Transonic Passage Turbine Blade Tip Clearance With Scalloped Shroud: Part II — Losses With and Without Scrubbing Effects in Engine Configuration

2004 ◽  
Author(s):  
Wayne S. Strasser ◽  
Gregory M. Feldman ◽  
F. Casey Wilkins ◽  
James H. Leylek
Keyword(s):  
Turbine Blade ◽  
Turbulence Model ◽  
Large Scale ◽  
Pressure Ratio ◽  
Viscous Sublayer ◽  
Tip Clearance ◽  
Tip Vortex ◽  
Tip Leakage ◽  
Passage Vortex ◽  
Blade Tip

Loss mechanisms in a scallop shrouded transonic power generation turbine blade passage at realistic engine conditions have been identified through a series of large-scale (typically 12 million finite volumes) simulations. All simulations are run with second-order discretization and viscous sublayer resolution, and they include the effects of viscous dissipation. The mesh (y+ near unity on all surfaces) is highly refined in the tip clearance region, casing recesses, and shroud region in order to fully capture complex interdependent flow physics and the associated losses. Aerodynamic losses, in order of their relative importance, are a result of the following: separation around the tip, recesses, and shroud; tip vortex creation; downstream mixing losses, localized shocks on the airfoil; and the passage vortex emanating from under the shroud. A number of helical lateral flows were established near the upper shroud surfaces as a result of lateral pressure gradients on the scalloped shroud. It was found that the tip leakage and passage losses increased approximately linearly with increasing tip clearance, both with and without the effect of the relative casing motion. For each tip clearance studied, scrubbing slightly reduced the tip leakage, but the overall production of entropy was increased by more than 50%. Also the overall passage mass flow rate, for a given inlet total pressure to exit static pressure ratio, increased almost linearly with increasing tip clearance. In addition, it was also found that there was slight positive and negative lift on the shroud, depending on the tip clearance. At the lowest tip clearance of 20 mils there was a negative lift on the shroud. In the 200-mil tip clearance case there was a positive lift on the shroud. The relative motion of the casing contributed positively to the lift at every tip clearance, affecting more at the lowest tip clearance where the casing is closest to the blade tip. Lastly, it was found that the computed entropy generation for the stationary 80-mils case using the SKE turbulence model was close to that of the 80-mils scrubbing case using the RKE turbulence model. In light of the proposed mechanisms and their relative contributions, suggested design considerations are posed.

Transonic Passage Turbine Blade Tip Clearance With Scalloped Shroud: Part III — Heat Transfer in Engine Configuration

2004 ◽  
Author(s):  
F. Casey Wilkins ◽  
Gregory M. Feldman ◽  
Wayne S. Strasser ◽  
James H. Leylek
Keyword(s):  
Heat Transfer ◽  
Turbine Blade ◽  
Large Scale ◽  
Numerical Study ◽  
Viscous Sublayer ◽  
Tip Clearance ◽  
Transfer Coefficients ◽  
Constant Wall Heat Flux ◽  
Blade Tip ◽  
Large Scale Land

This work presents a numerical study that was done to investigate the heat transfer characteristics of a transonic turbine blade with a scalloped shroud operating at realistic engine conditions typical of those found in a large scale, land-based gas turbine. The geometry under investigation was an infinite, linear cascade composed of the same blade and shroud design used in an experimental test rig by the research sponsor. This simulation was run for varying nominal tip clearances of 20, 80, and 5.08 mm. For each of these clearances, the simulation was run with and without the scrubbing effects of the outer casing, resulting in a total of six cases that could be used to determine the influence of tip clearance and relative casing motion on heat transfer. A high quality grid (ranging from approximately 10–12 million finite volumes depending on tip clearance) with y+ for first layer cells at or below 1.0 everywhere was used to resolve the flow down to the viscous sublayer. The “realizable” k-ε turbulence model was used for all cases. A constant wall heat flux was imposed on all the surrounding surfaces to obtain heat transfer data. Results produced include a full map of heat transfer coefficients for the suction and pressure surfaces of the blade as well as the tip, shroud, and outer casing for every case. Physical mechanisms responsible for the final heat transfer outcome for all six cases are documented.

