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3 edition of Heat transfer in a complex trailing edge passage for a high pressure turbine blade. found in the catalog.

Heat transfer in a complex trailing edge passage for a high pressure turbine blade.

Heat transfer in a complex trailing edge passage for a high pressure turbine blade.

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Published by National Aeronautics and Space Administration, Glenn Research Center, Available from NASA Center for Aerospace Information in [Cleveland, Ohio], Hanover, MD .
Written in English

    Subjects:
  • Trailing edges.,
  • Turbine blades.,
  • Computerized simulation.,
  • High pressure.,
  • Heat transfer coefficients.,
  • Three dimensional flow.

  • Edition Notes

    Other titlesSimulation results.
    StatementDavid L. Rigby, Ronald S. Bunker.
    Series[NASA contractor report] -- NASA/CR-2002-211701., NASA contractor report -- NASA CR-211701.
    ContributionsBunker, Ronald S., NASA Glenn Research Center.
    The Physical Object
    FormatMicroform
    Pagination1 v.
    ID Numbers
    Open LibraryOL16111312M

    This study presents experimental and numerical investigation for three-dimensional heat transfer characteristics in a turbine blade. An experimental setup was installed with a turbine cascade of five-blade channels. Blade heat transfer measurements were performed for the middle channel under uniform heat flux boundary conditions. Heat was supplied to the blades using twenty-nine electric. comparing the predictions of the heat transfer around the experimental high pressure turbine blade profile cascade of Arts et al. (). First, both LES predictions are compared to RANS modeling with a particular interest for the accuracy/cost ratio and improvement of the physical phenomena around the blade. LES’s are then detailed and further.

      Heat transfer in a complex trailing edge passage for a high pressure turbine blade—part 1: experimental measurements. ASME paper GT Coletti F, Armellini A, Arts T, et al. Aero-thermal investigation of a rib-roughened trailing edge channel with crossing-jets—part II: heat transfer analysis. ASME paper GT 1. Bunker, Ronald S., et al.: Heat Transfer in a Complex Trailing Edge Passage for a High Pressure Turbine Blade. ASME Paper GTPt-1, 2. Rigby, David L.; and Bunker, Ronald S.: Heat Transfer in a Complex Trailing Edge Passage for a High Pressure Turbine Blade. Part 2: Simulation Results, NASA/CR,

    This experimental study contains two points; part1 – turbine blade heat transfer under low Reynolds number flow conditions, and part 2 – trailing edge cooling and heat transfer. The effect of unsteady wake and free stream turbulence on heat transfer and pressure coefficients of a turbine blade was investigated in low Reynolds number flows. The heat transfer in the trailing edge passage is the weakest among the three passages. Affect by the turning and bleeding at trailing edge, the air flow into the passage deviate to the trailing edge. Part of airflow bleeds out from the hole in trailing edge. The backflow region near the division is .


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Heat transfer in a complex trailing edge passage for a high pressure turbine blade Download PDF EPUB FB2

Rigby, David L., and Bunker, Ronald S. "Heat Transfer in a Complex Trailing Edge Passage for a High Pressure Turbine Blade: Part 2 — Simulation Results." Proceedings of the ASME Turbo Expo Power for Land, Sea, and Air.

Volume 3: Turbo ExpoParts A and B. Amsterdam, The Netherlands. June 3–6, pp. by: 4. Bunker, Ronald S., Wetzel, Todd G., and Rigby, David L. "Heat Transfer in a Complex Trailing Edge Passage for a High Pressure Turbine Blade: Part 1 — Experimental Measurements." Proceedings of the ASME Turbo Expo Power for Land, Sea, and Air.

Volume 3: Turbo ExpoParts A and B. Amsterdam, The Netherlands. June 3–6, pp. Cited by: Request PDF | Heat Transfer in a Complex Trailing Edge Passage for a High Pressure Turbine Blade | A combined experimental and numerical study to investigate the heat transfer distribution in a.

The trailing edge region of gas turbine blades is generally subjected to extremely high external heat loads due to the combined effects of high mach numbers and gas temperatures.

same low-pressure turbine gh the heat loads for low- pressure turbine blades are not as critical as for high-pressure turbine blades, part durability is still a design consideration, espe. Heat transfer coefficients are experimentally measured in a rotating cooling channel used to model an internal cooling passage near the trailing edge of a gas turbine blade.

