Die Kühlung thermisch hoch belasteter Bauteile wird häufig mit Kühlverfahren realisiert, die auf dem Prinzip des Massetransports durch die Bauteilwand beruhen. Neben der Film- und Transpirationskühlung gehört die Effusionskühlung zu diesen Verfahren und basiert auf einer Reihe oder einem Raster von Bohrungen. Dieser Ansatz ermöglicht sowohl den Abtransport von Wärme aus dem Bauteil als auch die Ausbildung eines schützenden Kühlmittelfilms auf der Bauteiloberfläche. Viele Autoren beschäftigten sich in ihren Arbeiten mit den Auswirkungen der Filmkühlung auf den Wärmeübergang an der Bauteilwand und definierten einen Filmkühlwirkungsgrad, der die Effektivität der Kühlung widerspiegelt. Auch die Freistrahlen aus Effusionskühlbohrungen wurden mit diesen Mitteln untersucht und eine Vielzahl unterschiedlicher Einflussgrößen auf den Filmkühlwirkungsgrad identifiziert. Dazu gehören insbesondere geometrische Bedingungen, wie z.B. der Bohrungswinkel, das Verhältnis von Länge zu Durchmesser der Bohrung und die Austrittsgeometrie der Bohrungen. In späteren Beiträgen analysierten verschiedene Autoren die Einflüsse der Turbulenz sowie der Stoffwerte von Kühlmittel und Hauptströmung. Dabei kamen meist Luft und seltener Kohlendioxid oder Stickstoff als Kühlmittel zum Einsatz. In den letzten Jahren wurde das Verhalten des Kühlmittelstrahls vor allem numerisch untersucht. Dabei beschränkte sich das Berechnungsgebiet oftmals auf das direkte Umfeld der Effusionskühlbohrung und die Identifikation und Beschreibung auftretender Wirbelstrukturen. Der Bereich weiter stromab der Bohrung blieb oft unberücksichtigt. Die vorliegende Arbeit verfolgt den Ansatz, den Kühlmittelstrahl in der Hauptströmung zu beobachten. Das wird durch die Verwendung von Helium und Argon als Kühlmittel möglich, denn diese Gase können in der Luftströmung detektiert werden. Durch eine in zwei Richtungen bewegliche Kombisonde wird Gas aus der Grenzschicht abgesaugt und die Konzentration des Kühlmittels bestimmt. Die so an diskreten Punkten stromab der Effusionskühlbohrung erhaltenen Konzentrations- und Geschwindigkeitsprofile ermöglichen die Verfolgung des Kühlmittelstrahls und dessen Wechselwirkungen mit der Hauptströmung. Für eine vergleichende Analyse der gemessenen Profile entstand ein empirisches Verfahren zur Systematisierung der gesamten Messdaten. Die Definition einer mittleren Kühlmittelkonzentration innerhalb einer zweckmäßig festgelegten Höhe über der Wand und eines normierten Einblasparameters, der das Verhältnis der molaren Massen von Kühlmittel und Hauptströmung berücksichtigt, sind der Kern des empirischen Verfahrens. Für Vergleiche mit der Literatur erfolgte die Berechnung eines Filmkühlwirkungsgrads auf Basis der Massebilanz in der Grenzschicht und der mittleren Kühlmittelkonzentration. Während der Datenauswertung zeigte sich, dass der Bohrungswinkel einen geringen Einfluss auf die mittlere Kühlmittelkonzentration hat und so ein Bohrungswinkel von 30° ein guter Kompromiss zwischen Herstellungsaufwand und Kühlwirkung ist. Kühlmedien mit geringer molarer Masse und hoher spezifischer Wärmekapazität sollten bevorzugt werden, da deren Kühlwirkung hoch, der Einfluss auf die Grenzschicht aber gering ist. / The cooling of thermally heavily loaded components is commonly performed by injecting a mass flow through the component’s wall into the hot flow, which is called Film cooling. The main goal is to form a coolant film to reduce the hot side heat transfer and to absorb thermal energy in order to protect the component’s wall. There are different techniques available called film cooling, transpiration cooling and effusion cooling. By applying transpiration cooling, the cooling fluid is injected through a porous material into the hot gas flow. Unfortunately, these porous materials do not have the physical strength required to work within gas turbines. If the injection is done with a row or a pattern of holes so the cooling film is renewed at certain positions, the cooling technique is called effusion cooling. Film cooling means the injection of fluid through a slot without renewing the film. Many authors analyze the effect of the film or effusion cooling on the wall temperature, the heat transfer coefficient or the cooling effectiveness. Many influencing factors were identified, such as the length to diameter relation, the hole’s alignment, fluid properties as well as turbulence and vortices. Recent works use numerical simulations to investigate the turbulent flow and vortex development in the near field of the injection hole. Due to the complexity of the simulation, the effects far downstream area were not covered by these simulations. This work investigates the behavior of the cooling jet within the boundary layer above the wall. Therefore a