An Experimental Investigation on the Influence of Periodic Unsteady Inflow Conditions on Leading Edge Film Cooling

2002 ◽  
Author(s):  
Stefan Wolff ◽  
Leonhard Fottner ◽  
Sabine Ardey
Keyword(s):  
Film Cooling ◽  
Gas Turbines ◽  
Flow Behavior ◽  
Density Ratio ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Injection Zone ◽  
Unsteady Inflow ◽  
Inflow Conditions

In order to close the gap in knowledge concerning the influence of periodic unsteady inflow conditions on the mixing process of film cooling jets, time resolved flow velocities and turbulent fluctuations in the injection zone of a linear large scale high pressure turbine cascade with leading edge film cooling were measured by means of the 3D hot-wire anemometry. The periodic impinging wakes are generated by a wake generator consisting of moving bars upstream of the cascade inlet plane. Near the stagnation point the blades are equipped with one row of film cooling holes on the suction side and one row on the pressure side. Mach number and Reynolds number are set to values typically found in modern gas turbines. At a density ratio of unity the blowing ratio is set to M = 0.7 and the Strouhal number is set to Sr = 0.31. The general flow structures which were determined by the steady state measurements — i.e. normal jet in cross flow behavior with the kidney shape vortex on the suction side and a second counter rotating pair of vortices underneath the kidney vortex on the pressure side — have been detected by the unsteady measurements as well. The large recirculation zone behind the pressure side injection hole, caused by the strong adverse pressure gradient and the lift off of the coolant jet, is not suppressed by the passing wakes but rather reinforced promoting potential hot spots in this area. The suction side coolant jet has almost disappeared when the wake leaves the suction side region. It undergoes a recovery process until the next wake hits again on the suction side.

Turbulence Structures of Leading Edge Film Cooling Jets

2000 ◽  
Author(s):  
Sabine Ardey ◽  
Stefan Wolff ◽  
Leonhard Fottner
Keyword(s):  
Film Cooling ◽  
Large Scale ◽  
Leading Edge ◽  
Adverse Pressure Gradient ◽  
Suction Side ◽  
Pressure Side ◽  
Reynolds Stresses ◽  
Mean Flow ◽  
Turbulence Structures ◽  
Injection Zone

For a better understanding of the turbulence structures attached to film cooling jets, mean flow velocities and turbulent fluctuations were measured by means of 3D hot wire anemometry in the injection zone of a linear, large scale, high pressure turbine cascade with leading edge film cooling. Near the stagnation point, the blades are equipped with one row of film cooling holes each on the suction and pressure side. Two basically different coolant jet situations are compared: On the suction side the jet features the ordinary kidney vortex. On the pressure side, the jet separates completely from the blade surface since it is located above a large recirculation zone created by a locally adverse pressure gradient and a flow separation near the pressure side injection. The measured Reynolds stresses were analyzed with regard to turbulence production and diffusion. The Bousinesque Hypothesis was tested and could not be confirmed. It was found that the turbulence is highly anisotropic. In addition to the brief description of the experimental set up and the acquired data, given in this paper, the complete information are published as a test case (Ardey and Fottner, 1998) that is directly accessible via internet at: http://www.unibw-muenchen.de/campus/LRT12/welcome.html

Film Cooling Jet Injection Effect in Heat Transfer Coefficient Augmentation for the Pressure Side Cooling of Turbine Vane

2014 ◽  
Author(s):  
Hossein Nadali Najafabadi ◽  
Matts Karlsson ◽  
Mats Kinell ◽  
Esa Utriainen
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Time Resolved ◽  
Turbine Vane

Improving film cooling performance of turbine vanes and blades is often achieved through application of multiple arrays of cooling holes on the suction side, the showerhead region and the pressure side. This study investigates the pressure side cooling under the influence of single and multiple rows of cooling in the presence of a showerhead from a heat transfer coefficient augmentation perspective. Experiments are conducted on a prototype turbine vane working at engine representative conditions. Transient IR thermography is used to measure time-resolved surface temperature and the semi-infinite method is utilized to calculate the heat transfer coefficient on a low conductive material. Investigations are performed for cylindrical and fan-shaped holes covering blowing ratio 0.6 and 1.8 at density ratio of about unity. The freestream turbulence is approximately 5% close to the leading edge. The resulting heat transfer coefficient enhancement, the ratio of HTC with to that without film cooling, from different case scenarios have been compared to showerhead cooling only. Findings of the study highlight the importance of showerhead cooling to be used with additional row of cooling on the pressure side in order to reduce heat transfer coefficient enhancement. In addition, it is shown that extra rows of cooling will not significantly influence heat transfer augmentation, regardless of the cooling hole shape.