The r. An experimental study of the heat transfer distribution and pressure drop through a converging lattice-matrix structure has been performed. This structure represents a gas turbine blade trailing-edge cooling passage. Stationary tests were performed on a scaled up model under three Reynolds numbers (   The thermal analysis is based on a simplified conjugate heat transfer analysis of flow in a cooling passage of a turbine blade as shown in Figure 7 and assumes that the airofoil; (a) metal temperature is the average surface temperature at the airofoil midspan, (b) is exposed to the maximum hot gas temperature profile at the blade inlet, and (c.

The present study deals with trailing edge film cooling on the pressure side cutback of gas turbine airfoils. Before being ejected tangentially onto the inclined cut-back surface the coolant air passes a partly converging passage that is equipped with turbulators such as pin fins and ribs.

In this study, a high pressure turbine 1st stage blade was considered. A high pressure turbine nozzle geometry in this study, which has 56 nozzles in total, is explained by Seo et al.

The number and geometry of the blade is described in Table 1.A high pressure turbine 1st stage blade is fully designed with internal and external cooling scheme, shown in Fig.

Abstract An experimental investigation is conducted to obtain the heat transfer and pressure drop data for an integral trailing edge cavity test section that simulates a novel turbine blade’s internal cooling passage with bleed holes. Local heat transfer is measured on both the suction and pressure.

Add tags for "Heat transfer in a complex trailing edge passage for a high pressure turbine blade. 2, Simulation results". 2, Simulation results".

Be the first. passage of a turbine blade as shown in Figure 7 and assumes that the Heat transfer in trailing edge passages with di erent pin bank con high pressure transonic turbine including.

Turbine blade cooling channels are flow passages having multiple inlets and exits. The present in-house developed solver uses a network method for analyzing such a complicated flow pattern.

The heat transfer coefficient is defined as, (2) h = q T w-T ∞ where q is the heat flux on the blade tip, and T ∞ is the inflow temperature. In the heat transfer coefficient computations, the blade surface is set to K and the other walls (hub and shroud) are set to be adiabatic.

Details of local heat transfer on the leading and trai- ling edges of the blade are treated initially and this is followed by an analysis of the full axial and cir- cumferential heat transfer behaviour.

This is referred to as full field data. Heat transfer in a simulated turbine blade cooling passage. Kim et al. studied the analysis of conjugated heat transfer, stress and failure in a gas turbine blade with circular cooling passages. The result showed the highest heat transfer coefficients occurred at the stagnation point of the leading edge where as the lowest heat transfer coefficient occurred at the trailing edge.

A turbine blade is the individual component which makes up the turbine section of a gas turbine or steam blades are responsible for extracting energy from the high temperature, high pressure gas produced by the turbine blades are often the limiting component of gas turbines.

To survive in this difficult environment, turbine blades often use exotic materials like. Heat Transfer in a Complex Trailing Edge Passage for a High Pressure Turbine Blade A combined experimental and numerical study to investigate the heat transfer distribution in a complex blade trailing edge passage was conducted.

The geometry consists of a two pass serpentine passage with taper toward the trailing edge, as well as from hub to tip. Model configurations of turbine blade trailing-edge internal cooling passage with staggered elliptic pin-fins in streamwise and spanwise are adopted for numerical investigation using computational fluid dynamics (CFD).

Grid refinement study is performed at first to identify a baseline mesh, followed by validation study of passage total pressure loss, which gives 2% and 4% discrepancies.

Heat Transfer in a Complex Trailing Edge Passage for a High Pressure Turbine Blade: Part 1 — Experimental Measurements GT Large-Eddy Simulation and Conjugate Heat Transfer Around a Low-Mach Turbine Blade.HEAT TRANSFER IN A COMPLEX TRAILING EDGE PASSAGE FOR A HIGH PRESSURE TURBINE BLADE -PART 1: EXPERIMENTAL MEASUREMENTS Ronald S.

Bunker and Todd G. Wetzel General Electric R&D Center Schenectady, New York David L. Rigby ass, Inc. NASA Glenn Research Center Cleveland, Ohio 35 ABSTRACT.A combined experimental and computational study has been performed to investigate the detailed heat transfer coefficient distributions within a complex blade trailing edge passage.

The experimental measurements are made using a steady liquid crystal thermography technique applied to one major side of the passage.