passage more holistically. The experimental and numerical study of NGV aerodynamic loss characteristics presented in this work additionally provided the opportunity to consider a more fundamental aspect of performance evaluation in complex 3D flows. Performance parameters such as total pressure or kinetic energy loss coefficients can be represented in various different ways ranging from line and plane averages to fully mixed-out values. There are further considerations to be made for each of these representations such as weighting methods, mixing methods and so forth. The present work analyses the sensitivities of a range of different representations of performance parameter to axial measurement location, radial averaging range, and exit Mach number, in order to provide guidance on the accuracy of each method in a relevant, practical application. In particular, the area-weighted-average, despite its simplicity (which makes it a common choice in the literature), is demonstrated to be strongly sensitive to axial plane which makes it an extremely poor choice for comparing between CFD and experiment or between different experiments.

This thesis reports a combined experimental, numerical, and analytical work aimed to develop and optimize platform cooling for high-pressure nozzle guide vanes. The work is divided into a number of sections, each aimed to tackle different challenges in designing endwall cooling systems. In the frst section, a fundamental work on the scaling of flm cooling is presented and discussed. This work aims to give insights into the use of the correct parameters when scaling flm effectiveness data. With the aid of extensive RANS and LES predictions, a number of non-dimensional groups are independently varied whilst keeping the remaining groups constant. This enabled the independent effect of each group to be studied. An important observation that departs from previous literature is that there is a strong region dependence of the sensitivity of the flm effectiveness to each of these parameters. The second section involves an experimental investigation of the effect of the aerodynamic eld on the adiabatic lm eectiveness distribution on an engine-realistic endall geometry. The study is carried out in aerody- namic similarity of the flms between engine and laboratory conditions. It is shown that for relatively low values of momentum flux ratios (i.e. representative of the low portion of the engine range) jet lift-of is ob- served, with ineffcient use of coolant. This suggests that-for a given coolant mass flow rate-endwall flm cooling systems may be optimized by increasing internal pressure loss and therefore reducing the momentum flux ratio. The third section focuses on the optimization of external flm cooling. The aim is to develop a rapid optimization routine based on the scalar tracking model and multi-objective genetic algorithms. The individual contribution of each flm cooling hole is evaluated in CFD by solving an independent passive-scalar equation, and flm effectiveness sensitivity to coolant mass fow rate is dealt with linear superposition techniques. The latter is used as a proxy for the CFD and it is computationally inexpensive. The approach can be therefore used in standard optimization techniques| such as genetic algorithms-to perform rapid system optimizations. An optimization routine involving the flm cooling of a flat plate with external flow acceleration is performed by minimizing the temperature difference from a target distribution whilst minimizing mixing loss. Limitations of the proposed strategy are also reported and discussed in detail. The fourth section involves the design of a HP-NGV endwall cooling system based on the reverse-pass scheme. That is, a system in which the internal coolant flows substantially in the opposite direction to the main- stream flow. The HP-NGV platforms are designed and assessd by using bespoke conjugate heat transfer and aerodynamic models. The proposed designs are shown to outperform the baseline confguration by reducing the metal temperature peak and flattening the temperature distribution. Three designs are down-selected for manufacturing by laser-sintering using titanium alloy and tested at high speed conditions. The final section includes the aero-thermal characterization of novel HP-NGV platform designs based on the reverse-pass scheme and manufactured by laser-sintering. Tests are performed at engine-matched Reynolds number, Mach number, Biot number, and coolant-to-mainstream pressure ratio, on a full-annulus test section. Detailed flm effectiveness distributions on vane platforms are measured by means of infrared thermography. Engine baseline and reverse-pass-based designs are compared in back-to-back experimental tests at engine-representative conditions over a range of coolant-to-mainstream pressure ratios.

