Interaction Between Secondary Flow & Film Cooling Jets of a Realistic Annular Airfoil Cascade (high Mach Number) PDF Download
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Author: Cuong Quoc Nguyen
Publisher:
ISBN:
Category : Gas-turbines
Languages : en
Pages : 149
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Book Description
Film cooling is investigated on a flat plate both numerically and experimentally. Conical shaped film hole are investigated extensively and contribute to the current literature data, which is extremely rare in the open public domain. Both configuration of the cylindrical film holes, with and without a trench, are investigated in detail. Design of experiment technique was performed to find an optimum combination of both geometrical and fluid parameters to achieve the best film cooling performance. From this part of the study, it shows that film cooling performance can be enhanced up to 250% with the trenched film cooling versus non-trenched case provided the same amount of coolant. Since most of the relevant open literature is about film cooling on flat plate endwall cascade with linear extrusion airfoil, the purpose of the second part of this study is to examine the interaction of the secondary flow inside a 3D cascade and the injected film cooling jets. This is employed on the first stage of the aircraft gas turbine engine to protect the curvilinear (annular) endwall platform. The current study investigates the interaction between injected film jets and the secondary flow both experimentally and numerically at high Mach number (M=0.7). Validation shows good agreement between obtained data with the open literature. In general, it can be concluded that with an appropriate film coolant to mainstream blowing ratio, one can not only achieve the best film cooling effectiveness (FCE or [eta]) on the downstream endwall but also maintain almost the same aerodynamic loss as in the un-cooled baseline case. Film performance acts nonlinearly with respect to blowing ratios as with film cooling on flat plate, in the other hand, with a right blowing ratio, film cooling performance is not affect much by secondary flow. In turn, film cooling jets do not increase pressure loss at the downstream wake area of the blades.
Author: Cuong Quoc Nguyen
Publisher:
ISBN:
Category : Gas-turbines
Languages : en
Pages : 149
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Book Description
Film cooling is investigated on a flat plate both numerically and experimentally. Conical shaped film hole are investigated extensively and contribute to the current literature data, which is extremely rare in the open public domain. Both configuration of the cylindrical film holes, with and without a trench, are investigated in detail. Design of experiment technique was performed to find an optimum combination of both geometrical and fluid parameters to achieve the best film cooling performance. From this part of the study, it shows that film cooling performance can be enhanced up to 250% with the trenched film cooling versus non-trenched case provided the same amount of coolant. Since most of the relevant open literature is about film cooling on flat plate endwall cascade with linear extrusion airfoil, the purpose of the second part of this study is to examine the interaction of the secondary flow inside a 3D cascade and the injected film cooling jets. This is employed on the first stage of the aircraft gas turbine engine to protect the curvilinear (annular) endwall platform. The current study investigates the interaction between injected film jets and the secondary flow both experimentally and numerically at high Mach number (M=0.7). Validation shows good agreement between obtained data with the open literature. In general, it can be concluded that with an appropriate film coolant to mainstream blowing ratio, one can not only achieve the best film cooling effectiveness (FCE or [eta]) on the downstream endwall but also maintain almost the same aerodynamic loss as in the un-cooled baseline case. Film performance acts nonlinearly with respect to blowing ratios as with film cooling on flat plate, in the other hand, with a right blowing ratio, film cooling performance is not affect much by secondary flow. In turn, film cooling jets do not increase pressure loss at the downstream wake area of the blades.
Author:
Publisher:
ISBN:
Category :
Languages : en
Pages : 14
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Book Description
Author:
Publisher:
ISBN:
Category :
Languages : en
Pages : 0
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Book Description
A previously documented CFD methodology is further generalized and applied to realistic turbine airfoil film cooling test cases. First, a series of fundamental test cases are examined in order to document the ability of the robust and practical CFD methodology to deal with the effects of the individually isolated key physics mechanisms on film cooling. These include: (1) favorable and adverse pressure gradients; (2) convex curvature; (3) horseshoe vortex; (4) profile losses with and without coolant jets; (5) laminar-to-turbulent boundary layer transition; and (6) discontinuities. Second, in addition to the fundamental flow test cases, two turbine airfoil cascade, one low subsonic and another transonic, configurations were modeled. The subsonic cascade case was designed to study the aerodynamics losses with and without film cooling jets, as well as, the adiabatic effectiveness for a range of parameters. The transonic turbine cascade represents a truly modern design at realistic engine conditions. A total of 18 test configurations, corresponding to compound-round, axial-shaped, and compound-shaped film holes, are simulated in order to document in detail the current state of the readily available robust and practical CFD technology for use by the gas turbine design community.
