Data set for Conventional Rib Turbulators for SCO2 Turbine Internal Cooling - DOE/NETL-2021/2758

Experimental investigation on rotation and turning vane effects on heat transfer was performed in a two-pass rectangular internal cooling channel. The channel has an aspect ratio of AR = 2:1 and a 180 deg tip-turn, which is a scaled up model of a typical internal cooling passage of gas turbine airfoils. The leading surface (LS) and trailing surface (TS) are roughened with 45 deg angled parallel ribs (staggered P/e = 8, e/Dh = 0.1). Tests were performed in a pressurized vessel (570 kPa) where higher rotation numbers (Ro) can be achieved with a maximum Ro = 0.42. Five Reynolds numbers (Re) were examined (Re = 10,000–40,000). At each Reynolds number, five rotational speeds (Ω = 0–400 rpm) were considered. Results showed that rotation effects are stronger in the tip regions as compared to other surfaces. Heat transfer enhancement up to four times was observed on the tip wall at the highest rotation number. However, heat transfer enhancement is reduced to about 1.5 times with the presence of a tip turning vane at the highest rotation number. Generally, the tip turning vane reduces the effects of rotation, especially in the turn portion.

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Investigation of Innovative Trailing Edge Cooling Configurations With Enlarged Pedestals and Square or Semicircular Ribs: Part II—Numerical Results

Volume 4: Heat Transfer, Parts A and B ◽

10.1115/gt2008-51048 ◽

2008 ◽

Cited By ~ 4

Author(s):

Emiliano Di Carmine ◽

Bruno Facchini ◽

Luca Mangani

Keyword(s):

Heat Transfer ◽

Cooling System ◽

Convective Heat Transfer Coefficient ◽

Turbulence Models ◽

Internal Cooling ◽

Trailing Edge ◽

Pressure Losses ◽

Structural Resistance ◽

Pin Fins ◽

Rib Turbulators

Trailing edge is a critical region for turbine airfoils since this part of the blade has to match aerodynamic, cooling and structural requirements at the same time. In fact aerodynamic losses are strictly related to trailing edge thickness which, on the contrary, tends to be increased to implement an internal cooling system, in order to face high thermal loads. At the moment the most employed devices consist of pin fins of various shapes, which contribute to both heat transfer enhancement and structural resistance improvement. Enlarged pedestals decrease pressure losses in comparison with multirow pin fins, even if the heat transfer increase is limited. This work deals with the investigation of the usage of enlarged pedestals, inserted in a wedge shaped duct, in conjunction with square or semicircular rib turbulators. The aim of the analysis is the evaluation of the convective Heat Transfer Coefficient (HTC) distribution over the endwall surface and the pressure drop of the converging duct. Numerical analysis used 3D RANS calculations. An in-house modified object-oriented CFD code and a commercial one were used. Several turbulence models and mesh types were tested. Numerical calculations were compared with experimental results obtained on the same geometries using a transient Thermochromic Liquid Crystals (TLC) based technique. Goals of this comparison are both the evaluation of the accuracy of CFD packages with standard two equation turbulence models in heat transfer problems with complex geometries and the analysis of flow details to complete and support experimental activity.

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Three Dimensional Measurements of Instantaneous Flow Field in Corners With Smooth Surfaces Under a Zero Pressure Gradient

Volume 3: Turbo Expo 2004 ◽

10.1115/gt2004-54304 ◽

2004 ◽

Author(s):

Michael Phillips ◽

Steve Deutsch ◽

Arnie Fontaine ◽

Savas Yavuzkurt

Keyword(s):

Heat Transfer ◽

Pressure Gradient ◽

Turbine Blades ◽

Three Dimensional ◽

Instantaneous Velocity ◽

Stream Velocity ◽

Internal Cooling ◽

Mean Velocity ◽

Data Set ◽

Corner Flow

Three dimensional instantaneous velocity data were taken in a turbulent corner flow with smooth walls under a zero pressure gradient. Experiments were carried out in air with a free stream velocity of 13 m/s and an axial Reynolds number of about 10,000,000. The data were collected using a three-component LDV system that was configured in a nearly orthogonal setup. Measurements were made down to a y+ of approximately 5, and should provide a valuable data set in developing models and predictive codes. Data were collected at two axial locations, 0.93 and 1.26 m measured from the virtual origin. The boundary layer thickness was 20.90 mm and 24.91 mm respectively at these locations. At each position, instantaneous velocity profiles were measured at 6.35, 12.7, 20.6, 41.2, 82.3, 121.9, 164.5, 184.8, and 205.1 mm from the corner. The centerline profiles agree well with classical flat plate data. Three mean velocity and six Reynolds stress components have been calculated. The instantaneous velocity field data set is sufficient to compute higher order correlations. The data will be valuable for development of computer codes and models for heat transfer studies in the internal cooling channels of gas turbine blades and turbine end wall flow and heat transfer studies. An analysis of the data is presented. Future studies will concentrate on one smooth and one rough wall corner flow with favorable and adverse pressure gradients to provide a detailed database for corner flows in complex three dimensional flow fields.