Numerical Investigation on Film Cooling Effectiveness of Stator Blade with Different Density Ratio

2014 ◽  
Vol 521 ◽  
pp. 104-107
Author(s):  
Ling Zhang ◽  
Quan Heng Jin ◽  
Da Fei Guo
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Middle Section ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Stator Blade

The Realizable k-ε turbulence model was performed to investigate the film cooling effectiveness with different blowing ratio 1,1.5,2 and different density ratio 1,1.5,2.The results show that, cooling effectiveness increases with the augment of blowing ratio. On the pressure side, cooling effectiveness increases with the augment of density ratio. On the suction side, with higher density ratio the leading edge cooling increases, the middle section reduces, and the trailing edge cooling effectiveness increases first decreases.

Flow Field Measurements on a Large Scale Turbine Cascade With Leading Edge Film Cooling by Two Rows of Holes

1997 ◽  
Author(s):  
Sabine Ardey ◽  
Leonhard Fottner
Keyword(s):  
Flow Field ◽  
Film Cooling ◽  
Gas Turbines ◽  
Large Scale ◽  
Vortex Pair ◽  
Leading Edge ◽  
Field Measurements ◽  
Suction Side ◽  
Pressure Side ◽  
Turbine Cascade

To increase the understanding of the aerodynamic processes dominating the flow field of turbine bladings with leading edge film cooling, isothermal investigations were carried out on a large scale high pressure turbine cascade. Near the stagnation point the blades are equipped with one row of film cooling holes on the suction side and one on the pressure side. Blowing ratio, turbulence intensity, Mach number, and Reynolds number are set to values typically found in modern gas turbines. Experimental data of the cascade flow were obtained by pneumatic probes and static pressure tappings. The flow field was visualized by Schlieren and oil flow techniques. For detailed investigations near the blowing holes the Laser Transit Velocimetry and the three dimensional Hot Wire Anemometry were used. The flow field measurements in the near hole region of the suction side show the typical kidney shaped vortex pair. A local suction peak on the pressure side causes a large recirculation area behind the holes on the pressure side and induces separation bubbles in between the pressure side holes. This leads to the generation of two pairs of vortices: The kidney-vortex is located on top of a second vortex pair and a trough flow that fills up the deficit of the recirculation. Thus the film cooling air is detached from the pressure side surface. In addition to the mean flow vectors Reynolds stress components are a good means to judge the propagation of the jet. In spite of the complex flow pattern occurring on each single jet, the surveyed loss-increase due to the leading edge blowing can be predicted by the mixing layer model.

scholarly journals Aerothermal Investigations of Mixing Flow Phenomena in Case of Radially Inclined Ejection Holes at the Leading Edge

1999 ◽  
Author(s):  
Dieter E. Bohn ◽  
Karsten A. Kusterer
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Gas Turbines ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Blade Surface ◽  
Cooling Fluid ◽  
Flow Phenomena ◽  
Vortex Systems

A leading edge cooling configuration is investigated numerically by application of a 3-D conjugate fluid flow and heat transfer solver, CHT-Flow. The code has been developed at the Institute of Steam and Gas Turbines, Aachen University of Technology. It works on the basis of an implicit finite volume method combined with a multi-block technique. The cooling configuration is an axial turbine blade cascade with leading edge ejection through two rows of cooling holes. The rows are located in the vicinity of the stagnation line, one row is on the suction side, the other row is on the pressure side. The cooling holes have a radial ejection angle of 45°. This configuration has been investigated experimentally by other authors and the results have been documented as a test case for numerical calculations of ejection flow phenomena. The numerical domain includes the internal cooling fluid supply, the radially inclined holes and the complete external flow field of the turbine vane in a high resolution grid. Periodic boundary conditions have been used in the radial direction. Thus, end wall effects have been excluded. The numerical investigations focus on the aerothermal mixing process in the cooling jets and the impact on the temperature distribution on the blade surface. The radial ejection angles lead to a fully three dimensional and asymmetric jet flow field. Within a secondary flow analysis it can be shown that complex vortex systems are formed in the ejection holes and in the cooling fluid jets. The secondary flow fields include asymmetric kidney vortex systems with one dominating vortex on the back side of the jets. The numerical and experimental data show a good agreement concerning the vortex development. The phenomena on the suction side and the pressure side are principally the same. It can be found that the jets are barely touching the blade surface as the dominating vortex transports hot gas under the jets. Thus, the cooling efficiency is reduced.

