The present disclosure relates to a gas turbine engine and, more particularly, to a blade outer air seal therefor.
Gas turbine engines, such as those that power modern commercial and military aircraft, generally include a compressor to pressurize an airflow, a combustor to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine to extract energy from the resultant combustion gases. The compressor and turbine sections include rotatable blade and stationary vane arrays. Within an engine case structure, the radial outermost tips of each blade array are positioned in close proximity to a shroud assembly. Blade Outer Air Seals (BOAS) supported by the shroud assembly are located adjacent to the blade tips such that a radial tip clearance is defined therebetween.
When in operation, the thermal environment in the engine varies and may cause thermal expansion and contraction such that the radial tip clearance varies. The radial tip clearance is typically designed so that the blade tips do not rub against the BOAS under high power operations when the blade disk and blades expand as a result of thermal expansion and centrifugal loads. When engine power is reduced, the radial tip clearance increases. To facilitate engine performance, it is operationally advantageous to maintain a close radial tip clearance through the various engine operational conditions.
The assemblies are often segmented to accommodate the thermal gradients and permit expansion and contraction. The segmented interface, however, provides a potential cooling air leakage path that is often sealed with feather seals that may be prone to leakage and wear. Feather seals are also subjected to temperature limitations that may be lower than the Blade Outer Air Seal segments.
USPTO Backgrounds
USPTO Backgrounds