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Overall Cooling Effectiveness Measurements on Pressure Side Surface of the Nozzle Guide Vane With Optimized Film Cooling Hole Arrangements

Volume 5A: Heat Transfer ◽

10.1115/gt2017-63421 ◽

2017 ◽

Cited By ~ 2

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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Detailed Velocity and Heat Transfer Measurements in an Advanced Gas Turbine Vane Insert Using MRV and IR Thermometry

Volume 7A: Heat Transfer ◽

10.1115/gt2020-14984 ◽

2020 ◽

Author(s):

Michael J. Benson ◽

David Bindon ◽

Mattias Cooper ◽

F. Todd Davidson ◽

Benjamin Duhaime ◽

...

Keyword(s):

Heat Transfer ◽

Nusselt Number ◽

Gas Turbine ◽

Computational Models ◽

Internal Heat ◽

Internal Cooling ◽

Data Set ◽

Ir Camera ◽

Cooling Performance ◽

Turbine Vane

Abstract This work reports the results of paired experiments for a complex internal cooling flow within a gas turbine vane using Magnetic Resonance Velocimetry (MRV) and steady-state Infrared (IR) thermometry. A scaled model of the leading edge insert for a gas turbine vane with multi-pass impingement was designed, built using stereolithography (SLA) fabrication methods, and tested using MRV techniques to collect a three-dimensional, three-component velocity field data set for a fully turbulent test case. Stagnation and recirculation zones were identified and assessed in terms of impact on potential cooling performance. A paired experiment employed an IR camera to measure the temperature profile data of a thin, heated stainless steel impingement surface modeling the inside turbine blade wall cooled by the impingement from the vane cooling insert, providing complementary data sets. The temperature data allow for the calculation of wall heat transfer characteristics, including the Nusselt number distribution for cooling performance analysis to inform design and validate computational models. Quantitative and qualitative comparisons of the paired results show that the flow velocity and cooling performance are highly coupled. Module-to-module variation in the surface Nusselt number distributions are evident, attributable to the complex interaction between transverse and impinging flows within the apparatus. Finally, a comparison with internal heat transfer correlations is conducted using the data from Florschuetz [1]. Measurement uncertainty was assessed and estimated to be approximately +7% for velocity and ranging from +3% to +10% for Nusselt number.

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Sand Transport in a Two Pass Internal Cooling Duct With Rib Turbulators

Volume 1: Heat Transfer in Energy Systems; Theory and Fundamental Research; Aerospace Heat Transfer; Gas Turbine Heat Transfer; Transport Phenomena in Materials Processing and Manufacturing; Heat and ◽

10.1115/ht2012-58100 ◽

2012 ◽

Author(s):

Sukhjinder Singh ◽

Colin Reagle ◽

Jacob Delimont ◽

Danesh Tafti ◽

Wing Ng ◽

...

Keyword(s):

Sand Transport ◽

Internal Cooling ◽

Rib Turbulators

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Conjugate Heat Transfer Analysis in an Actual Gas Turbine Rotor Blade in Comparison With Pyrometer Data

Volume 5A: Heat Transfer ◽

10.1115/gt2014-26962 ◽

2014 ◽

Author(s):

Kazuhiro Tsukamoto ◽

Yasuhiro Horiuchi ◽

Kazuyuki Sugimura ◽

Shinichi Higuchi

Keyword(s):

Heat Transfer ◽

Gas Turbine ◽

Rotor Blade ◽

Conjugate Heat Transfer ◽

Rotor Blades ◽

Experimental Results ◽

Internal Cooling ◽

Pressure Side ◽

Temperature Distributions ◽

Rib Turbulators

Conjugate Heat Transfer (CHT) was analyzed in a first stage rotor blade in an actual gas turbine. The main objectives of this research were to simulate and validate improvements to the accuracy of predicting temperature on the surfaces of rotor blades in a gas turbine and compare these with experimental results. This simulation was carried out under similar conditions to those during gas turbine operation. Computational grids were generated based on CAD data obtained from the rotor blades with fully resolved rib turbulators and pin fins for both fluid and solid domains during CHT analysis. A tetrahedral mesh with prism layers was used and the y+ of the first mesh adjacent to the wall was kept at less than 1.0 over the whole surface. Thermal barrier coating was modeled by adding thermal resistance at the fluid-solid interfaces. Inlet boundary conditions for the external- and internal-gas-flow regions were defined based on one-dimensional analysis and measured results. Steady Reynolds-averaged Navier-Stokes simulation was carried out using the Shear Stress Transport (SST) turbulence model. The simulated results were compared with measured data obtained from a pyrometer and thermocouple. The temperature distributions predicted from CHT analysis agreed with those obtained from an experiment near the leading edge of the rotor blades. However, the temperature distribution at the center of the pressure side had a difference of 50 K with that obtained from the experimental data. The heat transfer coefficients on the surfaces of the blades were almost equal to those on the pressure side. Thus, we considered that the internal cooling flows contributed more to temperature distributions on the surfaces of the blades rather than the external gas flows. The main stream in the internal cooling flow passages leaned toward one side of the walls and the temperatures on this side became lower than those obtained from the experimental results. Therefore, we suspect CHT analysis underestimated the mixing effect generated by the rib turbulators. It is important to solve the complex flow phenomena in internal cooling passages to better predict the accuracy of temperature distributions on the surfaces of blades.

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