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The flow at the trailing edge of a turbine blade at transonic air velocities can be extremely complex. Solid trailing edges can shed vortices in a manner known as transonic vortex shedding, where vortices form very close to the trailing edge and cause large trailing edge shear layer deflection. This, in turn, results in shockwaves that can propagate upstream of the trailing edge. When this flow regime occurs, it is known to be the loss dominating mechanism for trailing edge flows. In this dissertation experiments and LES simulations are performed to increase the understanding of the flow mechanisms at the trailing edge, both for solid and cooled trailing edges. It was found that transonic vortex shedding is not the only flow regime possible behind round trailing edges at transonic air velocities. Under certain conditions the vortices are found to form approximately one trailing edge diameter downstream of the trailing edge, both the shear layer deflection and shockwave formation are dramatically reduced. This flow regime has been referred to as detached vortex shedding. The switch from detached to transonic vortex shedding is found to be the result of the transition of the pressure surface boundary layer and is characterised by a close to doubling in the mixed-out loss for the trailing edge tested. The effect of trailing edge wedge angle on the performance of solid, round trailing edges is also investigated. It is found that wedge angle is an important parameter governing trailing edge performance, with higher wedge angles offering reductions in loss. Switching from an 8◦ to a 14◦ wedge angle plate was found to reduce the loss by up to 29% when the plates were in the same flow regime. The effect of geometry variation on blown, through trailing edge holes geometries is investigated experimentally. These geometries undergo transonic vortex shedding when there is no trailing edge blowing, but the addition of coolant to the base region is able to suppress the vortex shedding and reduce the loss by up to 39%. In order to disrupt the vortex shedding, a threshold coolant mass flow ratio must be reached; the value of this threshold depends on Reynolds number, since this governs the strength of the vortex shedding. Holes of half the area, or holes drilled obliquely to the trailing edge are able to reduce the coolant mass flow for a given value of coolant stagnation pressure coefficient. Finally, the effects of manufacturing variation on trailing edge loss is investigated with the aid of GOM scans from real trailing edge geometries. For solid trailing edges, variations in geometry away from an ideal round trailing edge are found to have significant effects on the loss. This is almost always beneficial, since the performance is governed by the ability of the geometry to suppress transonic vortex shedding. Some variations are able to reduce the loss by over 50% from the baseline round geometry. For blown, cutback geometries, the differences in loss due to manufacturing variations are smaller (at most 5% reduction in loss). But, over the majority of operating points tested, there was no disadvantage and indeed sometimes a small advantage of real geometry.

The performance of turbomachines can be principally defined by differential pressure, electrical power and hydraulic efficiency at various volumetric flow rates; often known as the pump curve. To date the measurement of hydraulic efficiency has proved inaccurate, time consuming and cost prohibitive using contemporary methods. The conventional method relies on the measurement of differential pressure, flow rate, electrical power and an estimation of motor efficiency, where in-situ measurement of volumetric flow rate is the single biggest source of uncertainty using this method. An alternative to the conventional technique is the thermodynamic method, which requires only the measurement of differential pressure and differential temperature across a turbomachine to compute the hydraulic efficiency. The disadvantage is that, due to the remarkably high specific heat capacity of water, the temperature difference is very small, requiring thermometers of milli-Kelvin accuracy. Simulations demonstrated that pump efficiency accuracy is some thousand times more sensitive to differential temperature than absolute prompting a design focus on difference thermometer sets. No differential thermometer sets existed with the needed accuracy and the ability to function in an industrial environment for extended durations, so this technology was designed and developed. There was little research into difference thermometer set calibration techniques, this being currently being undertaken by absolute means, so a protocol for this also had to be developed. A suite of hardware was developed and tested in a throttling calorimeter and, together with the developed calibration methodology, proved capable of achieving an accuracy of 0.29mK. Such accuracy means on-site, in-situ measurement of turbomachinery efficiency to an accuracy of 0.5% is now made possible. Industrial environment case studies demonstrated the identification of tangible engineering intervention cost-saving measures for turbomachinery, such as water pumps and ventilation fans, made possible through this technology. Keywords: turbomachinery, pumps, fans, energy efficiency, thermodynamic method, Poirson, difference thermometry, pump efficiency.

