Effects of Inlet Swirl on Suction Side Phantom Cooling

2016 ◽  
Author(s):  
Yang Zhang ◽  
Yifei Li ◽  
Xin Yuan
Keyword(s):  
Film Cooling ◽  
Swirling Flow ◽  
Guide Vane ◽  
Density Ratio ◽  
Electric Energy ◽  
Side Surface ◽  
Suction Side ◽  
Rotating Flow ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness

Though the temperature of the coolant injected from the endwall increases after the mixing process in the main flow, the injections moving from the endwall to the airfoil suction side still have the potential of second order cooling. This part of the coolant is called “Phantom cooling” in the paper. This paper is focused on the function of the coolant from fan-shaped holes on endwall surface which brought by passage vortex to the airfoil suction side. The test cascades are based on the profile of General Electric Energy Efficient Engine (GE-E3), with the inlet Mach number of 0.1 and Reynolds number of 1.46×105. The scale ratio of the test vanes is 1.95 and the film cooling effectiveness is tested with Pressure Sensitive Painting (PSP) Technology. Nitrogen is used to simulate the coolant which can provide a density ratio near 1.0. Two adjacent passages are investigated simultaneously by which the film cooling effectiveness can be compared in the same case on the suction side surface. The inlet rotating flow is simulated by an upstream swirler at the inlet, fixed at 5 different positions along the pitchwise direction. They are Blade 0 aligned, Passage 1 aligned, Blade 1 aligned, Passage 2 aligned and Blade 2 aligned. According to the experimental results, the inlet rotating flow can dominate the film cooling effectiveness distribution at the Suction Side. The averaged film cooling effectiveness changes substantially with the change in the swirler position. The inlet swirling flow effects are compared on the adjacent vanes at upstream and downstream regions. The research shows that the different pitchwise position of the inlet swirler is a significant factor which can change the phantom cooling phenomenon on the suction side. The relative position of inlet swirler cannot be ignored for Nozzle Guide Vane design.

Effects of Inlet Swirl on Pressure Side Film Cooling of Neighboring Vane Surface

2016 ◽  
Author(s):  
Yifei Li ◽  
Yang Zhang ◽  
Xin Yuan
Keyword(s):  
Film Cooling ◽  
Incidence Angle ◽  
Density Ratio ◽  
Electric Energy ◽  
Suction Side ◽  
Rotating Flow ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Scale Ratio

The test cascades are based on the profile of General Electric Energy Efficient Engine (GE-E3) with the inlet Mach number of 0.1 and Reynolds number of 1.48×105. The scale ratio of the test vanes is 1.95 and the film cooling effectiveness is tested with Pressure Sensitive Painting (PSP) technology. Nitrogen is used to simulate the coolant which can provide a density ratio near to 1.0. The two adjacent passages are investigated simultaneously by which the film cooling effectiveness can be compared in the same case with pressure side and suction side respectively. The rotating flow is simulated by an upstream swirler at the inlet, fixed at 5 different positions along the pitchwise direction at the inlet. They are Blade 1 aligned, Passage 1 aligned, Blade 2 aligned, Passage 2 aligned and Blade 3 aligned. According to the experimental results, the inlet rotating flow can dominate the film cooling effectiveness distribution at Pressure Side. The averaged film cooling effectiveness changes significantly with the change in swirler position. The effect of the rotating flow at the Pressure Side region is mainly interacted with the main flow to cause the change in incidence angle. The influence of the inlet rotating flow is more obvious on the pressure side upstream part. Meanwhile the downstream part is not so sensitive to rotating flow as pressure side.

Effects of Inlet Swirl on Pressure Side Film Cooling of Neighboring Vane Surface

2019 ◽  
Vol 11 (6) ◽  
Author(s):  
Yang Zhang ◽  
Yifei Li ◽  
Xiutao Bian ◽  
Xin Yuan
Keyword(s):  
Film Cooling ◽  
Swirling Flow ◽  
Density Ratio ◽  
Suction Side ◽  
Scale Model ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Lean Combustion ◽  
Inlet Swirl ◽  
Nozzle Guide

