scholarly journals Turbine Nozzle Film Cooling Study Using the Pressure Sensitive Paint (PSP) Technique

1999 ◽  
Author(s):  
Luzeng Zhang ◽  
Michael Baltz ◽  
Ram Pudupatty ◽  
Michael Fox
Keyword(s):  
Reynolds Number ◽  
Mach Number ◽  
Film Cooling ◽  
High Speed ◽  
Surface Film ◽  
Pressure Sensitive Paint ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Sensitive

The use of pressure sensitive paint (PSP) to measure film cooling effectiveness on a turbine nozzle surface was demonstrated in a high speed wind tunnel. Film cooling effectiveness was measured from a single row of holes located on a turbine vane suction surface with a shaped exit. Nitrogen gas was used to simulate film cooling flow as well as a tracer gas to indicate oxygen concentration such that film effectiveness by the mass transfer analogy could be obtained. Three blowing ratios were studied for each of the five freestream conditions: a reference condition, a reduced and an increased Reynolds number condition, and a reduced and an increased Mach number condition. The freestream turbulence intensity was kept at 12.0% for all the tests. The PSP was calibrated at various temperatures and pressures to obtain better accuracy before being applied to the airfoil surface. The film effectiveness increased with blowing ratio for all the freestream conditions. The effects of secondary flow and freestream Mach number and Reynolds number on turbine nozzle suction surface film cooling are also discussed.

Unsteady Wake and Coolant Density Effects on Turbine Blade Film Cooling Using Pressure Sensitive Paint Technique

2012 ◽  
Vol 134 (8) ◽  
Author(s):  
Akhilesh P. Rallabandi ◽  
Shiou-Jiuan Li ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Suction Side ◽  
Pressure Sensitive Paint ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Sensitive ◽  
Unsteady Wake ◽  
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 midspan 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 and 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.

Study of Film Cooling Effectiveness on a Double-Walled Effusion-Cooled Turbine Blade in a High-Speed Flow Using Pressure Sensitive Paint

2019 ◽  
Author(s):  
Gladys C. Ngetich ◽  
Peter T. Ireland ◽  
Eduardo Romero
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
High Speed ◽  
High Porosity ◽  
Pressure Sensitive Paint ◽  
Blade Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Sensitive ◽  
Double Wall

Abstract A detailed analysis of film cooling performance on a double-walled effusion-cooled blade is essential for both the coolant consumption optimization and assessment of the film to offer the desired levels of the turbine blade protection. Yet there are hardly any film effectiveness studies on double-wall full-coverage film cooled turbine blades. This paper presents a detailed film cooling effectiveness study over the full surface of a double-walled effusion-cooled high-pressure turbine rotor blade using Pressure Sensitive Paint (PSP). PSP permitted a non-intrusive and conduction-errors-free means of obtaining clean and distinct local distribution of film effectiveness on the blade surface making it possible to extract valuable film cooling effectiveness performance data on the whole blade surface. Three large-scale circular pedestal double-wall blade designs with varying pedestal height, pedestal diameter and cooling hole diameter were tested in a high-speed stationary single-blade linear cascade running at engine-representative Mach and Reynolds numbers. All the blades were tested within a range of representative modern engine coolant mass flow, ṁc to mainstream, ṁg ratios; 1.6% < ṁc/ṁ∞ < 5.5%. High porosity blade exhibited a better flow distribution and was found to consistently perform the best.

scholarly journals Measurement Uncertainties Analysis in the Determination of Adiabatic Film Cooling Effectiveness by Using Pressure Sensitive Paint (PSP) Technique

2014 ◽  
Author(s):  
Blake Everett Johnson ◽  
Hui Hu
Keyword(s):  
Detailed Analysis ◽  
Film Cooling ◽  
High Speed ◽  
Sampling Rate ◽  
Pressure Sensitive Paint ◽  
Cooling Effectiveness ◽  
Total Uncertainty ◽  
Film Cooling Effectiveness ◽  
Low Speed ◽  
Pressure Sensitive

Pressure sensitive paint (PSP) is useful for measurements of wall pressure in high speed flows, but can be used in an alternative manner in low-speed flows as a gas species concentration detector. Film cooling technology studies have been greatly aided by this use of PSP through use of a mass transfer analogy to determine the adiabatic film cooling effectiveness. The PSP technique allows measurements that have high spatial resolution at high enough sampling rate that a good statistical mean can be determined rapidly. Due to the potential of this technique to deliver high quality adiabatic effectiveness measurements, a detailed analysis of its associated uncertainty is presented herein. In this study, an ambient temperature low speed wind tunnel drives air as the main flow while carbon dioxide (CO2, DR=1.5) is used as the “coolant” gas, though the experiments are done under isothermal conditions. A detailed analysis of the technique is performed here with focus on the measurement uncertainty and process uncertainty for a film cooling study using an array of five cylindrical holes spaced across the span of a flat test plate at a spacing of three diameters center-to-center. The final analysis indicates that the total uncertainty depends strongly on the local behavior of the coolant, with near-field uncertainty as high as 5% at isolated points. In the far-field, the total uncertainty is more uniform throughout the measurement domain and generally lower, at about 3%.

