Film Effectiveness Performance for a Shaped Hole on the Suction Side of a Scaled-up Turbine Blade PDF Download
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Author: Jacob Damian Moore
Publisher:
ISBN:
Category :
Languages : en
Pages : 324
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Book Description
Surface curvature has been shown to have significant effects on the film cooling performance of round holes, but the present literature includes very few studies dedicated to curvature’s effects on shaped hole geometries despite their prevalence in turbine blade and vane designs. Experiments were performed on two rows of holes placed on the suction side of a scaled-up gas turbine blade model in a low-Mach-number linear cascade wind tunnel. The test facility was set up to match a high-Mach-number pressure distribution without modifying the blade’s geometry or including contoured end walls to accelerate the flow. By adjusting the positions of the movable walls in the tunnel test section, the suction side pressure distribution could be matched to the design distribution. One row was placed in a region of high convex surface curvature; the other, in a region of low convex curvature. Other geometric and flow parameters near the rows were matched in the design of the experiment, including hole geometry and spacing. The hole geometry was a standard 7-7-7 shaped hole. In addition, local freestream conditions for the rows were measured and set to match as closely as possible. Comparison of the adiabatic effectiveness results from the two rows revealed trends similar to those seen in previous literature for round holes. The high curvature row outperformed the low curvature row at lower coolant injection rates, having wider jets and higher centerline effectiveness. But as the injection rate was increased, the low curvature row surpassed the high curvature row in effectiveness. The driver behind this behavior was the surface-normal pressure gradient that arose from the convex surface curvature. As flow traveled around the surface, centripetal acceleration produced a pressure gradient directed towards the surface, effectively pushing jets toward the blade wall. However, at higher blowing ratios, the jets’ high momenta overcame the effects of this pressure gradient. At these injection rates, the high curvature row’s jets’ trajectories did not follow the surface as it curved away. The high surface curvature exacerbated the adverse effects of jet separation on film cooling performance.
Author: Jacob Damian Moore
Publisher:
ISBN:
Category :
Languages : en
Pages : 324
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Book Description
Surface curvature has been shown to have significant effects on the film cooling performance of round holes, but the present literature includes very few studies dedicated to curvature’s effects on shaped hole geometries despite their prevalence in turbine blade and vane designs. Experiments were performed on two rows of holes placed on the suction side of a scaled-up gas turbine blade model in a low-Mach-number linear cascade wind tunnel. The test facility was set up to match a high-Mach-number pressure distribution without modifying the blade’s geometry or including contoured end walls to accelerate the flow. By adjusting the positions of the movable walls in the tunnel test section, the suction side pressure distribution could be matched to the design distribution. One row was placed in a region of high convex surface curvature; the other, in a region of low convex curvature. Other geometric and flow parameters near the rows were matched in the design of the experiment, including hole geometry and spacing. The hole geometry was a standard 7-7-7 shaped hole. In addition, local freestream conditions for the rows were measured and set to match as closely as possible. Comparison of the adiabatic effectiveness results from the two rows revealed trends similar to those seen in previous literature for round holes. The high curvature row outperformed the low curvature row at lower coolant injection rates, having wider jets and higher centerline effectiveness. But as the injection rate was increased, the low curvature row surpassed the high curvature row in effectiveness. The driver behind this behavior was the surface-normal pressure gradient that arose from the convex surface curvature. As flow traveled around the surface, centripetal acceleration produced a pressure gradient directed towards the surface, effectively pushing jets toward the blade wall. However, at higher blowing ratios, the jets’ high momenta overcame the effects of this pressure gradient. At these injection rates, the high curvature row’s jets’ trajectories did not follow the surface as it curved away. The high surface curvature exacerbated the adverse effects of jet separation on film cooling performance.
