Endwall contouring has been widely studied as an effective measure to improve aerodynamic performance by reducing secondary flow strength. The effects of endwall contouring with discrete holes and an upstream purge slot for a high turning (127") airfoil passage under transonic conditions are investigated. The total pressure loss and secondary flow field were measured for two endwall geometries. The non-axisymmetric endwall was developed through an optimization study [1] to minimize secondary losses and is compared to a baseline planar endwall. The blade inlet span increased by 13 degrees with respect to the inlet in order to match engine representative inlet/exit Mach number loading in a HP turbine.  The experiments were performed in a quasi-2D linear cascade with measurements at design exit Mach number 0.88 and incidence angle. Four cases were analyzed for each endwall -- the effect of slot presence (with/without coolant) and the effect of discrete holes (with/without coolant) without slot injection. The coolant to mainstream mass flow ratio was set at 1.0% and 0.25% for upstream purge slot and discrete holes, respectively.  Aerodynamic loss coefficient is calculated with the measured exit total pressure at 0.1 Cax downstream of the blade trailing edge. CFD studies were conducted in compliment. The aero-optimized endwall yielded lower losses than baseline without the presence of the slot. However, in presence of the slot, losses increased due to formation of additional vortices. For both endwall geometries, results reveal that the slot has increased losses, while the addition of coolant further influences secondary flow development. / Master of Science

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Endwall film cooling is a necessity in modern gas turbines for safe and reliable operation. Performance of endwall film cooling is strongly influenced by the hot gas flow field, among other factors. For example, aerodynamic design determines secondary flow vortices such as passage vortices and corner vortices in the endwall region. Moreover blockage presented by the leading edge of the airfoil subjects the incoming flow to a stagnating pressure gradient leading to roll-up of the approaching boundary layer and horseshoe vortices. In addition, for a number of heavy frame power generation gas turbines that use cannular combustors, the hot and turbulent gases exiting from the combustor are delivered to the first stage vane through transition ducts. Wakes induced by walls separating adjacent transition ducts located upstream of first row vanes also influence the entering main gas flow field. Furthermore, as hot gas enters vane passages, it accelerates around the vane airfoils. This flow acceleration causes significant streamline curvature and impacts lateral spreading endwall coolant films. Thus endwall flow field, especially those in utility gas turbines with cannular combustors, is quite complicated in the presence of vortices, wakes and strong favorable pressure gradient with resulting flow acceleration. These flow features can seriously impact film cooling performance and make difficult the prediction of film cooling in endwall. This study investigates endwall film cooling under the influence of pressure gradient effects due to stagnation region of an axisymmetric airfoil and in mainstream favorable pressure gradient. It also investigates the impact of wake on endwall film cooling near the stagnation region of an airfoil. The investigation consists of experimental testing and numerical simulation. Endwall film cooling effectiveness is investigated near the stagnation region on an airfoil by placing an axisymmetric airfoil downstream of a single row of inclined cylindrical holes. The holes are inclined at 35° with a length-to-diameter ratio of 7.5 and pitch-to-diameter ratio of 3. The ratio of leading edge radius to hole diameter and the ratio of maximum airfoil thickness to hole diameter are 6 and 20 respectively. The distance of the leading edge of the airfoil is varied along the streamwise direction to simulate the different film cooling rows preceding the leading edge of the airfoil. Wake effects are induced by placing a rectangular plate upstream of the injection point where the ratio of plate thickness to hole diameter is 6.4, and its distance is also varied to investigate the impact of strong and mild wake on endwall film cooling effectiveness. Blowing ratio ranged from 0.5 to 1.5. Film cooling effectiveness is also investigated under the presence of mainstream pressure gradient with converging main flow streamlines. The streamwise pressure distribution is attained by placing side inserts into the mainstream. The results are presented for five holes of staggered inclined cylindrical holes. The inclination angle is 30° and the tests were conducted at two Reynolds number, 5000 and 8000. Numerical analysis is employed to aid the understanding of the mainstream and coolant flow interaction. The solution of the computational domain is performed using FLUENT software package from Fluent, Inc. The use of second order schemes were used in this study to provide the highest accuracy available. This study employed the Realizable º-µ model with enhance wall treatment for all its cases. Endwall temperature distribution is measured using Temperature Sensitive Paint (TSP) technique and film cooling effectiveness is calculated from the measurements and compared against numerical predictions. Results show that the characteristics of average film effectiveness near the stagnation region do not change drastically. However, as the blowing ratio is increased jet to jet interaction is enhanced due to higher jet spreading resulting in higher jet coverage. In the presence of wake, mixing of the jet with the mainstream is enhanced particularly for low M. The velocity deficit created by the wake forms a pair of vortices offset from the wake centerline. These vortices lift the jet off the wall promoting the interaction of the jet with the mainstream resulting in a lower effectiveness. The jet interaction with the mainstream causes the jet to lose its cooling capabilities more rapidly which leads to a more sudden decay in film effectiveness. When film is discharged into accelerating main flow with converging streamlines, row-to-row coolant flow rate is not uniform leading to varying blowing ratios and cooling performance. Jet to jet interaction is reduced and jet lift off is observed for rows with high blowing ratio resulting in lower effectiveness.

