The present invention relates generally to gas turbine engines, and more particularly, to impingement cooling passages used in gas turbine engines.
A gas turbine engine commonly includes a fan, a compressor, a combustor, a turbine, and an exhaust nozzle. During engine operation, working medium gases, for example air, are drawn into the engine and compressed by the compressor. The compressed air is channeled to the combustor where fuel is added to the air and the air-fuel mixture ignited. The products of combustion are discharged to the turbine section, which extracts a portion of the energy from the combustion products to power the fan and the compressor.
The compressor and turbine often include alternating sections of rotating blades and stationary vanes. The operating temperatures of some engine stages, such as in the high pressure turbine rotor and stator stages, may exceed the material limits of the airfoils and therefore necessitate cooling of the airfoils. Cooled airfoils may include cooling channels, sometimes referred to as passages through which a coolant, such as compressor bleed air, is directed to convectively cool the airfoil. Airfoil cooling channels may be oriented spanwise from the base to the tip of the airfoil or axially between leading and trailing edges. The channels may be fed by one or more supply channels toward the airfoil base, where the coolant flows radially into the cooling channels. In some configurations, the cooling channels include small cooling passages, referred to as impingent cooling passages, which connect the cooling channel with an adjacent cavity or channel. The impingement cooling passages are sized and placed to direct jets of coolant on to interior airfoil surfaces such as the interior surfaces of the leading and trailing edges.
Prior airfoil designs have continually sought to decrease airfoil temperatures through cooling. A particular challenge in prior impingement cooled airfoil designs is with respect to a region affected by the thermal boundary layer. The thermal boundary layer of an impinging coolant jet is the flow region near the interior surface of the airfoil distorted by the effects of the coolant interacting with the surface. Because the thermal boundary layer distortion redirects a portion of the impinging coolant jet away from the interior airfoil surfaces, the cooling efficiency of the impingement jet decreases. However, due to the relatively high temperatures encountered during operation, a need still exists to improve impingement cooling of turbine blade and vane airfoils.