# US 10,480,335 B2 · Compressor turbine vane airfoil profile

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- Number: US 10,480,335 B2
- Document ID: US10480335B2
- Status: Granted
- Kind: B2
- Application number: 15/693,755
- Filed: 2017-09-01
- Published: 2019-03-07
- Granted: 2019-11-19
- Earliest claimed: 2017-09-01
- Assignees: Pratt & Whitney Canada Corp.
- Applicants: Pratt & Whitney Canada Corp.
- Inventors: Constantinos Ravanis
- Web page: https://patentgrep.com/patent/US10480335B2

## CPC

- F01D 9/041: {using blades (F01D5/148 takes precedence)}
- F05D 2220/32: in gas turbines
- F05D 2220/3212: the first stage of a turbine
- 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 compressor turbine includes a 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 compressor turbine vane for a gas turbine engine comprising an airfoil having a portion defined by a nominal profile in accordance with Cartesian coordinate values of X, Y, and Z of Sections 3 to 9 set forth in Table 2, wherein a point of origin of 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 values are coordinate values defining the profile at each distance Z.

### Claim 2 (depends on 1)

The compressor turbine vane as defined in claim 1, wherein the compressor turbine vane is a high pressure turbine stage vane of the gas turbine engine.

### Claim 3 (depends on 2)

The compressor turbine vane as defined in claim 2, wherein the high pressure turbine stage vane is a first stage compressor turbine vane.

### Claim 4 (depends on 1)

The compressor turbine vane as defined in claim 1, wherein the turbine vane has a manufacturing tolerance of ±0.030 inches in a direction perpendicular to the airfoil.

### Claim 5 (depends on 1)

The compressor 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, each profile section at each Z distance being joined smoothly with one another to form an airfoil shape of the portion.

### Claim 6

A compressor turbine vane for a gas turbine engine, the compressor turbine vane having a cold coated intermediate airfoil portion defined by a nominal profile in accordance with Cartesian coordinate values of X, Y, and Z of Sections 3 to 9 set forth in Table 2, wherein a point of origin of 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 values are coordinate values defining the profile at each distance Z.

### Claim 7 (depends on 6)

The compressor turbine vane as defined in claim 6 wherein the compressor turbine vane is a high pressure turbine vane of the gas turbine engine.

### Claim 8 (depends on 7)

The compressor turbine vane as defined in claim 7, wherein the high pressure turbine vane is a first stage compressor turbine vane.

### Claim 9 (depends on 6)

The compressor turbine vane as defined in claim 6, wherein the turbine vane has a manufacturing tolerance of ±0.030 inches.

### Claim 10 (depends on 6)

The compressor turbine vane as defined in claim 6, 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, each profile section at each Z distance being joined smoothly with one another to form an airfoil shape of the intermediate portion.

### Claim 11

A compressor 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 in accordance with Cartesian coordinate values of X, Y, and Z of Sections 3 to 9 set forth in Table 2, wherein a point of origin of 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 values are coordinate values defining the profile at each distance Z.

### Claim 12

A compressor turbine vane comprising at least one airfoil having a surface defined by coordinate values given in Table 2, the at least one airfoil extending between platforms defined by coordinate values given in Table 1, wherein a fillet radius is applied around the at least one airfoil between the at least one airfoil and the platforms.