Aero-Thermodynamic Aspects of Film Cooling in Regions of Separated Flow on the Pressure Side of a High-Lift HPT Blade

2004 ◽  
Author(s):  
Lars Homeier ◽  
Ewald Lutum ◽  
Erik Janke ◽  
Frank Haselbach
Keyword(s):  
Film Cooling ◽  
High Speed ◽  
Turbine Blades ◽  
Turbulence Models ◽  
Side Surface ◽  
Pressure Side ◽  
High Lift ◽  
Numerical Investigations ◽  
Coolant Injection ◽  
Highly Loaded

Due to the ever increasing demand for cost-optimised designs, modern engine design concepts lead to more and more highly loaded HP turbine blades. In order to achieve the high lift required, turbine airfoils will have to cope with main flow diffusion up to separation both on suction and pressure side. Thus, for film cooled HP turbine blades and vanes, the possible aerodynamic and aero-thermal interaction of highly loaded blade rows and film cooling needs to be addressed. The first results to be presented from this ongoing work within the European 5th Frame-Work-Project AITEB jointly comprises experimental high-speed cascade wind-tunnel as well as numerical investigations with state-of-the-art 3D-RANS CFD. Steady and unsteady experimental results detailing the row characteristic of the highly-loaded T120 HP-turbine cascade set the stage for detailed numerical investigations with and without coolant injection from rows of holes on the pressure side surface as well as comparative numerical calculations with different codes and turbulence models. Despite the current focus of the experimental work on aerodynamic topics, the numerical results to be presented comprise thermodynamic investigations and detailed studies on optimised coolant injection geometries as well.

Effects of Periodic Unsteady Inflow on Film Cooling and Heat Transfer on Highly Loaded High Pressure Turbine Blades With Flow Separation

2012 ◽  
Author(s):  
Reinaldo A. Gomes ◽  
Reinhard Niehuis
Keyword(s):  
Heat Transfer ◽  
Flow Separation ◽  
Film Cooling ◽  
High Speed ◽  
Turbine Blades ◽  
Suction Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Unsteady Inflow ◽  
Highly Loaded

Film cooling experiments were run at the High-Speed Cascade Wind Tunnel of the University of the Federal Armed Forces Munich. The investigations were carried out on a linear cascade of highly loaded turbine blades. The main targets of the tests were to assess the film cooling effectiveness and the heat transfer in zones with main flow separation. The previous cascade was designed to have a large zone with flow separation on the pressure side starting at the leading edge and reaching up to approximately half of the axial chord. This cascade was changed for a new design with a larger pitch to chord ratio in order to set the focus on flow separation on the suction side. This increased pitch forces a massive separation on the suction side due to strong shocks. The flow separation is controlled with aid of vortex generating jets in order to reduce the total pressure loss caused by it. Film cooling is provided on the suction side upstream of the vortex generating jets. The measurements comprise of blade loading, profile loss, adiabatic film cooling effectiveness and heat transfer coefficient under two Mach numbers at a Reynolds number of 390,000. In a previous publication detailed results with homogeneous inflow where shown. Now, the focus is set on the effects of periodic unsteady wakes resulting from bars moving upstream of the cascade. These moving bars create a periodic unsteady inflow similar to the interaction between stator and rotor in the machine. It is shown how these wakes have significant influence on the heat transfer in the acceleration region of the suction side and affect the adiabatic film cooling effectiveness upstream of the shock.

