The present invention relates generally to gas turbine engines, and, more specifically, to turbine blade cooling therein.
In a gas turbine engine, air is pressurized in a multistage compressor and mixed with fuel for generating hot combustion gases in a combustor. The gases are discharged through a high pressure turbine (HPT) which powers the compressor, typically followed by a low pressure turbine (LPT) which provides output power by typically powering a fan at the upstream end of the engine. This turbofan configuration is used for powering commercial or military aircraft.
Engine performance or efficiency may be increased by increasing the maximum allowed operating temperature of the combustion gases that are discharged to the HPT which extracts energy therefrom. Furthermore, engines are continually being developed for increasing cruise duration and distance, for one exemplary commercial application for a supersonic business jet and for an exemplary military application such as a long range strike aircraft.
Increasing turbine inlet temperature and cruise duration correspondingly increases the cooling requirements for the hot engine components, such as the high pressure turbine rotor blades. The first stage rotor blades receive the hottest combustion gases from the combustor and are presently manufactured with state-of-the-art superalloy materials having enhanced strength and durability at elevated temperature. These blades may be configured from a myriad of different cooling features for differently cooling the various portions of the blades against the corresponding differences in heat loads thereto during operation.
The presently known cooling configurations for first stage turbine blades presently limit the maximum allowed turbine inlet temperature for obtaining a suitable useful life of the blades. Correspondingly, the superalloy blades are typically manufactured as directionally solidified materials or monocrystal materials for maximizing the strength and life capability thereof under the hostile hot temperature environment in the gas turbine engine.
The intricate cooling configurations found in the blades are typically manufactured using common casting techniques in which one or more ceramic cores are utilized. The complexity of the cooling circuits in the rotor blades are limited by the ability of conventional casting processes in order to achieve suitable yield in blade casting for maintaining competitive costs.
Accordingly, it is desired to provide an improved turbine rotor blade cooling configuration for further advancing temperature and duration capability thereof in a gas turbine engine.
A turbine blade includes a hollow airfoil integrally joined to a dovetail. The airfoil includes a perforate first bridge defining a flow channel behind the airfoil leading edge. A second bridge is spaced behind the first bridge and extends from a pressure sidewall of the airfoil short of the airfoil trailing edge. A third bridge has opposite ends joined to the pressure sidewall and the second bridge to define with the first bridge a supply channel for the leading edge channel, and defines with the second bridge a louver channel extending aft along the second bridge to its distal end at the pressure sidewall.