HP TURBINE VANE AIRFOIL PROFILE

• US 20080124223A1
• Filed: 11/28/2006
• Published: 05/29/2008
• Est. Priority Date: 11/28/2006
• Status: Active Grant

• Current Assignee: Pratt & Whitney Canada Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Original Assignee: Pratt & Whitney Canada Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Inventors: Girgis, Sami, Vlasic, Edward, Marini, Remo
• Granted Patent: US 7,566,200 B2
• Time in Patent Office: Days
• US Class Current: 416/241R
• CPC Class Codes: F01D 5/141 Shape, i.e. outer, aerodynamic form F01D5/148 - F01D5/20 take precedence; blade construction F01D5/147
• CPC Class Codes: F05D 2250/70 Shape
• CPC Class Codes: F05D 2250/74 given by a set or table of xyz-coordinates

Abstract

A single stage high pressure turbine vane includes an airfoil having a profile substantially in accordance with at least an intermediate portion of the Cartesian coordinate values of X, Y and Z set forth in Table 2. The X and Y values are distances, which when smoothly connected by an appropriate continuing curve, define airfoil profile sections at each distance Z. The profile sections at each distance Z are joined smoothly to one another to form a complete airfoil shape.

Claims

1. A turbine vane for a gas turbine engine comprising an airfoil having an intermediate portion defined by a nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z of Sections 5 to 10 set forth in Table 2, wherein the point of origin of the orthogonally related axes X, Y and Z is located at an intersection of a centerline of the gas turbine engine and a stacking line of the turbine vane, the Z values are radial distances measured along the stacking line, the X and Y are coordinate values defining the profile at each distance Z.

2. The turbine vane as defined in claim 1 forming part of a high pressure turbine stage of the gas turbine engine.

3. The turbine vane as defined in claim 2, wherein the vane forms part of a single stage high pressure turbine.

4. The turbine vane as defined in claim 1, wherein the X and Y values are scalable as a function of the same constant or number.

5. The turbine vane as defined in claim 1, wherein the X and Y coordinate values have a manufacturing tolerance of ±

0.003 inch.

6. The turbine vane as defined in claim 5, wherein the nominal profile defining the intermediate portion is for an uncoated airfoil, and wherein a coating having a thickness of 0.001 to 0.002 inch is applied to the airfoil.

7. The turbine vane as defined in claim 1, wherein X and Y values define a set of points for each Z value which when connected by smooth continuing arcs define an airfoil profile section, the profile sections at the Z distances being joined smoothly with one another to form an airfoil shape of the intermediate portion.

8. A turbine vane for a gas turbine engine, the turbine vane having an uncoated intermediate airfoil portion defined by a nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z of Sections 5 to 10 set forth in Table 2, wherein the point of origin of the orthogonally related axes X, Y and Z is located at an intersection of a centerline of the gas turbine engine and a stacking line of the turbine vane, the Z values are radial distances measured along the stacking line, the X and Y are coordinate values defining the profile at each distance Z, and wherein the X and Y values are scalable as a function of the same constant or number.

9. The turbine vane as defined in claim 8 forming part of a vane of a high pressure turbine stage of the gas turbine engine.

10. The turbine vane as defined in claim 9, wherein the vane is part of a single stage high pressure turbine.

11. The turbine vane as defined in claim 8, wherein the X, and Y coordinate values have a manufacturing tolerance of ±

0.003 inch.

12. The turbine vane as defined in claim 11, wherein a coating having a thickness of 0.001 to 0.002 inch is applied to the vane.

13. The turbine vane as defined in claim 8, wherein X and Y values define a set of points for each Z value which when connected by smooth continuing arcs define an airfoil profile section, the profile sections at the Z distances being joined smoothly with one another to form an airfoil shape of the intermediate portion.

14. A turbine stator assembly for a gas turbine engine comprising a plurality of vanes, each vanes including an airfoil having an intermediate portion defined by a nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z of Sections 5 to 10 set forth in Table 2, wherein the point of origin of the orthogonally related axes X, Y and Z is located at an intersection of a centerline of the gas turbine engine and a stacking line of the turbine vane, the Z values are radial distances measured along the stacking line, the X and Y are coordinate values defining the profile at each distance Z.

15. A high pressure turbine vane comprising at least one airfoil having a surface lying substantially on the points of Table 2, the airfoil extending between platforms defined generally by Table 1, wherein a fillet radius is applied around the airfoil between the airfoil and platforms, and wherein the values of Table 2 are subject to relevant tolerance.