Compact axial compressors are of interest to the domestic appliance industry. The associated low Reynolds number leads to high losses compared to large-scale compressors due to a transitional flow field with large regions of separation. This thesis investigates how Reynolds number variations affect the three-dimensional and unsteady flow field in a compact axial compressor both pre-stall and in stall. An experimental study has been conducted using a scaled-up single-stage axial compressor across a Reynolds number range of 10⁴ to 10⁵. Steady and unsteady casing static pressure measurements, along with rotor upstream and downstream unsteady velocity measurements, have been used to observe the rotor flow field. As the Reynolds number is reduced below a critical value, 60,000 in the case of the compressor studied, the pressure rise coefficient of the compressor rapidly decreases. This fall-off in performance, and the critical Reynolds number, is determined by the point at which the compressor endwall separations begin to grow in size. Compact axial compressors have relatively large tip clearances, typically greater than 3% of the chord length. This results in unsteady movement of the tip leakage jet, which impinges on the subsequent blade, and causes leading edge spillage of the flow when operating towards stall. This is particularly prevalent when there is a large casing endwall separation, which can be caused by operating below the compressor's critical Reynolds number. This spillage leads to a broadband hump and a pattern of equally spaced peaks appearing in the near-field casing static pressure spectra. The frequency spacing of these peaks has been found to equal the measured stall cell speed once rotating stall is established. With the use of an analytical model, and detailed unsteady measurements of the tip leakage jet, it is shown that these flow structures, which decay over a short circumferential distance and rotate at the subsequent stall cell speed, are capable of generating this spectral pattern. Previous research into axial compressor stall has mainly focused on stall inception and methods to extend the stable operating range. This thesis considers the performance of an axial compressor beyond stall and investigates how the characteristics of stall cells depend on Reynolds number. Detailed unsteady measurements have been used to measure the behaviour across a range of in-stall flow coefficients. These measurements have been used to extract the stall hysteresis and to determine the size, speed, number, and spanwise extent of the stall cells. The results show that for the stalled compressor, as the Reynolds number increases, the size of the minimum stall cell that can be sustained decreases. This means that a larger change in throttle area is needed to reduce the stall cell down to a size where the compressor can recover from stall. The stall cells increase in size and slow down as the Reynolds number decreases. The size and shape of the stall cells that form are related to the extent of the three-dimensional flow field present prior to stall. In all cases, as the number of stall cells increases, so do the speed and the total size of the stall cells. Further reductions in the flow coefficient cause an increase in the total size of the stall cells and a decrease in the speed. The value of the critical Reynolds number varies with the compressor blade geometry. A redistribution of the rotor spanwise loading away from the endwalls, with an increased loading at the midspan, can recover some of the loss in performance at the lower Reynolds numbers, and decrease the critical Reynolds number. Increasing the rotor clearance gap in a compact axial compressor results in a significant loss in pressure rise. However, when operating below the critical Reynolds number, this loss is lower, because the casing endwall flow is already separated at lower clearances. Increasing the rotor clearance reduces the stall hysteresis by increasing the minimum stable stall cell size. Below a critical value of Reynolds number, which is dependent on the specific geometry used, the pre-stall and in-stall flow features are highly sensitive to Reynolds number, becoming increasingly three-dimensional and unsteady.

By ingesting the fuselage Boundary Layer (BL) and re-energising it, Boundary Layer Ingestion (BLI) offers the potential for considerable fuel savings. However, it remains unclear which the most promising aircraft configuration for BLI is. Furthermore, the tight integration between the propulsor and the airframe poses great challenges to design and analysis. The aim of this thesis is to provide the first detailed assessment of the flow field and BLI performance of a novel distributed aft-fuselage BLI aircraft. In particular, the focus is on the link between the flow features to the mechanical power losses within the flow field and their consequent impact on BLI performance. A fully-coupled computational test case featuring the aft-fuselage installation and propulsor geometries was set up to resolve the external and propulsor internal flow fields. The unsteady simulations were carried out using the Unsteady Reynolds-Averaged Navier-Stokes (URANS) approach. Using the computed flow field, a mechanical power balance was drawn up to analyse its performance and identify the loss sources. By varying the boundary conditions of the simulation and/or BLI fan rotational speed, the effects of changes in fuselage BL thickness, aircraft cruise Mach number and BLI propulsor Fan Pressure Ratio (FPR) were also investigated. The baseline flow field of the BLI aircraft at cruise shows that the exhaust region incurs a significant amount of loss due to the mixing of the exhaust jet and the BL wake. With a thinner incoming BL, a different mode of flow field was observed, with a significant increase in loss caused by supersonic flow and separation over the cowl. This highlights the challenges associated with the robustness of the installation design and operability of the aircraft. As the flight Mach number is reduced, the flow over the cowl is improved, but with the same FPR, this leads to increased loss in the exhaust region due to the higher ratio of jet to flight speeds, which has a strong effect on the mixing loss. At reduced FPRs, the over-speed in the jet relative to the freestream is reduced, but mixing with the wake can leave behind a larger velocity non-uniformity and therefore energy outflow downstream if the FPR is too low. Reducing the fan speed also leads to a redistribution of the losses within the fan stage, as the rotor shock loss is decreased, whilst the changes in incidence onto the rotor and stator see larger variations. Besides identifying the important flow features specific to this novel BLI concept and showing their links to mechanical power loss, this thesis makes a contribution to the simulations and analysis of highly integrated BLI systems. Through the URANS simulations, the interaction between the fuselage BL turbulence and the BLI propulsor was identified as an area that warrants further research. It is hoped that this research will be applicable to not only BLI aircraft, but also future aircraft concepts that feature tightly integrated propulsion systems.

