This disclosure relates generally to heat transfer in gas turbine engines and more particularly to apparatus for cooling structures in such engines.
A gas turbine engine includes a turbomachinery core having a high pressure compressor, combustor, and high pressure turbine (“HPT”) in serial flow relationship. The core is operable in a known manner to generate a primary gas flow. The high pressure turbine includes annular arrays (“rows”) of stationary vanes or nozzles that direct the gases exiting the combustor into rotating blades or buckets. Collectively one row of nozzles and one row of blades make up a “stage”. Typically two or more stages are used in serial flow relationship. The combustor and HPT components operate in an extremely high temperature environment, and must be cooled by air flow to ensure adequate service life.
Cooling air flow is typically provided by utilizing relatively lower-temperature “bleed” air extracted from an upstream part of the engine, for example the high pressure compressor, and then feeding that bleed air to high-temperature downstream components. The bleed air may be applied in numerous ways, for example through internal convection cooling or through film cooling or both. Preexisting usage of bleed air and other cooling air flows the air over rib rougheners, trip strips, and pin fins. When used for convection cooling, the bleed air is often routed through serpentine passages or other structures which generate a pressure loss as the cooling air passes through them. Because bleed air represents a loss to the engine cycle and reduces efficiency, it is desired to maximize heat transfer rates and thereby use the minimum amount of cooling flow possible. For this reason heat transfer improvement structures, such as pin fins or turbulators may be on cooled subsurfaces.
Conventional turbulators are elongated strips or ribs having a square, rectangular, or other symmetric cross-section, and are generally aligned transverse to the direction of flow. The turbulators serve to “trip” the boundary layer across the entire width of a flow passage at the component subsurface and create turbulence which increases heat transfer. Cooling effectiveness is thereby increased. One problem with the use of conventional turbulators is that a flow stagnation zone is present downstream of each turbulator. This zone causes dust, which is naturally entrained in the cooling air, to be deposited and build up behind the turbulator. This build-up is an insulating layer which reduces heat transfer also can cause undesirable wear.
An example of a particular gas turbine engine structure requiring effective cooling is an HPT blade. HPT blades are configured as an array or stage of airfoils connected to, or an integral part of, the HPT rotor and located within the hot gas path annular flow region immediately after the combustor exit HPT nozzle. The HPT blades operate within extremely high gas temperatures while also experiencing high rotational loads and mechanical stresses. These blades are conventionally cooled by one or more mechanisms such as internal cooling passages with turbulators, cooling cavities with arrays of pin fins, impingement jet cooling, and film cooling. Within the very complex cooling passages and features contained in them, there exist many locations where the cooling flow experiences separation and recirculation zones. For example, the regions immediately following each conventional turbulator experience such flow recirculations to varying degrees. As a consequence, particulates carried with the flow have a longer residence time in these regions and may have a higher probability of accumulating and depositing on the cooled surfaces leading to increased undesirable thermal resistance. It is therefore desirable to incorporate alternate geometries of turbulators that can alleviate or minimize these flow recirculation and stagnation regions.