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Title: Unshrouded turbine blade tip heat transfer and film cooling
Author: Tang, Brian M. T.
Awarding Body: University of Oxford
Current Institution: University of Oxford
Date of Award: 2011
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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.
Supervisor: Gillespie, David R. H. Sponsor: Not available
Qualification Name: Thesis (Ph.D.) Qualification Level: Doctoral
EThOS ID:  DOI: Not available
Keywords: Aero engines ; Aerodynamics and heat transfer ; Turbomachinery hypersonics and aerodynamics ; Mechanical engineering ; Engineering & allied sciences ; gas turbine ; jet engine ; turbomachinery ; heat transfer ; film cooling ; aerodynamics ; fluid flow ; computational fluid dynamics ; turbine ; blade tip