Parametrical Investigation of Turbine Stages With Open Tip Clearance for the Purpose of Stage Efficiency Increase

2010 ◽  
Author(s):  
Andrei Granovskiy ◽  
Mikhail Kostege ◽  
Nikolay Lomakin
Keyword(s):  
Pressure Ratio ◽  
Axial Flow ◽  
Tip Clearance ◽  
Tip Vortex ◽  
Turbine Stage ◽  
Suction Surface ◽  
Tip Leakage ◽  
Degree Of Reaction ◽  
Blade Tip ◽  
Stage Efficiency

The aerodynamic loss due to tip leakage vortex of rotor blades represents a significant part of viscous losses in axial flow turbines. The mixing of leakage flow with the main rotor passage flow causes losses and reduces turbine stage efficiency. Many measures have been proposed to reduce the loss in the tip clearance area. In this paper the reduction of the tip clearance loss due to changes made to the blade tip section profile is presented. The blade tip profile was modified to decrease the pressure gradient between pressure surface and suction surface. This approach allows the reduction of tip leakage and tip vortex strength and consequently the reduction of tip clearance losses. A 3D Navier-Stokes solver with q-ω turbulence model is used to analyze the flow in the turbine with various tip section profiles. Test data of three single-stage experimental turbines have been used to validate analytical predictions: • Highly loaded turbine stage with a pressure ratio π0T = 3.2 and reaction degree ρmean = 0.5. • Two turbines with a pressure ratio π0T = 3.9. (One with high degree of reaction ρmean = 0.55; the other with low degree of reaction ρmean = 0.26). The numerical investigation of the influence of various tip section profiles on stage efficiency was carried out in the range of relative tip clearance 0.5%–2.4% with the objective of a decreasing the influence of the tip clearance on the stage efficiency.

Effect of Turbine Blade Tip Cooling Configuration on Tip Leakage Flow and Heat Transfer

2020 ◽  
Vol 142 (2) ◽  
Author(s):  
Sergen Sakaoglu ◽  
Harika S. Kahveci
Keyword(s):  
Heat Transfer ◽  
Turbine Blade ◽  
Gas Turbines ◽  
Thermal Response ◽  
Thermal Conditions ◽  
Tip Clearance ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Tip Leakage ◽  
Blade Tip

Abstract The pressure difference between suction and pressure sides of a turbine blade leads to tip leakage flow, which adversely affects the first-stage high-pressure (HP) turbine blade tip aerodynamics. In modern gas turbines, HP turbine blade tips are exposed to extreme thermal conditions requiring cooling. If the coolant jet directed into the blade tip gap cannot counter the leakage flow, it will simply add up to the pressure losses due to leakage. Therefore, the compromise between the aerodynamic loss and the gain in tip-cooling effectiveness must be optimized. In this paper, the effect of tip-cooling configuration on the turbine blade tip is investigated numerically from both aerodynamics and thermal aspects to determine the optimum configuration. Computations are performed using the tip cross section of GE-E3 HP turbine first-stage blade for squealer and flat tips, where the number, location, and diameter of holes are varied. The study presents a discussion on the overall loss coefficient, total pressure loss across the tip clearance, and variation in heat transfer on the blade tip. Increasing the coolant mass flow rate using more holes or by increasing the hole diameter results in a decrease in the area-averaged Nusselt number on the tip floor. Both aerodynamic and thermal response of squealer tips to the implementation of cooling holes is superior to their flat counterparts. Among the studied configurations, the squealer tip with a larger number of cooling holes located toward the pressure side is highlighted to have the best cooling performance.

scholarly journals Evaluating the Aerodynamic Impact of Circumferentially Grooved Fan Casing Treatments With Integrated Acoustic Liners on a Turbofan Rotor

2019 ◽  
Author(s):  
Richard F. Bozak ◽  
Gary G. Podboy
Keyword(s):  
Pressure Ratio ◽  
Aerodynamic Performance ◽  
Tip Clearance ◽  
Tip Vortex ◽  
Fan Noise ◽  
Tip Leakage ◽  
Fan Performance ◽  
Acoustic Liners ◽  
Rotor Stator Interaction ◽  
Real Gas Effects

Abstract NASA is investigating the potential of integrating acoustic liners into fan cases to reduce fan noise, while maintaining the fans aerodynamic performance. An experiment was conducted to quantify the aerodynamic impact of circumferentially grooved fan cases with integrated acoustic liners on a 1.5 pressure ratio turbofan rotor. In order to improve the ability to measure small performance changes, fan performance calculations were updated to include real gas effects including the effect of humidity. For all fan cases tested, the measured difference in fan isentropic efficiency was found to be less than the measurement repeatability for a torque-based efficiency calculation (≈ 0.2%), however, an unintended tip clearance difference between configurations makes it difficult to determine if circumferentially grooved fan cases degraded fan performance. Fan exit turbulence measurements showed a 1.5% reduction in total turbulence intensity between hardwall and circumferentially grooved fan cases in the tip vortex region, which is attributed to a disruption in the formation of the tip leakage vortex. This decrease in fan exit turbulence could potentially lead to a 1–2dB reduction in broadband rotor-stator interaction noise. Reduced aerodynamic performance losses associated with over-the-rotor liners could enable further fan noise reduction.

