Effects of Tip Clearance Gap and Exit Mach Number on Turbine Blade Tip and Near-Tip Heat Transfer

This study focuses on local heat transfer characteristics on the tip and near-tip regions of a turbine blade with a flat tip, tested under transonic conditions in a stationary, 2-D linear cascade with high freestream turbulence. The experiments were conducted at the Virginia Tech transonic blow-down wind tunnel facility. The effects of tip clearance and exit Mach number on heat transfer distribution were investigated on the tip surface using a transient infrared thermography technique. In addition, thin film gages were used to study similar effects in heat transfer on the near-tip regions at 94% height based on engine blade span of the pressure and suction sides. Surface oil flow visualizations on the blade tip region were carried-out to shed some light on the leakage flow structure. Experiments were performed at three exit Mach numbers of 0.7, 0.85, and 1.05 for two different tip clearances of 0.9% and 1.8% based on turbine blade span. The exit Mach numbers tested correspond to exit Reynolds numbers of 7.6 × 105, 9.0 × 105, and 1.1 × 106 based on blade true chord. The tests were performed with a high freestream turbulence intensity of 12% at the cascade inlet. Results at 0.85 exit Mach showed that an increase in the tip gap clearance from 0.9% to 1.8% translates into a 3% increase in the average heat transfer coefficients on the blade tip surface. At 0.9% tip clearance, an increase in exit Mach number from 0.85 to 1.05 led to a 39% increase in average heat transfer on the tip. High heat transfer was observed on the blade tip surface near the leading edge, and an increase in the tip clearance gap and exit Mach number augmented this near-leading edge tip heat transfer. At 94% of engine blade height on the suction side near the tip, a peak in heat transfer was observed in all test cases at s/C = 0.66, due to the onset of a downstream leakage vortex, originating from the pressure side. An increase in both the tip gap and exit Mach number resulted in an increase, followed by a decrease in the near-tip suction side heat transfer. On the near-tip pressure side, a slight increase in heat transfer was observed with increased tip gap and exit Mach number. In general, the suction side heat transfer is greater than the pressure side heat transfer, as a result of the suction side leakage vortices.

Temperature-Based Optimization of Film Cooling in Gas Turbine Hot Gas Path Components

To achieve high efficiency, modern gas turbines operate at temperatures that exceed melting points of metal alloys used in turbine hot gas path parts. Parts exposed to hot gas are actively cooled with a portion of the compressor discharge air (e.g., through film cooling) to keep the metal temperature at levels needed to meet durability requirements. However, to preserve efficiency, it is important to optimize the cooling system to use the least amount of cooling flow. In this study, film cooling optimization is achieved by varying cooling hole diameters, hole to hole spacing, and film row placements so that the specified targets for maximum metal temperature are met while preserving (or saving) cooling flow. The computational cost of the high-fidelity physics models, the large number of design variables, the large number and nonlinearity of responses impose severe challenges to numerical optimization. Design of experiments and cheap-to-evaluate approximations (radial basis functions) are used to alleviate the computational burden. Then, the goal attainment method is used for optimizing of film cooling configuration. The results for a turbine blade design show significant improvements in temperature distribution while maintaining/reducing the amount of used cooling flow.

Review of the von Karman Institute Compression Tube Facility for Turbine Research

Advances in turbine-based engine efficiency and reliability are achieved through better knowledge of the mechanical interaction with the flow. The life-limiting component of a modern gas turbine engine is the high-pressure (HP) turbine stage due to the arduous environment. For the same reason, real gas turbine engine operation prevents fundamental research. Various types of experimental approaches have been developed to study the flow and in particular the heat transfer, cooling, materials, aero-elastic issues and forced response in turbines. Over the last 30 years short duration facilities have dominated the research in the study of turbine heat transfer and cooling. Two decades after the development of the von Karman Institute compression tube facility (built in the 90s), one could reconsider the design choices in view of the modern technology in compression, heating, control and electronics. The present paper provides first the history of the development and then how the wind tunnel is operated. Additionally the paper disseminates the experience and best practices in specifically designed measurement techniques to both experimentalists and experts in data processing. The final section overviews the turbine research capabilities, providing details on the required upgrades to the test section.