Author: National Aeronautics and Space Administration (NASA)
Publisher: Createspace Independent Publishing Platform
ISBN: 9781720482284
Category :
Languages : en
Pages : 64
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Book Description
The detailed heat transfer coefficient and film cooling effectiveness distributions as well as tile detailed coolant jet temperature profiles on the suction side of a gas turbine blade A,ere measured using a transient liquid crystal image method and a traversing cold wire and a traversing thermocouple probe, respectively. The blade has only one row of film holes near the gill hole portion on the suction side of the blade. The hole geometries studied include standard cylindrical holes and holes with diffuser shaped exit portion (i.e. fanshaped holes and laidback fanshaped holes). Tests were performed on a five-blade linear cascade in a low-speed wind tunnel. The mainstream Reynolds number based on cascade exit velocity was 5.3 x 10(exp 5). Upstream unsteady wakes were simulated using a spoke-wheel type wake generator. The wake Strouhal number was kept at 0 or 0.1. Coolant blowing ratio was varied from 0.4 to 1.2. Results show that both expanded holes have significantly improved thermal protection over the surface downstream of the ejection location, particularly at high blowing ratios. However, the expanded hole injections induce earlier boundary layer transition to turbulence and enhance heat transfer coefficients at the latter part of the blade suction surface. In general, the unsteady wake tends to reduce film cooling effectiveness.Han, J. C. and Teng, S.Glenn Research CenterHEAT TRANSFER COEFFICIENTS; COOLANTS; TEMPERATURE PROFILES; SUCTION; TURBINE BLADES; HEAT MEASUREMENT; FILM COOLING; BOUNDARY LAYER TRANSITION; CASCADE WIND TUNNELS; CYLINDRICAL BODIES; EJECTION; GAS TURBINES; HOLE DISTRIBUTION (MECHANICS); LIQUID CRYSTALS; LOW SPEED; THERMAL PROTECTION; THERMOCOUPLES; WIND TUNNELS
Author:
Publisher:
ISBN:
Category :
Languages : en
Pages : 66
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Book Description
Author: Jacob Damian Moore
Publisher:
ISBN:
Category :
Languages : en
Pages : 484
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Book Description
The leading-edge regions of turbine vanes and blades require careful attention to their cooling designs because of the high heat loads. External cooling is typically accomplished with dense "showerhead" arrangements of film cooling holes surrounding the stagnation point at the airfoil leading edge. In modern film cooling studies, shaped holes are prevalent in downstream areas of turbine airfoils; however, the literature contains few studies of shaped holes in the showerhead. This leads to a lack of physics-based insight that would lead to the design of high-performing showerhead arrays. This study examined the performance and physical behavior of several showerhead arrangements at the leading edge of a scaled-up turbine blade. A low-speed linear cascade test section was used to simulate the blade environment, and experiments were conducted at scaled engine-realistic conditions. First, the cooling performances of baseline cylindrical and shaped hole designs were compared. The shaped hole design mimicked a standard design in the literature for flat plate studies but with some modifications expected to improve performance specifically at the leading edge. The result was a novel off-center, elliptically-expanding hole. Adiabatic effectiveness and thermal field measurements revealed that the baseline shaped hole had 20-100% performance due to better jet attachment, stemming from its diffuser, which effectively decreased the exit momenta of the coolant jets. The expansion area ratio was increased by 40% for a subsequent design to gauge sensitivity to this parameter; but, surprisingly, the performances of the new design and of the baseline one were nearly identical. A third shaped hole design with a 45% larger breakout area but an identical expansion area resulted in slightly worse performance than either, highlighting the detrimental effect of increasing breakout area and expansion angle. These experiments informed a new proposed scaling parameter incorporating both of these areas and their counteracting effects to predict shaped hole performance in the showerhead. The highest performing design of the group was then tested with an engine-realistic impingement coolant feed, for which performance was overall similar. Supplemental thermal fields using this configuration were performed to construct a 3D representation of the flow field in the showerhead region
Author: Diganta P. Narzary
Publisher:
ISBN:
Category :
Languages : en
Pages :
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Book Description
