Turbine blade shroud for gas turbine engine with power turbine and method of manufacturing same
A blade for a gas turbine engine includes an airfoil that includes an internal passage. A shroud is arranged at an end of the airfoil and has a shroud perimeter. Axially spaced knife edges extend radially from the shroud. An area is provided between the knife edges. A pocket is recessed into the area and is circumscribed by a perimeter edge that is arranged interiorly of the shroud perimeter. An outlet fluidly connects the internal passage to the pocket.
1. A blade for a gas turbine engine, comprising:
an airfoil including an internal passage;
a shroud is arranged at an end of the airfoil and has a shroud perimeter;
axially spaced knife edges extend radially from the shroud, and an area is provided between the knife edges;
wherein a pocket is recessed into the area and is circumscribed by a perimeter edge that is arranged interiorly of the shroud perimeter, and an outlet fluidly connects the internal passage to the pocket; and
wherein the shroud perimeter provides a first area, the knife edges include fillets, and a second area is the area between and including the fillets facing the pocket, the second area is in a range of 40% to 60% of the first area, the pocket has a third area defined by the perimeter edge, the third area is in a range of 80% to 90% of the second area.
2. The blade of claim 1 , comprising a root supporting a platform, and the airfoil extends radially from the platform to the shroud, and the internal passage extends from the root to the shroud.
3. The blade of claim 2 , wherein the airfoil extends from the platform to the shroud a span, the span is less than 2.6 inches (66.0 mm).
4. The blade of claim 1 , wherein the shroud includes axially spaced apart axial faces and circumferentially spaced apart circumferential faces, the axial and circumferential faces define the shroud perimeter.
5. The blade of claim 4 , wherein the knife edges extend to the circumferential faces.
6. The blade of claim 1 , wherein the outlet has a fourth area, and the fourth area is in a range of 5% to 25% of the third area.
7. A blade for a gas turbine engine, comprising:
an airfoil including an internal passage;
a shroud is arranged at an end of the airfoil and has a shroud perimeter;
axially spaced knife edges extend radially from the shroud, and an area is provided between the knife edges; and
wherein a pocket is recessed into the area and is circumscribed by a perimeter edge that is arranged interiorly of the shroud perimeter, and an outlet fluidly connects the internal passage to the pocket, wherein the pocket has first and second surfaces joined at a peak, the outlet arranged on at least one side of the peak.
8. The blade of claim 7 , wherein the shroud includes axially spaced apart axial faces and circumferentially spaced apart circumferential faces, the axial and circumferential faces define the shroud perimeter, the shroud perimeter provides a first area, the knife edges include fillets, and a second area is the area between and including the fillets facing the pocket, the second area is in a range of 40% to 60% of the first area.
9. The blade of claim 8 , wherein the pocket has a third area defined by the perimeter edge, the third area is in a range of 80% to 90% of the second area.
10. The blade of claim 9 , wherein the outlet has a fourth area, and the fourth area is in a range of 5% to 25% of the third area.
11. A blade for a gas turbine engine, comprising:
an airfoil including an internal passage;
a shroud is arranged at an end of the airfoil and has a shroud perimeter;
axially spaced knife edges extend radially from the shroud, and an area is provided between the knife edges; and
wherein a pocket is recessed into the area and is circumscribed by a perimeter edge that is arranged interiorly of the shroud perimeter, and an outlet fluidly connects the internal passage to the pocket, wherein the pocket has first and second surfaces joined at a peak, the shroud has first and second thicknesses spaced in a direction perpendicular to the peak on at least one of the first and second surfaces, the first and second thicknesses extending in a radial direction, the first thickness provided at a tangent point to a fillet joining the shroud and the airfoil, and the second thickness provided at the perimeter edge, the ratio of the first thickness to the second thickness in a range of 1.1 to 2.1.
12. The blade of claim 11 , wherein a depth of the pocket corresponds to the location of the second thickness at the perimeter edge is 0.060 inch (1.52 mm) or less.
13. The blade of claim 11 , wherein the shroud includes axially spaced apart axial faces and circumferentially spaced apart circumferential faces, the axial and circumferential faces define the shroud perimeter, the shroud perimeter provides a first area, the knife edges include fillets, and the second area is the area between and including the fillets facing the pocket, the second area is in a range of 40% to 60% of the first area.
14. The blade of claim 13 , wherein the pocket has a third area defined by the perimeter edge, the third area is in a range of 80% to 90% of the second area.
15. The blade of claim 14 , wherein the outlet has a fourth area, and the fourth area is in a range of 5% to 25% of the third area.
16. A gas turbine engine comprising:
a gas generator portion providing an air source;
a power turbine arranged fluidly downstream from the gas generator portion, the power turbine mechanically disconnected from the gas generator portion, the power turbine including at least one stage of blades, each of the blades including:
an airfoil including an internal passage;
a shroud is arranged at an end of the airfoil and has a shroud perimeter;
axially spaced knife edges extend radially from the shroud, and an area is provided between the knife edges; and
wherein a pocket is recessed into the area and is circumscribed by a perimeter edge that is arranged interiorly of the shroud perimeter, and an outlet fluidly connects the internal passage to the pocket, wherein the pocket has first and second surfaces joined at a peak, the shroud has first and second thicknesses spaced in a direction perpendicular to the peak on at least one of the first and second surfaces, the first and second thicknesses extending in a radial direction, the first thickness provided at a tangent point to a fillet joining the shroud and the airfoil, and the second thickness provided at the perimeter edge, the ratio of the first thickness to the second thickness in a range of 1.1 to 2.1.
17. The gas turbine engine of claim 16 , wherein the shroud perimeter provides a first area, the knife edges include fillets, and a second area is the area between the fillets facing the pocket, the pocket has a third area defined by the perimeter edge, the third area is in a range of 80% to 90% of the second area.
18. The gas turbine engine of claim 16 , wherein a depth of the pocket corresponds to the location of the second thickness at the perimeter edge is 0.060 inch (1.52 mm) or less.
19. The gas turbine engine of claim 16 , comprising a root supporting a platform, and the airfoil extends radially from the platform to the shroud, and the internal passage extends from the root to the shroud, wherein the airfoil extends from the platform to the shroud a span, the span is less than 2.6 inches (66.0 mm).
20. The gas turbine engine of claim 16 , wherein the gas generator portion includes a compressor section, the power turbine is mechanically disconnected from the compressor section.