Exhausting the thermal power from a fusion tokamak is a critical engineering challenge. The life of components designed for these conditions has a strong influence on the availability of the machine. For a fusion power plant this dependence becomes increasingly important, as it will influence the cost of electricity. The most extreme thermal loading for a fusion power plant will occur in the divertor region, where components will be expected to survive heat fluxes in excess of 10 MW/m2 over a number of years. This research focussed on the development of a heat sink module for operation under such conditions, drawing on advanced cooling strategies from the aerospace industry. A reference concept was developed using conjugate Computational Fluid Dynamics. The results were experimentally validated by matching Reynolds numbers on a scaled model. Heat transfer data was captured using a transient thermochromic liquid crystal technique. The results showed excellent agreement with the corresponding numerical simulations. To facilitate comparison against other divertor heat sink proposals, a nondimensional figure of merit for cooling performance was developed. When plotted against a non-dimensional mass flow rate, the reference heat sink was shown to have superior cooling performance to all other divertor proposals to date. Results from Finite Element Analysis were used in conjunction with the ITER structural design criteria to life the heat sink. The sensitivity of life to both boundary conditions, and local geometric features, were explored. The reference design was shown to be capable of exceeding the life requirements for heat fluxes in excess of 15 MW/m2. A number of heat sinks, based on the reference design, were fabricated. These underwent non-destructive testing, before experimentation in a high-heat flux facility developed by the author. The heat transfer performance of the tested modules was found to exceed that predicted by numerical modelling, which was concluded to be caused by the fabrication processes used.

Gas turbine combustors employing a single lean combustion stage represent the next generation of design for reduced NOXemissions. These lean-burn combustors rely on swirl-stabilised flames resulting in highly non-uniform outflows. Non-uniform conditions adversely affect high-pressure turbine performance. 3D numerical simulations provide a means to understand and optimise engine design, however, the modelling of turbulence means experimental validation is crucial. Turbine test facilities operating at scaled, non-reacting conditions, with simulated combustor flows are an important source of validation data. This thesis presents advances in combustor simulator design, testing and instrumentation relevant to the challenges of modern, highly-integrated turbine testing. The design of a lean-burn combustor simulator, characterised by swirl and non-uniform temperature, is presented. The design was based on measurements and predictions of engine conditions. Unsteady numerical simulations were employed as a predictive design tool. An engine-scale combustor simulator was manufactured and characterised experimentally in a bespoke facility. Surveys of flow structure are presented, focusing on experimental turbine inlet data. These data confirmed that the combustor simulator reproduces the important features of a lean-burn combustor; e.g. swirling mainstream flow and high turbulence intensity. The lean-burn combustor simulator will be the first of its kind to be implemented in a rotating turbine test facility, and will provide important validation data. Measurement techniques were developed alongside the core work. Miniaturised five-hole probe rakes for turbine inlet measurements were developed using additive manufacturing (AM). Building on this work, an open source AM five-hole probe design is presented with experimental validation. The problem of estimating pressure probe bandwidth was also addressed, and a simplified model is presented. These tools have direct applications in turbomachinery research.

This thesis describes the development of heated double-sided thin film gauge configurations for transient heat transfer measurements. By heating the substrate it is possible to measure the heat flux over a range of wall temperatures, which improves the resulting regression and leads to a more accurate determination of the adiabatic wall temperature and the heat transfer coefficient. The heated gauges can be used to measure the heat transfer of a film-cooled high-pressure turbine stage with engine-realistic combustor exit flows (hot-streaks and swirl) in order to study turbine aero-thermal performance in the presence of cooling flows. This thesis also presents the aerodynamic and heat transfer design as well as the mechanical assessment of the first fully cooled turbine stage for the Oxford Turbine Research Facility (OTRF). The high-pressure turbine stage is designed for lean-burn combustor interaction effects and has engine-state external cooling features. Platinum thin film resistive gauges are the most advanced instrumentation for transient heat transfer measurement where high frequency response is required. Double-sided thin film gauges, consisting of two thin film temperature gauges mounted on either side of an insulating layer, have the advantage that the driving temperature difference is directly determined at the point of interest. A new gauge design is presented, which combines double-sided thin film gauges with the possibility to heat the substrate layer. Two arrangements of heated gauges have been developed, namely double-sided gauges with an underside thin film heater and self-heating double-sided gauges. These gauges yield an improved regression without the need for complex heating systems. The lateral conduction effects have been investigated using analytical models. The performance of different gauge types is quantified based on the quality of the regression. The heated gauge arrangements have been tested in the OTRF under engine-representative conditions on uncooled and cooled MT1 high-pressure nozzle guide vanes. Measurement strategies have been implemented to obtain the time-dependent temperature trace over a range of heating powers and calculate the convective heat transfer. This improved measurement technique allows for an accurate determination of non-dimensional parameters such as the Nusselt number and the film cooling effectiveness.