Effect of Turbine Blade Tip Cooling Configuration on Tip Leakage Flow and Heat Transfer

2019 ◽  
Author(s):  
Sergen Sakaoglu ◽  
Harika S. Kahveci
Keyword(s):  
Heat Transfer ◽  
Turbine Blade ◽  
Gas Turbines ◽  
Thermal Response ◽  
Thermal Conditions ◽  
Tip Clearance ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Tip Leakage ◽  
Blade Tip

Abstract The pressure difference between suction and pressure sides of a turbine blade leads to the so-called phenomenon, the tip leakage flow, which most adversely affects the first-stage high-pressure (HP) turbine blade tip aerodynamics. In modern gas turbines, HP turbine blade tips are also exposed to extreme thermal conditions requiring the use of tip cooling. If the coolant jet directed into the blade tip gap cannot counter the leakage flow, it will simply add up to the pressure losses due to this leakage flow. Therefore, it is necessary to handle the design of tip cooling in such a way that the compromise between the aerodynamic loss and the gain in the tip cooling effectiveness is optimized. In this paper, the effect of tip cooling configuration on the turbine blade tip is investigated numerically both from the aerodynamics and thermal aspects in order to determine the optimum tip cooling configuration. The studies are carried out using the tip cross-section of General Electric E3 (Energy Efficient Engine) HP turbine first-stage blade for two different tip geometries, squealer tip and flat tip, where the number, location, and diameter of the cooling holes are varied. The study presents a discussion on the overall loss coefficient, the total pressure loss across the tip clearance, and the variation of heat transfer on the blade tip. The aerodynamic and heat transfer results are compared with the experimental data from literature. It is observed that increasing the coolant mass flow rate by using more holes or by increasing the hole diameter results in a decrease in the area-averaged Nusselt number on the tip floor, as expected. The findings show that both aerodynamic and thermal response of the squealer tips to the implementation of cooling holes is superior to their flat counterparts. Among the studied configurations, the squealer tip with larger number of cooling holes located towards the pressure side is highlighted as the configuration having the best cooling performance.

Detailed Heat Transfer Coefficient Distributions on a Large-Scale Gas Turbine Blade Tip

2000 ◽  
Vol 123 (4) ◽  
pp. 803-809 ◽  
Author(s):  
Shuye Teng ◽  
Je-Chin Han ◽  
G. M. S. Azad
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Wind Tunnel ◽  
Transfer Coefficient ◽  
Turbine Blade ◽  
Large Scale ◽  
Tip Clearance ◽  
Low Speed ◽  
Blade Tip ◽  
Low Speed Wind Tunnel

Measurements of detailed heat transfer coefficient distributions on a turbine blade tip were performed in a large-scale, low-speed wind tunnel facility. Tests were made on a five-blade linear cascade. The low-speed wind tunnel is designed to accommodate the 107.49 deg turn of the blade cascade. The mainstream Reynolds number based on cascade exit velocity was 5.3×105. Upstream unsteady wakes were simulated using a spoke-wheel type wake generator. The wake Strouhal number was kept at 0 or 0.1. The central blade had a variable tip gap clearance. Measurements were made at three different tip gap clearances of about 1.1 percent, 2.1 percent, and 3 percent of the blade span. Static pressure distributions were measured in the blade mid-span and on the shroud surface. Detailed heat transfer coefficient distributions were measured on the blade tip surface using a transient liquid crystal technique. Results show that reduced tip clearance leads to reduced heat transfer coefficient over the blade tip surface. Results also show that reduced tip clearance tends to weaken the unsteady wake effect on blade tip heat transfer.

