METHOD FOR ESTABLISHING A LOCATION OF AN ELEVATED STRESS REGION

• US 20090038151A1
• Filed: 08/08/2007
• Published: 02/12/2009
• Est. Priority Date: 08/08/2007
• Status: Active Grant

• Current Assignee: Rtx Corporation
• Original Assignee: United Technologies Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Inventors: Davis, Ephraim C., Bischof, Christopher J.
• Granted Patent: US 8,291,589 B2
• Time in Patent Office: Days
• US Class Current: 29/889.100
• CPC Class Codes: B23P 6/002 Repairing turbine components, e.g. moving or stationary blades, rotors, B23P6/045 takes precedence
• CPC Class Codes: F01D 5/005 Repairing methods or devices
• CPC Class Codes: F05D 2230/80 Repairing, retrofitting or upgrading methods
• CPC Class Codes: F05D 2260/94 Functionality given by mechanical stress related aspects such as low cycle fatigue [LCF] of high cycle fatigue [HCF]
• CPC Class Codes: F05D 2270/44 active, predictive, or anticipative
• CPC Class Codes: Y10T 29/49318 Repairing or disassembling
• CPC Class Codes: Y10T 29/49336 Blade making
• CPC Class Codes: Y10T 29/49718 Repairing
• CPC Class Codes: Y10T 29/49725 by shaping
• CPC Class Codes: Y10T 29/49771 Quantitative measuring or gauging

Abstract

A method for use in repairing gas turbine engine components includes applying a stress to a first gas turbine engine component to cause surface cracking on the first gas turbine engine component and establishing a location of an elevated stress region of a second gas turbine engine component based upon the location of the surface cracking on the first gas turbine engine component.

Claims

1. A method for use in repairing gas turbine engine components, comprising:

(a) applying a stress to a first gas turbine engine component to cause surface cracking on the first gas turbine engine component; and

(b) establishing a location of an elevated stress region of a second gas turbine engine component based upon the location of the surface cracking on the first gas turbine engine component.

2. The method as recited in claim 1, wherein said step (a) includes applying a torque stress.

3. The method as recited in claim 1, wherein said step (a) includes applying a magnitude of the stress that is greater than a yield strength and less than an ultimate strength of the gas turbine engine component and greater than an ultimate strength of a protective coating on the gas turbine engine component.

4. The method as recited in claim 3, wherein the gas turbine engine component comprises a nickel alloy and the protective coating comprises aluminum.

5. The method as recited in claim 4, wherein the protective coating comprises a diffused aluminide coating.

6. The method as recited in claim 4, wherein the protective coating comprises MCrAlY, where the M includes at least one of nickel, cobalt, iron, or a combination thereof, Cr is chromium, Al is aluminum and Y is yttrium

7. The method as recited in claim 1, wherein said step (a) includes applying the stress to cause the surface cracking in a protective coating on the first gas turbine engine component.

8. The method as recited in claim 1, wherein said step (a) includes establishing a maximum magnitude of a stress that is applied in a repair process to the second gas turbine engine component based upon a magnitude of the stress applied to the first gas turbine engine component.

9. The method as recited in claim 1, further including establishing a predicted location of the elevated stress region using finite element analysis and comparing the predicted location to the location established in step (b) as verification of the predicted location.

10. The method as recited in claim 1, wherein step (b) includes visually identifying the location of the surface cracking using a dye penetrant.

11. A method for use in repairing gas turbine engine components, comprising:

(a) establishing a predicted location of an elevated stress region on a gas turbine engine component;

(b) applying a stress to the gas turbine engine component to determine an actual location of the elevated stress region; and

(c) comparing the predicted location and the actual location to determine an accuracy of the predicted location.

12. The method as recited in claim 11, wherein said step (a) includes using finite element analysis to establish the predicted location.

13. The method as recited in claim 11, wherein said step (b) includes applying a torque stress.

14. The method as recited in claim 11, wherein said step (b) includes identifying the actual location of the elevated stress region based upon surface cracking in a protective coating of the gas turbine engine component.