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Title: State-of-the-Art Cooling Technology for a Turbine Rotor Blade

Abstract

Effective internal and external cooling of airfoils is key to maintaining component life for efficient gas turbines. Cooling designs have spanned the range from simple internal convective channels to more advanced double-walls with shaped film-cooling holes. This paper describes the development of an internal and external cooling concept for a state-of-the-art cooled turbine blade. These cooling concepts are based on a review of literature and patents, as well as, interactions with academic and industry turbine cooling experts. The cooling configuration selected and described in this paper is referred to as the “baseline” design, since this design will simultaneously be tested with other more advanced blade cooling designs in a rotating turbine test facility using a “rainbow turbine wheel” configuration. For the baseline design, the leading edge is cooled by internal jet impingement and showerhead film cooling. The midchord region of the blade contains a three-pass serpentine passage with internal discrete V-shaped trip strips to enhance the internal heat transfer coefficient (HTC). The film cooling along the midchord of the blade uses multiple rows of shaped diffusion holes. Furthermore, the trailing edge is internally cooled using jet impingement and externally film cooled through partitioned cuts on the pressure side of themore » blade.« less

Authors:
 [1];  [2];  [3];  [1];  [4]
  1. Pennsylvania State Univ., University Park, PA (United States)
  2. National Energy Technology Lab. (NETL), Morgantown, WV (United States)
  3. National Energy Technology Lab. (NETL), Pittsburgh, PA, (United States)
  4. Purdue Univ., West Lafayette, IN (United States)
Publication Date:
Research Org.:
National Energy Technology Lab. (NETL), Morgantown, WV (United States)
Sponsoring Org.:
USDOE
OSTI Identifier:
1461194
Report Number(s):
NETL-PUB-21762
Journal ID: ISSN 0889-504X
Resource Type:
Accepted Manuscript
Journal Name:
Journal of Turbomachinery
Additional Journal Information:
Journal Volume: 140; Journal Issue: 7; Journal ID: ISSN 0889-504X
Publisher:
ASME
Country of Publication:
United States
Language:
English
Subject:
01 COAL, LIGNITE, AND PEAT; 03 NATURAL GAS; 20 FOSSIL-FUELED POWER PLANTS; 42 ENGINEERING; gas turbines; film cooling; leading edge cooling; trailing edge cooling; blade cooling

Citation Formats

Town, Jason, Straub, Douglas, Black, James, Thole, Karen A., and Shih, Tom I-P. State-of-the-Art Cooling Technology for a Turbine Rotor Blade. United States: N. p., 2018. Web. doi:10.1115/1.4039942.
Town, Jason, Straub, Douglas, Black, James, Thole, Karen A., & Shih, Tom I-P. State-of-the-Art Cooling Technology for a Turbine Rotor Blade. United States. doi:10.1115/1.4039942.
Town, Jason, Straub, Douglas, Black, James, Thole, Karen A., and Shih, Tom I-P. Thu . "State-of-the-Art Cooling Technology for a Turbine Rotor Blade". United States. doi:10.1115/1.4039942. https://www.osti.gov/servlets/purl/1461194.
@article{osti_1461194,
title = {State-of-the-Art Cooling Technology for a Turbine Rotor Blade},
author = {Town, Jason and Straub, Douglas and Black, James and Thole, Karen A. and Shih, Tom I-P.},
abstractNote = {Effective internal and external cooling of airfoils is key to maintaining component life for efficient gas turbines. Cooling designs have spanned the range from simple internal convective channels to more advanced double-walls with shaped film-cooling holes. This paper describes the development of an internal and external cooling concept for a state-of-the-art cooled turbine blade. These cooling concepts are based on a review of literature and patents, as well as, interactions with academic and industry turbine cooling experts. The cooling configuration selected and described in this paper is referred to as the “baseline” design, since this design will simultaneously be tested with other more advanced blade cooling designs in a rotating turbine test facility using a “rainbow turbine wheel” configuration. For the baseline design, the leading edge is cooled by internal jet impingement and showerhead film cooling. The midchord region of the blade contains a three-pass serpentine passage with internal discrete V-shaped trip strips to enhance the internal heat transfer coefficient (HTC). The film cooling along the midchord of the blade uses multiple rows of shaped diffusion holes. Furthermore, the trailing edge is internally cooled using jet impingement and externally film cooled through partitioned cuts on the pressure side of the blade.},
doi = {10.1115/1.4039942},
journal = {Journal of Turbomachinery},
number = 7,
volume = 140,
place = {United States},
year = {2018},
month = {6}
}

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