Effects of Periodic Unsteady Inflow on Film Cooling and Heat Transfer on Highly Loaded High Pressure Turbine Blades With Flow Separation

2013 ◽  
Vol 136 (2) ◽  
Author(s):  
Reinaldo A. Gomes ◽  
Reinhard Niehuis
Keyword(s):  
Heat Transfer ◽  
Flow Separation ◽  
Film Cooling ◽  
High Speed ◽  
Turbine Blades ◽  
Suction Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Unsteady Inflow ◽  
Highly Loaded

Film cooling experiments were run at the high-speed cascade wind tunnel of the University of the Federal Armed Forces Munich. The investigations were carried out on a linear cascade of highly loaded turbine blades. The main targets of the tests were to assess the film cooling effectiveness and the heat transfer in zones with main flow separation. The previous cascade was designed to have a large zone with flow separation on the pressure side starting at the leading edge and reaching up to approximately half of the axial chord. This cascade was changed for a new design with a larger pitch to chord ratio in order to set the focus on flow separation on the suction side. This increased pitch forces a massive separation on the suction side due to strong shocks. The flow separation is controlled with aid of vortex generating jets in order to reduce the total pressure loss caused by it. Film cooling is provided on the suction side upstream of the vortex generating jets. The measurements comprise of blade loading, profile loss, adiabatic film cooling effectiveness, and heat transfer coefficient under two Mach numbers at a Reynolds number of 390,000. In a previous publication detailed results with homogeneous inflow where shown. Now, the focus is set on the effects of periodic unsteady wakes resulting from bars moving upstream of the cascade. These moving bars create a periodic unsteady inflow similar to the interaction between stator and rotor in the machine. It is shown how these wakes have significant influence on the heat transfer in the acceleration region of the suction side and affect the adiabatic film cooling effectiveness upstream of the shock.

scholarly journals Comparison of Two-Equation Turbulence Models for Prediction of Heat Transfer on Film-Cooled Turbine Blades

1997 ◽  
Author(s):  
Vijay K. Garg ◽  
Ali A. Ameri
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Turbine Blades ◽  
Three Dimensional ◽  
Turbulence Models ◽  
Computer Time ◽  
Suction Surface ◽  
The Heat Transfer Coefficient

A three-dimensional Navier-Stokes code has been used to compute the heat transfer coefficient on two film-cooled turbine blades, namely the VKI rotor with six rows of cooling holes including three rows on the shower head, and the C3X vane with nine rows of holes including five rows on the shower head. Predictions of heat transfer coefficient at the blade surface using three two-equation turbulence models, specifically, Coakley’s q-ω model, Chien’s k-ε model and Wilcox’s k-ω model with Menter’s modifications, have been compared with the experimental data of Camci and Arts (1990) for the VKI rotor, and of Hylton et al. (1988) for the C3X vane along with predictions using the Baldwin-Lomax (B-L) model taken from Garg and Gaugler (1995). It is found that for the cases considered here the two-equation models predict the blade heat transfer somewhat better than the B-L model except immediately downstream of the film-cooling holes on the suction surface of the VKI rotor, and over most of the suction surface of the C3X vane. However, all two-equation models require 40% more computer core than the B-L model for solution, and while the q-ω and k-ε models need 40% more computer time than the B-L model, the k-ω model requires at least 65% more time due to slower rate of convergence. It is found that the heat transfer coefficient exhibits a strong spanwise as well as streamwise variation for both blades and all turbulence models.

Turbine Nozzle Endwall Phantom Cooling With Compound Angled Pressure Side Injection

2018 ◽  
Author(s):  
Kevin Liu ◽  
Hongzhou Xu ◽  
Michael Fox
Keyword(s):  
Film Cooling ◽  
High Speed ◽  
Axial Direction ◽  
Secondary Flows ◽  
Test Case ◽  
Pressure Side ◽  
Complex Flow ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Extended Area