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.

As an axial compressor is throttled three-dimensional separations develop in the corners between the blades and annulus endwall. Surprisingly, little is understood about the unsteady topology of these separations. One of the problems with studying corner separations is that it is often difficult to understand whether a particular flow structure in the separation is inherent to the separation itself, or due to the response of the separation to changes in the inlet flow. In this thesis a novel experimental approach is taken with the aim of isolating the corner separation from external influences. A cascade is designed with the specific aim of precisely controlling the inlet flow. Contrary to previous work, it is shown that the key saddle and focus pair, which describes the time-mean topology of the corner separation on the endwall, moves smoothly and continuously as the incidence of the flow is raised. This behavior is shown to be the result of the time-resolved topology of the flow field, which comprises numerous saddle and focus pairs which are produced stochastically in regions of high streamline curvature. Most importantly, the separation is shown to exhibit an extremely long timescale behavior, changing in topology over timescales many times longer than the convection time through the blade passage. The behavior is shown to be intrinsic to the separation and causes the separation, for periods, to completely disappear from the endwall. This underlying unsteady structure is shown to have implications for the ability of RANS-based design codes to accurately predict corner separations.

High fidelity simulations of three compressor cascades and a fundamental trailing edge flow are used to interrogate, and improve upon, the errors present in conventional RANS simulations. Using the optimum turbulence viscosity, calculated from the high fidelity simulations, it is found that error in moderate Reynolds number (Re_c > 300,000) compressor cascades is dominated by errors due to the turbulence model, as opposed to errors inherent to the use of a turbulence viscosity or the RANS method itself. Significant errors due to using turbulence viscosity based RANS to model turbulence are found in a compressor cascade operated at a chord-wise Reynolds number of 148,000. Focusing on a compressor cascade's suction surface boundary layer, it is found that much of the error when using the k-ω SST turbulence model is due to incorrect non-equilibrium boundary layer modelling. Using a Design of Experiments based approach, it is found that by reducing β_1, a turbulence model constant affecting near wall modelling of ω, significant improvements in non-equilibrium modelling can be achieved. These improvements are found to be robust with respect to blade loading style and Mach number (within the subsonic regime), and also lead to improvements in blade performance parameter prediction. These results are used to investigate the impact of non-equilibrium boundary layer effects on compressor cascade performance. It is found that vortex shedding trailing edge flows cannot be satisfactorily modelled using conventional steady RANS simulations. Using the Selective Frequency Damping technique, two reasons for this are found: errors arising from the Boussinesq approximation, and the fact that the correct time-averaged flow is an unstable solution of the RANS equations, leading to convergence issues. It is shown that, despite not being able to stabilise the RANS equations on the correct mean flow, Reynolds stresses calculated from the high fidelity simulation are the optimum RANS equation source terms. By modelling the vortex shedding process while simulating turbulence, Unsteady RANS simulations are shown to qualitatively predict all the main features of the fundamental trailing edge flow.

Further improvements in aero-engine efficiencies require better understanding of loss mechanisms. The rise of high performance computing is unlocking the potential of scale-resolving simulations for industrially relevant cases thus allowing new levels of simulation fidelity. As a result, previously unexplored effects of unsteadiness can be simulated and their impact on loss understood. The work undertaken in this thesis aims to establish a framework for accurate loss predictions using scale-resolving simulations and inform the field with regards to the effects of unsteadiness on loss mechanisms within the multi-stage compressors. The lack of computational requirements for accurate industrial simulations lead to inconsistent loss predictions even for scale-resolving simulations depending on the chosen convergence criteria. This work studies aspects of loss generation by employing two test-cases: Taylor-Green vortex and compressor cascade subjected to freestream turbulence. The results show that both resolving local entropy generation rate and capturing the inception and growth of instabilities are critical to accuracy of loss prediction. In particular, the interaction of free-stream turbulence at the leading-edge and development of instabilities in the laminar region of the boundary layer are found to be important. These two outcomes allow for a formulation of resolution criteria that ensure accurate loss predictions for compressor flows. One of the major sources of uncertainty in the current simulation methods for compressor flows is the level of unsteadiness and its impact on loss This work makes a series of steps towards understanding the nature and the origin of unsteadiness within multi-stage machines and investiages the impact of gapping on mid-span compressor loss. It is found that freestream turbulence levels rise significantly as the size of the rotor-stator axial gap is reduced. This is because of the effect of axial gap on turbulence production mechanisms, which amplify at smaller axial gaps and drive increases in dissipation and loss. This effect is found to raise loss by between 5.5-9.5% over the range of conditions tested here. This effect was found to significantly outweigh the beneficial effects of wake recovery on loss.

