James D. Heidmann NASA Glenn Research Center, Cleveland, OH 44135 Srinath Ekkad Louisiana State University, Baton Rouge, LA 70803 A Novel Antivortex Turbine Film-Cooling Hole Concept A novel turbine film-cooling hole shape has been conceived and designed at NASA Glenn Research Center. This “antivortex” design is unique in that it requires only easily ma- chinable round holes, unlike shaped film-cooling holes and other advanced concepts. The hole design is intended to counteract the detrimental vorticity associated with standard circular cross-section film-cooling holes. This vorticity typically entrains hot freestream gas and is associated with jet separation from the turbine blade surface. The antivortex film-cooling hole concept has been modeled computationally for a single row of 30 deg angled holes on a flat surface using the 3D Navier–Stokes solver GLENN-HT. A blowing ratio of 1.0 and density ratios of 1.05 and 2.0 are studied. Both film effectiveness and heat transfer coefficient values are computed and compared to standard round hole cases for the same blowing rates. A net heat flux reduction is also determined using both the film effectiveness and heat transfer coefficient values to ascertain the overall effectiveness of the concept. An improvement in film effectiveness of about 0.2 and in net heat flux reduction of about 0.2 is demonstrated for the antivortex concept compared to the stan- dard round hole for both blowing ratios. Detailed flow visualization shows that as ex- pected, the design counteracts the detrimental vorticity of the round hole flow, allowing it to remain attached to the surface. DOI: 10.1115/1.2777194 Introduction Film cooling is commonly used on high pressure turbine vanes and blades to enable increased turbine inlet temperatures for im- proved engine cycle performance. Relatively cool air is bled from the compressor to supply this film-cooling flow to the turbine. However, this compressor bleed represents a loss to the system and should be minimized. There has thus been a substantial effort to reduce the film-cooling flow rate required to provide the nec- essary cooling. Many new film-cooling hole shapes and concepts have been proposed in the literature to address this issue. Goldstein 1 summarized early studies in the area of film cool- ing, primarily focusing on cooling flows issuing from slots. In a gas turbine engine, slots are typically not practical, so the flow must be introduced through discrete holes. Many published stud- ies have discussed the physics and presented data for discrete hole film cooling in various geometries. Kercher 2,3 presents an ex- haustive list of film-cooling references from the literature. The most basic film-cooling geometry consists of a row of round holes in a flat plate. There is a relatively large body of experimental data for 30–35 deg round holes with a pitch of 3d Pedersen et al. 4, Foster and Lampard 5, Pietrzyk et al. 6, Pietrzyk et al. 7, and Sinha et al. 8. This geometry allows for a study of jet lift-off behavior at various blowing ratios and is perhaps the most realis- tic simplified geometry for turbine film cooling. In addition, the computational study of Leylek and Zerkle 9 uses this geometry and gives an excellent description of the vortical flows associated with this geometry. The present study will use the data of Dhungel et al. 10 as the base line case, which consists of 30 deg round holes and a pitch of 3d. A well-known and extensively studied property of the flow from a round film-cooling hole is the counter-rotating vortex pair that causes the cooling jet to separate from the surface at suffi- ciently high blowing ratio. This phenomenon is shown in Fig. 1. The counter-rotating vortex pair entrains the hot freestream gas and lifts the coolant away from the surface, drastically reducing its effectiveness. The counter-rotating vortex pair is described by Haven et al. 11. Lemmon et al. 12 showed that this vorticity is caused by the bending of the jet by the freestream and not by viscous wall effects in the hole or plenum. For cases with varying density ratio, the momentum ratio is considered to be an even better predictor of jet lift-off than blowing ratio since higher den- sity ratio jets will tend to remain attached to the surface at higher blowing ratios. The jet lift-off phenomenon typically occurs at momentum ratios above about 0.5. For very low momentum ratios such as 0.25, the circular cross-section hole produces lower vor- ticity levels and stays attached to the surface, providing excellent cooling. However, this momentum ratio is typically unachievable in an engine, as the available coolant pressure unavoidably pro- duces higher coolant flow rates. Shaped film-cooling holes have typically been used to combat the jet lift-off behavior. Bunker 13 provides an excellent over- view of the shaped film-cooling hole literature. The primary effect of the hole shaping is to expand the exit area of the hole, thereby reducing jet velocity. However, shaped film-cooling holes can be expensive to manufacture and can be limited by flow and thermal barrier coating TBC limitations in their expansion angles and other parameters. Many other unique film-cooling designs have been proposed over the years to improve film-cooling effective- ness. Besides the typical shaped hole with spanwise and/or streamwise expanded exits, various exotic hole shapes have been proposed. Ideas such as struts within the holes Shih et al. 14, cusp-shaped holes Papell 15, triangular tabs at the hole exit Zaman and Foss 16, and trenched holes Bunker 17 have been proposed and studied. A common theme in these designs is the desire to offset the detrimental vorticity caused by the round hole jet. Many of the proposed designs are shown to be effective in this regard, but the majority of these ideas prove to be difficult to manufacture and/or produce unwanted features such as addi- tional solid surfaces that must be cooled or additional sharp edges that are aerodynamic liabilities. Another area of recent research has been in the area of pulsed film-cooling flows e.g., Ekkad et al. 18. The film-cooling flow is pulsed intentionally with the goal of producing an improvement in time-averaged cooling effectiveness. This concept is perhaps most effective for applications where a high blowing ratio would be present for the nonpulsed case. If such a film-cooling flow is Contributed by the International Gas Turbine Institute of ASME for publication in the JOURNAL OF TURBOMACHINERY. Manuscript received June 7, 2007; final manuscript received June 18, 2007; published online May 6, 2008. Review conducted by David Wisler. Paper presented at the ASME Turbo Expo 2007: Land, Sea and Air GT2007, Montreal, Quebec, Canada, May 14–17, 2007, Paper No. GT2007-27528. Journal of Turbomachinery JULY 2008, Vol. 130 / 031020-1 Copyright © 2008 by ASME Downloaded 24 Apr 2009 to 128.156.10.80. Redistribution subject to ASME license or copyright; see http://www.asme.org/terms/Terms_Use.cfm