Cooling of the turbine nozzle endwall is challenging due to its complex flow field involving strong secondary flows. Increasingly-effective cooling schemes are required to meet the higher turbine inlet temperatures required by today’s gas turbine applications. Therefore, in order to cool the endwall surface near the pressure side of the airfoil and the trailing edge extended area, the spent cooling air from the airfoil film cooling and pressure side discharge slots, referred to as “phantom cooling” is utilized. This paper studies the effect of compound angled pressure side injection on nozzle endwall surface. The measurements were conducted in a high speed linear cascade, which consists of three nozzle vanes and four flow passages. Two nozzle test models with a similar film cooling design were investigated, one with an axial pressure side film cooling row and trailing edge slots; the other with the same cooling features but with compound angled injection, aiming at the test endwall. Phantom cooling effectiveness on the endwall was measured using a Pressure Sensitive Paint (PSP) technique through the mass transfer analogy. Two-dimensional phantom cooling effectiveness distributions on the endwall surface are presented for four MFR (Mass Flow Ratio) values in each test case. Then the phantom cooling effectiveness distributions are pitchwise-averaged along the axial direction and comparisons were made to show the effect of the compound angled injection. The results indicated that the endwall phantom cooling effectiveness increases with the MFR significantly. A compound angle of the pressure side slots also enhanced the endwall phantom cooling significantly. For combined injections, the phantom cooling effectiveness is much higher than the pressure side slots injection only in the endwall downstream extended area.

CFD Investigation of the Flow of Trailing Edge Cooling Slots

2018 ◽  
Author(s):  
Y. Jiang ◽  
N. Gurram ◽  
E. Romero ◽  
P. T. Ireland ◽  
L. di Mare
Keyword(s):  
Mach Number ◽  
Kinetic Energy ◽  
Flow Separation ◽  
Film Cooling ◽  
High Speed ◽  
Turbine Blades ◽  
Cross Flow ◽  
Turbulent Energy ◽  
Trailing Edge ◽  
Flow Interactions

Slot film cooling is a popular choice for trailing edge cooling in high pressure (HP) turbine blades because it can provide more uniform film coverage compared to discrete film cooling holes. The slot geometry consists of a cut back in the blade pressure side connected through rectangular openings to the internal coolant feed passage. The numerical simulation of this kind of film cooling flows is challenging due to the presence of flow interactions like step flow separation, coolant-mainstream mixing and heat transfer. The geometry under consideration is a cutback surface at the trailing edge of a constant cross-section aerofoil. The cutback surface is divided into three sections separated by narrow lands. The experiments are conducted in a high speed cascade in Oxford Osney Thermo-Fluids Laboratory at Reynolds and Mach number distributions representative of engine conditions. The capability of CFD methods to capture these flow phenomena is investigated in this paper. The isentropic Mach number and film effectiveness are compared between CFD and pressure sensitive paint (PSP) data. Compared to steady k–ω SST method, Scale Adaptive Simulation (SAS) can agree better with the measurement. Furthermore, the profiles of kinetic energy, production and shear stress obtained by the steady and SAS methods are compared to identify the main source of inaccuracy in RANS simulations. The SAS method is better to capture the unsteady coolant-hot gas mixing and vortex shedding at the slot lip. The cross flow is found to affect the film significantly as it triggers flow separation near the lands and reduces the effectiveness. The film is non-symmetric with respect to the half-span plane and different flow features are present in each slot. The effect of mass flow ratio (MFR) on flow pattern and coolant distribution is also studied. The profiles of velocity, kinetic energy and production of turbulent energy are compared among the slots in detail. The MFR not only affects the magnitude but also changes the sign of production.

Cooling of Turbine Blades With Expanded Exit Holes: Computational Analyses of Leading Edge and Pressure-Side of a Turbine Blade

2017 ◽  
Vol 139 (4) ◽  
Author(s):  
Fariborz Forghan ◽  
Omid Askari ◽  
Uichiro Narusawa ◽  
Hameed Metghalchi
Keyword(s):  
High Speed ◽  
Turbine Blades ◽  
Leading Edge ◽  
Weak Shock ◽  
Suction Side ◽  
Pressure Side ◽  
Blade Surface ◽  
Cooling Effectiveness ◽  
Hot Gas ◽  
Computational Analyses