Numerical Investigation of Coolant-to-Mainstream Scaling Parameters With Film Cooling on Pressure and Suction Side of a Gas Turbine Blade

2017 ◽  
Author(s):  
Lingyu Zeng ◽  
Xueying Li ◽  
Jing Ren ◽  
Hongde Jiang
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Momentum Flux ◽  
Velocity Ratio ◽  
Density Ratio ◽  
Flux Ratio ◽  
Suction Side ◽  
Pressure Side ◽  
Blowing Ratio ◽  
Adiabatic Effectiveness

Most experiments of blade film cooling are conducted with density ratio lower than that of turbine conditions. In order to accurately model the performance of film cooling under a high density ratio, choosing an appropriate coolant to mainstream scaling parameter is necessary. The effect of density ratio on film cooling effectiveness on the surface of a gas turbine twisted blade is investigated from a numerical point of view. One row of film holes are arranged in the pressure side and two rows in the suction side. All the film holes are cylindrical holes with a pitch to diameter ratio P/d = 8.4. The inclined angle is 30°on the pressure side and 34° on the suction side. The steady solutions are obtained by solving Reynolds-Averaged-Navier-Stokes equations with a finite volume method. The SST turbulence model coupled with γ-θ transition model is applied for the present simulations. A film cooling experiment of a turbine vane was done to validate the turbulence model. Four different density ratios (DR) from 0.97 to 2.5 are studied. To independently vary the blowing ratio (M), momentum flux ratio (I) and velocity ratio (VR) of the coolant to the mainstream, seven conditions (M varying from 0.25 to 1.6 on the pressure side and from 0.25 to 1.4 on the suction side) are simulated for each density ratio. The results indicate that the adiabatic effectiveness increases with the increase of density ratio for a certain blowing ratio or a certain momentum flux ratio. Both on the pressure side and suction side, none of the three parameters listed above can serve as a scaling parameter independent of density ratio in the full range. The velocity ratio provides a relative better collapse of the adiabatic effectiveness than M and I for larger VRs. A new parameter describing the performance of film cooling is introduced. The new parameter is found to be scaled with VR for nearly the whole range.

Evaluation of Superposition Predictions for Showerhead Film Cooling on a Vane

2014 ◽  
Author(s):  
Joshua B. Anderson ◽  
James R. Winka ◽  
David G. Bogard ◽  
Michael E. Crawford
Keyword(s):  
Film Cooling ◽  
Momentum Flux ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Edge Region ◽  
Current Investigation ◽  
Engine Design ◽  
Turbine Vane ◽  
Adiabatic Effectiveness

The leading edge of a turbine vane is subject to some of the highest temperature loading within an engine, and an accurate understanding of leading edge film coolant behavior is essential for modern engine design. Although there have been many investigations of the adiabatic effectiveness for showerhead film cooling of a vane leading edge region, there have been no previous studies in which individual rows of the showerhead were tested with the explicit intent of validating superposition models. For the current investigation, a series of adiabatic effectiveness experiments were performed with a five-row and three-row showerhead. The experiments were repeated separately with each individual row of holes active. This allowed evaluation of superposition methods on both the suction side of the vane, which was moderately convex, and the pressure side of the vane, which was mildly concave. Superposition was found to accurately predict performance on the suction side of the vane at lower momentum flux ratios, but not at higher momentum flux ratios. On the pressure side of the vane the superposition predictions were consistently lower than measured values, with significant errors occurring at the higher momentum flux ratios. Reasons for the under-prediction by superposition analysis are presented.