Film Cooling Effectiveness on a Rotating Turbine Platform Using Pressure Sensitive Paint Technique

2007 ◽  
Author(s):  
A. Suryanarayanan ◽  
B. Ozturk ◽  
M. T. Schobeiri ◽  
J. C. Han
Keyword(s):  
Film Cooling ◽  
Gas Injection ◽  
Turbine Rotor ◽  
Pressure Sensitive Paint ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Sensitive ◽  
Coolant Injection ◽  
Purge Flow ◽  
Effects Of Rotation

Film cooling effectiveness is measured on a rotating turbine blade platform for coolant injection through discrete holes using pressure sensitive paint technique (PSP). Most of the existing literatures provide information only for stationary end-walls. The effects of rotation on the platform film cooling effectiveness are not well documented. Hence, the existing 3-stage turbine research facility at TPFL, Texas A&M University was re-designed and installed to enable coolant gas injection on the 1st stage rotor platform. Two distinct coolant supply loops were incorporated into the rotor to facilitate separate feeds for upstream cooling using stator-rotor gap purge flow and downstream discrete-hole film cooling. As a continuation of the previously published work involving stator-rotor gap purge cooling, this study investigates film cooling effectiveness on the 1st stage rotor platform due to coolant gas injection through nine discrete holes located downstream within the passage region. Film cooling effectiveness is measured for turbine rotor frequencies of 2400rpm, 2550rpm and 3000rpm corresponding to rotation numbers of Ro = 0.18, 0.19 and 0.23 respectively. For each of the turbine rotational frequencies, film cooling effectiveness is determined for average film-hole blowing ratios of Mholes = 0.5, 0.75, 1.0, 1.25, 1.5 and 2.0. To provide a complete picture of hub cooling under rotating conditions, simultaneous injection of coolant gas through upstream stator-rotor purge gap and downstream discrete film-hole is also studied. The combined tests are conducted for gap purge flow corresponding to coolant to mainstream mass flow ratio of MFR = 1% with three downstream film-hole blowing ratios of Mholes = 0.75, 1.0 and 1.25 for each of the three turbine speeds. The results for combined upstream stator-rotor gap purge flow and downstream discrete holes provide information about the optimum purge flow coolant mass, average coolant hole blowing ratios for each rotational speed and coolant injection location along the passage to obtain efficient platform film cooling.

Influence of Coolant Density on Turbine Blade Film Cooling at Transonic Cascade Flow Conditions Using the Pressure Sensitive Paint Technique

2021 ◽  
Author(s):  
Izhar Ullah ◽  
Sulaiman M. Alsaleem ◽  
Lesley M. Wright ◽  
Chao-Cheng Shiau ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Pressure Side ◽  
Pressure Sensitive Paint ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Pressure Sensitive ◽  
Blade Tip ◽  
Film Cooling Holes

Abstract This work is an experimental study of film cooling effectiveness on a blade tip in a stationary, linear cascade. The cascade is mounted in a blowdown facility with controlled inlet and exit Mach numbers of 0.29 and 0.75, respectively. The free stream turbulence intensity is measured to be 13.5 % upstream of the blade’s leading edge. A flat tip design is studied, having a tip gap of 1.6%. The blade tip is designed to have 15 shaped film cooling holes along the near-tip pressure side (PS) surface. Fifteen vertical film cooling holes are placed on the tip near the pressure side. The cooling holes are divided into a 2-zone plenum to locally maintain the desired blowing ratios based on the external pressure field. Two coolant injection scenarios are considered by injecting coolant through the tip holes only and both tip and PS surface holes together. The blowing ratio (M) and density ratio (DR) effects are studied by testing at blowing ratios of 0.5, 1.0, and 1.5 and three density ratios of 1.0, 1.5, and 2.0. Three different foreign gases are used to create density ratio effect. Over-tip flow leakage is also studied by measuring the static pressure distributions on the blade tip using the pressure sensitive paint (PSP) measurement technique. In addition, detailed film cooling effectiveness is acquired to quantify the parametric effect of blowing ratio and density ratio on a plane tip design. Increasing the blowing ratio and density ratio resulted in increased film cooling effectiveness at all injection scenarios. Injecting coolant on the PS and the tip surface also resulted in reduced leakage over the tip. The conclusions from this study will provide the gas turbine designer with additional insight on controlling different parameters and strategically placing the holes during the design process.