Turbine blades are cooled by a jet flow from expanded exit holes (EEH) forming a low-temperature film over the blade surface. Subsequent to our report on the suction-side (low-pressure, high-speed region), computational analyses are performed to examine the cooling effectiveness of the flow from EEH located at the leading edge as well as at the pressure-side (high-pressure, low-speed region). Unlike the case of the suction-side, the flow through EEH on the pressure-side is either subsonic or transonic with a weak shock front. The cooling effectiveness, η (defined as the temperature difference between the hot gas and the blade surface as a fraction of that between the hot gas and the cooling jet), is higher than the suction-side along the surface near the exit of EEH. However, its magnitude declines sharply with an increase in the distance from EEH. Significant effects on the magnitude of η are observed and discussed in detail of (1) the coolant mass flow rate (0.001, 0.002, and 0.004 (kg/s)), (2) EEH configurations at the leading edge (vertical EEH at the stagnation point, 50 deg into the leading-edge suction-side, and 50 deg into the leading-edge pressure-side), (3) EEH configurations in the midregion of the pressure-side (90 deg (perpendicular to the mainstream flow), 30 deg EEH tilt toward upstream, and 30 deg tilt toward downstream), and (4) the inclination angle of EEH.

Experimental Study on Film Cooling Flow Characteristics of Turbine Stator Blade at Different Setting Angles

2012 ◽  
Vol 614-615 ◽  
pp. 592-595
Author(s):  
Ling Zhang ◽  
Hai Rui Dong ◽  
Guo Liang Wen
Keyword(s):  
Film Cooling ◽  
Effective Means ◽  
Turbine Blades ◽  
Flow Characteristics ◽  
Suction Side ◽  
Pressure Side ◽  
Blowing Ratio ◽  
Turbine Stator ◽  
Setting Angle ◽  
Stator Blade

The technology of film cooling is one of the most effective means of protecting the turbine blades. In this paper, flow structures of the turbine stator blade with six hole-rows at different blowing ratio（M=0.5, 1.0 and 1.5）and setting angles（β=40°, 50°, 60°, 70°, 80° and 90°） was measured by PIV in piston flow type of low-speed wind tunnel laboratory. Velocity was analyzed. Results show that: velocity gradient of suction side was much higher than pressure side and increased with setting angle reduction; Adherence of film is influenced by setting angle and blowing ratio, when M=1.0 and β=70° anchorage dependent is best and suction side is greater than pressure side.

Computational Study of Jet-in-Crossflow and Film Cooling Using a New Unsteady-Based Turbulence Model

2005 ◽  
Author(s):  
D. Scott Holloway ◽  
D. Keith Walters ◽  
James H. Leylek
Keyword(s):  
Turbulence Model ◽  
Film Cooling ◽  
Computational Study ◽  
Velocity Ratio ◽  
Density Ratio ◽  
Turbulence Models ◽  
Streamwise Velocity ◽  
Pressure Side ◽  
Good Test ◽  
Jet In Crossflow

This paper documents a computational investigation of the unsteady behavior of jet-in-crossflow applications. Improved prediction of fundamental physics is achieved by implementing a new unsteady, RANS-based turbulence model developed by the authors. Two test cases are examined that match experimental efforts previously documented in the open literature. One is the well-documented normal jet-in-crossflow, and the other is film cooling on the pressure side of a turbine blade. All simulations are three-dimensional, fully converged, and grid-independent. High-quality and high-density grids are constructed using multiple topologies and an unstructured, super-block approach to ensure that numerical viscosity is minimized. Computational domains include the passage, film hole, and coolant supply plenum. Results for the normal jet-in-crossflow are for a density ratio of 1 and velocity ratio of 0.5 and include streamwise velocity profiles and injected flow or “coolant” distribution. The Reynolds number based on the average jet exit velocity and jet diameter is 20,500. This represents a good test case since normal injection is known to exaggerate the key flow mechanisms seen in film-cooling applications. Results for the pressure side film-cooling case include coolant distribution and adiabatic effectiveness for a density and blowing ratio of 2. In addition to the in-house model that incorporates new unsteady physics, CFD simulations utilize standard, RANS-based turbulence models, such as the “realizable” k-ε model. The present study demonstrates the importance of unsteady physics in the prediction of jet-in-crossflow interactions and for film cooling flows that exhibit jet liftoff.

