GAS TURBINE ENGINE TURBINE BLADE TIP WITH COATED RECESS
A blade for a gas turbine engine includes an airfoil that has a tip with a terminal end surface. The terminal end surface includes a recess that has a depth of less than 40 mils (1.016 mm). The recess is filled with a thermal barrier coating. The recess is without any cooling holes.
1 . A blade for a gas turbine engine comprising:
an airfoil having a tip with a terminal end surface, the terminal end surface includes a recess having a depth of less than 40 mils (1.016 mm), the recess filled with a thermal barrier coating, the recess without any cooling holes.
2 . The blade according to claim 1 , wherein the airfoil includes pressure and suction sides joined at leading and trailing edges, the pressure and suction sides terminating at the terminal end surface.
3 . The blade according to claim 2 , wherein the recess is located near a leading edge of the airfoil.
4 . The blade according to claim 2 , wherein the recess is located near a trailing edge of the airfoil.
5 . The blade according to claim 2 , wherein the recess extends to at least one of the pressure and suction sides.
6 . The blade according to claim 2 , wherein the recess extends to at least one of the leading and trailing edges.
7 . The blade according to claim 2 , wherein the recess is arranged inward of the pressure and suction sides.
8 . The blade according to claim 1 , wherein the depth is between 5 to 15 mils (0.127 to 0.381 mm).
9 . The blade according to claim 1 , wherein the thermal barrier coating is generally flush with the terminal end surface or radially below the terminal end surface.
10 . The blade according to claim 1 , wherein thermal barrier coating extends radially below the terminal end surface.
11 . The blade according to claim 1 , wherein the airfoil includes a cooling passage, the recess in non-communication with and fluidly isolated from the cooling passage.
12 . The blade according to claim 1 , wherein the terminal end surface includes multiple recesses.
13 . A gas turbine engine comprising:
compressor and turbine sections, one of the compressor and turbine sections including a blade outer air seal;
a blade including an airfoil having a tip with a terminal end surface, the terminal end surface includes a recess having a depth of less than 40 mils (1.016 mm), the recess filled with a thermal barrier coating, the recess without any cooling holes.
14 . The gas turbine engine according to claim 13 , wherein the airfoil includes a cooling passage, the recess in non-communication with and fluidly isolated from the cooling passage.
15 . The gas turbine engine according to claim 13 , wherein the airfoil includes pressure and suction sides joined at leading and trailing edges, the pressure and suction sides terminating at the terminal end surface, wherein the depth is between 5 to 15 mils (0.127 to 0.381 mm).
16 . The gas turbine engine according to claim 15 , wherein the recess is located near one of the leading and trailing edges.
17 . The gas turbine engine according to claim 15 , wherein the recess extends to at least one of the pressure and suction sides.
18 . The gas turbine engine according to claim 15 , wherein the recess extends to at least one of the leading and trailing edges.
19 . The gas turbine engine according to claim 15 , wherein the recess is arranged inward of the pressure and suction sides.
20 . The gas turbine engine according to claim 13 , wherein the thermal barrier coating is generally flush with the terminal end surface or radially beyond the terminal end surface.
21 . The gas turbine engine according to claim 13 , wherein thermal barrier coating extends radially below the terminal end surface.
22 . The gas turbine engine according to claim 13 , wherein the terminal end surface includes multiple recesses.