scholarly journals Turbine Blade Tip External Cooling Technologies

Aerospace ◽  
2018 ◽  
Vol 5 (3) ◽  
pp. 90 ◽  
Author(s):  
Song Xue ◽  
Wing Ng
Keyword(s):  
Turbine Blade ◽  
Recent Progress ◽  
Tip Clearance ◽  
New Techniques ◽  
Cooling Performance ◽  
Tip Leakage ◽  
Blowing Ratio ◽  
External Cooling ◽  
Blade Tip ◽  
Cooling Technology

This article provides an overview of gas turbine blade tip external cooling technologies. It is not the intention to comprehensively review all the publications from past to present. Instead, selected reports, which represent the most recent progress in tip cooling technology in open publications, are reviewed. The cooling performance on flat tip and squealer tip blades from reports are compared and discussed. As a generation conclusion, tip clearance dimension and coolant flow rate are found as the most important factors that significant influence the blade tip thermal performance was well as the over tip leakage (OTL) flow aerodynamics. However, some controversial trends are reported by different researchers, which could be attributed to various reasons. One of the causes of this disagreement between different reports is the lacking of unified parametric definition. Therefore, a more appropriate formula of blowing ratio definition has been proposed for comparison across different studies. The last part of the article is an outlook of the new techniques that are promising for future tip cooling research. As a new trend, the implementation of artificial intelligence techniques, such as genetic algorithm and neural network, have become more popular in tip cooling optimization, and they will bring significantly changes to the future turbine tip cooling development.

scholarly journals Theoretical and Experimental Investigation of a 34 Watt Radial-Inflow Steam Turbine with Partial-Admission

2020 ◽  
Author(s):  
Patrick Wagner ◽  
Jan Van herle ◽  
Jurg Schiffmann
Keyword(s):  
Steam Turbine ◽  
Pressure Ratio ◽  
Turbine Blades ◽  
Turbine Rotor ◽  
Tip Clearance ◽  
Experimental Results ◽  
Tip Leakage ◽  
Hot Air ◽  
Partial Admission ◽  
Blade Tip

Abstract A micro steam turbine with a tip diameter of 15 mm was designed and experimentally characterized. At the nominal mass flow rate and total-to-total pressure ratio of 2.3 kg/h and 2, respectively, the turbine yields a power of 34 W and a total-to-static isentropic efficiency of 37 %. The steam turbine is conceived as a radial-inflow, low-reaction (15 %), and partial admission (21 %) machine. Since the steam is limited in the system (solid oxide fuel cell), a low-reaction and high-power-density design is preferred. The partial-admission design allows for reduced losses: The turbine rotor and stator blades are prismatic, have a radial chord length of 1 mm and a height of 0.59 mm. Since the relative rotor blade tip clearance (0.24) is high, the blade tip leakage losses are significant. Considering a fixed steam supply, this design allows to increase the blade height, and thus reducing the losses. The steam turbine drives a fan, which operates at low Mach numbers. The rotor is supported on dynamic steam-lubricated bearings; the nominal rotational speed is 175 krpm. A numerical simulation of the steam turbine is in good agreement with the experimental results. Furthermore, a novel test rig setup, featuring extremely-thin thermocouples (0.15 mm) is investigated for an operation with ambient and hot air at 220 °C. Conventional zero and one-dimensional pre-design models correlate well to the experimental results, despite the small size of the turbine blades.

Numerical Modeling of Heat Transfer and Pressure Losses for Uncooled Gas Turbine Blade Tip: Effect of Tip Clearance and Tip Geometry

2007 ◽  
Author(s):  
Lamyaa A. El-Gabry
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Gas Turbine ◽  
Computational Study ◽  
Numerical Models ◽  
Pressure Ratio ◽  
Tip Clearance ◽  
Pressure Losses ◽  
Blade Tip ◽  
Exit Mach Number

A computational study has been performed to predict the heat transfer distribution on the blade tip surface for a representative gas turbine first stage blade. CFD predictions of blade tip heat transfer are compared to test measurements taken in a linear cascade, when available. The blade geometry has an inlet Mach number of 0.3 and an exit Mach number of 0.75, pressure ratio of 1.5, exit Reynolds number based on axial chord of 2.57×106, and total turning of 110 deg. Three blade tip configurations were considered; they are flat tip, a full perimeter squealer, and an offset squealer where the rim is offset to the interior of the tip perimeter. These three tip geometries were modeled at three tip clearances of 1.25, 2.0, and 2.75% of blade span. The tip heat transfer results of the numerical models agree fairly well with the data and are comparable to other CFD predictions in the open literature.

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