The interaction between rotor blade tip leakage vortex and inflow disturbances, such as encountered in shrouded marine propulsors, was simulated in the Virginia Tech Linear Cascade Wind Tunnel equipped with a moving endwall system. Upstream of the blade row, idealized periodic inflow unsteadiness was generated using vortex generator pairs attached to the endwall at the same spacing as the blade spacing. At three tip gap settings, 1.7%c, 3.3%c and 5.7%c, the flow near the lower endwall of the center blade passage was investigated through three-component mean velocity and turbulence distributions measured by four-sensor hotwires. Besides time-averaged data, the measurements were processed for phase-locked analysis, with respect to pitchwise locations of the vortex generators relative to the blade passage. Moreover, surface pressure distributions at the blade tip were acquired at eight tip gaps from 0.87%c to 12.9%c. Measurements of pressure-velocity correlation were also performed with wall motion but without inflow disturbances. Achieved in this study is an understanding of the characteristics and structures of the tip leakage vortex at its initial formation. The mechanism of the tip leakage vortex formation seems to be independent of the tip gap setting. The tip leakage vortex consists of a vortical structure and a region of low streamwise-momentum fluid next to the endwall. The vortical structure is initially attached to the blade tip that creates it. This structure picks up circulation shed from that blade tip, as well as those from the endwall boundary layer, and becomes stronger with downstream distance. Partially induced by the mirror images in the endwall, the vortical structure starts to move across the passage resulting in a reduction in its rotational strength as the cross sectional area of the vortex increases but little circulation is added. The larger the tip gap, the longer the vortical structure stays attached to the blade tip, and the stronger the structure when it reaches downstream of the passage. Phased-averaged data show that the inflow disturbances cause small-scale responses and large-scale responses upstream and downstream of the vortex shedding location, respectively. This difference in scale is possibly dictated by a variation in the shedding location since the amount of circulation in the vortex is dependent on this location. The inflow disturbances possibly cause a variation in the shedding location by manipulating the separation of the tip leakage flow from the endwall and consequently the flow's roll-up process. Even though this manipulation only perturbs the leakage flow in a small scale, the shedding mechanism of the tip leakage vortex amplifies the outcome. / Ph. D.