foreign gas (Helium, Argon) was injected as coolant into a cross flow and a pitot probe was used to get gas samples out of the boundary layer and the coolant gas fraction was measured. The measured concentration was empirically systematized by comparative data analysis. Therefore, a mean concentration within a certain height above the wall was calculated. Also a normed blowing rate was used to include the molar masses of coolant and cross flow. With this mean concentration a cooling effectiveness is calculated based on a balance model and compared to the results in the literature. As a result of the data evaluation, the hole’s angle was found to have a minor influence on the mean coolant concentration. An angle of about 30° is a good compromise between production effort and cooling efficiency. Also coolant fluids with a low molar mass and high specific heat capacity should be preferred because of their low impact on the boundary layer.

A body in a hypersonic flow field will experience very high heating especially during re-entry. Conventionally this problem is tackled to some extent by the use of large angle blunt cones. At the cost of increased drag, the heat transfer rate is lower over most parts of the blunt body, except in a region around the stagnation point. Thus even with blunt cones, management of heat transfer rates and drag on bodies at hypersonic speeds continues to be an interesting research area. Various thermal protection systems have been proposed in the past, like heat sink cooling, ablation cooling and aerospikes. The ablative cooling system becomes extremely costly when reusability is the major concern. Also the shape change due to ablation can lead to issues with the vehicle control. The aerospikes themselves may become hot and ablate at hypersonic speeds. Hence an alternate form of cooling system is necessary for hypersonic flows, which is more feasible, cost effective and efficient than the conventional cooling systems. Injection of a mass of cold fluid into the boundary layer through the surface is one of the potential cooling techniques in the hypersonic flight corridors. These kinds of thermal protection systems are called mass transfer cooling systems. The injection of the mass may be through discrete slots or through a porous media. When the coolant is injected through a porous media over the entire surface, the coolant comes out as a continuous mass. Such a cooling system is also referred as “transpiration cooling system”. When the fluid is injected through discrete slots, the system is called as “film cooling system”. In either case, the coolant absorbs the incoming heat through its rise in enthalpy and thus modifies the boundary layer characteristics in such a way that the heat flow rate to the surface is less. Injection of a forward facing jet (opposite to the freestream direction) from the stagnation point of a blunt body can be used for mitigating both the aerodynamic drag and heat transfer rates at hypersonic Mach numbers. If the jet has enough momentum it can push the bow shock forward, resulting in reduced drag. This will also reduce heat transfer rate over most part of the body except around the jet re-attachment region. A reattachment shock impinging on the blunt body invariably increases the local heat flux. At lower momentum fluxes the forward facing jet cannot push the bow shock ahead of the blunt body and spreads easily over the boundary layer, resulting in reduced heat transfer rates. While the film cooling performance improves with mass flow rate of the jet, higher momentum flow rates can lead to a stronger reattachment leading to higher heat transfer rate at the reattachment zone. If we are able to reduce the momentum flux of the coolant for the same mass flow rate, the gas coming out can easily spread over the boundary layer and it is possible to improve the film cooling performance. In all the reported literature, the mass flow rate and the momentum flux are not varied independently. This means, if the mass flow rate is increased, there is a corresponding increase in the momentum flux. This is because the injection (from a particular orifice and for a particular coolant gas) is controlled only by the total pressure of injection and free stream conditions. The present investigation is mainly aimed at demonstrating the effect of reduction in momentum of the coolant (injected opposing a hypersonic freestream from the stagnation point of a blunt cone), keeping the mass flow rate the same, on the film cooling performance. This is achieved by splitting a single jet into a number of smaller jets of same injection area (for same injection total pressure and same free stream conditions). To the best of our knowledge there is no report on the use of forward facing