Unsteady Wake and Coolant Density Effects on Turbine Blade Film Cooling Using PSP Technique

2010 ◽  
Author(s):  
Akhilesh P. Rallabandi ◽  
Shiou-Jiuan Li ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Suction Side ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Unsteady Wake ◽  
Lift Off ◽  
Film Cooling Holes

The effect of an unsteady stator wake (simulated by wake rods mounted on a spoke wheel wake generator) on the modeled rotor blade is studied using the Pressure Sensitive Paint (PSP) mass transfer analogy method. Emphasis of the current study is on the mid-span region of the blade. The flow is in the low Mach number (incompressible) regime. The suction (convex) side has simple angled cylindrical film-cooling holes; the pressure (concave) side has compound angled cylindrical film cooling holes. The blade also has radial shower-head leading edge film cooling holes. Strouhal numbers studied range from 0 to 0.36; the exit Reynolds Number based on the axial chord is 530,000. Blowing ratios range from 0.5 to 2.0 on the suction side; 0.5 to 4.0 on the pressure side. Density ratios studied range from 1.0 to 2.5, to simulate actual engine conditions. The convex suction surface experiences film-cooling jet lift-off at higher blowing ratios, resulting in low effectiveness values. The film coolant is found to reattach downstream on the concave pressure surface, increasing effectiveness at higher blowing ratios. Results show deterioration in film cooling effectiveness due to increased local turbulence caused by the unsteady wake, especially on the suction side. Results also show a monotonic increase in film-cooling effectiveness on increasing the coolant to mainstream density ratio.

An Experimental Investigation of Flow Characteristics Downstream of Discrete Film Cooling Holes on Turbine Blade Leading Edge

2011 ◽  
Vol 383-390 ◽  
pp. 5553-5560
Author(s):  
Shao Hua Li ◽  
Hong Wei Qu ◽  
Mei Li Wang ◽  
Ting Ting Guo
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Flow Characteristics ◽  
Leading Edge ◽  
Suction Side ◽  
Wall Region ◽  
Pressure Side ◽  
Blowing Ratio ◽  
Curvature Gradient ◽  
Film Cooling Holes

The gas turbine blade was studied on the condition that the mainstream velocity was 10m/s and the Renolds number based on the chord length of the blade was 160000.The Hot-film anemometer was used to measure the two-dimension speed distribution along the downstream of the film cooling holes on the suction side and the pressure side. The conclusions are as follows: When the blowing ratio of the suction side and the pressure side increasing, the the mainstream and the jet injection mixing center raising. Entrainment flow occurs at the position where the blade surface with great curvature gradient, simultaneously the mixing flow has a wicked adhere to the wall. The velocity gradient of the u direction that on the suction side increase obviously, also the level of the wall adherence is better than the pressure side. With the x/d increasing, the velocity u that on the pressure side gradually become irregularly, also the secondary flow emerged near the wall region where the curvature is great. The blowing ratio on the suction side has a little influence on velocity v than that on the pressure side.

A New Test Facility to Investigate Film Cooling on a Non-Axisymmetric Contoured Turbine Endwall: Part I — Introduction and Aerodynamic Measurements

2015 ◽  
Author(s):  
Franz Puetz ◽  
Johannes Kneer ◽  
Achmed Schulz ◽  
Hans-Joerg Bauer
Keyword(s):  
Heat Transfer ◽  
Flow Field ◽  
Pressure Distribution ◽  
Film Cooling ◽  
Gas Turbines ◽  
Guide Vane ◽  
Leading Edge ◽  
Suction Side ◽  
Test Facility ◽  
Endwall Contouring

An increased demand for lower emission of stationary gas turbines as well as civil aircraft engines has led to new, low emission combustor designs with less liner cooling and a flattened temperature profile at the outlet. As a consequence, the heat load on the endwall of the first nozzle guide vane is increased. The secondary flow field dominates the endwall heat transfer, which also contributes to aerodynamic losses. A promising approach to reduce these losses is non-axisymmetric endwall contouring. The effects of non-axisymmetric endwall contouring on heat transfer and film cooling are yet to be investigated. Therefore, a new cascade test rig has been set up in order to investigate endwall heat transfer and film cooling on both a flat and a non-axisymmetric contoured endwall. Aerodynamic measurements that have been made prior to the upcoming heat transfer investigation are shown. Periodicity and detailed vane Mach number distributions ranging from 0 to 50% span together with the static pressure distribution on the endwall give detailed information about the aerodynamic behavior and influence of the endwall contouring. The aerodynamic study is backed by an oil paint study, which reveals qualitative information on the effect of the contouring on the endwall flow field. Results show that the contouring has a pronounced effect on vane and endwall pressure distribution and on the endwall flow field. The local increase and decrease of velocity and the reduced blade loading towards the endwall is the expected behavior of the 3d contouring. So are the results of the oil paint visualization, which show a strong change of flow field in the leading edge region as well as that the contouring delays the horse shoe vortex hitting the suction side.

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