# US 7,611,326 B2 · HP turbine vane airfoil profile

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- Number: US 7,611,326 B2
- Document ID: US7611326B2
- Status: Granted
- Kind: B2
- Application number: 11/470,416
- Filed: 2006-09-06
- Published: 2008-03-06
- Granted: 2009-11-03
- Earliest claimed: 2006-09-06
- Assignees: Pratt & Whitney Canada Corp.
- Inventors: Ricardo Trindade; Edward Vlasic; Sami Girgis
- Web page: https://patentgrep.com/patent/US7611326B2

## CPC

- F01D 5/141: {Shape, i.e. outer, aerodynamic form (F01D5/148 - F01D5/20 take precedence; blade construction F01D5/147)}
- F05D 2250/70: Shape
- F05D 2250/74: given by a set or table of xyz-coordinates
- F01D: NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES (machines or engines for liquids F03; non-positive displacement pumps F04D)
- F05D: INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS

## Abstract

A two-stage high pressure turbine includes a first stage vane having an airfoil with 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

### Claim 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.

### Claim 2 (depends on 1)

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

### Claim 3 (depends on 2)

The turbine vane as defined in claim 2, wherein the vane forms part of a first stage of a two-stage high pressure turbine.

### Claim 4 (depends on 1)

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.

### Claim 5 (depends on 1)

The turbine vane as defined in claim 1, wherein the turbine vane has a manufacturing tolerance of ±0.003 inch in a direction perpendicular to the airfoil.

### Claim 6 (depends on 5)

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.

### Claim 7 (depends on 1)

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.

### Claim 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.

### Claim 9 (depends on 8)

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

### Claim 10 (depends on 9)

The turbine vane as defined in claim 9, wherein the vane is part of a first stage of a two-stage high pressure turbine.

### Claim 11 (depends on 8)

The turbine vane as defined in claim 8, wherein the turbine vane has a manufacturing tolerance of ±0.003 inch.

### Claim 12 (depends on 11)

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.

### Claim 13 (depends on 8)

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.

### Claim 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.

### Claim 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.

### Claim 16 (depends on 15)

The high pressure turbine vane of claim 15 wherein the surface is lying within a +/−0.003 inch profile tolerance of the points of Table 2.