micro-jet array for film cooling at hypersonic Mach numbers. In this backdrop the main objectives of the present study are: • To experimentally demonstrate the effect of splitting a single jet into an array of closely spaced smaller micro-jets of same effective area of injection (injected opposite to a hypersonic freestream from the stagnation zone), on the reduction in surface heat transfer rates on a large angle blunt cone. · Identifying various parameters that affect the flow phenomenon and doing a systematic investigation of the effect of the different parameters on the surface heat transfer rates and drag. Experimental investigations are carried out in the IISc hypersonic shock tunnel on the film cooling effectiveness. Coolant gas (nitrogen and helium) is injected opposing hypersonic freestream as a single jet (diameter 2 mm and 0.9 mm), and as an array of iv micro jets (diameter 300 micron each) of same effective area (corresponding to the respective single jet). The coolant gas is injected from the stagnation zone of a blunt cone model (58o apex angle and nose radius of 35 mm). Experiments are performed at a flow freestream Mach number of 5.9 at 0o angle of attack, with a stagnation enthalpy of 1.84 MJ/Kg, with and without injections. The ratios of the jet stagnation pressure to the pitot pressure (stagnation pressure ratio) used in the present study are 1.2 and 1.45. Surface convective heat transfer measurements using platinum thin film sensors, time resolved schlieren flow visualization and aerodynamic drag measurements using accelerometer force balance are used as flow diagnostics in the present study. The theoretical stagnation point heat transfer rate without injection for the given freestream conditions for the test model is 79 W/cm2 and the corresponding aerodynamic drag from Newtonian theory is 143 N. The measured drag value without injection (125 N) shows a reasonable match with theory. As the injection is from stagnation zone it is not possible to measure the surface heat transfer rates at the stagnation point. The sensors thus are placed from the nearest possible location from the stagnation point (from 16 mm from stagnation point on the surface). The sensors near the stagnation point measures a heat transfer rate of 65 W/cm2 on an average without any injection. Some of the important conclusions from the study are: • Up to 40% reduction in surface heat transfer rate has been measured near the stagnation point with the array of micro jets, nitrogen being the coolant, while the corresponding reduction was up to 30% for helium injection. Considering the single jet injection, near the stagnation point there is either no reduction in heat transfer rate or a slight increase up to 10%. · Far away from stagnation point the reduction in heat transfer with array of micro-jets is only slightly higher than corresponding single jet for the same pressure ratio. Thus the cooling performance of the array of closely spaced micro jets is better than the corresponding single jet almost over the entire surface. • The time resolved flow visualization studies show no major change in the shock standoff distance with the low momentum gas injection, indicating no major changes in other aerodynamic aspects such as drag. · The drag measurements also indicate that there is virtually no change in the overall aerodynamic drag with gas injection from the micro-orifice array. · The spreading of the jets injected from the closely spaced micro-orifice array over the surface is also seen in the visualization, indicating the absence of a region of strong reattachment. · The reduction in momentum flux of the injected mass due to the interaction between individual jets in the case of closely spaced micro-jet array appears to be the main reason for better performance when compared to a single jet. The thesis is organized in six chapters. The importance of film cooling at hypersonic speeds and the objectives of the investigation are concisely presented in Chapter 1. From the knowledge of the flow field with counter-flow injection obtained from the literature, the important variables governing the flow phenomena are organized as non-dimensional parameters using dimensional analysis in Chapter 2. The description of the shock tunnel facility, diagnostics and the test model used in the present study is given in Chapter 3. Chapter 4 describes the results of drag measurements and flow visualization studies. The heat transfer measurements and the observed trends in heat transfer rates with and without coolant injection are then discussed in detail in Chapter 5. Based on the obtained results the possible physical picture of the flow field is discussed in Chapter 6, followed by the important conclusions of the investigation.

