Gas turbine engine truncated airfoil fillet

Patent metadata

Field	Value
Publication number	US-9890641-B2
Application number	US-201514597633-A
Country	US
Kind code	B2
Filing date	Jan 15, 2015
Priority date	Jan 15, 2015
Publication date	Feb 13, 2018
Grant date	Feb 13, 2018

Abstract

An integrally bladed rotor includes a rim integral with a web that extends radially inward to a bore. The rim provides an end wall from which integral blades extend radially outward to a tip. The blades have an airfoil that extends in a chord-wise direction from a leading edge to a trailing edge. A fillet circumscribes the airfoil and joins the airfoil to the end wall. The fillet is at one of the leading edge and the trailing edge is truncated at least 50% in an axial direction to provide a face of the rim.

First claim

What is claimed is:

1. An integrally bladed rotor comprising:
a rim integral with a web that extends radially inward to a bore, the rim providing an end wall from which integral blades extend radially outward to a tip, the blades have an airfoil that extends in a chord-wise direction from a leading edge to a trailing edge, a fillet circumscribes the airfoil and joins the airfoil to the end wall, the fillet at the trailing edge truncated at least 50% in an axial direction to provide a face of the rim.

2. The integrally bladed rotor according to claim 1 , wherein the trailing edge is truncated at least 65% in the axial direction.

3. The integrally bladed rotor according to claim 2 , wherein the trailing edge is truncated at least 80% in the axial direction.

4. The integrally bladed rotor according to claim 3 , wherein the trailing edge is truncated at least 90% in the axial direction.

5. The integrally bladed rotor according to claim 1 , where an aft side of the web includes an annular groove configured to receive a hub.

6. A gas turbine engine compressor section comprising:
a stack of rotor disks, one of the rotor disks includes a rim supporting a blade, the blade includes an end wall and has an airfoil that extends in a chord-wise direction from a leading edge to a trailing edge, a fillet circumscribes the airfoil and joins the airfoil to the end wall, the fillet at the trailing edge truncated at least 50% in an axial direction to provide a face of the rim.

7. The compressor section according to claim 6 , wherein the stack of rotor disks provides multiple rotating stages, the rotor disk provides a last stage of the stack.

8. The compressor section according to claim 7 , comprising an exit guide vane arranged downstream from the rotor disk, an axial gap provided between the face and the exit guide vane.

9. The compressor section according to claim 8 , wherein the trailing edge is truncated at least 65% in the axial direction.

10. The compressor section according to claim 9 , wherein the trailing edge is truncated at least 80% in the axial direction.

11. The compressor section according to claim 10 , wherein the trailing edge is truncated at least 90% in the axial direction.

12. The compressor section according to claim 8 , where an aft side of the web includes an annular groove configured to receive a hub.

13. The compressor section according to claim 6 , wherein the blades are integral with the rim and a web that extends radially inward to a bore.

14. A gas turbine engine comprising:
a compressor section;
a turbine section;
a combustor section arranged between the compressor and turbine sections; and
wherein one of the compressor and turbine sections includes a rotor disk, the rotor disk has a rim supporting a blade, the blade includes an end wall and has an airfoil that extends in a chord-wise direction from a leading edge to a trailing edge, a fillet circumscribes the airfoil and joins the airfoil to the end wall, the fillet at the trailing edge truncated at least 50% in an axial direction to provide a face of the rim.

15. The engine according to claim 14 , wherein the rotor disk is arranged in the compressor section, and comprising an exit guide vane arranged downstream from the rotor disk and upstream from the combustor section, an axial gap provided between the face and the exit guide vane.

16. The engine according to claim 14 , wherein the trailing edge is truncated at least 65% in the axial direction.

17. The engine according to claim 16 , wherein the trailing edge is truncated at least 80% in the axial direction.

18. The engine according to claim 17 , wherein the trailing edge is truncated at least 90% in the axial direction.

19. The engine according to claim 14 , wherein the blades are integral with the rim and a web that extends radially inward to a bore.

Assignees

• United Technologies Corp

Inventors

• Potter Christopher L
• Espinoza Michael

Classifications

• Y02T50/671

  Cross-Sectional Technologies · mapped topic

• F01D5/34 Primary

  Rotor-blade aggregates of unitary construction {, e.g. formed of sheet laminae; (discs formed of sheet laminae F01D5/028; ceramic materials F01D5/284, composite materials F01D5/282)} · CPC title

• F05D2240/304

  related to the trailing edge of a rotor blade · CPC title

• F05D2260/941

  particularly aimed at mechanical or thermal stress reduction · CPC title

• Y02T50/673

  Cross-Sectional Technologies · mapped topic