In the past few decades, the operating temperatures of gas turbine engines have increased significantly with a view towards increasing the overall thermal efficiency and specific power output. As a result of increased turbine inlet temperatures, the hot gas path components downstream of the combustor section are subjected to high heat loads. Though materials with improved temperature capabilities are used in the construction of the hot gas path components, in order to ensure safe and durable operation, the hot gas path components are additionally supplemented with thermal barrier coatings (TBCs) and sophisticated cooling techniques. The present study focusses on two aspects of gas turbine cooling, namely augmented internal cooling and external film cooling. One of the commonly used methods for cooling the vanes involves passing coolant air bled from the compressor through serpentine passages inside the airfoils. The walls of the internal cooling passages are usually roughened with turbulence promoters like ribs to enhance heat transfer. Though the ribs help in augmenting the heat transfer, they have an associated pressure penalty as well. Therefore, it is important to study the thermal-hydraulic performance of ribbed internal cooling passages. The first section of the thesis deals with the numerical investigation of flow and heat transfer characteristics in a ribbed two-pass channel. Four different rib shapes- 45° angled, V-shaped, W-shaped and M-shaped, were studied. This study further aims at exploring the performance of different rib-shapes at a large rib pitch-to-height ratio (p/e=16) which has potential applications in land-based gas turbines operating at high Reynolds numbers. Detailed flow and heat transfer analysis have been presented to illustrate how the innate flow physics associated with the bend region and the different rib shapes contribute to heat transfer enhancement in the two-pass channel. The bend-induced secondary flows were observed to significantly affect the flow and heat transfer distribution in the 2nd pass. The thermal-hydraulic performance of V-shaped and 45° angled ribs were better than W-shaped and M-shaped ribs. The second section of the study deals with the analysis of film cooling performance of different hole configurations on the endwall upstream of a first stage nozzle guide vane. The flow along the endwall of the airfoils is highly complex, dominated by 3-dimensional secondary flows. The presence of complex secondary flows makes the cooling of the airfoil endwalls challenging. These secondary flows strongly influence endwall film cooling and the associated heat transfer. In this study, three different cooling configurations- slot, cylindrical holes and tripod holes were studied. Steady-state experiments were conducted in a low speed, linear cascade wind tunnel. The adiabatic film cooling effectiveness on the endwall was computed based on the spatially resolved temperature data obtained from the infrared camera. The effect of mass flow ratio on the film cooling performance of the different configurations was also explored. For all the configurations, the coolant jets were unable to overcome the strong secondary flows inside the passage at low mass flow ratios. However, the coolant jets were observed to provide much better film coverage at higher mass flow ratios. In case of cylindrical ejection, the effectiveness values were observed to be very low which could be because of jet lift-off. The effectiveness of tripod ejection was comparable to slot ejection at mass flow ratios between 0.5-1.5, while at higher mass flow ratios, slot ejection was observed to outperform tripod ejection. / Master of Science

The growing need for increased performance from gas turbines has fueled the drive to raise turbine inlet temperatures. This results in high thermal stresses especially along the first stage nozzle guide vane cascade as the hot combustion products exiting modern day gas turbine combustors generally reach temperatures that could endanger the structural stability of these vanes and greatly reduce the vane life. The highest heat transfer coefficients in the vane passage occurs near the endwall, particularly in the leading edge-endwall junction where vortical flows cause the flow of hotter fluid in the mainstream to mix with relatively lower temperature boundary layer fluid. This work documents the computational investigation of air injection at the end wall through a cylindrical hole placed upstream of the nozzle guide vane leading edge-end wall junction. The effect of the secondary jet on the formation of the leading edge horseshoe