Modern gas turbine engines require higher turbine-entry gas temperature to improve their thermal efficiency and thereby their performance. A major accompanying concern is the heat-up of the turbine components which are already subject to high thermal and mechanical stresses. This heat-up can be reduced by: (i) applying thermal barrier coating (TBC) on the surface, and (ii) providing coolant to the surface by injecting secondary air discharged from the compressor. However, as the bleeding off of compressor discharge air exacts a penalty on engine performance, the cooling functions must be accomplished with the smallest possible secondary air injection. This necessitates a detailed and systematic study of the various flow and geometrical parameters that may have a bearing on the cooling pattern. In the present study, experiments were performed in three regions of a non-rotating gas turbine blade cascade: blade platform, blade span, and blade tip. The blade platform and blade span studies were carried out on a high pressure turbine rotor blade cascade in medium flow conditions. Film-cooling effectiveness or degree of cooling was assessed in terms of cooling hole geometry, blowing ratio, freestream turbulence, coolant-to-mainstream density ratio, purge flow rate, upstream vortex for blade platform cooling and blowing ratio, and upstream vortex for blade span cooling. The blade tip study was performed in a blow-down flow loop in a transonic flow environment. The degree of cooling was assessed in terms of blowing ratio and tip clearance. Limited heat transfer coefficient measurements were also carried out. Mainstream pressure loss was also measured for blade platform and blade tip film-cooling with the help of pitot-static probes. The pressure sensitive paint (PSP) and temperature sensitive paint (TSP) techniques were used for measuring film-cooling effectiveness whereas for heat transfer coefficient measurement, temperature sensitive paint (TSP) technique was employed. Results indicated that the blade platform cooling requires a combination of upstream purge flow and downstream discrete film-cooling holes to cool the entire platform. The shaped cooling holes provided wider film coverage and higher film-cooling effectiveness than the cylindrical holes while also creating lesser mainstream pressure losses. Higher coolant-to-mainstream density ratio resulted in higher effectiveness levels from the cooling holes. On the blade span, at any given blowing ratio, the suction side showed better coolant coverage than the pressure side even though the former had two fewer rows of holes. Film-cooling effectiveness increased with blowing ratio on both sides of the blade. Whereas the pressure side effectiveness continued to increase with blowing ratio, the increase in suction side effectiveness slowed down at higher blowing ratios (M=0.9 and 1.2). Upstream wake had a detrimental effect on film coverage. 0% and 25% wake phase positions significantly decreased film-cooling effectiveness magnitude. Comparison between the compound shaped hole and the compound cylindrical hole design showed higher effectiveness values for shaped holes on the suction side. The cylindrical holes performed marginally better in the curved portion of the pressure side. Finally, the concept tip proved to be better than the baseline tip in terms of reducing mainstream flow leakage and mainstream pressure loss. The film-cooling effectiveness on the concept blade increased with increasing blowing ratio and tip gap. However, the film-coverage on the leading tip portion was almost negligible.
Author:
Publisher:
ISBN:
Category :
Languages : en
Pages : 230
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Book Description
Author: Je-Chin Han
Publisher: CRC Press
ISBN: 1439855684
Category : Science
Languages : en
Pages : 892
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Book Description
A comprehensive reference for engineers and researchers, Gas Turbine Heat Transfer and Cooling Technology, Second Edition has been completely revised and updated to reflect advances in the field made during the past ten years. The second edition retains the format that made the first edition so popular and adds new information mainly based on selected published papers in the open literature. See What’s New in the Second Edition: State-of-the-art cooling technologies such as advanced turbine blade film cooling and internal cooling Modern experimental methods for gas turbine heat transfer and cooling research Advanced computational models for gas turbine heat transfer and cooling performance predictions Suggestions for future research in this critical technology The book discusses the need for turbine cooling, gas turbine heat-transfer problems, and cooling methodology and covers turbine rotor and stator heat-transfer issues, including endwall and blade tip regions under engine conditions, as well as under simulated engine conditions. It then examines turbine rotor and stator blade film cooling and discusses the unsteady high free-stream turbulence effect on simulated cascade airfoils. From here, the book explores impingement cooling, rib-turbulent cooling, pin-fin cooling, and compound and new cooling techniques. It also highlights the effect of rotation on rotor coolant passage heat transfer. Coverage of experimental methods includes heat-transfer and mass-transfer techniques, liquid crystal thermography, optical techniques, as well as flow and thermal measurement techniques. The book concludes with discussions of governing equations and turbulence models and their applications for predicting turbine blade heat transfer and film cooling, and turbine blade internal cooling.
Author: James D. Heidmann
Publisher:
ISBN:
Category :
Languages : en
Pages : 14
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Book Description
Presented at the International Gas Turbine & Aeroengine Congress & Exhibition, Orlando, FL, Jun 2 - Jun 5, 1997.
Author: James D. Heidmann
Publisher:
ISBN:
Category : Airplanes
Languages : en
Pages : 272
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Book Description
Author: Vijay K. Garg
Publisher:
ISBN:
Category :
Languages : en
Pages : 26
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Book Description
Presented at the International Gas Turbine and Aeroengine Congress and Exposition, Houston, Texas - June 5-8, 1995.