Validation of Surface Pressure Predictions for Nozzle Airfoil and Endwall Film Cooling Design Using Transonic Cascade Measurements

This study compares surface pressure measurements and predictions for a high pressure turbine first-stage nozzle vane. The surface pressure measurements were taken in a 3D annular cascade, consisting of four airfoils and five passages. The cascade was uncooled, axisymmetric at both inner and outer endwalls, and reproduced the design intent Reynolds and Mach numbers of the real engine component. Static pressure measurements were taken along the airfoil profile at 15, 50, and 85% span, with duplicate midspan measurements across the four airfoils for assessing the tangential periodicity of the flow. Static pressure measurements were also taken on the inner and outer endwall surfaces of the center airfoil passage, with 40 measurement points uniformly distributed over each endwall. Three methods of surface pressure prediction were compared with the data: (1) a 2D inviscid CFD solution of a single airfoil passage at fixed spanwise locations, (2) a 3D RANS CFD solution of a single airfoil passage, and (3) a 3D RANS CFD solution of the full five-passage cascade domain. Both of the single-passage solutions assumed flowfield periodicity in the tangential direction and compared favorably to the center passage airfoil data. This finding suggested that the cascade center passage was sufficiently representative of the full-annulus turbomachine environment and validated the cascade for further experimental studies. The adjacent airfoil pressure measurements quantified the passage-to-passage variation in the cascade flowfield, and the 3D full-cascade CFD compared favorably with the peripheral airfoil data. The full-cascade CFD also compared favorably with the data on both endwalls: with an average and maximum deviation of 0.5 and 2%, respectively. These findings provide confidence in the 3D CFD methods for use in determining local flow rates from cooling/leakage geometry, and serve as an important first step toward validating the methods for real-engine blockage effects like coolant and endwall contouring.

Direct Measurements of Overall Effectiveness and Heat Flux on a Film Cooled Test Article at High Temperatures and Pressures

The energy requirements associated with recovering greenhouse gases from Integrated Gasification Combined Cycle (IGCC) or Natural Gas Combined Cycle (NGCC) power plants are significant. The subsequent reductions in overall plant efficiency also result in a higher cost of electricity. In order to meet the future demand for cleaner energy production, this research is focused on improving gas turbine efficiency through advancements in gas turbine cooling capabilities. For this study, an experimental approach was developed to quantify overall effectiveness and net heat flux reduction for a film-cooled test article at high temperature and pressure conditions. A major part of this study focused on validating an advanced optical thermography technique capable of distinguishing between emitted and reflected radiation from film-cooled test articles exposed to exhaust gases in excess of 1000°C and 5 bar. The optical thermography method was used to acquire temperature maps of both external and internal wall temperatures on a test article with fan-shaped film cooling holes. The overall effectiveness and heat flux were quantified with one experiment. The optical temperature measurement technique was capable of measuring wall temperatures to within ±7.2°C. Uncertainty estimates showed that the methods developed for this study were capable of quantifying improvements in overall effectiveness necessary to meet DOE program goals. Results showed that overall effectiveness increased with an increase in blowing ratio and a decrease in mainstream gas pressure while heat flux contours indicated consistent trends.