There is an economical and ecological incentive for jet engine manufacturers to reduce fuel consumption of its engines. Reducing the pressure ratio of the fan system is one route through which fuel consumption can be reduced by increasing propulsive efficiency. However, reducing the fan total pressure ratio encourages a move to increase the stage loading, which can enable engine weight reduction. This moves the fan into an area of the design space that is highly challenging in terms of stall. Stall inception of a modern, high stage loading, low pressure ratio transonic was investigated using single-passage steady and full annulus unsteady computational fluid dynamics. The computational method was validated using experimental data and approached as a blind test case. A key post-processing method used in this work to expose the route to stall in the full annulus unsteady results was the circumferential averaging of incidence and blockage for each passage separately. Sudden growth of the suction surface separation behind the shock around the mid-span of the mis-staggered blade passage has been found to trigger stall. The baseline fan geometry has a mid-loaded radial work distribution, so a combination of high work loading and high incidence made the suction surface boundary layer behind the strong shock sensitive to separation. After stall has been triggered around mid-span, the stall cell embryo moved to the casing and develops into a stable stall cell. Two design alterations to the baseline fan geometry confirmed that stall was triggered around mid-span: the effect of changing tip gap is small and the introduction of a slip patch on the suction surface between 25-75\% span increased stall margin by 25\%. These showed there is significant potential to increase stall margin by delaying the onset of the suction surface boundary layer separation separation behind the shock around the mid-span of the baseline rotor geometry. It was shown that redesigning the baseline fan rotor geometry between 25-75\% span only can be beneficial for stall margin. The rotor blade leading edge was recambered in the direction to decrease incidence near the stability limit relative to the baseline. This decreased the pre-shock Mach number by 13\% near stall, which reduced the suction surface boundary layer separation behind the shock and in this way increased the stall margin by 6.7\%. This design change had minimal effect on the design point efficiency, which stayed within 0.1\% of the datum geometry. At off-design rotational speeds, the behaviour of the suction surface boundary layer changed. At 100\% speed, the blade relative Mach number increased by 9\%, which caused the radial work distribution of the fan to increase towards the tip of the rotor blade compared to design speed. The trigger of stall inception, which was still driven by suction surface separation behind the shock, followed this radial distribution change and moved away from mid-span towards the tip region of the rotor blade. At part speed conditions, the tip relative Mach number dropped to become subsonic, which altered the behaviour of the interaction between the shock and suction surface boundary layer. The suction surface boundary layer now separated close to the leading edge near stall (x/c_x < 1.5%) and induced axial oscillations prior to stall. Eventually, these stable periodic oscillations became unstable and pushed the fan into stall. The implications of this work is that a fan designer who aims to maximise flow range for a high stage loading low pressure ratio transonic fan needs to pay close attention to the behaviour of the suction surface boundary layer and its interaction with the shock along the whole span of the rotor blade since increasing stage loading coefficient make the suction surface boundary layer more prone to separate behind the shock. It has been shown that the radial location of the stall trigger changes concurrently with changes in pre-shock Mach number and radial work loading distribution along the span. Therefore, it is insufficient to solely focus on the tip region when investigating stall inception in high stage loading low pressure ratio transonic fans.