vortex and the consequent formation of the passage vortex has been studied. For the computations, the Reynolds averaged Navier–Stokes (RANS) equations were solved with the commercial software ANSYS Fluent using the SST k-ω model. Total pressure loss coefficient and kinetic energy loss Coefficient contour plots at the exit of the cascade to estimate the effect of the endwall fluid injection on loss profiles at the vane cascade exit. Swirling strength contours were plotted at several axial chord locations in order to track the path of the passage vortex in and downstream of the vane cascade. Two different hole-positions (located at 1 hole diameter and 2 hole diameters from the leading edge) along a plane parallel to the incident flow were considered in order to study the effect of the hole position with respect to the vane leading edge-endwall junction. Three different streamwise hole inclination angles with respect to the mainstream flow direction were studied to identify the best angle for the injection of fluid through the endwall. This angle was combined with five different compound angles (0°, 30°, 45°, 60° and 90°) in order to study the effect of varying the compound angle on the leading edge vortex and the passage vortex. Each of the above studies were conducted at two different injected fluid-to-mainstream mass flow ratios (0.5% and 1%) in order to study the effect of varying injected flow rate on the formation of the leading edge vortex and the vane passage vortex. From the results it was observed that suitable selection of the secondary injection mass flow rate, injection angle and hole-position caused an absence of the leading edge horseshoe vortex and delayed migration of the passage vortex across the guide vane passage. Heat Transfer studies were also conducted to observe the absence/weakening of the leading edge vortex and the delayed pitch-wise movement of the passage vortex. / Master of Science

The present study investigates mixed out aerodynamic loss coefficient measurements for a high turning, contoured endwall passage under transonic operating conditions in presence of upstream purge slot and mateface gap. The upstream purge slot represents the gap between stator-rotor interface and the mateface gap simulates the assembly feature between adjacent airfoils in an actual high pressure turbine stage. While the performance of the mateface and upstream slot has been studied for lower Mach number, no studies exist in literature for transonic flow conditions. Experiments were performed at the Virginia Tech's linear, transonic blow down cascade facility. Measurements were carried out at design conditions (isentropic exit Mach number of 0.87, design incidence) without and with coolant blowing. Upstream leakage flow of 1.0% coolant to mainstream mass flow ratio (MFR) was considered with the presence of mateface gap. There was no coolant blowing through the mateface gap itself. Cascade exit pressure measurements were carried out using a 5-hole probe traverse at a plane 1.0Cax downstream of the trailing edge for a planar geometry and two contoured endwalls. Spanwise measurements were performed to complete the entire 2D loss plane from endwall to midspan, which were used to plot pitchwise averaged losses for different span locations and loss contours for the passage. Results reveal significant reduction in aerodynamic losses using the contoured endwalls due to the modification of flow physics compared to a non contoured planar endwall. / Master of Science

Η θερμοδυναμική ανάλυση του κύκλου Brayton υποδεικνύει ότι η θερμική απόδοση και το ειδικό έργο εξόδου ενός αεριοστρόβιλου μπορούν να βελτιωθούν με την αύξηση της θερμοκρασίας εισόδου των αεριών της καύσης στον στρόβιλο. Επιπλέον, οι αυξημένες θερμοκρασίες εισόδου στον στρόβιλο συνοδεύονται και από μείωση της κατανάλωσης καυσίμου, ενώ σε αεροπορικές εφαρμογές οι υψηλότερες θερμοκρασίες έχουν ώς αποτέλεσμα την αύξηση της ώσης του κινητήρα. Δυστυχώς όμως, οι υψηλές αυτές θερμοκρασίες θέτουν σε κίνδυνο την ακεραιότητα των εξαρτημάτων του στροβίλου υψηλής πίεσης και ειδικότερα τα πτερύγια (blades) του στροβίλου και το δάπεδο (endwall) στο οποίο τα πτερύγια αυτά είναι προσκολλημένα. Στους μοντέρνους κινητήρες, η θερμοκρασία εισόδου στον στρόβιλο μπορεί να φτάνει και στα επίπεδα των 1900Κ, θερμοκρασία η οποία υπερβαίνει το σημείο τήξης των υλικών από τα οποία είναι κατασκευασμένα τα εξαρτήματα του στροβίλου. Αυτό έχει ως αποτέλεσμα τα εξαρτήματα του στροβίλου να λειτουργούν σε πολύ σκληρότερο περιβάλλον απ' ότι στο παρελθόν. Η διατήρηση επαρκούς διάρκειας ζωής στις υψηλές αυτές θερμοκρασίες απαιτεί την