Aerodynamic Loss in a Gas Turbine Stage with Film Cooling

Ito, S.; Eckert, E. R. G.; Goldstein, R. J.

doi:10.1115/1.3230368

Cited by 31 publications

(8 citation statements)

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“…This is caused by the superposition of two effects: one is the well-known fact, e.g. Ito et al (1980) and Arts (1992), that the presence of the holes may both trigger the transition of the boundary layer and increase the turbulence of a turbulent boundary layer; the other is due to the fraction of the mainstream entering the hollow cavity in the leading edge region and exiting from the suction side and from the trailing edge holes, thus experiencing an additional dissipation. In order to separate these two effects the holes should be sealed from the inside; however, this was not done in the course of the present work.…”

Section: Film Cooling Test Conditionsmentioning

confidence: 99%

“…Several previously published papers investigated this problem on the base of cascade testing; most of them examined the effects of individual injection rows, depending on their position along the blade surface. Ito et al(1980) reported on the influence of early injection on suction and pressure side at low speed. Kost and Holmes (1985) investigated different trailing edge injection configurations for a transonic rotor blade; they separated the injection loss from the profile loss as a function of coolant flow rate, by considering the base pressure changes.…”

Section: Introductionmentioning

confidence: 99%

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The Influence of Film-Cooling on the Aerodynamic Performance of a Turbine Nozzle Guide Vane

Osnaghi

Perdichizzi

Savini

et al. 1997

Volume 1: Aircraft Engine; Marine; Turbomachinery; Microturbines and Small Turbomachinery

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The paper presents the results of an investigation on the aerodynamic performance of a full coverage film-cooled nozzle guide vane. The blading is a typical high pressure turbine vane of advanced design, working in the high subsonic regime. Tests have been carried out for a wide range of conditions, including variations in Mach number, coolant to mainstream mass flow rate ratio and location of the coolant injection. Both air and carbon dioxide at ambient conditions have been utilized, as coolant flow. Measurements have been performed in a plane located at 0.5 axial chord downstream of the trailing edge by means of a miniaturized five-hole pressure probe. Performances, in terms of losses, flow angles and profile pressure distributions, for different cooling mass flow rates are presented and compared to the results of the solid blade tests (i.e. with no cooling holes). The results showed a significant increase of the losses with blowing. Test with air and carbon dioxide provided almost equal losses if carried out at the same global momentum flux ratio; however the density ratio was found to influence slightly the share of the coolant fluid among the injection rows and the local momentum flux ratio as well. In order to define the individual contributions of groups of cooling rows on the performance of the blade, three different modes of injection have been tested, namely full, trailing edge and shower head injection. The main trend observed is that trailing edge injection produces the least amount of additional losses at high blowing rates. Full-coverage film-cooling injection did not lead to marked variations in the blade pressure distribution and/or outlet flow angle.

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Section: Film Cooling Test Conditionsmentioning

confidence: 99%

Section: Introductionmentioning

confidence: 99%

The Influence of Film-Cooling on the Aerodynamic Performance of a Turbine Nozzle Guide Vane

Osnaghi

Perdichizzi

Savini

et al. 1997

Volume 1: Aircraft Engine; Marine; Turbomachinery; Microturbines and Small Turbomachinery

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“…Willett and Fottner (1994) performed an experiment about the mixing process at turbine blades, especially on the suction side. Ito et al (1980) measured the local pressure losses in a cascade with air ejection in order to simulate the film cooling. They also predicted the mixing pressure loss analytically by assuming a mixing at constant static pressure and made a comparison between experiment and prediction.…”

Section: Nomenclature Amentioning

confidence: 99%

Performance Analysis of a Turbine Stage Having Cooled Nozzle Blades With Trailing Edge Ejection

Kim

Lee

et al. 1996

ASME 1996 Turbo Asia Conference

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This work presents an aerothermodynamic modeling of a cooled turbine blade and the performance analysis of a turbine stage having cooled nozzle blades with trailing edge coolant ejection. A mean line analysis, based on the well-known Ainley-Mathieson scheme, is adopted for the basic loss prediction of the blade rows without cooling. A unique model regarding the interaction between coolant and main gas is proposed. The interactions considered are the heat transfer from main gas to coolant and the temperature and pressure losses by the mixing of two streams due to the trailing edge coolant ejection. For a model turbine stage with nozzle cooling, parametric analyses are carried out to investigate the effect of main design variables (amount of coolant flow, coolant temperature and coolant ejection area) on the stage performance. The influences of coolant mass flow ratio and temperature on the mixing loss and specific work are investigated. The results are also rearranged to investigate the effect of blade temperature on the specific work. Analysis is also carried out by varying the ejection area, which may give useful criteria in determining the coolant condition and ejection hole size of real gas turbine engines.

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“…However, the simulated mass flow and efficiency underline the need for a cooling loss model. Subsequent development efforts will tend to introduce an one-dimensional mixing model for the evaluation of the cooling losses [39,40].…”

Section: Discussionmentioning

confidence: 99%

A quasi-one-dimensional CFD model for multistage turbomachines

Léonard

Adam

2008

J. Therm. Sci.

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The objective of this paper is to present a fast and reliable CFD model that is able to simulate stationary and transient operations of multistage compressors and turbines. This analysis tool is based on an adapted version of the Euler equations solved by a time-marching, finite-volume method. The Euler equations have been extended by including source terms expressing the blade-flow interactions. These source terms are determined using the velocity triangles and a row-by-row representation of the blading at mid-span. The losses and deviations undergone by the fluid across each blade row are supplied by correlations. The resulting flow solver is a performance prediction tool based only on the machine geometry, offering the possibility of exploring the entire characteristic map of a multistage compressor or turbine. Its efficiency in terms of CPU time makes it possible to couple it to an optimization algorithm or to a gas turbine performance tool. Different test-cases are presented for which the calculated characteristic maps are compared to experimental ones.

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Aerodynamic Loss in a Gas Turbine Stage with Film Cooling

Cited by 31 publications

References 0 publications

The Influence of Film-Cooling on the Aerodynamic Performance of a Turbine Nozzle Guide Vane

The Influence of Film-Cooling on the Aerodynamic Performance of a Turbine Nozzle Guide Vane

Performance Analysis of a Turbine Stage Having Cooled Nozzle Blades With Trailing Edge Ejection

A quasi-one-dimensional CFD model for multistage turbomachines

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