Method of forming gas turbine engine components

Patent metadata

Field	Value
Publication number	US-11148221-B2
Application number	US-201916555223-A
Country	US
Kind code	B2
Filing date	Aug 29, 2019
Priority date	Aug 29, 2019
Publication date	Oct 19, 2021
Grant date	Oct 19, 2021

Abstract

A method of forming a gas turbine engine component according to an example of the present disclosure includes, among other things, attaching a cover skin to an airfoil body, the airfoil body and the cover skin cooperating to establish pressure and suction sides of an airfoil, positioning the airfoil between first and second dies of a deforming station, heating the airfoil body to a first predefined temperature threshold between the first and second dies, and moving the first die relative to the second die to hold the airfoil between the first and second dies subsequent to the heating step, and then deforming the airfoil between the first and second dies.

First claim

What is claimed is:

1. A method of forming a gas turbine engine component comprising:
attaching a cover skin to an airfoil body, the airfoil body and the cover skin cooperating to establish pressure and suction sides of an airfoil;
positioning the airfoil between first and second dies of a deforming station;

heating the airfoil body to a first predefined temperature threshold between the first and second dies;
moving the first die relative to the second die to hold the airfoil between the first and second dies subsequent to the heating step, and then deforming the airfoil between the first and second dies;
moving the airfoil from the deforming station to a cooling chamber of a cool down station subsequent to the deforming step;
cooling the airfoil in the cooling chamber to a second predefined temperature threshold less than the first predefined temperature threshold;
wherein walls of the cooling chamber are twisted about the longitudinal axis such that a first end portion of the cooling chamber is offset from a second, opposed end portion of the cooling chamber.

2. The method as recited in claim 1 , wherein the airfoil is a fan blade.

3. The method as recited in claim 1 , wherein the airfoil is metallic.

4. The method as recited in claim 1 , wherein at least a majority of surfaces of the airfoil are spaced apart from the first and second dies during the heating step.

5. The method as recited in claim 1 , wherein the positioning step includes suspending the airfoil in a vertical direction from a support fixture and moving the airfoil in the vertical direction between the first and second dies.

6. The method as recited in claim 1 , wherein the step of moving the airfoil from the deforming station to the cooling chamber includes moving the airfoil along a substantially arcuate path extending between the first and second dies and the cooling chamber.

7. The method as recited in claim 1 , wherein the step of moving the airfoil includes translating the airfoil axially along a longitudinal axis of the cooling chamber between first and second positions, and rotating the airfoil about the longitudinal axis between the first and second positions.

8. The method as recited in claim 1 , further comprising conveying cooling flow to a plurality of cooling regions along the longitudinal axis of the cooling chamber such that a predefined temperature gradient is established between the first and second end portions during the cooling step.

9. The method as recited in claim 1 , wherein the airfoil body extends from a root section to a tip portion, the tip portion defines a stagger angle relative to the root section, and the stagger angle is greater than or equal to 10 degrees, absolute, prior to the attaching step.

10. The method as recited in claim 9 , wherein the deforming step occurs such that a change in the stagger angle of the airfoil presented to the deforming station is no more than 2 degrees, absolute.

11. The method as recited in claim 9 , wherein the first predefined temperature threshold is equal to or greater than 1200 degrees Fahrenheit.

12. The method as recited in claim 9 , wherein the attaching step includes welding at least a perimeter of the cover skin to the airfoil body, and the cover skin is dimensioned to enclose at least one internal cavity in the airfoil body.

13. A method of forming a gas turbine engine component comprising:
welding a cover skin to an airfoil body to define an airfoil such that the airfoil body is twisted along a spanwise axis to define a stagger angle;
heating the airfoil body and the cover skin to a predefined temperature threshold between first and second dies of a deforming station while the airfoil body and the cover skin are spaced apart from the first and second dies;
deforming the airfoil between the first and second dies subsequent to the heating step;
moving the airfoil from the first and second dies to a cooling chamber subsequent to the deforming step, wherein walls of the cooling chamber are twisted along a longitudinal axis of the cooling chamber between first and second end portions of the cooling chamber such that a perimeter of the cooling chamber at the first end portion is substantially offset from a perimeter of the cooling chamber at the second end portion; and
cooling the airfoil in the cooling chamber.

14. The method as recited in claim 13 , further comprising moving the first die towards the second die to hold the airfoil body between the first and second dies subsequent to the heating step, but prior to the deforming step.

15. The method as recited in claim 13 , wherein:
the airfoil body extends from a root section to a tip portion, and the stagger angle is greater than or equal to 10 degrees, absolute, at the tip portion relative to the root section prior to the welding step; and
the deforming step occurs such that a change in the stagger angle of the airfoil presented to the deforming station is no more than 1 degree, absolute.

16. The method as recited in claim 15 , wherein the welding step includes enclosing at least one internal cavity between the airfoil body and the cover skin.

17. The method as recited in claim 13 , further comprising:
rotating the airfoil about the longitudinal axis between first and second positions.

18. The method as recited in claim 17 , wherein the cooling step includes conveying cooling flow to at least three separate and distinct cooling regions along the longitudinal axis of the cooling chamber such that a predefined temperature gradient is established between the first and second end portions.

Assignees

• United Technologies Corp
• Raytheon Tech Corp

Inventors

• Bales Daniel A
• Demichael Thomas
• Dolan Robert C
• Parisi Francis B

Classifications

• B22F5/04

  of turbine blades · CPC title

• B23P15/04 Primary

  turbine or like blades from several pieces · CPC title

• F05D2230/25

  by forging · CPC title

• F05D2230/234

  Laser welding · CPC title

• B21D22/022

  by heating the blank or stamping associated with heat treatment (C21D takes precedence) · CPC title