ανάπτυξη νέων υλικών κατασκευής και αποτελεσματικών μεθόδων ψύξης για τα εξαρτήματα του στροβίλου. Για την αντιμετώπιση και την αποφυγή της αστοχίας των πτερυγίων (blades) και των δάπεδων (endwall) των πτερυγικών διακένων στους στροβίλους, η μέθοδος του "film cooling" έχει ενσωματωθεί στον σχεδιασμό τους. Κατά την διεργασία της ψύξης των εξαρτημάτων με την μέθοδο αυτή, ψυχρός αέρας αφαιμάσσεται από το στάδιο του συμπιεστή, διοχετεύεται μέσω εσωτερικών θαλάμων του κινητήρα στα εξαρτήματα του στροβίλου και εγχέεται μέσω διακριτών οπών στα τοιχώματα των πτερυγίων και των δαπεδικών τοιχωμάτων. Μετά την έξοδο του από τις οπές, ο ψυκτικός αέρας σχηματίζει ένα λεπτό, προστατευτικό στρώμα-φιλμ μεταξύ των θερμών αερίων της καύσης και της μεταλλικής επιφάνειας των εξαρτημάτων. Μια εκ των κρίσιμων περιοχών οι οποίες υποβάλλονται σε αυξημένους ρυθμούς μετάδοσης θερμότητας είναι και η περιοχή γύρω από την περιφέρεια σύνδεσης των πτερυγίων (blades) με τα δάπεδα (endwalls) του στροβίλου. Η περιοχή αυτή κυριαρχείται από την παρουσία ισχυρών τρισδιάστατων δευτερογενών ροών (γνωστές και ώς junction flows) οι οποίες προκαλούν αύξηση των τοπικών ρυθμών μετάδοσης θερμότητας στην περιοχή της τάξης του 350%. Επιπλέον, οι ροές αυτές, εμποδίζουν την διείσδυση ψυκτικού ρευστού στην προβληματική περιοχή εκτοπίζοντας το μακριά από την επιφάνεια του δαπέδου πριν αυτό προλάβει να παράσχει ικανοποιητική ψύξη. Αντικείμενο της παρούσας διδακτορικής διατριβής, είναι η μελέτη, ανάπτυξη και δοκιμή (τόσο πειραματικά όσο και υπολογιστικά) μιας πρωτότυπης γεωμετρίας ψύξης (με την μέθοδο του film cooling), για την αποτελεσματική αντιμετώπιση του προβλήματος της υπερθέρμανσης της περιοχής σύνδεσης του πτερυγίου – δαπέδου κυρίως γύρω από το επίπεδο του χείλους προσβολής. Το κύριο χαρακτηριστικό της πρωτότυπης μεθόδου έγχυσης είναι ότι το ψυκτικό εκχέεται κατά τέτοιο τρόπο έτσι ώστε οι ροϊκές γραμμές του ψυκτικού να υποβοηθούνται από την περιστροφική κίνηση των τοπικών τρισδιάστατων ροών. Η πολυπλοκότητα του προβλήματος ψύξης της συγκεκριμένης περιοχής προκύπτει από δύο στοιχεία. Πρώτον, όπως αναφέρθηκε και προηγουμένως, η ροή στην περιοχή σύνδεσης κοντά στο τοίχωμα χαρακτηρίζεται από πολύπλοκη τρισδιάστατη δομή. Δεύτερον, το πρόβλημα χαρακτηρίζεται από τρείς θερμοκρασίες: την θερμοκρασία της κύριας ροής, την θερμοκρασία του τοιχώματος και την θερμοκρασία του ψυκτικού αέρα. Για την πλήρη διερεύνηση των χαρακτηριστικών της προτεινόμενης μεθόδου ψύξης η εργασία περιλαμβάνει τόσο πειραματικό όσο και υπολογιστικό σκέλος: Υπολογιστικό Σκέλος (Computational part): Ο επιτυχής σχεδιασμός μιας πιθανής γεωμετρίας ψύξης για την συγκεκριμένη περιοχή του δαπέδου (endwall) απαιτεί την γνώση και κατανόηση της τοπικής ροής μέσα στην οποία το τζετ του ψυκτικού πρόκειται να εισέλθει. Επιπλέον, είναι σημαντική η κατανόηση της αλληλεπίδρασης που αναμένεται μεταξύ του ψυκτικού αέρα με την τοπική τρισδιάστατη ροή. Για τον σκοπό αυτό, χρησιμοποιήθηκε η μέθοδος της υπολογιστικής ρευστοδυναμικής (Computational Fluid Dynamics) για την πρόβλεψη του σχετικού τρισδιάστατου βασικού πεδίου ροής στην περιοχή σύνδεσης του πτερυγίου (blade) - δαπέδου (endwall). Έγιναν προσομοιώσεις τόσο για την βασική γεωμετρία απουσία έγχυσης (οι οποίες χρησιμοποιήθηκαν ως πεδίο αναφοράς) όσο και προσομοιώσεις παρουσία της πρωτότυπης έγχυσης οι οποίες αφορούσαν την επίδραση διαφόρων παραμέτρων στην αποτελεσματικότητα της ψύξης της προβληματικής περιοχής. Στις προσομοιώσεις υιοθετήθηκε η εξής θερμοκρασιακή κατανομή: Θερμό δάπεδο (endwall) - Θερμότερη κύρια ροή (mainstream) - Ψυχρός αέρας έγχυσης, η οποία είναι και αντίστοιχη με αυτήν που εμφανίζεται σε πραγματικές εφαρμογές. Τα αποτελέσματα των προσομοιώσεων βοήθησαν στην κατανόηση του ροϊκού πεδίου στην περιοχή σύνδεσης τόσο ποιοτικά όσο και ποσοτικά σε ότι αφορά τα σχετικά μεγέθη των ροϊκών δομών και των αεροδυναμικών χαρακτηριστικών τις περιοχής. Αυτό είχε ώς αποτέλεσμα τον αποτελεσματικό σχεδιασμό της πρωτότυπης γεωμετρίας έγχυσης. Επιπλέον, οι υπολογιστικές προβλέψεις ήταν πολύ βοηθητικές προς την κατεύθυνση κατανόησης και ερμηνείας των πειραματικών αποτελεσμάτων, αφού παρείχαν την δυνατότητα συσχέτισης της προκύπτουσας κατανομής της θερμοκρασίας στο δάπεδο (endwall) με τις τοπικές τρισδιάστατες ροές. Πειραματικό Σκέλος (Experimental part): Για την πειραματική διερεύνηση της αποτελεσματικότητας της προτεινόμενης μεθόδου ψύξης, χρησιμοποιήθηκε μια νέα τεχνική η οποία αναπτύχθηκε ως μέρος της παρούσας εργασίας, υιοθετώντας θερμοκρασιακή κατανομή αντίστροφη από αυτήν που χρησιμοποιήθηκε για τις υπολογιστικές προβλέψεις, π.