1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to an air cooled turbine rotor blade with near wall cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine, such as an industrial gas turbine (IGT) engine, includes a turbine with multiple rows or stages or stator vanes that guide a high temperature gas flow through adjacent rotors of rotor blades to produce mechanical power and drive a bypass fan, in the case of an aero engine, or an electric generator, in the case of an IGT. In both cases, the turbine is also used to drive the compressor.
The efficiency of the engine can be increased by passing a higher temperature gas flow into the turbine section. However, the highest temperature gas than can be passed into the turbine is limited to the material properties of the turbine, especially the first stage stator vanes and rotor blades since these airfoils are exposed to the highest temperature gas flow. To allow for temperatures high enough to melt these airfoils, complex airfoil internal cooling circuits have been proposed to provide convection, impingement and film cooling for the airfoils to allow even higher temperatures. However, the pressurized cooling air used for cooling of the airfoils is typically bled off from the compressor. The cooling air thus is not used for producing mechanical work but reduces the efficiency of the engine. It is therefore useful to also minimize the amount of cooling air used while at the same time maximizing the cooling capability of this minimized cooling air.
For an airfoil used in a turbine of a gas turbine engine, the airfoil leading edge, the airfoil suction side immediately downstream of the leading edge, as the airfoil trailing edge region experiences a higher hot gas side external heat transfer coefficient than the mid-chord section of the pressure side and downstream of the suction side surfaces. The heat load for the airfoil aft section is higher than the forward section. Also, due to a hot gas leakage cross flow effect, the blade tip section will also experience high heat load. Cooling of the blade leading edge, trailing edge and tip peripheral edge becomes the most difficult region for blade cooling designs. Without a good cooling circuit design, high cooling flow consumption is required for the blade edge cooling. As the TBC technology improves, more industrial gas turbine blades are applied with a relatively thick or low conductivity TBC. The cooling air flow demand will then be greatly reduced while allowing for higher turbine inlet temperatures. As a result, the cooling flow demand for these high heat load regions of the blade needs to be eliminated.
Composite turbine blades have been proposed in the past in order to take advantage of the high temperature resistant properties of ceramic materials. Blade or vanes have been made using metal and ceramic materials (CMC or Carbon-Carbon materials) to form a single piece airfoil. However, one major problem while these composite airfoils have not been used is due to the large difference between the coefficient of thermal of expansion of metal and ceramic. The metal material will expand much more than the ceramic material, and thus very high stress loads are formed at the bonded surfaces. This results in cracks or complete breaks.