The continuing maturation of metal laser-sintering technology has presented the opportunity to de-risk the engine design process by experimentally down-selecting high pressure nozzle guide vane (HPNGV) cooling designs using laboratory tests of laser-sintered—instead of cast—parts to assess thermal performance. Such tests are very promising as a reliable predictor of the thermal-paint-engine-test, which is used during certification to validate cooling system designs. In this thesis, conventionally cast and laser-sintered parts are compared in back-to-back experimental tests at engine-representative conditions over a range of coolant mass flow rates. Tests were performed in the University of Oxford Annular Sector Heat Transfer Facility. The aerothermal performance of the cast and laser-sintered parts is shown to be very similar, demonstrating the utility of laser-sintered parts for preliminary engine thermal assessments. It can be shown that in most situations counter-current heat exchanger arrangements outperform co-current arrangements. This concept, though familiar in the heat exchanger community, has not yet been applied to hot-section gas turbine cooling. In this thesis, the performance benefit of novel reverse-pass cooling systems—that is, systems in which the internal coolant flows substantially in the opposite direction to the mainstream flow—is demonstrated numerically and experimentally in film-cooled HPNGVs. It is shown numerically that reverse-pass cooling systems always act to flatten lateral wall temperature variation and to reduce peak metal temperature by maximising internal convective cooling at the point of minimum film cooling effectiveness. Reverse-pass cooling systems therefore require less coolant than other internal flow arrangements to maintain acceptable metal temperatures. The benefits of reverse-pass cooling can be fully realised in systems with long, undisturbed surface length, such as the suction-side (SS) of a HPNGV, afterburner liners, HPNGV platforms, and combustor liners. Three engine-scale HPNGVs with SS reverse-pass cooling systems were subsequently designed using bespoke numerical conjugate heat transfer and aerodynamic models to satisfy engine-realistic aerothermal and manufacturing constraints. The reverse-pass HPNGVs were metal laser-sintered and tested in back-to-back experiments with conventionally cooled HPNGVs in the Annular Sector Heat Transfer Facility. The reverse-pass HPNGVs are shown to reduce peak engine metal temperature by 30 K and reduce mean SS engine metal temperature by 60 K compared to conventionally cooled HPNGVs with the same cooling mass flow. A physically-based infra-red thermography procedure was implemented which takes into account the transmittance of the external optics, the surface emissivity of the object, the black-body temperature-radiometric characteristics of the camera, and the time-varying surrounding radiance. Failure to account for surrounding radiance is shown to result in an absolute error in overall cooling effectiveness of 0.05. A new experimental facility—the Coolant Capacity Rig—was developed in order to measure row-by-row, compartmental and total coolant capacity of HPNGVs to a precision of 0.03%, over a large range of pressure ratios and mass flows using a differential mass flow measurement technique, bypass system, and calibrated mass flow orifice. A novel method for estimating internal loss coefficients from the coolant capacity measurements has been devised which, uniquely, does not require internal pressure measurement.

The design, modelling and testing of a film cooling system intended for the casing of an unshrouded HP turbine rotor is described in this thesis. Due to the dense network of small film cooling holes employed in such a system, this is often referred to as a casing effusion cooling scheme. Though there are patent references to such systems, there is as yet very limited published material on the aero thermal performance of such film cooling schemes. The casing of an unshrouded HP rotor is an incredibly hostile environment, witnessing the periodic passing of the HP rotor tips within close proximity at a frequency of ∼10 kHz. These blade passing events subject the casing to extremely large amplitude fluctuations of pressure and heat load, which may at first seem to preclude the use of a film cooling scheme. This thesis details many theoretical, computational and experimental advancements related to the research topic. Highlights include: The introduction of a new fundamental mechanism to the field of film cooling, the propagation and reflection of pressure waves within film cooling holes and the impact on film cooling performance. The development of new miniature thin film heat flux gauges manufactured using a new process. Sensor resolution is improved by a factor of seven. The first published computational model reporting heat transfer data on a film cooled rotor casing. Improvements to heat transfer data processing techniques and theory. These are applied to experimental work to produce the highest resolution heat transfer data obtained on the casing of a scaled rotating transonic HP rotor for both uncooled and cooled geometries. Computational models are used to demonstrate that coolant injection on the rotor casing reduces the over-tip leakage mass flow, offsetting the spoiling and mixing losses that film cooling schemes introduce. Much of the work in this thesis is based on papers that have been submitted to or are pending submission. To date three papers have been presented at conference with two published in journals and the third recommended and pending journal publication. Two other papers are pending submission. A patent has also been filed with the European and American patent office regarding novel film cooling hole shapes designed to make use of acoustic effects.