Film Cooling Effectiveness on the Leading Edge Region of a Rotating Turbine Blade With Two Rows of Film Cooling Holes Using Pressure Sensitive Paint

2006 ◽  
Vol 128 (9) ◽  
pp. 879-888 ◽  
Author(s):  
Jaeyong Ahn ◽  
M. T. Schobeiri ◽  
Je-Chin Han ◽  
Hee-Koo Moon
Keyword(s):  
Rotational Speed ◽  
Film Cooling ◽  
Rotor Blade ◽  
Incidence Angle ◽  
Leading Edge ◽  
Edge Region ◽  
Pressure Sensitive Paint ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Sensitive

Detailed film cooling effectiveness distributions are measured on the leading edge of a rotating gas turbine blade with two rows (pressure-side row and suction-side row from the stagnation line) of holes aligned to the radial axis using the pressure sensitive paint (PSP) technique. Film cooling effectiveness distributions are obtained by comparing the difference of the measured oxygen concentration distributions with air and nitrogen as film cooling gas respectively and by applying the mass transfer analogy. Measurements are conducted on the first-stage rotor blade of a three-stage axial turbine at 2400rpm (positive off-design), 2550rpm (design), and 3000rpm (negative off-design), respectively. The effect of three blowing ratios is also studied. The blade Reynolds number based on the axial chord length and the exit velocity is 200,000 and the total to exit pressure ratio was 1.12 for the first-stage rotor blade. The corresponding rotor blade inlet and outlet Mach numbers are 0.1 and 0.3, respectively. The film cooling effectiveness distributions are presented along with discussions on the influence of rotational speed (off design incidence angle), blowing ratio, and upstream nozzle wakes around the leading edge region. Results show that rotation has a significant impact on the leading edge film cooling distributions with the average film cooling effectiveness in the leading edge region decreasing with an increase in the rotational speed (negative incidence angle).

Film-Cooling Effectiveness on a Gas Turbine Blade Tip Using Pressure-Sensitive Paint

2005 ◽  
Vol 127 (5) ◽  
pp. 521-530 ◽  
Author(s):  
Jaeyong Ahn ◽  
Shantanu Mhetras ◽  
Je-Chin Han
Keyword(s):  
Oxygen Concentration ◽  
Film Cooling ◽  
Pressure Sensitive Paint ◽  
Nitrogen Gas ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Pressure Sensitive ◽  
Blade Tip ◽  
Film Cooling Holes

Effects of the presence of squealer, the locations of the film-cooling holes, and the tip-gap clearance on the film-cooling effectiveness were studied and compared to those for a plane (flat) tip. The film-cooling effectiveness distributions were measured on the blade tip using the pressure-sensitive paint technique. Air and nitrogen gas were used as the film-cooling gases, and the oxygen concentration distribution for each case was measured. The film-cooling effectiveness information was obtained from the difference of the oxygen concentration between air and nitrogen gas cases by applying the mass transfer analogy. Plane tip and squealer tip blades were used while the film-cooling holes were located (a) along the camber line on the tip or (b) along the tip of the pressure side. The average blowing ratio of the cooling gas was 0.5, 1.0, and 2.0. Tests were conducted with a stationary, five-bladed linear cascade in a blow-down facility. The free-stream Reynolds number, based on the axial chord length and the exit velocity, was 1,138,000, and the inlet and the exit Mach numbers were 0.25 and 0.6, respectively. Turbulence intensity level at the cascade inlet was 9.7%. All measurements were made at three different tip-gap clearances of 1%, 1.5%, and 2.5% of blade span. Results show that the locations of the film-cooling holes and the presence of squealer have significant effects on surface static pressure and film-cooling effectiveness, with film-cooling effectiveness increasing with increasing blowing ratio.

scholarly journals Showerhead Film Cooling Performance of a Turbine Vane at High Freestream Turbulence in a Transonic Cascade

2012 ◽  
Vol 134 (5) ◽  
Author(s):  
Shakeel Nasir ◽  
Trey Bolchoz ◽  
Wing-Fai Ng ◽  
Luzeng J. Zhang ◽  
Hee Koo Moon ◽  
...  
Keyword(s):  
Reynolds Number ◽  
Nusselt Number ◽  
Mach Number ◽  
Film Cooling ◽  
Boundary Layer Transition ◽  
Suction Surface ◽  
Film Cooling Effectiveness ◽  
Pressure Surface ◽  
Blowing Ratio ◽  
Adiabatic Effectiveness