Using a series of direct numerical simulations, the individual and cumulative effects of various disturbance environments existing in a low pressure turbine (LPT) are investigated. In particular, the effects of free-stream turbulence (FST), unsteady wakes, roughness and blade oscillations on the separation-induced transition on the suction surface of a low pressure turbine blade are analyzed. Two configurations are considered: (i) a flat plate subjected to streamwise pressure gradient representative of the suction surface of a low pressure turbine blade, (ii) a flat plate subjected to a convecting free-stream vortex of fixed strength and at a fixed height over the plate. The first configuration represents the ‘ultra high-lift’ blades for the next generation low pressure turbine. The local pressure gradient induced by the convecting vortex in the second configuration is representative of the adverse pressure gradient on the suction surface of a low pressure turbine blade. The results are validated against existing experimental or numerical data and it is demonstrated that the numerical framework has captured most of the phenomena to a reasonable level of accuracy. A kernel experiment for bypass transition is simulated for the vortex-induced instability. The effect of the convection speed and strength of the vortex are discussed and the paths of transition adopted are distinguished. At low disturbance levels, the transition to turbulence is primarily due to the breakdown of ‘Kelvin-Helmholtz’ roll up vortices. In the presence of aeroelastic blade oscillations, unsteady wakes, free-stream turbulence and roughness, transition takes the bypass route and the results show evidence of streamwise streaks. These streaks impart spanwise waviness to the separated shear layer and cause early destabilisation. The blade oscillation has an effect in reducing the separated region and hence, the profile loss, which is further accentuated in the presence of free-stream turbulence. A blade fluctuating at higher reduced frequency is found to be more effective in shrinking the separation region. Blade vibration is found to increase the level of pre-transitional fluctuations, without having a significant influence on the growth beyond separation. There is a cumulative effect in suppressing the separation region when blade oscillation and free-stream turbulence are studied in conjunction, although the additional effect of free-stream turbulence is marginal. A secondary separation bubble, noted in the unperturbed flow, is reduced in size with blade oscillation and further reduced in the presence of free-stream turbulence. The vortex-induced instability has been proposed to be a unit process of free-stream turbulence, the effect of which is studied in the presence of a discrete roughness element (similar in functionality to a trip wire). The roughness element triggers early transition by destabilizing the mean flow. Streaks are observed in the presence of the convecting and rotating cylinder, generating a vortex of fixed strength, and are enhanced by the presence of roughness in the pre-transitional zone. Enhanced spanwise waviness is noted with the roughness, leading to earlier breakdown to turbulence. The route of transition and the origin of three-dimensionality marked by the prominence of the vortex stretching is shown. An optimum range of convection speeds of the free-stream vortex is obtained and the maximum receptivity is noted at a speed of 0.386, which concurs with prior experiments on a periodic convecting vortex (Kendall, 1987). The unsteady wake has a direct effect on the velocity profile. A lag is noted between the wake passing and transition. While the wake convects at the local free-stream velocity, its impression in the boundary layer convects much slower, between 50% and 70% of the local free-stream velocity. Both unsteady wakes and blade oscillation promote near-wall mixing. The unsteady wakes and blade oscillations have a conjunctive effect on reducing the size of separation bubble. The secondary separation bubble observed in the unperturbed flow is reduced with the presence of wakes and is completely suppressed with the addition of the blade oscillation. Turbulent kinetic energy production increases with increasing perturbation levels, with the maximum effect seen for the combination of wakes and oscillation. A receptivity method based on disturbance enstrophy transport equation (DETE) is proposed. The disturbance enstrophy is evaluated as the difference between the instantaneous and mean enstrophies, and while these are positive definite quantities, the difference may not be so. This aspect of disturbance enstrophy has been used meaningfully to obtain new information about the flow instability, such as segmenting regions of flow instabilities into those dictated by positive growth rates of disturbance enstrophy, and those due to its negative counterpart. The positive growth rates are associated with large scale coherence in the free-stream whereas the negative growth rates are found to be arising in near-wall viscous structures (Sengupta et al., 2019a,b). The extension of this method to structure detection and comparison against existing vortex identification methods, indicates the ability of the DETE method to capture small-scale structures induced by the viscous term of the Navier- Stokes equation. DETE has been used for the two configurations operating with varying perturbation levels and valuable insight pertaining to the flow dynamics has been attained with respect to its budget terms. In particular, the role of vortex stretching in leading the flow to three-dimensionality is highlighted. The global and local spatio-temporal receptivity analysis of flows perturbed by plate oscillations, wakes and free-stream excitation yields a spatio-temporal wave front to be the causal mechanism for flow transition. This is essentially representative of the nonmodal part of the disturbance spectrum. While this flow instability has already been established for geophysical flows such as tsunami and other fluid dynamical problems such as for zero pressure gradient boundary layer formed over a semi-infinite flat plate excited from inside the shear layer, its role for pressure gradient dominated flows (such as in LPTs) is shown for the first time. The importance of nonlinearity and its dispersion effects is shown, specifically for flows excited at the free stream, even when the onset of disturbances follows a linear mechanism. The need for using a global, nonlinear, spatio-temporal setup is established in order to capture all pertinent flow physics.