Blade Tip Shape Optimization for Enhanced Turbine Aerothermal Performance

In high-speed unshrouded turbines tip leakage flows generate large aerodynamic losses and intense unsteady thermal loads over the rotor blade tip and casing. The stage loading and rotational speeds are steadily increased to achieve higher turbine efficiency, and hence the overtip leakage flow may exceed the transonic regime. However, conventional blade tip geometries are not designed to cope with supersonic tip flow velocities. A great potential lays in the modification and optimization of the blade tip shape as a means to control the tip leakage flow aerodynamics, limit the entropy production in the overtip gap, manage the heat load distribution over the blade tip and improve the turbine efficiency at high stage loading coefficients. The present paper develops an optimization strategy to produce a set of blade tip profiles with enhanced aerothermal performance for a number of tip gap flow conditions. The tip clearance flow was numerically simulated through two-dimensional compressible Reynolds-Averaged Navier-Stokes (RANS) calculations that reproduce an idealized overtip flow along streamlines. A multi-objective optimization tool, based on differential evolution combined with surrogate models (artificial neural networks), was used to obtain optimized 2D tip profiles with reduced aerodynamic losses and minimum heat transfer variations and mean levels over the blade tip and casing. Optimized tip shapes were obtained for relevant tip gap flow conditions in terms of blade thickness to tip gap height ratios (between 5 and 25), and blade pressure loads (from subsonic to supersonic tip leakage flow regimes) imposing fixed inlet conditions. We demonstrated that tip geometries which perform superior in subsonic conditions are not optimal for supersonic tip gap flows. Prime tip profiles exist depending on the tip flow conditions. The numerical study yielded a deeper insight on the physics of tip leakage flows of unshrouded rotors with arbitrary tip shapes, providing the necessary knowledge to guide the design and optimization strategy of a full blade tip surface in a real 3D turbine environment.

Design of a Non-Reacting Combustor Simulator With Swirl and Temperature Distortion With Experimental Validation

Nonuniform combustor outlet flows have been demonstrated to have significant impact on the first and second stage turbine aerothermal performance. Rich-burn combustors, which generally have pronounced temperature profiles and weak swirl profiles, primarily affect the heat load in the vane but both the heat load and aerodynamics of the rotor. Lean burn combustors, in contrast, generally have a strong swirl profile which has an additional significant impact on the vane aerodynamics which should be accounted for in the design process. There has been a move towards lean burn combustor designs to reduce NOx emissions. There is also increasing interest in fully integrated design processes which consider the impact of the combustor flow on the design of the HP vane and rotor aerodynamics and cooling. There are a number of current large research projects in scaled (low temperature and pressure) turbine facilities which aim to provided validation data and physical understanding to support this design philosophy. There is a small body of literature devoted to rich burn combustor simulator design but no open literature on the topic of lean burn simulator design. The particular problem is that in non-reacting, highly swirling and diffusing flows, vortex instability in the form of a precessing vortex core or vortex breakdown is unlikely to be well matched to the reacting case. In reacting combustors the flow is stabilised by heat release, but in low temperature simulators other methods for stabilising the flow must be employed. Unsteady Reynolds-averaged Navier-Stokes and Large eddy simulation have shown promise in modelling swirling flows with unstable features. These design issues form the subject of this paper.

A Study of the Heat Transfer in Winglet and Squealer Rotor Tip Configurations for a Non-Cooled HPT Blade Based on CFD Calculations

HPT operate at high pressure and temperatures. One of the most important loss sources is the tip leakage flow on the rotor tip region. The flow that leaks in this region does not participate in the energy transfer process between the hot gas and rotor blade row. Hence, the main flow suffers a penalty to maintain the energy conservation. To try decreasing this mass flow leakage some techniques can be applied. The most common are the winglet and squealer rotor tip configuration. These techniques improve the turbine performance, but some attention should be taken into account because the temperature distribution changes on this region for different tip configurations. In this work, the winglet and squealer tip geometries are compared with the common flat tip configuration. The analysis was performed for design and off-design conditions. The HPT developed in the E3 program was used as baseline turbine to explore the differences of the flowfield on the rotor tip region. The results are compared and discussed in detail.

Endwall Heat Transfer Measurements of an Outlet Guide Vane at On and Off Design Conditions

The endwall heat transfer characteristics of forced flow past outlet guide vanes (OGVs) in a linear cascade have been investigated by using a liquid crystal thermography (LCT) method. The Reynolds number is kept at 250,000 and both on- and off-design conditions are tested. For the on-design condition where the incidence angle of OGVs is 30°, no obvious flow separation phenomenon was observed; on the contrary, for the off-design conditions where the incidence angle of OGVs is 0° and −31°, respectively, remarkable flow separation was noticed. The results indicate that the incidence angle of OGVs has a significant effect on the endwall heat transfer. In general, the endwall heat transfer coefficients for the off-design conditions are higher than the corresponding on-design condition. In addition, a preliminary CFD analysis was performed and presented. Basically, the results are consistent with the experiments but further investigations are needed in the future work.