In the United States, natural gas turbine generators account for approximately 7% of the total primary energy consumed. A one percent increase in gas turbine efficiency could result in savings of approximately 30 million dollars for operators and, subsequently, electricity end-users. The efficiency of a gas turbine engine is tied directly to the temperature at which the products of combustion enter the first stage, high-pressure turbine. The maximum operating temperature of the turbine components’ materials is the major limiting factor in increasing the turbine inlet temperature. In fact, current turbine inlet temperatures regularly exceed the melting temperature of the turbine vanes through advanced vane cooling techniques. These cooling techniques include vane surface film cooling, internal vane cooling, and the addition of a thermal barrier coating (TBC) to the exterior of the turbine vane. Typically, the performance of vane cooling techniques is evaluated using the adiabatic film effectiveness. However, the adiabatic film effectiveness, by definition, does not consider conjugate heat transfer effects. In order to evaluate the performance of internal vane cooling and a TBC it is necessary to consider conjugate heat transfer effects. The goal of this study was to provide insight into the conjugate heat transfer behavior of actual turbine vanes and various vane cooling techniques through experimental and analytical modeling in the pursuit of higher turbine inlet temperatures resulting in higher overall turbine efficiencies. The primary focus of this study was to experimentally characterize the combined effects of a TBC and film cooling. Vane model experiments were performed using a 10x scaled first stage inlet guide vane model that was designed using the Matched Biot Method to properly scale both the geometrical and thermal properties of an actual turbine vane. Two different TBC thicknesses were evaluated in this study. Along with the TBCs, six different film cooling configurations were evaluated which included pressure side round holes with a showerhead, round holes only, craters, a novel trench design called the modified trench, an ideal trench, and a realistic trench that takes manufacturing abilities into account. These film cooling geometries were created within the TBC layer. Each of the vane configurations was evaluated by monitoring a variety of temperatures, including the temperature of the exterior vane wall and the exterior surface of the TBC. This study found that the presence of a TBC decreased the sensitivity of the thermal barrier coating and vane wall interface temperature to changes in film coolant flow rates and changes in film cooling geometry. Therefore, research into improved film cooling geometries may not be valuable when a TBC is incorporated. This study also developed an analytical model which was used to predict the performance of the TBCs as a design tool. The analytical prediction model provided reasonable agreement with experimental data when using baseline data from an experiment with another TBC. However, the analytical prediction model performed poorly when predicting a TBC’s performance using baseline data collected from an experiment without a TBC. / text

Developments regarding internal cooling techniques have allowed the operation of modern gas turbine engines at turbine inlet temperatures which exceed the metallurgical capability of the turbine blade. This has allowed the operation of engines at a higher thermal efficiency and lower specific fuel consumption. Modern turbine blade-cooling techniques rely on external film cooling to protect the outer surface of the blade from the hot gas path and internal cooling to remove thermal energy from the blade. Optimization of coolant performance and blade-life estimation require knowledge regarding the influence of cooling application on the blade inner and outer surface heat transfer. The following study describes a combined experimental and computational study of heat transfer augmentation near the entrance to a film-cooling hole. Steady-state heat transfer results were acquired by using a transient measurement technique in an 80 x actual rectangular channel, representing an internal cooling channel of a turbine blade. Platinum thin-film gauges were used to measure the inner surface heat transfer augmentation as a result of thermal boundary layer renewal and impingement near the entrance of a film-cooling hole. Measurements were taken at various suction ratios, extraction angles and wall temperature ratios with a main duct Reynolds number of 25×103. A numerical technique, based on the resolution of the unsteady conduction equation, using a Crank-Nicholson scheme, was used to obtain the surface heat flux from the measured surface temperature history. Computational data was generated with the use of a commercial CFD solver. / Dissertation (MEng)--University of Pretoria, 2008. / Mechanical and Aeronautical Engineering / unrestricted

