IPC분류정보
국가/구분 |
United States(US) Patent
등록
|
국제특허분류(IPC7판) |
|
출원번호 |
UP-0470416
(2006-09-06)
|
등록번호 |
US-7611326
(2009-11-16)
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발명자
/ 주소 |
- Trindade, Ricardo
- Vlasic, Edward
- Girgis, Sami
|
출원인 / 주소 |
- Pratt & Whitney Canada Corp.
|
대리인 / 주소 |
|
인용정보 |
피인용 횟수 :
16 인용 특허 :
41 |
초록
▼
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 a
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.
대표청구항
▼
What is claimed is: 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 o
What is claimed is: 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 first stage of a two-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 turbine vane has a manufacturing tolerance of ±0.003 inch in a direction perpendicular to the airfoil. 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 first stage of a two-stage high pressure turbine. 11. The turbine vane as defined in claim 8, wherein the turbine vane has 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. 16. 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.
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