Author: Publisher:
ISBN:
Category :
Languages : en
Pages : 290
Book Description
Film-cooling of a Turbine Nozzle Guide Vane Endwall
Author: Measurements in a Gas Turbine First Stage Nozzle Guide Vane Cascade with Film Cooling and Endwall Contouring
Author: Rohit Anand OkeThe Influence of Film-cooling on the Aerodynamic Performance of a Turbine Nozzle Guide Vane
Author: C. OsnaghiPublisher:
ISBN:
Category :
Languages : en
Pages : 0
Book Description
Presented at the International Gas Turbine & Aeroegine Congress & Exhibition, Orlando, FL, Jun 2-Jun 5, 1997.
The Influence of Film Cooling and Inlet Temperature Profile on Heat Transfer for the Vane Row of a 1-1/2 Stage Transonic High-pressure Turbine
Author: Harika Senem KahveciPublisher:
ISBN:
Category :
Languages : en
Pages : 269
Book Description
Abstract: The goal of this research was to establish an extensive database for typical engine hardware with a film-cooled first stage vane, which represents the foundation for future turbomachinery film cooling modeling and component heat transfer studies. Until this time, such a database was not available within the gas turbine industry. Accordingly, the study focuses on determination of the local heat flux for the airfoil and endwall surfaces of the vane row of a fully-cooled turbine stage. The measurements were performed at the Ohio State University Gas Turbine Laboratory using the Turbine Test Facility. The full-scale rotating 1 and 1/2 turbine stage is operated at the proper corrected engine design conditions: Flow Function (FF), corrected speed, stage Pressure Ratio (PR), and temperature ratios of gas to wall and gas to coolant. The primary measurements of temperature, pressure, and heat flux are repeated for different vane inlet temperature profiles and different vane cooling flows to establish an understanding of the influence of film cooling on local heat transfer. Double-sided Kapton heat-flux gauges are used for heat-flux measurements at different span locations along the airfoil surfaces and along the inner endwall. The cooling scheme consists of numerous cooling holes located on the endwalls, at the airfoil leading edge, on the airfoil pressure and suction surfaces, and at the trailing edge, resulting in a fully cooled first stage vane. The unique film-cooled endwall heat transfer data demonstrated in contour plots reveals insight to the complex flow behavior that is dominant in this region, which becomes even more complicated with the addition of coolant. Varying profile shapes resulted in significant heat transfer variations in a growing fashion towards the trailing edge region, which increased in magnitude when there is no coolant supply. The largest cooling effect is observed on 5% span pressure surface and at the inner endwall region. Heat transfer decreases from tip towards hub with addition of cooling. However, a similar decrease is not observed at the inner endwall region by doing so, which suggests excess coolant once beyond an optimum blowing ratio. Cooling flow rate and temperature profile shape affect the distributions on the airfoil surface very similarly, the latter observed more clearly at the endwall region. The vane outer cooling effect is comparable to the combined coolant effect at all surfaces, while no impact of purge flow is observed. Aligning the hot streaks with the vane leading edge lowered heat transfer compared to mid-passage alignment at the mid-span suction surface and through the endwall passage, and increased it at the endwall exit, while the pressure surface is found to be insensitive to this switch. Comparison with a previous research program with the un-cooled version of the vane gave good agreement on the pressure surface and at the endwall, but significantly lower heat transfer on the suction surface due to ingestion of the hot flow through the cooling holes when there is no cooling.
Flow and Heat Transfer in a Turbine Nozzle Guide Vane with Endwall Contouring
Author: Y.-L. LinFilm Cooling and End Wall Heat Transfer in Small Turbine Blade Passages
Author: Publisher:
ISBN:
Category : Heat
Languages : en
Pages : 56
Book Description
Two topics have been studied related to the cooling of the end wall of a turbine passage. The first concerns the development of a method for measuring the adiabatic wall effectiveness and heat transfer coefficient of a film cooling system for protecting a surface from high heating derived from a hot compressible flow. The second concerns the measurement of the heat transfer rate distribution to a turbine cascade end wall in order to choose an appropriate film cooling system. These are related to providing the background to the final phase of the study in which the effectiveness of a film cooling system to cool a turbine end wall will be made combined with the measurement of the aerodynamic losses incurred by such a system. (Author).
