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Research Papers

Aerothermal Impact of Stator-Rim Purge Flow and Rotor-Platform Film Cooling on a Transonic Turbine Stage

[+] Author and Article Information
M. Pau, G. Paniagua, D. Delhaye, A. de la Loma

Department of Turbomachinery and Propulsion, von Karman Institute for Fluid Dynamics, 1640 Rhode Saint Genèse, Belgium

P. Ginibre

Department of Turbine Aerodynamics, Snecma-Groupe Safran, 77550 Moissy Cramayel, France

J. Turbomach 132(2), 021006 (Jan 11, 2010) (12 pages) doi:10.1115/1.3142859 History: Received May 02, 2008; Revised February 14, 2009; Published January 11, 2010; Online January 11, 2010

The sealing of the stator-rotor gap and rotor-platform cooling are vital to the safe operation of the high-pressure turbine. Contrary to the experience in subsonic turbines, this paper demonstrates the potential to improve the efficiency in transonic turbines at certain rim seal rates. Two types of cooling techniques were investigated: purge gas ejected out of the cavity between the stator rim and the rotor disk, and cooling at the rotor-platform. The tests were carried out in a full annular stage fed by a compression tube at M2is=1.1, Re=1.1×106, and at temperature ratios reproducing engine conditions. The stator outlet was instrumented to allow the aerothermal characterization of the purge flow. The rotor blade was heavily instrumented with fast-response pressure sensors and double-layer thin film gauges. The tests are coupled with numerical calculations performed using the ONERA’s code ELSA . The results indicate that the stator-rotor interaction is significantly affected by the stator-rim seal, both in terms of heat transfer and pressure fluctuations. The flow exchange between the rotor disk cavity and the mainstream passage is mainly governed by the vane trailing edge shock patterns. The purge flow leads to the appearance of a large coherent vortex structure on the suction side of the blade, which enhances the overall heat transfer coefficient due to the blockage effect created. The impact of the platform cooling is observed to be restricted to the platform, with negligible effects on the blade suction side. The platform cooling results in a clear attenuation of pressure pulsations at some specific locations. The experimental and computational fluid dynamics results show an increase in the turbine performance compared with the no rim seal case. A detailed loss breakdown analysis helped to identify the shock loss as the major loss source. The presented results should help designers improve the protection of the rotor platform while minimizing the amount of coolant used.

Copyright © 2010 by American Society of Mechanical Engineers
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References

Figures

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Figure 3

(a) Coolant flow paths and (b) pressure signals along the coolant path during a typical test

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Figure 2

(a) Meridional view of the turbine rig, (b) turbine measurement planes, (c) total pressure sensors located at the rotor leading edge, and (d) Heat transfer instrumentation at 15% span, 7.5% span, and on the rotor platform

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Figure 1

Turbine stage and rotor-platform coolant path

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Figure 13

Turbine exit flow field, experiments, and CFD

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Figure 14

Measured turbine efficiency and windage effects compared with the ingestion −1%

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Figure 15

Turbine efficiency predicted using Kacker and Okapuu (43) correlations, with reference to the ingestion −1%

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Figure 4

Structure of the blocks used to mesh the rotor

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Figure 5

(a) CFD static pressure distribution on the stator-rim cavity, (b) measurement locations on the stator rim at R/R0=0.944, (c) experimental pulsations of static pressure for all the conditions, and (d) enlarged detail of the unsteady static pressure amplitudes at the blade passing frequency

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Figure 6

Pitchwise static pressure variation at the stator rim

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Figure 7

Interaction of the hot mainstream flow with the cold purge flow: (a) stream traces of the cold rim seal and hot mainstream flow on the 3D rotor static pressure field, (b) isotemperature contours at blade midpitch, and (c) isotemperature contours in an axial plane aligned with the rotor leading edge

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Figure 8

(a) Time-averaged total pressure in the relative frame on the rotor leading edge and (b) time-resolved total relative pressure at 15% span

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Figure 9

Nusselt number distribution along the rotor airfoil: (top) at 15% span and (bottom) at 7.5% span

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Figure 10

Static pressure field on the rotor platform without cavity flow and for ejection +0.8% (streamlines through the blade passage do not line-up due to the computational-domain)

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Figure 11

(a) Pressure measurement locations on the rotor platform, (b) photograph of the instrumented blade, (c) time-averaged static pressure on the rotor platform, and (d) time-resolved static pressure downstream of the coolant holes

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Figure 12

(Left) Detailed view of the heat transfer sensor located downstream of the cooling holes and (right) Nusselt number at the same sensor

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