Cette recherche s’articule en deux parties. L’objectif de la première partie est d’analyser par voie expérimentale et numérique la phénoménologie du décollement interne, dit décollement de jet (en regimes transitoire et établi) dans les tuyères supersoniques refroidies par film fluide. La deuxième partie porte sur la réinterprétation des concepts de tuyère existants pour aboutir à la proposition d’un nouveau dispositif de détente supersonique, qui offre une résistance accrue au décollement de jet. La première partie de cette thèse est basée sur l’analyse des résultats expérimentaux obtenus lors de la campagne d’essais réalisée à l’ONERA. Ces essais, ont mis en évidence des spécificités de comportement de la tuyère, inhérentes à la manière d’amorcer le jet supersonique principal par rapport à l’établissement du film pariétal. Ces mêmes expériences ont permis d’étudier le comportement instationnaire du décollement de jet lorsque les conditions d’alimentation sont maintenues en régime établi. L’apparition de fréquences caractéristiques a été mise en évidence et leur origine a été étudiée à l’aide de simulations numériques. En nous appuyant sur les considérations issues de la première partie de l’étude, une revue critique des concepts de tuyère existants a été menée. Ce travail a permis d’identifier une lacune majeure dans la définition des tuyères à écoulement interne, à savoir l’absence d’une « barrière » qui puisse prévenir l’occurrence du décollement de jet. Ainsi, nous avons proposé la conjonction d’un dispositif à écoulement externe (aerospike) et d’une tuyère classique afin de résoudre cette problématique in nuce, en créant une barrière fluidique continue tout autour du plan de sortie de la tuyère principale. L’efficacité de ce concept a donc été prouvée par calcul, puis une campagne expérimentale a été organisée afin de valider les résultats obtenus. / This research is in two parts. The objective of the first part is to analyse by experimental and numerical means the phenomenology of nozzle flow separation in transient and steady state conditions. The second part of this research work focuses on the reinterpretation of existing concepts of converging-diverging nozzles, leading to the proposal of a new supersonic expansion device, with improved flow separation characteristics.Experimental data, collected during the test campaign conducted at ONERA, have been analysed and are presented in the first part of this thesis. Obtained results highlight some peculiarities of the transient behavior of the nozzle, mostly dependent on the synchronisation between the start-up phase of the main jet and the grow-up of the wall film. These same experiments have been also used to investigate the unsteadiness of the flow separation, when nozzle feeding conditions are maintained constant. Appearance of characteristic frequencies has been highlighted and their origin has been investigated by CFD simulations.In the second part, a critical review of existing nozzle concepts was conducted. This allowed identifying a major gap in the definition of traditional supersonic nozzles, namely the absence of a "barrier" that can prevent the occurrence of the flow separation. Thus, in the second part of this thesis we propose a new nozzle concept. It is based on the combination of a small aerospike and a conventional nozzle (main flow). Such an arrangement allows solving the flow separation problem in nuce. The effectiveness of this concept has been proved by calculation and by an experimental test campaign.

Die Kühlung thermisch hoch belasteter Bauteile wird häufig mit Kühlverfahren realisiert, die auf dem Prinzip des Massetransports durch die Bauteilwand beruhen. Neben der Film- und Transpirationskühlung gehört die Effusionskühlung zu diesen Verfahren und basiert auf einer Reihe oder einem Raster von Bohrungen. Dieser Ansatz ermöglicht sowohl den Abtransport von Wärme aus dem Bauteil als auch die Ausbildung eines schützenden Kühlmittelfilms auf der Bauteiloberfläche. Viele Autoren beschäftigten sich in ihren Arbeiten mit den Auswirkungen der Filmkühlung auf den Wärmeübergang an der Bauteilwand und definierten einen Filmkühlwirkungsgrad, der die Effektivität der Kühlung widerspiegelt. Auch die Freistrahlen aus Effusionskühlbohrungen wurden mit diesen Mitteln untersucht und eine Vielzahl unterschiedlicher Einflussgrößen auf den Filmkühlwirkungsgrad identifiziert. Dazu gehören insbesondere geometrische Bedingungen, wie z.B. der Bohrungswinkel, das Verhältnis von Länge zu Durchmesser der Bohrung und die Austrittsgeometrie der Bohrungen. In späteren Beiträgen analysierten verschiedene Autoren die Einflüsse der Turbulenz sowie der Stoffwerte von Kühlmittel und Hauptströmung. Dabei kamen meist Luft und seltener Kohlendioxid oder Stickstoff als Kühlmittel zum Einsatz. In den letzten Jahren wurde das Verhalten des Kühlmittelstrahls vor allem numerisch untersucht. Dabei beschränkte sich das Berechnungsgebiet oftmals auf das direkte Umfeld der Effusionskühlbohrung und die Identifikation und Beschreibung auftretender Wirbelstrukturen. Der Bereich weiter stromab der Bohrung blieb oft unberücksichtigt. Die vorliegende Arbeit verfolgt