Discharge Coefficients of Nozzle Guide Vane Film Cooling Holes
Author: Dave RowburyAxial Turbine Aerodynamics for Aero-engines
Author: Zhengping ZouPublisher: Springer
ISBN: 9811057508
Category : Technology & Engineering
Languages : en
Pages : 572
Book Description
This book is a monograph on aerodynamics of aero-engine gas turbines focusing on the new progresses on flow mechanism and design methods in the recent 20 years. Starting with basic principles in aerodynamics and thermodynamics, this book systematically expounds the recent research on mechanisms of flows in axial gas turbines, including high pressure and low pressure turbines, inter-turbine ducts and turbine rear frame ducts, and introduces the classical and innovative numerical evaluation methods in different dimensions. This book also summarizes the latest research achievements in the field of gas turbine aerodynamic design and flow control, and the multidisciplinary conjugate problems involved with gas turbines. This book should be helpful for scientific and technical staffs, college teachers, graduate students, and senior college students, who are involved in research and design of gas turbines.
Aerodynamic Loss Reduction in a Vane Cascade with Leading Edge Fillet and Upstream Endwall Film Cooling
Author: Keenesh ArnachellanPublisher:
ISBN:
Category : Turbines
Languages : en
Pages : 228
Book Description
Secondary flow structures account for nearly 50% of aerodynamic losses experienced in the turbine blade passages. The adverse effects of these vortex structures transport the hot mainstream fluid towards the endwall blade surfaces, which enhances thermal stresses and leads to blade failure. The effects of leading-edge fillets and film-cooling with flush slots located upstream near the leading-edge region were investigated experimentally in the study in a large-scale linear vane cascade in which the aerodynamic flow field was considered. The introduction of slot film flow and fillet aimed to reduce the effects of the secondary flow structures from the leading edge through the passage towards the exit in an effort to decrease the pressure losses, improve film-cooling coverage and flow field uniformity for the next blade row. The two-dimensional vane profile was obtained from the hub-side airfoil of the GE-E3 engine nozzle guide vane. The slots were configured for two experimental cases to evaluate the influence of coolant flow rate and momentum; first, the effects of slot film injection from all four slots were observed and then compared with the second case injecting coolant only through the two central slots. Further effects were investigated by combining slot film-cooling with the leading-edge fillets employed on the endwall blade junction. The flow field measurements were quantified with spatial distributions of axial vorticity, total pressure loss, endwall static pressure and flow angle deviations taken across the cascade passage. The measurements were obtained at a Reynolds number of 2.0E+05 based on the cascade inlet velocity and vane chord length. Film-cooling inlet blowing ratios between 1.1 and 2.3 were investigated with the supply of coolant provided by a secondary channel. Film-cooling results were compared with the baseline case without slot film flow and fillet. The results indicated substantial improvement in the passage and exit planes with high inlet blowing ratios. The introduction of high momentum coolant flow from the central slots was seen to create laterally reversed axial vorticity, thereby counteracting the cross-flow tendency in the passage. The effects at the passage exit showed suppressed vortex structures with slot film injection from the two central slots only, with further improvements in the flow angle deviations. The leading-edge slots were seen to contribute positive axial vorticity, which enhanced the passage vortex that was pushed away from the endwall at the exit. When the fillet was introduced, it had favourable effects in reducing the pitchwise pressure gradients along the endwall. Filleted film-cooling then resulted in a faint passage vortex system (50-80% size and 20-50% strength reduction) with a restored endwall boundary layer at high film flow rates. The leading-edge fillet was highly effective at the inlet of the blade passage because it weakened the horseshoe vortex formation. Thus, upstream slot film-cooling has great potential to decrease the aerodynamic losses and is further compounded with the leading-edge fillet.
Film Injection
Author: Michel Robert AyachePublisher:
ISBN:
Category :
Languages : en
Pages : 382
Book Description