Aerodynamic and thermal performance of a three dimensional annular transonic nozzle guide vane

Arts, Tony; Heider, R.

doi:10.2514/6.1994-2929

Cited by 12 publications

(10 citation statements)

References 6 publications

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“…Here, IAL values increase as the exit freestream Mach number increases for each value of k s /cx. This is consistent with results from Arts [34] and Xu and Denton [9], whose experimental and analytical results show that total pressure losses increase approximately with the square of the Mach number. Figure 7 also shows that the largest IAL magnitude increases are present with the largesized roughness (k s /cx = 0.00164).…”

Section: Integrated Aerodynamic Lossessupporting

confidence: 91%

“…Radomsky and Thole [33] present measurements of time-averaged velocity components and Reynolds stresses along a turbine stator vane at elevated freestream turbulence levels and present data which show that transition occurs further upstream on the suction side, as the freestream turbulence level increases. Arts [34] describes experimental aerodynamic performance data for a three-dimensional annular transonic nozzle guide vane. Coton et al [35] investigate the effects of Reynolds number and Mach number on the profile losses of a conventional low-pressure turbine rotor cascade and report that the exit Mach number affects the losses through a modification of the pressure gradient imposed on the boundary layer.…”

Section: Introductionmentioning

confidence: 99%

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Aerodynamic Losses in Turbines with and without Film Cooling, as Influenced by Mainstream Turbulence, Surface Roughness, Airfoil Shape, and Mach Number

Ligrani

2012

International Journal of Rotating Machinery

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The influences of a variety of different physical phenomena are described as they affect the aerodynamic performance of turbine airfoils in compressible, high-speed flows with either subsonic or transonic Mach number distributions. The presented experimental and numerically predicted results are from a series of investigations which have taken place over the past 32 years. Considered are (i) symmetric airfoils with no film cooling, (ii) symmetric airfoils with film cooling, (iii) cambered vanes with no film cooling, and (iv) cambered vanes with film cooling. When no film cooling is employed on the symmetric airfoils and cambered vanes, experimentally measured and numerically predicted variations of freestream turbulence intensity, surface roughness, exit Mach number, and airfoil camber are considered as they influence local and integrated total pressure losses, deficits of local kinetic energy, Mach number deficits, area-averaged loss coefficients, mass-averaged total pressure loss coefficients, omega loss coefficients, second law loss parameters, and distributions of integrated aerodynamic loss. Similar quantities are measured, and similar parameters are considered when film-cooling is employed on airfoil suction surfaces, along with film cooling density ratio, blowing ratio, Mach number ratio, hole orientation, hole shape, and number of rows of holes.

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Section: Integrated Aerodynamic Lossessupporting

confidence: 91%

Section: Introductionmentioning

confidence: 99%

Aerodynamic Losses in Turbines with and without Film Cooling, as Influenced by Mainstream Turbulence, Surface Roughness, Airfoil Shape, and Mach Number

Ligrani

2012

International Journal of Rotating Machinery

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“…Dunn and Hause (1982) have developed pressure and heat flux measurement techniques for full-scale rotating hardware in the shock tunnel environment, with test times of tens of milliseconds. More recently, combined aerodynamic and heat transfer measurements have been gathered in annular cascades on inlet guide vanes (Arts 1994;Chana 1992).…”

Section: Introductionmentioning

confidence: 99%

High Pressure Turbine Vane Annular Cascade Heat Flux and Aerodynamic Measurements With Comparisons to Predictions

Joe

Montesdeoca

Soechting

et al. 1998

Volume 4: Heat Transfer; Electric Power; Industrial and Cogeneration

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Experimental tests were performed at the USAF Turbine Research Facility (TRF) to obtain heat transfer and aerodynamic data on a first stage vane of a modern high pressure turbine. This is a transient blowdown facility that provides data from short duration tests. Data for a matrix of test conditions were obtained to document the effect of inlet Reynolds number, the stage pressure ratio across the vane, and the gas-to-wall temperature ratio. The objectives of these tests were to assess the capability of obtaining accurate aerodynamic total pressure loss measurements and airfoil static pressure measurements as well as determine the heat transfer coefficient distributions on the vanes. Results from these tests were compared to analytical predictions and are presented. The unique contribution of the work presented herein is: 1) demonstration of circumferential traversing temperature and pressure data in a short duration facility test, and 2) heat loss closure during a short duration test using heat flux gauges and the measured temperature loss. The transient heat loss during a short duration test is a fundamental requirement to determine turbine efficiency when work extraction is determined from the temperature drop across the turbine stage. Heat transfer data were acquired from heat flux gauges that were fabricated using thin-film sputtering techniques and placed on the airfoil surfaces. The surface temperature of the gauge was measured and heat flux was determined from a closed form transient semi-infinite solution that included the resistance of the heat flux gauge and the underlying metal substrate. Circumferentially, pressure measurements were obtained on the airfoil surfaces and on traversing rakes at the inlet and exit of the vane test section. Total and differential pressure rake instrumentation was required to obtain accurate aerodynamic loss measurements over a range of gas-to-wall temperature ratios.

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“…York et al(1984) obtained data in a linear cascade for a Table I. -Description of cases used in analysis Vanes Source of data Re3cX10 -' MEXIT TWIT Cascade Test approach York et al(1984) 2.1-18.0 0.3-1.1 0.75 Linear Steady state Arts and Heider(1994) 22.5 0.92-1.15 0.73 Annular Shock tube Boyle and Russell(1990) 3.4-20 0.1-0.7 1.1 Linear Liquid Crystals Harasagama and Wedlake(1990) 17-52 0.94-1.29 0.67 Annular Shock tube Rotors Source of data Re2, X10-5 MExyr TwIT; Cascade Test approach Blair(1994) 2.7-7.1 0.06-0.15 1.1 Rotating Steady state Goldstein and Spores(1988) 1.4-2.3 0.03-0.04 1 Linear Napthalene Chen and Goldstein(1992) 1.2-2.0 0.02-0.03 1 Linear Napthalene Graziani et al(1980) 10.7 0.1 1.1 Linear Steady state Giel et al(1996) 13-26 1.0-1.3 1.1 Linear Steady state York et al(1984) 7.0 Y N 1.0 1.56 4 Arts and Heider(1994) 4.5 Y Y 0.73 1.56 1 Boyle and Russell(1990) 1.0 Y N 1.0 1.51 2 Harasagama and Wedlake(1990) 6.5 Y Y 0.67 1.47 5 Rotors Source of data Tu% Heat Transfer (T) 1 /2; Ec No. of Endwall Blade Cases Blair(1994) High Y Y 1.0 0.15 3 Goldstein and Spores(1988) 1.2 Y N 1.0 N.A.…”

Section: Subscriptsmentioning

confidence: 99%

Predicted Turbine Heat Transfer for a Range of Test Conditions

Boyle

Lucci

1996

Volume 4: Heat Transfer; Electric Power; Industrial and Cogeneration

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Aerodynamic and thermal performance of a three dimensional annular transonic nozzle guide vane

Cited by 12 publications

References 6 publications

Aerodynamic Losses in Turbines with and without Film Cooling, as Influenced by Mainstream Turbulence, Surface Roughness, Airfoil Shape, and Mach Number

Aerodynamic Losses in Turbines with and without Film Cooling, as Influenced by Mainstream Turbulence, Surface Roughness, Airfoil Shape, and Mach Number

High Pressure Turbine Vane Annular Cascade Heat Flux and Aerodynamic Measurements With Comparisons to Predictions

Predicted Turbine Heat Transfer for a Range of Test Conditions

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