This paper experimentally investigates the effect of blowing ratio and exit Reynolds number/Mach number on the film cooling performance of a showerhead film cooled first stage turbine vane. The vane midspan was instrumented with single-sided platinum thin film gauges to experimentally characterize the Nusselt number and film cooling effectiveness distributions over the surface. The vane was arranged in a two-dimensional, linear cascade in a heated, transonic, blow-down wind tunnel. Three different exit Mach numbers of Mex = 0.57, 0.76 and 1.0—corresponding to exit Reynolds numbers based on vane chord of 9.7 × 105, 1.1 × 106 and 1.5 × 106, respectively—were tested with an inlet free stream turbulence intensity (Tu) of 16% and an integral length scale normalized by vane pitch (Λx/P) of 0.23. A showerhead cooling scheme with five rows of cooling holes was tested at blowing ratios of BR = 0, 1.5, 2.0, and 2.5 and a density ratio of DR = 1.3. Nusselt number and adiabatic film cooling effectiveness distributions were presented on the vane surface over a range of s/C = −0.58 on the pressure side to s/C = 0.72 on the suction side of the vane. The primary effects of coolant injection were to augment the Nusselt number and reduce the adiabatic wall temperature downstream of the injection on the vane surface as compared to no film injection case (BR = 0) at all exit Mach number conditions. In general, an increase in blowing ratio (BR = 1.5 to 2.5) showed noticeable Nusselt number augmentation on pressure surface as compared to suction surface at exit Mach 0.57 and 0.75; however, the Nusselt number augmentation for these blowing ratios was found to be negligible on the vane surface for exit Mach 1.0 case. At exit Mach 1.0, an increase in blowing ratio (BR = 1.5 to 2.5) was observed to have an adverse effect on the adiabatic effectiveness on the pressure surface but had negligible effect on suction surface. The effectiveness trend on the suction surface was also found to be influenced by a favorable pressure gradient due to Mach number and boundary layer transition in the region s/C = 0.28 to s/C = 0.45 at all blowing ratio and exit Mach number conditions. An increase in Reynolds number from exit Mach 0.76 to 1.0 increased heat transfer levels on the vane surface at all blowing ratio conditions. A large increase in Reynolds number adversely affected adiabatic effectiveness on the pressure surface at all blowing ratio conditions. On the suction surface, a large increase in Reynolds number also affected adiabatic effectiveness in the favorable pressure gradient and boundary layer transition region.

Full-Scale Turbine Vane Endwall Film-Cooling Effectiveness Distribution Using Pressure-Sensitive Paint Technique

2016 ◽  
Vol 138 (5) ◽  
Author(s):  
Chao-Cheng Shiau ◽  
Andrew F Chen ◽  
Je-Chin Han ◽  
Salam Azad ◽  
Ching-Pang Lee
Keyword(s):  
Film Cooling ◽  
Full Scale ◽  
Pressure Sensitive Paint ◽  
Complex Flow ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Pressure Sensitive ◽  
Turbine Vane ◽  
Tunnel Design

Researchers in gas turbine field take great interest in the cooling performance on the first-stage vane because of the complex flow characteristics and intensive heat load that comes from the exit of the combustion chamber. A better understanding is needed on how the coolant flow interacts with the mainstream and the resulting cooling effect in the real engine especially for the first-stage vane. An authentic flow channel and condition should be achieved. In this study, three full-scale turbine vanes are used to construct an annular-sector cascade. The film-cooling design is attained through numerous layback fan-shaped and cylindrical holes dispersed on the vane and both endwalls. With the three-dimensional vane geometry and corresponding wind tunnel design, the true flow field can thus be simulated as in the engine. This study targets the film-cooling effectiveness on the inner endwall (hub) of turbine vane. Tests are performed under the mainstream Reynolds number 350,000; the related inlet Mach number is 0.09; and the freestream turbulence intensity is 8%. Two variables, coolant-to-mainstream mass flow ratios (MFR = 2%, 3%, and 4%) and density ratios (DR = 1.0 and 1.5), are examined. Pressure-sensitive paint (PSP) technique is utilized to capture the detail contour of film-cooling effectiveness on the inner endwall and demonstrate the coolant trace. The presented results serve as a comparison basis for other sets of vanes with different cooling designs. The results are expected to strengthen the promise of PSP technique on evaluating the film-cooling performance of the engine geometries.

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