Gas turbines are widely used for aircraft propulsion, land-base power generation, and other industrial applications like trains, marines, automobiles, etc. To satisfy the fast development of advanced gas turbines, the operating temperature must be increased to improve the thermal efficiency and output work of the gas turbine engine. However, the heat transferred to the turbine blade is substantially increased as the turbine inlet temperature is continuously increased. Thus, it is very important to cool the turbine blades for a long durability and safe operation. Cooling the blade must include cooling of the key regions being exposed to the hot gas. The blade tip region is such a critical area and is indeed difficult to cool. This results from the tip clearance gap where the complex tip leakage flow occurs and thereby local high heat loads prevail. This paper presents a literature survey of blade tip leakage flow and heat transfer, as well as research of external and internal cooling technologies. The pres...

Gas Turbine Blade Tip Heat Transfer and Cooling: A Literature Survey

Dominant flow structure in the squealer tip gap and its impact on turbine aerodynamic performance

This paper presents an experimental investigation of the aero-thermal performance of a cooled winglet tip, under transonic conditions (exit Mach number of 1.0, and an exit Reynolds number of 1.27×106 , based on axial chord). Spatially-resolved heat transfer data and film cooling effectiveness data are obtained using the transient infrared thermography technique in the Oxford High-Speed Linear Cascade test facility. Aerodynamic loss data are obtained by traversing a specially-made and calibrated three-hole pressure probe and a single-hole probe one axial chord downstream of the blade. Detailed contours of Nusselt number show that for an increase in tip clearance, with and without film cooling, and for coolant injection, for both tip clearances, the Nusselt number increases. Also the smaller tip clearance observes higher film cooling effectiveness overall. Detailed distributions of kinetic energy losses as well as pitch-wise averaged loss coefficients and loss coefficients at a mixed-out plane indicate that the size of the loss core associated with the over-tip leakage vortex decreases with cooling injection.Copyright © 2011 by Rolls-Royce plc

Aero-Thermal Performance of a Winglet at Engine Representative Mach and Reynolds Numbers

A turbine blade (10) includes an airfoil (12) having pressure and suction sidewalls (24,26) extending between leading and trailing edges (38,30) and root (32) and tip (34). The tip (34) includes squealer ribs (36,38) extending from a tip floor (40) forming an open tip cavity (42). The rib (36) along the pressure sidewall (24) has a squared external corner, and a flute (44) extends along the base of the rib at the tip floor.

Fluted tip turbine blade

Effects of turbine blade tip shape on total pressure loss and secondary flow of a linear turbine cascade

In axial turbine the tip clearance flow occurring in rotor blade rows is responsible for about one third of the aerodynamic losses in the blade row and in many cases is the limiting factor for the blade lifetime. The tip leakage vortex forms when the leaking fluid crosses the gap between the rotor blade tip and the casing from pressure to suction side and rolls up into a vortex on the blade suction side. The flow through the tip gap is both of high velocity and high temperature, with the heat transfer to the blade from the hot fluid being very high in the blade tip area. In order to avoid blade tip burnout and a failure of the machine, blade tip cooling is commonly used. This paper presents the physical study and an improved design of a recessed blade tip for a highly loaded axial turbine rotor blade with application in high pressure axial turbines in aero engine or power generation. With use of three-dimensional Computational Fluid Dynamics (CFD), the flow field near the tip of the blade for different shapes of the recess cavities is investigated. Through better understanding and control of cavity vortical structures, an improved design is presented and the differences to the generic flat tip blade are highlighted. It is observed that by an appropriate profiling of the recess shape, the total tip heat transfer Nusselt Number was significantly reduced, being 15% lower than the flat tip and 7% lower than the baseline recess shape. Experimental results also showed an overall improvement of 0.2% in the one-and-1/2-stage turbine total efficiency with the improved recess design compared to the flat tip case. The CFD analysis conducted on single rotor row configurations predicted a 0.38% total efficiency increase for the rotor equipped with the new recess design compared to the flat tip rotor.Copyright © 2006 by ASME