χ. Ψυχρή κύρια ροή (mainstream) - Θερμό πλαϊνό τοίχωμα (endwall) - Θερμότερος αέρας έγχυσης. Χρησιμοποιώντας την μέθοδο αυτή και με την χρήση υπέρυθρης θερμογραφίας (infrared thermography), ποσοτικοποιείται η αποτελεσματικότητα στην ψύξη του πλαϊνού τοιχώματος και προσδιορίζεται η περιοχή στην οποία η ψύξη είναι αποτελεσματική. Επιπλέον της ποιοτικής και ποσοτικής αποτίμησης της αποτελεσματικότητας της ψύξης, ήταν αναγκαίες αεροδυναμικές μετρήσεις για τον καθορισμό του αεροδυναμικού κόστους της προτεινόμενης μεθόδου ψύξης. Οι μετρήσεις αυτές, δίνουν μια ένδειξη του κατά πόσον η μέθοδος επηρεάζει την μεγέθυνση και ένταση των δευτερογενών ροών (π.χ. δίνη διακένου (passage vortex)) στην περιοχή κατάντη της ζώνης αλληλεπίδρασης του ψυκτικού τζετ με την τοπική τρισδιάστατη ροή. / The thermodynamic analysis of the Brayton cycle designates that the thermal efficiency and the specific work output of a Gas Turbine can be improved by increasing the Turbine Inlet Temperatures. Furthermore, increment of the turbine inlet temperatures also results into lower fuel consumption rates, while, if the gas turbine is meant for propulsion purposes, increment of the turbine inlet temperatures also results into increased engine thrust. Unfortunately, these high gas temperatures jeopardize the integrity of the high pressure turbine components and more particular, the turbine blades and the endwall on which the blades are attached. In modern turbines, the turbine inlet temperature may reach the level of 1900K, exceeding by far the melting temperature of the metal walls. As a result, the turbine components operate at much harsher environments than in the past. Maintaining adequate life in these high temperatures requires the development of new materials and manufacturing processes, as well as efficient cooling methods for the components of the turbine. In order to address and avoid the failure of the blades and endwall of a turbine cascade, the method of "film cooling" has been incorporated as part of the components design process. In the latter method, air is bled from the compressor stage, passed through internal chambers of the engine to the turbine components and is injected through discrete holes in the walls of the blades and the endwall, forming a thin protective layer film between the hot combustion gases and the metal surfaces of the parts. A critical region that is subjected into increased thermal stresses is the area around the leading edge - endwall juncture, which is inherently dominated by the presence of strong three dimensional secondary flows (also known as juncture flows) responsible for the increment of the local heat transfer rates to the order of 350%. Moreover, these flows, prevent the penetration of the fluid in the problematic area, displacing the coolant mass flux away from the surface of the endwall before providing adequate cooling. The subject of the current thesis, is the design, development and testing (both experimental and computational) of a prototype cooling scheme (with the method of film cooling), in order to effectively address the endwall overheating problem around the leading edge - endwall juncture, especially around the stagnation plane area. The main feature of the novel injection method is that the coolant air is ejected in such a way that the cooling effectiveness in the area is assisted by the rotational sense of local three-dimensional flows. The complexity of film cooling for the problematic area arises from two facts. Firstly, as mentioned previously the flow around the leading edge junction is characterized by complex three dimensional flows. Secondly, the problem is characterized by three temperatures: the temperature of the main flow, the endwall temperature and the temperature of the coolant air. In order to fully investigate the features and characteristics of the proposed cooling method, the work of the current thesis includes both, an experimental and a computational part: Computational part: The successful design of a possible cooling scheme for the particular region of the endwall requires the knowledge and understanding of local flow in which the coolant jet is to be entrained. Furthermore, it is important to understand the expected interaction between the coolant air and the local three-dimensional flow. For this purpose, the method of Computational Fluid Dynamics was employed for predicting the relevant three-dimensional flow field around the blade-endwall junction area. Simulations were made for both, the basic geometry in the absence of any coolant injection (which were used as a reference point) and simulations during the employment of the proposed coolant injection method which concerned the effect