In gas turbines, film cooling is required to protect metal parts from hot combustion gases. Reduction in coolant mass flow increases cycle efficiency, and hence reduces greenhouse gas emissions. However, the lifespan of a cooled component is sensitive to the metal temperature within the part. A designer requires predictions of cooling performance to make this compromise, yet present design methods are subject to uncertainty and are not viable without empirical input. Flow through a turbine is inherently unsteady due to relative motion of stators and rotors, termed blade row interaction. Blade row interaction is not captured in flat-plate and cascade testing, or present design methods, contributing to uncertainty in predicted cooling performance. The aims of this thesis are to establish the mechanisms by which blade row interaction affects rotor film cooling, and quantify their influence on cooling performance in a representative case. A new experimental rig is developed to facilitate aerodynamic and heat transfer measurements of cooling holes subject to unsteady main-stream boundary conditions. The effect of unsteadiness is set by non-linearity in the hole response. Unsteadiness reduces film effectiveness by up to 31% with cylindrical holes at a low momentum flux ratio, because the response to perturbations is non-linear. Cylindrical holes at a high momentum flux ratio, and fan-shaped holes, are robust to unsteadiness because they respond linearly. Non-film-resolved computations are used to identify the blade row interaction mechanisms generating unsteady main-stream boundary conditions in a turbine rotor. A quasi-steady model is used to predict instantaneous excursions in cooling hole momentum flux ratio. Fluctuations of at least ±30% are present for all hole locations, due to both upstream vane wake and potential field interaction. A hybrid URANS-LES computational approach is implemented, validated against experimental data, and applied to a turbine stage cascade model. Compared to steady conditions, blade row interaction reduces rotor film effectiveness: by up to 18% on the pressure side, due to migration of vane coolant across the passage; and by up to 30% on the suction side, due to wake interactions increasing the film mixing rate.

Modern gas turbines employ Endwall Contouring (EWC) to improve the efficiency of gas turbines by reducing the losses associated with secondary flow. The secondary flow features are influenced by the egress flow exiting the wheel-space between stator and rotor. Previous studies have demonstrated that endwall contouring designs developed without egress flow may show a poorer performance once this is taken into account. This reveals the need to conduct a combined design of contoured endwalls in the presence of egress flow. CFD computations can be employed to generate these designs, but require experimental validation. This thesis describes the development of a novel experimental technique, Volumetric Velocimetry, in steps of increasing complexity. The development can be divided into three main stages: (i) implementation of VV in testing a stationary facility for studying film cooling, (ii) the development of borescope probes to introduce the laser beam in rigs with restricted access, and (iii) implementation of VV inside a 1-stage constant duration turbine facility, and testing with different purge flow cases. These measurements show the full three-dimensional flow field including secondary flow features. The novelty of the experimental technique required the design of a new facility, named the Film Cooling Wind Tunnel, to develop and debug the technique in a stationary turbomachinery scenario. The design was targeted towards optical techniques and included a modular design to accommodate a large range of flow and geometry parameters for studies of film cooling. Volumetric Velocimetry successfully captured the flowfield including the counter-rotating kidney vortices. The measurements were validated against prior studies available in the literature. The effect of the momentum flux ratio, up to IR = 6.50, was assessed by investigating the 3D vortex structures of film cooling. The circulation of the kidney vortices was observed to increase significantly with increasing momentum flux ratio. For each momentum flux ratio case, the circulation was shown to increase to a peak and then decay along the streamwise direction. Vortex identification and tracking showed that the lateral separation of the counter-rotating vortex pair correlated well with the self-induced velocity dipole for high momentum flux ratios. The results demonstrated that VV is a powerful technique for capturing 3D flows. The experimental processes developed during this phase were taken forward to the rotating turbine. Prior to implementation of Volumetric Velocimetry in a turbine facility it was necessary to develop borescope probes to introduce the laser beam into areas of limited optical access. To achieve this objective a modular software program was developed, capable of predicting the path of the laser beam through an array of lenses. The program features include: modularity of optical elements, laser power distribution, window refraction calculation, and an evaluation of laser power density at inlet and outlet. The program was employed to develop several designs, and also helped identify the challenges associated with this application, namely regarding the laser damage threshold of optical materials and the applicability of different optical elements. Several probe designs were manufactured in-house and tested. One of the probes was successfully employed to conduct Volumetric Velocimetry measurements in the Film Cooling Wind Tunnel, demonstrating good agreement with the results obtained previously. The program has been made open source for other researchers to benefit. Volumetric Velocimetry was implemented for the first time in a rotating turbomachinery facility in the Large Annulus Rig, a new facility designed for studying both egress flow and endwall contouring. This implementation built upon the testing and borescope development carried out beforehand in the Film Cooling Wind Tunnel. The campaign in this facility required: (i) the development of a mathematical model of the wind tunnel, used for laser beam path prediction and for computing the clock angle of the turbine during the measurements; (ii) bespoke seeding; (iii) laser delivery using a borescope inserted in a windowed vane; and (iv) a method to transform the results from the camera reference frame into the absolute and the relative reference frames of the turbine. Measurements were performed without purge flow and with a sealed wheel-space. The results demonstrated that the presence of egress accentuates the pressure side horseshoe vortex, increasing the strength and penetration depth of the secondary flow features in the rotor passage. The purge flow was shown to enhance the velocity in the boundary layer just downstream of the rim-seal. Schreiner et al. (2019) conducted CFD computations for the same geometry and flow conditions and also concluded that the secondary flow features were enhanced by purge flow. The VV results validate the conclusions obtained with CFD computations. This PhD thesis has demonstrated the first implementation of Volumetric Velocimetry for measurements in both static and rotating turbomachinery facilities. This has opened the door for future researchers to investigate the highly complex three-dimensional flowfields associated with egress flow and mainstream gas path interactions, as well as the impact of EWC.