den Ansatz, den Kühlmittelstrahl in der Hauptströmung zu beobachten. Das wird durch die Verwendung von Helium und Argon als Kühlmittel möglich, denn diese Gase können in der Luftströmung detektiert werden. Durch eine in zwei Richtungen bewegliche Kombisonde wird Gas aus der Grenzschicht abgesaugt und die Konzentration des Kühlmittels bestimmt. Die so an diskreten Punkten stromab der Effusionskühlbohrung erhaltenen Konzentrations- und Geschwindigkeitsprofile ermöglichen die Verfolgung des Kühlmittelstrahls und dessen Wechselwirkungen mit der Hauptströmung. Für eine vergleichende Analyse der gemessenen Profile entstand ein empirisches Verfahren zur Systematisierung der gesamten Messdaten. Die Definition einer mittleren Kühlmittelkonzentration innerhalb einer zweckmäßig festgelegten Höhe über der Wand und eines normierten Einblasparameters, der das Verhältnis der molaren Massen von Kühlmittel und Hauptströmung berücksichtigt, sind der Kern des empirischen Verfahrens. Für Vergleiche mit der Literatur erfolgte die Berechnung eines Filmkühlwirkungsgrads auf Basis der Massebilanz in der Grenzschicht und der mittleren Kühlmittelkonzentration. Während der Datenauswertung zeigte sich, dass der Bohrungswinkel einen geringen Einfluss auf die mittlere Kühlmittelkonzentration hat und so ein Bohrungswinkel von 30° ein guter Kompromiss zwischen Herstellungsaufwand und Kühlwirkung ist. Kühlmedien mit geringer molarer Masse und hoher spezifischer Wärmekapazität sollten bevorzugt werden, da deren Kühlwirkung hoch, der Einfluss auf die Grenzschicht aber gering ist. / The cooling of thermally heavily loaded components is commonly performed by injecting a mass flow through the component’s wall into the hot flow, which is called Film cooling. The main goal is to form a coolant film to reduce the hot side heat transfer and to absorb thermal energy in order to protect the component’s wall. There are different techniques available called film cooling, transpiration cooling and effusion cooling. By applying transpiration cooling, the cooling fluid is injected through a porous material into the hot gas flow. Unfortunately, these porous materials do not have the physical strength required to work within gas turbines. If the injection is done with a row or a pattern of holes so the cooling film is renewed at certain positions, the cooling technique is called effusion cooling. Film cooling means the injection of fluid through a slot without renewing the film. Many authors analyze the effect of the film or effusion cooling on the wall temperature, the heat transfer coefficient or the cooling effectiveness. Many influencing factors were identified, such as the length to diameter relation, the hole’s alignment, fluid properties as well as turbulence and vortices. Recent works use numerical simulations to investigate the turbulent flow and vortex development in the near field of the injection hole. Due to the complexity of the simulation, the effects far downstream area were not covered by these simulations. This work investigates the behavior of the cooling jet within the boundary layer above the wall. Therefore a foreign gas (Helium, Argon) was injected as coolant into a cross flow and a pitot probe was used to get gas samples out of the boundary layer and the coolant gas fraction was measured. The measured concentration was empirically systematized by comparative data analysis. Therefore, a mean concentration within a certain height above the wall was calculated. Also a normed blowing rate was used to include the molar masses of coolant and cross flow. With this mean concentration a cooling effectiveness is calculated based on a balance model and compared to the results in the literature. As a result of the data evaluation, the hole’s angle was found to have a minor influence on the mean coolant concentration. An angle of about 30° is a good compromise between production effort and cooling efficiency. Also coolant fluids with a low molar mass and high specific heat capacity should be preferred because of their low impact on the boundary layer.

High fidelity measurements are necessary to validate existing and future turbulence models for the purpose of producing the next generation of more efficient gas turbines. The objective of the present study is to conduct several different measurements of multi-row film cooling arrays in order to better understand the physics involved with injection of coolant through multiple rows of discrete holes into a flat plate turbulent boundary layer. Adiabatic effectiveness distributions are measured for several multi-row film cooling geometries. The geometries are designed with two different hole spacings and two different hole types to yield four total geometries. One of the four geometries tested for adiabatic effectiveness was selected for flowfield measurements. The wall and flowfield are studied with several testing techniques, including: particle image velocimetry, hot wire anemometry, pressure sensitive paint and discrete gas sampling.

No description available.

No description available.