Author: James C. Kullberg
Publisher:
ISBN:
Category :
Languages : en
Pages : 105
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Book Description
Many studies have been published on the subject of film cooling, but because of the difficulty and expense of simulating turbine realistic conditions, many authors introduce vast simplifications such as low speed conditions or linear cascades. These simplifications do not adequately represent the behavior of a turbine and therefore their results are of limited use. This study attempts to eliminate many of those simplifications.
Author:
Publisher:
ISBN:
Category : Aeronautics
Languages : en
Pages : 580
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Book Description
Author: S. Stephen Papell
Publisher:
ISBN:
Category : Mach number
Languages : en
Pages : 68
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Book Description
Author: Christopher Yoon
Publisher:
ISBN:
Category :
Languages : en
Pages : 126
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Book Description
Film cooling is often used for turbine airfoil cooling, and there are numerous studies of the performance of a single row of holes. Typically, blades and vanes in gas turbine engines have multiple rows of holes that interact. Consequently, there is a need to develop techniques to predict film cooling performance with multiple rows of holes. One of the method is to superposition single row cooling effectiveness to predict combined effectiveness. Although there have been many studies of superposition techniques with multiple rows of cylindrical holes, there have been very few in which shaped holes were used with a typical turbine airfoil model. In this study, film effectiveness was measured on the suction side of a turbine blade model using two rows of 7-7-7 shaped holes, with pitch to diameter ratio of 6, and the two rows were more than 40 diameters apart. Measurements were made with each row operating independently, which provided the experimental data for superposition predictions. These predictions were evaluated with effectiveness measurements with both rows operational. For these combined row tests, two different upstream blowing ratios and a wide range of downstream blowing ratios were selected. The superposition predictions were reasonably accurate when the upstream blowing ratio was high with a corresponding smaller film effectiveness downstream (due to jet separation). However, when the upstream coolant holes were operated at optimum blowing ratio with maximum film effectiveness downstream, the superposition analysis predicted film effectiveness levels slightly lower than actual levels. These results indicate that there was an interaction between jets that resulted in higher film effectiveness than what the superposition method had predicted
Author:
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ISBN:
Category : Aeronautics
Languages : en
Pages : 712
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Book Description
A selection of annotated references to unclassified reports and journal articles that were introduced into the NASA scientific and technical information system and announced in Scientific and technical aerospace reports (STAR) and International aerospace abstracts (IAA)
Author: D. Granser
Publisher:
ISBN:
Category : Cooling
Languages : en
Pages : 0
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Book Description
After compressor discharge air has initially been used to cool the heat shields of the hot gas inlet casing, it can subsequently be employed for film cooling of the first-stage vane shrouds. Since the flow field near these shrouds is three-dimensional, the film cooling effectiveness cannot be predicted correctly by common two-dimensional codes. The secondary flow transports the film from the pressure side to the suction side where it can even climb up the airfoil to cool its trailing section. Such film cooling effectiveness was first investigated experimentally in a linear vane cascade at atmospheric pressure. The temperatures and static pressure levels at the adiabatic shrouds, as well as the temperature measurements within the vane cascade, are reported for different cooling film blowing rates. In addition, the secondary flow was analysed numerically using a partially-parabolic computer code for 3D viscous flows. It involves mutual interaction of the boundary layer with the mainstream. The secondary flow can also be modelled with this algorithm, which requires less numerical effort than solving the fully 3D elliptic flow equations. The numerical results of the experiment and numerical predictions are compared. In addition, the application of these results to a high-temperature gas turbine is presented.
Author: Christian Liess
Publisher:
ISBN:
Category : Cooling
Languages : en
Pages : 116
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Book Description
