Blade outer air seal

Abstract

An example blade outer air seal assembly for a gas turbine engine includes a main body portion extending along an axis. The main body portion has a radially inward facing surface and at least one radially outward facing surface. A passage is provided in the main body portion between the at least one radially inward facing surface and the at least one radially outward facing surface. The passage has a first portion and a second portion transverse to the first portion, and at least one rib disposed along the axis radially outward of the second portion.

We claim: 
     
         1 . A blade outer air seal assembly for a gas turbine engine, comprising:
 a main body portion extending along an axis, the main body portion having a radially inward facing surface and at least one radially outward facing surface;   a passage provided in the main body portion between the at least one radially inward facing surface and the at least one radially outward facing surface, the passage having a first portion and a second portion transverse to the first portion; and   at least one rib disposed along the axis radially outward of the second portion.   
     
     
         2 . The assembly of  claim 1 , wherein the at least one rib is at least partially in a cavity that opens in a radially outward facing direction, the at least one rib comprising one of the radially outward facing surfaces. 
     
     
         3 . The assembly of  claim 1 , including at least one airflow inlet provided by one of the radially outward facing surfaces. 
     
     
         4 . The assembly of  claim 3 , wherein the at least one airflow inlet includes a first airflow inlet opening to a cavity that opens in a radially outward facing direction. 
     
     
         5 . The assembly of  claim 1 , including at least one attachment portion disposed at both a leading edge of the main body portion and a trailing edge of the main body portion, each of the at least one attachment portions having a flange extending axially aft. 
     
     
         6 . The assembly of  claim 5 , wherein the at least one rib is circumferentially aligned with one of the attachment portions disposed at the leading edge. 
     
     
         7 . The assembly of  claim 1 , wherein at least one fluid outlet is disposed at a circumferential end of the main body portion radially outward of the radially inward facing surface. 
     
     
         8 . The assembly of  claim 1 , including at least one airflow inlet provided by one of the radially outward facing surfaces, wherein the at least one airflow inlet includes at least a first airflow inlet having a cross-section greater than a second airflow inlet. 
     
     
         9 . The assembly of  claim 1 , wherein the main body portion comprises a single crystal nickel alloy having a sulfur content that is equal to or below 1 part per million. 
     
     
         10 . The assembly of  claim 1 , wherein a thermal barrier coating is disposed adjacent the radially inward facing surface, wherein a thickness of the thermal barrier coating is substantially equal to a thickness of an inner wall of the main body portion defined between a floor of the passage and the radially inward facing surface. 
     
     
         11 . The assembly of  claim 10 , wherein the passage is entirely radially outward of the thermal barrier coating. 
     
     
         12 . The assembly of  claim 1 , wherein the second portion is radially inward of the first portion. 
     
     
         13 . The assembly of  claim 1 , wherein at least one of the radially outward facing surfaces is exposed. 
     
     
         14 . A gas turbine engine assembly, comprising:
 a plurality of blade outer air seal assemblies each configured to attach to a casing and disposed circumferentially about an engine axis, the blade outer air seal assemblies each having a main body portion with a radially inward facing surface and a radially outward facing surface, wherein at least one passage is defined in the main body portion between the radially inward facing surface and the radially outward facing surface, the main body portion having an inner wall having a first thickness defined between a floor of each of the at least one passages and the radially inward facing surface; and   a thermal barrier coating adjacent to the radially inward facing surface, the thermal barrier coating having a second thickness that is substantially equal to the first thickness.   
     
     
         15 . The gas turbine engine assembly of  claim 14 , wherein at least one airflow inlet is opening to the radially outward facing surface, the at least one of passage generally perpendicular to the engine axis, the at least one passage in fluid communication with the at least one inlet and configured to receive cooling air flow. 
     
     
         16 . The gas turbine engine assembly of  claim 15 , wherein the at least one passage includes a first passage in fluid communication with a first airflow inlet of the at least one airflow inlet having airflow greater than a second passage in fluid communication with a second airflow inlet. 
     
     
         17 . The gas turbine engine assembly of  claim 16 , wherein the first airflow inlet has a cross-section greater than the second airflow inlet. 
     
     
         18 . The gas turbine engine assembly of  claim 15 , wherein an airflow source communicates compressor bleed air to the at least one inlet. 
     
     
         19 . The gas turbine engine assembly of  claim 14 , wherein each of the plurality of blade outer air seal assemblies includes at least one attachment portion disposed at each of a leading edge of the main body portion and a trailing edge of the main body portion, each of the at least one attachment portions having a flange extending axially aft, each of the at least one attachment portions aligned with corresponding receiving portion of the casing to attach the plurality of blade outer air seal assemblies to the casing. 
     
     
         20 . The gas turbine engine assembly of  claim 14 , including at least at least one rib is disposed along the engine axis radially outward of the second portion. 
     
     
         21 . A method of forming a blade outer air seal for a gas turbine engine comprising:
 providing a main body portion having a radially inward facing surface and a radially outward facing surface that axially extend along an axis, wherein at least one passage is defined in the main body portion between the radially inward facing surface and the radially outward facing surface, the at least one passage having a first portion extending generally radially and a second portion extending generally circumferentially transverse to the first portion, wherein the second portion has a floor, wherein at least one rib is disposed along the axis on the radially outward facing surface, the at least one rib radially outward of the circumferential second portion;   machining an inner wall defined between the floor and the radially inward facing surface to reduce a thickness of the inner wall; and   depositing a thermal barrier coating adjacent the radially inward facing surface, wherein the thermal barrier coating is substantially equal to the reduced thickness of the inner wall.   
     
     
         22 . The method of  claim 21 , including the step of forming at least one airflow inlet opening to the radially outward facing surface, wherein the at least one passage is in fluid communication with the at least one inlet and configured to receive cooling air flow.