Despite the blades downstream of the combustor being cooled in most large gas turbines, the effect this cooling has on their performance is still unclear. Fundamental questions, such as how turbine efficiency should be defined and which flow features reduce performance, remain unanswered. This thesis answers these questions and addresses the problem of how heat transfer affects turbine performance. Currently, there is a direct contradiction between industrial experience and the academic methods used to evaluate turbine performance. In a cooled turbine, these entropy-centric (exergy) methods predict that the irreversible heat transfer due to cooling would cause an extremely large drop in turbine efficiency (around 4-6%). In practice, this drop is not observed. By applying a new method called euergy, this thesis demonstrates that the performance of a cooled turbine can be defined in a way that agrees with industrial experience. This method shares the view of the practical device that the ideal work is defined by a reversible adiabatic turbine. A key consequence of this method is the value placed on heat, relative to work, becomes set by the Joule (Brayton) cycle efficiency. This means whenever heat is transferred, or when viscous reheat occurs, the value of this heat should be set by the Joule cycle efficiency. This new understanding finally allows both the efficiency to be defined and the flow features that change it to be identified. The method also provides cooling designers with a new way of raising turbine efficiency, a form of thermal recuperation in the flow. This mechanism offers the exciting potential that future cooling systems, when added to a blade profile, could reduce profile loss by up to 9%. Furthermore, the generality of the new method allows all cooled components, however complex, to be systematically analysed for the first time. This is demonstrated using a computational approach. The large-scale effect of heat transfer on performance is captured in newly developed loss models which are then compared with conjugate heat transfer simulations. The small-scale effect of heat transfer on performance is investigated by examining the near-wall region of cooled boundary layers using direct numerical simulation (DNS). This thesis establishes a new framework and illustrates how heat transfer affects turbine performance. This new understanding allows performance to be communicated across all levels of turbine design. The new method offers the prospect of future innovative cooling schemes which, in addition to cooling the blade, act to raise turbine efficiency.

Boundary Layer Ingestion (BLI) technology could achieve a fuel burn reduction of up to 15%. However, the fans need to operate in severely distorted flow, which affects the blades' mechanical integrity, aerodynamic efficiency and stability. The aim of this thesis is to understand the impact of strengthening rotor blades on aerodynamic performance, investigate how the performance can be improved by modifying the fan rotor radial work profile and explain the mechanism that allows BLI fans to operate stably even when there are regions where the fan operates beyond the clean flow stability limit. Two test cases were used: a low speed rig fan and a representative transonic fan. The transonic fan rotor design was thickened to satisfy aeromechanical criteria and the flow field was simulated with a representative BLI inlet stagnation profile to investigate the impact of the design change. A combination of computational and experimental techniques was utilised to redesign and test a new low speed fan rotor, for a different purpose. The blade metal angles, chord and positions of maximum camber and thickness were modified to reduce the peak in loading, and hence improve the aerodynamic performance and the stability margin in the same BLI inlet flow field as for the transonic fan. Unsteady casing static pressure measurements were used to track the development of disturbances in the experimental rig at the rotor tip to analyse stall inception in distortion. The necessary transonic blade design modification led to a loss of efficiency of 0.5% and reduction of mass flow of 2%, which were attributed to a rise in blockage due to blade thickening and an increase in profile loss in the hub region. This is a promising result, as it indicates that the necessary mechanical design changes only impacted slightly on the aerodynamic performance of the BLI fan. The low speed fan redesign reduced the radial loading at the tip and hub and increased at the mid span to keep the same operating point. This helped alleviate the losses at the blade tip in distorted flow and improved blade performance, also increasing the stability margin. Casing static pressure measurements demonstrated that disturbances are created where the incidence is higher than the critical level at a near stall point in clean flow, but they decay as they propagate into the region where the incidence drops below the critical level. Stall was found to occur once the disturbances were able to propagate around the entire annulus without being suppressed, which is why the loss of stability margin due to distortion is small in BLI fans. The results are promising as they indicate that the impact of necessary mechanical design changes is not severe and that adjusting the blade design to target the regions of peak loss in distorted flow can recover some efficiency lost due to a combination of operating in distorted flow and requiring a more robust mechanical design. Moreover, the stall inception mechanism identified supports the previous findings that loss of stability margin due to distortion is minor, confirming that stability is not an issue for a fan operating in BLI.