The lean combustion chamber of low NOx emission engines has a short distance between combustion outlet and nozzle guide vanes (NGVs), with strong swirlers located upstream of the turbine inlet to from steady circulation in the combustion region. Although the lean combustion design benefits emission control, it complicates the turbine’s aerodynamics and heat transfer. The strong swirling flow will influence the near-wall flow field where film cooling acts. This research investigates the influence of inlet swirl on the film cooling of cascades. The test cascades are a 1.95 scale model based on the GE-E3 profile, with an inlet Mach number of 0.1 and Reynolds number of 1.48 × 105. Film cooling effectiveness is measured with pressure-sensitive paint (PSP) technology, with nitrogen simulating coolant at a density ratio of near to 1.0. Two neighboring passages are investigated simultaneously, so that pressure and suction side the film cooling effectiveness can be compared. The inlet swirl is produced by a swirler placed upstream, near the inlet, with five positions along the pitchwise direction. These are as follows: blade 1 aligned, passage 1–2 aligned, blade 2 aligned, passage 2–3 aligned and blade 3 aligned. According to the experimental results, the near-hub region is strongly influenced by inlet swirl, where the averaged film cooling effectiveness can differ by up to 12% between the neighboring blades. At the spanwise location Z/Span = 0.7, when the inlet swirl is moved from blade 1 aligned (position 5) to blade 2 aligned (position 3), the film cooling effectiveness in a small area near the endwall can change by up to 100%.

Film Cooling Effectiveness of Transonic Turbine Vane Suction Side With Compound-Angle Shaped Hole Configuration

2015 ◽  
Author(s):  
Shang-Feng Yang ◽  
Je-Chin Han ◽  
Alexander MirzaMoghadam ◽  
Ardeshir Riahi
Keyword(s):  
Transonic Flow ◽  
Film Cooling ◽  
Surface Film ◽  
Density Ratio ◽  
Suction Side ◽  
Flow Loop ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Turbine Vane ◽  
Compound Angle

This paper studies the effect of transonic flow velocity on local film cooling effectiveness distribution of turbine vane suction side, experimentally. A conduction-free Pressure Sensitive Paint (PSP) method is used to determine the local film cooling effectiveness. Tests were performed in a five-vane annular cascade at Texas A&M Turbomachinery laboratory blow-down flow loop facility. The exit Mach numbers are controlled to be 0.7, 0.9, and 1.1, from subsonic to transonic flow conditions. Three foreign gases N2, CO2 and Argon/SF6 mixture are selected to study the effects of three coolant-to-mainstream density ratios, 1.0, 1.5, and 2.0 on film cooling. Four averaged coolant blowing ratios in the range, 0.7, 1.0, 1.3 and 1.6 are investigated. The test vane features 3 rows of radial-angle cylindrical holes around the leading edge, and 2 rows of compound-angle shaped holes on the suction side. Results suggest that the PSP technique is capable of producing clear and detailed film cooling effectiveness contours at transonic condition. The effects of coolant to mainstream blowing ratio, density ratio, and exit Mach number on the vane suction-surface film cooling distribution are obtained, and the consequence results are presented and explained in this investigation.

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.

High-Resolution Film Cooling Effectiveness Measurements of Axial Holes Embedded in a Transverse Trench With Various Trench Configurations

2006 ◽  
Vol 129 (2) ◽  
pp. 294-302 ◽  
Author(s):  
Scot K. Waye ◽  
David G. Bogard
Keyword(s):  
High Resolution ◽  
Film Cooling ◽  
Density Ratio ◽  
Field Measurements ◽  
Suction Side ◽  
Lateral Spreading ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Flow Parameters ◽  
Adiabatic Effectiveness

Adiabatic film cooling effectiveness of axial holes embedded within a transverse trench on the suction side of a turbine vane was investigated. High-resolution two-dimensional data obtained from infrared thermography and corrected for local conduction provided spatial adiabatic effectiveness data. Flow parameters of blowing ratio, density ratio, and turbulence intensity were independently varied. In addition to a baseline geometry, nine trench configurations were tested, all with a depth of 1∕2 hole diameter, with varying widths, and with perpendicular and inclined trench walls. A perpendicular trench wall at the very downstream edge of the coolant hole was found to be the key trench characteristic that yielded much improved adiabatic effectiveness performance. This configuration increased adiabatic effectiveness up to 100% near the hole and 40% downstream. All other trench configurations had little effect on the adiabatic effectiveness. Thermal field measurements confirmed that the improved adiabatic effectiveness that occurred for a narrow trench with perpendicular walls was due to a lateral spreading of the coolant and reduced coolant jet separation. The cooling levels exhibited by these particular geometries are comparable to shaped holes, but much easier and cheaper to manufacture.

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.