This thesis presents a joint computational and experimental investigation into the heat transfer to unshrouded turbine blade tips suitable for use in high bypass ratio, large civil aviation turbofan engines. Both the heat transfer to the blade tip and the over-tip leakage flow over the blade tip are characterised, as each has a profound influence on overall engine efficiency. The study is divided into two sections; in the first, computational simulations of a very large scale, low speed linear cascade with a flat blade tip were conducted. These simulations were validated against experimental data collected by Palafox (2006). A thorough assessment of turbulence models and minimum meshing requirements was performed. The standard k-ω and standard k-ϵ turbulence models significantly overpredicted the turbulence levels within the tip gap. The other models were very similar in performance; the SST k-ω and realisable k-ϵ models were found to be the most suitable for the flow environment. The second section documents the development and testing of a novel hybrid blade tip design, the squealet tip, which seeks to combine the known benefits of winglet and double squealer tips. The development of the external geometry was performed primarily through engine-representative CFD simulations at a range of tip gaps from 0.45% to 1.34% blade chord. The squealet tip was found to have a similar aerodynamic sensitivity to tip clearance as a baseline double squealer tip, with a tip gap efficiency exchange rate of 2.03, although this was 18% greater than the alternative winglet tip. The squealet tip displayed higher predicted stage efficiency than the winglet tip over the majority of the range of tip clearances investigated, however. The overall heat load was reduced by 14% compared with the winglet tip but increased by 28% over the double squealer tip, primarily due to the change in wetted surface area. The predicted local heat transfer coefficients were similar across all geometries. A realistic internal cooling plenum and an array of blade tip cooling holes were subsequently added to the squealet tip geometry and the cooling configuration refined by the selective sealing of cooling holes. Film cooling performance was largely assessed by the predicted adiabatic wall temperature distributions. A viable cooling scheme which reduced the cooling air requirement by 38% was achieved, compared to the initial case which had all cooling holes open. This was associated with just a 7% increase in blade tip heat flux and no penalty in peak temperature on the blade tip. Film cooling air ejected from holes on the blade suction side was swept away from the blade tip region, making the squealet rim at the crown of the blade particularly challenging to cool. It was demonstrated that this region could be cooled effectively by ballistic cooling from holes located on the blade tip cavity floor, although this was expensive in terms of the mass flow rate of cooling air required. The computational results were reinforced with experimental data collected in a transonic linear cascade. Downstream aerodynamic loss measurements were taken for a linearised version of the squealet tip design without cooling at nominal tip gaps of 0.45%, 0.89% and 1.34% blade chord, which was compared to similar data taken by O’Dowd (2010) for flat and winglet tips. The squealet was seen to have a similar aerodynamic loss to the flat tip and a reduced loss compared with the winglet tip. Full surface heat transfer measurements were taken for the uncooled squealet tip, at tip gaps of 0.89% and 1.34% blade chord, and for two configurations of the cooled squealet tip, at a tip clearance of 0.89% blade chord. The qualitative similarity between the measured heat transfer distributions and the those predicted by the engine-representative CFD simulations was good. A CFD simulation of the uncooled linear cascade environment at the 1.34% blade chord tip clearance was performed using a single blade with translationally periodic boundary conditions. The predicted size of the over-tip leakage vortex was smaller than had been measured, resulting in a large underprediction in the magnitude of the downstream area-averaged aerodynamic loss. The magnitudes of the predicted blade tip Nusselt number distribution were similar to those produced by the engine-representative CFD simulations and lower than that measured experimentally. Differences in the shape of the Nusselt number distribution were observed in the vicinity of regions of separated and reattaching flow, but other salient features were replicated in the computational data. The squealet tip has been shown to be a promising, viable unshrouded blade tip design with an aerodynamic performance similar to the double squealer tip but is more amenable to film cooling. It is significantly lighter than a winglet tip and incurs a reduced thermal load. The squealet tip design can now be developed into a blade tip geometry for use in real engines to provide an alternative to shrouded turbine blades and current unshrouded blade tip designs. A commercial CFD solver, Fluent 6.3, was shown to capture blade tip heat transfer and over-tip leakage flow sufficiently well to be a useful design guide. However, the sensitivity of the flow structure (and hence, heat transfer) in the forward part of the blade tip cavity suggests that physical testing cannot be eliminated from the design process entirely.