Influence of Coolant Flow Rate on Aero-Thermal Performance of a Rotor Blade Cascade With Endwall Film Cooling

Abstract

This paper investigates the influence of coolant injection on the aerodynamic and thermal performance of a rotor blade cascade with endwall film cooling. A seven blade cascade of a high-pressure-rotor stage of a real gas turbine has been tested in a low speed wind tunnel for linear cascades. Coolant is injected through 10 cylindrical holes distributed along the blade pressure side. Tests have been preliminarily carried out at low Mach number (Ma2is = 0.3). Coolant-to-mainstream mass flow ratio has been varied in a range of values corresponding to inlet blowing ratios M1 = 0–4.0. Secondary flows have been surveyed by traversing a five-hole miniaturized aerodynamic probe in two downstream planes. Local and overall mixed-out secondary loss coefficient and vorticity distributions have been calculated from measured data. The thermal behavior has been also analyzed by using thermochromic liquid crystals technique to obtain film cooling effectiveness distributions. All this information, including overall loss production for variable injection conditions, allows us to draw a comprehensive picture of the aero-thermal flow field in the endwall region of a high pressure rotor blade cascade.

The Aero-Thermal Performance of Purge Flow and Discrete Holes Film Cooling of Rotor Blade Platform in Modern High Pressure Gas Turbines: A Review

“…Measurements highlighted that for this positioning, the impact on losses was negligible because the coolant flow did not interact with the horseshoe vortex and passage vortex generation, but instead it only energised the platform boundary layer. As also shown in [21,24], testing different hole patterns, still located inside of the passage, the platform flow field is mainly influenced by the passage vortex and its interaction with coolant flow. When the injected cool air is not captured by the vortex, only the end wall crossflow acts, and secondary flows are barely affected.…”

“…Indeed, the arrangement of cooling holes within the platform is a key issue. At Bergamo University a research activity was carried out over some years to design and verify the introduction of discrete holes at different locations: close to the pressure side, to exploit the passage vortex cross flow [21], along the inter-platform gap [24] and rear in the passage, downstream the throat [26]. The aim was always to effectively protect specific regions of the platform particularly exposed to hot gas contact, while, in the meantime, considering the constraints imposed by the secondary air system.…”

“…The former family of experimental facilities may simulate rotation effects through swirl motion or fin introduction. Most of these researches were carried out at Texas A and M University on a five-blade cascade test facility [13][14][15][16][17][18][19][20], at the University of Bergamo on a seven-blade cascade test facility [21][22][23][24][25][26] and at Whittle Laboratory of the University of Cambridge [27,28]. Meanwhile, the latter family of experiments (rotating platforms) were carried out in the TPFL-research turbine facility at Texas A and M University [29,30], in the CT3 compression tube turbine test rig at the von Karman Institute [31], at Ohio State University Turbine Test Facility (TTF) [32][33][34], at the LISA research turbine available at ETH Zurich [35,36] and, more recently, at the Large Annulus Rig (LAR) of the University of Bath [37,38].…”

“…Similarly, in Figure 3, some of the considered discrete cooling hole designs are gathered to illustrate their positioning within the blade channel, e.g., holes located on the fileted end wall along the blade pressure side as in [21] and distributed along the interplatform gap [24], downstream half of the passage towards the trailing edge as in [20,26,30] and over the whole platform, as in [13,14]. [22], (b) S2 [13] and (c) S1s [18].…”

Meander-Like Shape Optimization of the Stator-Rotor Platform Cavity

“…They determined that rotor platform cooling has minimal effects on the suction side and trailing edge of the rotor. Barigozzi et al [6] also suggested that this is due to the coolant flow from the hub disk being captured by the rotor hub passage vortex which moves the coolant away from the rotor endwalls. Suryanarayanan et al [12,13] showed using pressuresensitive paint that coolant supplied by purge flow did not provide sufficient coolant to the pressure side of the rotor.…”