of various parameters on the cooling efficiency of the problematic area. For the simulations, the following temperature step was adopted: Warm endwall - Warmer main flow (mainstream) - Cold air injection, which is similar to that seen in real applications. The CFD predictions were very helpful towards understanding the relevant flow field in the junction area, both qualitatively and quantitatively in terms of the relative magnitudes of the flow structures and the aerodynamic characteristics of the flow in the region. Experimental part: For the experimental investigation regarding the effectiveness of the proposed cooling method, a new experimental technique was employed which was developed as part of the current thesis. In the latter technique, a reversed temperature step is adopted (when compared to the relevant temperature step adopted for the numerical simulations), e.g. Cold main flow (mainstream) - Warm endwall - Warmer air injection. Along with the use of infrared thermography, the endwall film cooling effectiveness is quantified and the region that the injection is effective is determined. In addition to the qualitative and quantitative evaluation of the cooling effectiveness, extensive aerodynamic measurements were necessary in order to evaluate the aerodynamic costs of the proposed cooling method. These measurements provide an indication of whether the cooling process affects the growth and intensity of secondary flows (e.g. passage vortex) in the region downstream of the coolant jet-local three-dimensional flow interaction.

This thesis aims to explore the potential for using large eddy simulation (LES) as a predictive tool for gas-turbine compressor flows. Compressors present a significant challenge for the Reynolds Averaged Navier-Stokes (RANS) based CFD methods commonly used in industry. RANS models require extensive calibration to experimental data, and thus cannot be used predictively. This thesis explores how LES can offer a more predictive alternative, by exploring the sensitivity of LES to sources of uncertainty. Specifically, the importance of the numerical scheme, the Sub-Grid Scale (SGS) model, and the correct specification of inflow turbulence is examined. The sensitivity of LES to the numerical scheme is explored using the Taylor-Green vortex test case. The numerical smoothing, controlled by a user defined smoothing constant, is found to be important. To avoid tuning the numerical scheme, a locally adaptive smoothing (LAS) scheme is implemented. But, this is found to perform poorly in a forced isotropic turbulence test case, due to the intermittency of the dispersive error. A novel scheme, the LAS with windowing (LASW) scheme, is thus introduced. The LASW scheme is shown to be more suitable for predictive LES, as it does not require tuning to a known solution. The LASW scheme is used to perform LES on a compressor cascade, and results are found to be in close agreement with direct numerical simulations. Complex transition mechanisms, combining characteristics of both natural and bypass modes, are observed on the pressure surface. These mechanisms are found to be sensitive to numerical smoothing, emphasising the importance of the LASW scheme, which returns only the minimum smoothing required to prevent dispersion. On the suction surface, separation induced transition occurs. The flow here is seen to be relatively insensitive to numerical smoothing and the choice of SGS model, as long as the Smagorinsky-Lilly SGS model is not used. These findings are encouraging, as they show that, with the LASW scheme and a suitable SGS model, LES can be used predictively in compressor flows. In order to be predictive, the accurate specification of inflow conditions was shown to be just as important as the numerics. RANS models are shown to over-predict the extent of the three dimensional separation in the endwall - suction surface corner. LES is used to examine the challenges for RANS in this region. The LES shows that it is important to accurately capture the suction surface transition location, with early transition leading to a larger endwall separation. Large scale aperiodic unsteadiness is also observed in the endwall region. Additionally, turbulent anisotropy in the endwall - suction surface corner is found to be important. Adding a non-linear term to the RANS model leads to turbulent stresses that are in better agreement with the LES. This results in a stronger corner vortex which is thought to delay the corner separation. The addition of a corner fillet reduces the importance of anisotropy, thereby reducing the uncertainty in the RANS prediction.