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.

This research work investigates the thermal performance of a film-cooled gas turbine endwall under two different mainstream flow conditions. In the first part of the research investigation, the effect of unsteady passing wakes on a film-cooled pitchwise-curved surface (representing an endwall without airfoils) was experimentally studied for heat transfer characteristics on a time-averaged basis. The temperature sensitive paint technique was used to obtain the local temperatures on the test surface. The required heat flux input was provided using foil heaters. Discrete film injection was implemented on the test surface using cylindrical holes with a streamwise inclination angle of 35? and no compound angle relative to the mean approach velocity vector. The passing wakes increased the heat transfer coefficients at both the wake passing frequencies that were experimented. Due to the increasing film cooling jet turbulence and strong jet-mainstream interaction at higher blowing ratios, the heat transfer coefficients were amplified. A combination of film injection and unsteady passing wakes resulted in a maximum pitch-averaged and centerline heat transfer augmentation of ? 28% and 31.7% relative to the no wake and no film injection case. The second part of the research study involves an experimental and numerical analysis of secondary flow and coolant film interaction in a high subsonic annular cascade with a maximum isentropic throat Mach number of ? 0.68. Endwall (platform) thermal protection is provided using discrete cylindrical holes with a streamwise inclination angle of 30? and no compound angle relative to the mean approach velocity vector. The surface flow visualization on the inner endwall provided the location of the saddle point and the three-dimensional separation lines. Computational predictions showed that the leading-edge horseshoe vortex was confined to approximately 1.5% of the airfoil span for the no film injection case and intensified with low momentum film injection. At the highest blowing ratio, the film cooling jet weakened the horseshoe vortex at the leading-edge plane. The passage vortex was intensified with coolant injection at all blowing ratios. It was seen that increasing average blowing ratio improved the film effectiveness on the endwall. The discharge coefficients calculated for each film cooling hole indicated significant non-uniformity in the coolant discharge at lower blowing ratios and the strong dependence of discharge coefficients on the mainstream static pressure and the location of three-dimensional separation lines. Near the airfoil suction side, a region of coalesced film cooling jets providing close to uniform film coverage was observed, indicative of the mainstream acceleration and the influence of three-dimensional separation lines.

The study presented herein describes and analyzes a detailed experimental investigation of the effects of density ratio on endwall thermal performance at varying blowing rates for a typical nozzle guide vane platform purge jet cooling scheme. An axisymmetric converging endwall with an upstream doublet staggered cylindrical hole purge jet cooling scheme was employed. Nominal exit flow conditions were engine representative and as follows: {rm Ma}_{Exit} = 0.85, {rm Re}_{Exit,C_{ax}} = 1.5 times {10}^6, and large-scale freestream Tu = 16%. Two blowing ratios were investigated corresponding to the upper and lower engine extrema. Each blowing ratio was investigated amid two density ratios; one representing typical experimental neglect of density ratio, at DR = 1.2, and another engine representative density ratio achieved by mixing foreign gases, DR = 1.95. All tests were conducted on a linear cascade in the Virginia Tech Transonic Blowdown Wind Tunnel using IR thermography and transient data reduction techniques. Oil paint flow visualization techniques were used to gather quantitative information regarding the alteration of endwall flow physics due two different blowing rates of high-density coolant. High resolution endwall adiabatic film cooling effectiveness, Nusselt number, and Net Heat Flux Reduction contour plots were used to analyze the thermal effects. The effect of density is dependent on the coolant blowing rate and varies greatly from the high to low blowing condition. At the low blowing condition better near-hole film cooling performance and heat transfer reduction is facilitated with increasing density. However, high density coolant at low blowing rates isn't adequately equipped to penetrate and suppress secondary flows, leaving the SS and PS largely exposed to high velocity and temperature mainstream gases. Conversely, it is observed that density ratio only marginally affects the high blowing condition, as the momentum effects become increasingly dominant. Overall it is concluded density ratio has a first order impact on the secondary flow alterations and subsequent heat transfer distributions that occur as a result of coolant injection and should be accounted for in purge jet cooling scheme design and analysis. Additionally, the effect of increasing high density coolant blowing rate was analyzed. Oil paint flow visualization indicated that significant secondary flow suppression occurs as a result of increasing the blowing rate of high-density coolant. Endwall adiabatic film cooling effectiveness, Nusselt number, and NHFR comparisons confirm this. Low blowing rate coolant has a more favorable thermal impact in the upstream region of the passage, especially near injection. The low momentum of the coolant is eventually dominated and entrained by secondary flows, providing less effectiveness near PS, near SS, and into the throat of the passage. The high momentum present for the high blowing rate, high-density coolant suppresses these secondary flows and provides enhanced cooling in the throat and in high secondary flow regions. However, the increased turbulence impartation due to lift off has an adverse effect on the heat load in the upstream region of the passage. It is concluded that only marginal gains near the throat of the passage are observed with an increase in high density coolant blowing rate, but severe thermal penalty is observed near the passage onset. / Master of Science / Gas turbine technology is used frequently in the burning of natural gas for power production. Increases in engine efficiency are observed with increasing firing temperatures, however this leads to the potential of overheating in the stages following. To prevent failure or melting of components, cooler air is extracted from the upstream compressor section and used to cool these components through various highly complex cooling schemes. The design and operational adequacy of these schemes is highly subject to the mainstream and coolant flow conditions, which are hard to represent in a laboratory setting. This experimental study explores the effects of various coolant conditions, and their respective response, for a purge jet cooling scheme commonly found in engine. This scheme utilizes two rows of staggered cylindrical holes to inject air into the mainstream from platform, upstream of the nozzle guide vane. It is the hope that this air forms a protective layer, effectively shielding the platform from the hostile mainstream conditions. Currently, little research has been done to quantify these effects of purge flow cooling scheme while mimicking engine geometry, mainstream and coolant conditions. For this study, an endwall geometry like that found in engine with a purge jet cooling scheme is studied. Commonly, an upstream gap is formed between the combustor lining and first stage vane platform, which is accounted for in this testing. Mainstream and coolant flow conditions can have large impacts on the results gathered, so both were matched to engine conditions. Varying of coolant density and injection rate is studied and quantitative results are gathered. Results indicate coolant fluid density plays a large role in purge jet cooling, and with neglection of this, potential thermal failure points could be overlooked This is exacerbated with less coolant injection. Interestingly, increasing the amount of coolant injected decreases performance across much of the passage, with only marginal gains in regions of complex flow. These results help to better explain the impacts of experimental neglect of coolant density, and aid in the understanding of purge jet coolant injection.

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.