3D Heat Transfer Assessment of Full-Scale Inlet Vanes With Surface-Optimized Film Cooling: Part 2 — Conjugate CFD Simulations

2019 ◽  
Author(s):  
J. J. Johnson ◽  
J. P. Clark ◽  
R. A. Anthony ◽  
M. K. Ooten ◽  
R. H. Ni ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Side Surface ◽  
Full Scale ◽  
Pressure Side ◽  
Computationally Efficient ◽  
Cfd Simulations ◽  
Experimental Heat ◽  
Cooling Hole ◽  
Cooling Design

Abstract An investigation of the experimental heat transfer and cooling effectiveness for a modern fully-cooled high-pressure turbine (HPT) inlet vane is presented. Conjugate Heat Transfer (CHT) Computational Fluid Dynamics (CFD) is conducted to simulate experiments using thin-film heat-flux gauges on full-scale 3D vanes at engine-representative conditions from Part 1 of this paper. Pressure side (PS) film cooling performance is compared for a baseline and optimized configuration, in which the latter was previously developed using genetic algorithm (GA) optimization. The optimized vane was iterated using hundreds of computationally efficient 3D Reynolds Averaged Navier Stokes (RANS) CFD simulations with a transpiration boundary condition to simulate film cooling. This combination of CFD and GAs determined surface-optimized cooling hole orientations and placement. Steady-state flat plate infrared thermography experiments that followed also determined the best cooling hole shapes to use on different sections of the vane pressure side surface. This ultimately generated the cooling design to be fabricated using realistic materials and experimentally tested in Part 1 and simulated using CHT CFD in the current work (Part 2). Here, spanwise and streamwise heat transfer distributions for the baseline and optimized cooling design are validated against experimental data. 3D CHT CFD results are then assessed at the same conditions, providing relevance and credence to the overall cooling design methods. Ultimately, surface-optimized film cooling designs can be used to reduce the adverse effects of sub-optimal heat distribution on critical high temperature engine parts, increasing the life of the part. Alternatively, such a design could lead to increases in engine efficiency since less cooling air is required from the mainstream per part.

Overall Cooling Effectiveness Measurements on Pressure Side Surface of the Nozzle Guide Vane With Optimized Film Cooling Hole Arrangements

2017 ◽  
Author(s):  
Dong-Ho Rhee ◽  
Young Seok Kang ◽  
Bong Jun Cha ◽  
Sanga Lee
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Side Surface ◽  
Internal Cooling ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Nozzle Guide Vane ◽  
Rib Turbulators ◽  
Nozzle Guide

Most of the optimization researches on film cooling have dealt with adiabatic film cooling effectiveness on the surface. However, the information on the overall cooling effectiveness is required to estimate exact performance of the optimization configuration since hot components such as nozzle guide vane have not only film cooling but also internal cooling features such as rib turbulators, jet impingement and pin-fins on the inner surface. Our previous studies [1,2] conducted the hole arrangement optimization to improve adiabatic film cooling effectiveness values and uniformity on the pressure side surface of the nozzle guide vane. In this study, the overall cooling effectiveness values were obtained at various cooling mass flow rates experimentally for the baseline and the optimized hole arrangements proposed by the previous study [1] and compared with the adiabatic film cooling effectiveness results. The tests were conducted at mainstream exit Reynolds number based on the chord of 2.2 × 106 and the coolant mass flow rate from 5 to 10% of the mainstream. For the experimental measurements, a set of tests were conducted using an annular sector transonic turbine cascade test facility in Korea Aerospace Research Institute. To obtain the overall cooling effectiveness values on the pressure side surface, the additive manufactured nozzle guide vane made of polymer material and Inconel 718 were installed and the surface temperature was measured using a FLIR infrared camera system. Since the optimization was based on the adiabatic film cooling effectiveness, the regions with rib turbulators and film cooling holes show locally higher overall cooling effectiveness due to internal convection and conduction, which can cause non-uniform temperature distributions. Therefore, the optimization of film cooling configuration should consider the effect of the internal cooling to avoid undesirable non-uniform cooling.

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