Influence of Coolant Density on Turbine Blade Film-Cooling With Compound-Angle Shaped Holes

2012 ◽  
Author(s):  
Kevin Liu ◽  
Shang-Feng Yang ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Momentum Flux ◽  
Density Ratio ◽  
Flux Ratio ◽  
Suction Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Shaped Hole ◽  
Compound Angle

Adiabatic film-cooling effectiveness is examined systematically on a typical high pressure turbine blade by varying three critical flow parameters: coolant blowing ratio, coolant-to-mainstream density ratio, and freestream turbulence intensity. Three average coolant blowing ratios 1.0, 1.5, and 2.0; three coolant density ratios 1.0, 1.5, and 2.0; two turbulence intensities 4.2% and 10.5%, are chosen for this study. Conduction-free pressure sensitive paint (PSP) technique is used to measure film-cooling effectiveness. Three foreign gases — N2 for low density, CO2 for medium density, and a mixture of SF6 and Argon for high density are selected to study the effect of coolant density. The test blade features 45° compound-angle shaped holes on the suction side and pressure side, and 3 rows of 30° radial-angle cylindrical holes around the leading edge region. The inlet and the exit Mach number are 0.27 and 0.44, respectively. Reynolds number based on the exit velocity and blade axial chord length is 750,000. Results reveal that the PSP is a powerful technique capable of producing clear and detailed film effectiveness contours with diverse foreign gases. As blowing ratio exceeds the optimum value, it induces more mixing of coolant and mainstream. Thus film-cooling effectiveness reduces. Greater coolant-to-mainstream density ratio results in lower coolant-to-mainstream momentum and prevents coolant to lift-off; as a result, film-cooling increases. Higher freestream turbulence causes effectiveness to drop everywhere except in the region downstream of suction side. Results are also correlated with momentum flux ratio and compared with previous studies. It shows that compound shaped hole has the greatest optimum momentum flux ratio, and then followed by axial shaped hole, compound cylindrical hole, and axial cylindrical hole.

High Resolution Film Cooling Effectiveness Comparison of Axial and Compound Angle Holes on the Suction Side of a Turbine Vane

2006 ◽  
Author(s):  
Scot K. Waye ◽  
David G. Bogard
Keyword(s):  
High Resolution ◽  
Flat Plate ◽  
Film Cooling ◽  
Density Ratio ◽  
Suction Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Turbine Vane ◽  
Adiabatic Effectiveness ◽  
Compound Angle

Film cooling adiabatic effectiveness for axial and compound angle holes on the suction side of a simulated turbine vane was investigated to determine the relative performance of these configurations. The effect of the surface curvature was also evaluated by comparing to previous curvature studies and flat plate film cooling results. Experiments were conducted for varying coolant density ratio, mainstream turbulence levels, and hole spacing. Results from these measurements showed that for mild curvature, 2r/d ≈ 160, flat plate results are sufficient to predict the cooling effectiveness. Furthermore, the compound angle injection improves adiabatic effectiveness for higher blowing ratios, similar to previous studies using flat plate facilities.

Turbine Vane Endwall Film Cooling With Slashface Leakage and Discrete Hole Configuration

2017 ◽  
Vol 139 (6) ◽  
Author(s):  
Nafiz H. K. Chowdhury ◽  
Chao-Cheng Shiau ◽  
Je-Chin Han ◽  
Luzeng Zhang ◽  
Hee-Koo Moon
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Density Ratio ◽  
Secondary Flows ◽  
Blow Down ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Free Stream Turbulence ◽  
Leakage Flows ◽  
Turbine Vane

Turbine vanes are typically assembled as a section containing single or double airfoil units in an annular pattern. First stage guide vane assembly results in two common mating interfaces: a gap between combustor and vane endwall and another resulted from the adjacent sections, called slashface. High pressure coolant could leak through these gaps to reduce the ingestion of hot gas and achieve certain cooling benefit. As vane endwall region flow field is already very complicated due to highly three-dimensional secondary flows, then a significant influence on endwall cooling can be expected due to the gap leakage flows. To determine the effect of leakage flows from those gaps, film cooling effectiveness distributions were measured using pressure sensitive paint (PSP) technique on the endwall of a scaled up, midrange industrial turbine vane geometry with the multiple rows of discrete film cooling (DFC) holes inside the passages. Experiments were performed in a blow-down wind tunnel cascade facility at the exit Mach number of 0.5 corresponding to Reynolds number of 3.8 × 105 based on inlet conditions and axial chord length. Passive turbulence grid was used to generate free-stream turbulence (FST) level about 19% with an integral length scale of 1.7 cm. Two parameters, coolant-to-mainstream mass flow ratio (MFR) and density ratio (DR), were studied. The results are presented as two-dimensional film cooling effectiveness distribution on the vane endwall surface with the corresponding spanwise averaged values along the axial direction.

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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