There is a continued need to further capabilities of turbomachinery flow design and analysis method tools, both in terms of the flow modeling fidelity (e.g. moving towards DES/LES) and temporal and spatial accuracy of the numerical discretization methods. To enhance solution accuracy with high-order methods, an implicit solver using the Discontinuous Galerkin (DG) discretization on unstructured meshes has been developed. The DG scheme is favored chiefly due to its distinctive feature that a higher-order accuracy being achieved through simple internal sub-divisions of a given mesh cell. It thus can considerably relieve the burden of mesh generation for local refinement of complex configurations. The implicit formulation is well suited for an unstructured mesh solver, for which some well-established solution techniques for explicit methods (e.g. multigrid methods) are less effective. The RANS equations in conjunction with the Sparlart-Allmaras one equation turbulence model are discretized using a high-order DG method on unstructured meshes. The all-speed AUSM+-up scheme is applied for the inviscid flux calculations and the BR2 scheme is implemented for the viscous flux discretization. An implicit time integration is adopted and the system linear equations are solved at each time step by using a preconditioned GMRES iterative algorithm with the ILU0. A modification of the turbulence model is also adopted to enhance the robustness for high-order turbulent flow simulations. The validity and effectiveness of the present developed method have been examined in several test cases, for both validating the method implementation and illustrating salient features and characteristics of the methodology. The flows around a cylinder solved up to 5th-order of accuracy on an extremely coarse mesh illustrate the need for high-order geometrical representation in relation to the high-order flow field discretization. A laminar boundary layer and the vortex shedding after a cylinder validate both the steady and unsteady modelling capabilities with the implicit time integration method. A calculated turbulent boundary layer on flat plate demonstrates the capability of a high-order DG in capturing the laminar sub-layer in a coarse mesh (y+>20) without a wall function, in a clear contrast to a conventional 2nd-order scheme. The all-speed flows past an airfoil show the distinct capability of AUSM+-up scheme on solving a wide range of flow regimes. A main focal test case of the present work is a high pressure turbine cascade, where the flow losses are strongly influenced by the boundary layer transition on the blade surface and the secondary flow development in the endwall region. In this case the RANS solutions consistently over-predict the losses, prompting further modelling considerations. A higher-order solution does indicate higher resolution for more detailed 3D secondary flow structures. It is however more remarkable that both overall and distributed losses of the unsteady laminar solutions on a coarse mesh are consistently better than the RANS solutions. The results indicate that for the turbine blade flows similar to the one presently tested, a direct unsteady laminar solution on a coarse mesh (far coarser than those typically required for LES) may be further explored.

The continuing rapid growth of civil aviation and the associated greenhouse gas emissions necessitate further reduction of gas turbine engine specific fuel consumption. Since the use of high pressure compressor bleed air to cool combustor and turbine components is parasitic to engine efficiency, one of the key drivers of this progress is the ongoing improvement of the efficiency with which the coolant air is used. Because it is exposed to the undiluted and highly turbulent combustor outlet gases, the leading edge of the nozzle guide vane requires a sizeable proportion of the overall turbine coolant. Moreover, the low mainstream flow momentum in the leading edge stagnation region means that film coolant jets are highly prone to lift off of the surface, so are often poorly utilised. This study investigates the performance of this film coolant both experimentally and computationally, and proposes a novel film cooling geometry in which jets initially travel against the mainstream flow before crossing the geometric stagnation line. This is shown to achieve vastly more effective film coolant usage near the stagnation line than can conventional designs, without compromising performance farther downstream on the vane surface. Additionally, the importance of combustor dilution port flows in determining the performance of conventional leading edge film cooling designs is quantified, as is the importance of infrared calibrations which take into account temperature inhomogeneities which can occur in advanced aerothermal experimental apparatuses. A novel infrared thermography technique has been developed for determining both of the ubiquitous film cooling performance parameters in a single experiment, where usually two are required. Also proposed is a leading edge coolant supply passage geometric modification which reduces the risk of destructive ingestion of hot combustion gases into film cooling holes. This innovation is also conducive to the improvement of gas turbine engine efficiency, primarily through its potential to facilitate reduced parasitic combustor pressure losses.