Flow Physics and Profiling of Recessed Blade Tips: Impact on Performance and Heat Load

This paper investigates the influence of stator-rotor interaction on the stage performance of three blade tip geometries. A reference flat tip is used to assess two different recess blade geometries. The study is made in the context of the realistic turbine stage configuration provided by the ETHZ 1.5 stage LISA turbine research facility. This numerical investigation describes the details of unsteady recess cavity flow structure and confirms the beneficial effects of the improved recess geometry over the flat tip and the nominal recess design both in terms of stage efficiency and tip heat load. The tip flow field obtained from the improved recess design combines the advantages of a nominal recess design (aerodynamic sealing) and the flat tip configuration. The turbine stage capacity is almost unchanged between the flat tip and the improved recess tip cases, which simplifies the design procedure when using the improved recess design. Overall heat load in the improved recess case is reduced by 26% compared to the flat tip and by 12% compared to the nominal recess. A key finding of this study is the difference in effects of the upstream stator wake on the recess cavity flow. Where cavity flow in the nominal design is only moderately influenced, the improved recess cavity flow shows enhanced flow unsteadiness. The tip Nusselt number from a purely steady state prediction in the nominal recess case is nearly identical to the time-average prediction. The improved design shows a 6% difference between steady state and time average tip Nusselt number. This is due to the strong influence of the wake passing on the recess cavity flow. In fact, the wake enhances a small flow difference at the leading edge of the recess cavity between the nominal and improved recess cavities, which results in a completely different flow field further downstream in the recess cavity.Copyright © 2008 by ASME and Alstom

Influence of Stator-Rotor Interaction on the Aerothermal Performance of Recess Blade Tips

A turbine blade (1) with a side surface (7) having an aerodynamic profile. The turbine blade (1) further comprises an end surface (6) and is mounted in a turbine, whereby the end surface (6) is delimited by a gap from a casing (8) of the turbine. The end surface (6) comprises a recess (2) which is shaped such that it acts as an improved aerodynamic seal.

Turbine Blade With Recessed Tip

\n Unshrouded industrial centrifugal compressor impellers operate at high rotational speeds and volume flow rates. Under such conditions, impeller blade excitation is dominated by high frequency interaction with stationary parts, i.e., vaned diffusers or inlet guide vanes. In a previous study conducted on two full compression units of the original equipment manufacturer (OEM), the authors identified, characterized, and quantified resonant blade vibration caused by the interaction of the impeller blades with rotating stall cells during severe off-design conditions. This caused significant dynamic stress in the blades. In a follow-up study, this phenomenon was reproduced successfully experimentally under representative off-design conditions in a downscaled test rig and numerically with unsteady computational fluid dynamics (CFD) and structural mechanical finite element method (FEM) analysis. The gained knowledge was translated into a new diffuser design philosophy, based on sectorwise circumferential variation of the leading edge angle. This paper presents the patented philosophy, which is experimentally verified on the same test rig configuration in terms of flow path geometry and measurement equipment that was used in the mentioned prior study to assess resonant blade interaction. The results confirm the design aims: rotating stall onset was delayed without affecting the aerodynamic performance of the stage. Resonant blade interaction with rotating stall observed in the baseline diffuser could not be avoided with the two new diffuser designs. However, with the two new diffusers, the induced mechanical stresses in the impeller and the excitability were reduced by up to 12%.

Bob Mischo

Papers

Flow Physics and Profiling of Recessed Blade Tips: Impact on Performance and Heat Load

Influence of Stator-Rotor Interaction on the Aerothermal Performance of Recess Blade Tips

Turbine Blade With Recessed Tip

Experimental Investigation for Enhanced Control of Rotating Unsteady Flow Instabilities in an Unshrouded Centrifugal Compressor Impeller