Gas turbine engine mid turbine frame with flow turning features
원문보기
IPC분류정보
국가/구분
United States(US) Patent
등록
국제특허분류(IPC7판)
F02D-001/00
F01D-009/04
F01D-025/28
F02K-003/06
출원번호
US-0221450
(2014-03-21)
등록번호
US-8915090
(2014-12-23)
발명자
/ 주소
Praisner, Thomas J.
Magge, Shankar S.
Estes, Matthew B.
출원인 / 주소
United Technologies Corporation
대리인 / 주소
Carlson, Gaskey & Olds, P.C.
인용정보
피인용 횟수 :
2인용 특허 :
18
초록▼
A gas turbine engine includes first and second stages having a rotational axis. A circumferential array of airfoils is arranged axially between the first stage and the second stage. At least one of the airfoils have a curvature provided equidistantly between pressure and suction sides. The airfoils
A gas turbine engine includes first and second stages having a rotational axis. A circumferential array of airfoils is arranged axially between the first stage and the second stage. At least one of the airfoils have a curvature provided equidistantly between pressure and suction sides. The airfoils extend from a leading edge to a trailing edge at a midspan plane along the airfoil. An angle is defined between first and second lines respectively tangent to the intersection of the midspan plane and the curvature at airfoil leading and trailing edges. The angle is equal to or greater than about 10°.
대표청구항▼
1. A gas turbine engine comprising: a turbine section disposed about a rotational axis; anda circumferential array of stationary airfoils arranged within the turbine section with at least one of the airfoils having a curvature provided equidistantly between pressure and suction sides and extending f
1. A gas turbine engine comprising: a turbine section disposed about a rotational axis; anda circumferential array of stationary airfoils arranged within the turbine section with at least one of the airfoils having a curvature provided equidistantly between pressure and suction sides and extending from a leading edge to a trailing edge at a midspan plane along the airfoil, and an angle defined between first and second lines respectively tangent to the curvature at the airfoil leading and trailing edges, the angle being equal to or greater than 10°, wherein a rotational axis plane extends through the rotational axis and intersects the trailing edge and the curvature with a first angle provided between the rotational axis plane and the second line that is greater than 20°. 2. The gas turbine engine according to claim 1, wherein the midspan plane is oriented at a flow path angle relative to the rotational axis between 20° and 60°. 3. The gas turbine engine according to claim 2, comprising an inner case and an outer case joined by the airfoils, the leading and trailing edges respectively extending in a generally radial direction from the inner case and the outer case, and the airfoils extend in an axial direction an axial chord length between the leading and trailing edges, the at least one of airfoils having an aspect ratio of less than 1.5, wherein the aspect ratio is an average of the sum of the leading and trailing edge spans divided by the axial chord length. 4. The gas turbine engine according to claim 1, wherein the array includes twenty or fewer airfoils. 5. The gas turbine engine according to claim 2, wherein the array of stationary airfoils are supported between an inner case and an outer case. 6. The gas turbine engine according to claim 5, wherein at least one of the array of stationary airfoils defines a cavity through which a support structure for a bearing structure extends. 7. The gas turbine engine according to claim 5, wherein at least one of the array of stationary airfoils defines a cavity through which a fluid is communicated between the outer case and the inner case. 8. The gas turbine engine according to claim 1, including a compressor section comprising a first compressor and a second compressor;a combustor in communication with the compressor section; wherein the turbine section is in communication with the combustor section, the turbine section including a first turbine and a second turbine and the circumferential array of airfoils is positioned between the first turbine and the second turbine. 9. The gas turbine engine according to claim 8, further comprising a fan driven by the turbine section. 10. The gas turbine engine according to claim 9, including a geared architecture configured to drive the fan, wherein one of the first turbine and the second turbine is configured to drive the geared architecture. 11. The gas turbine engine according to claim 10, wherein the geared architecture is configured to provide a speed reduction greater than 2.5:1. 12. The gas turbine engine according to claim 11, wherein the geared architecture comprises an epicyclic gear train. 13. The gas turbine engine according to claim 12, wherein the epicyclic gear train comprises a planetary gear system. 14. The gas turbine engine according to claim 10, wherein the gas turbine engine is a high bypass geared aircraft engine having a bypass ratio of greater than six (6). 15. The gas turbine engine according to claim 14, wherein the bypass ratio is greater than ten (10). 16. The gas turbine engine according to claim 14, wherein the gas turbine engine includes a Fan Pressure Ratio of less than 1.45. 17. The gas turbine engine according to claim 16, wherein a fan tip speed is less than 1150 ft/second. 18. The gas turbine engine according to claim 17, wherein the second turbine is configured to drive the geared architecture and has a pressure ratio that is greater than 5. 19. The gas turbine engine according to claim 18, wherein the first turbine rotates in a direction opposite the second turbine. 20. A turbine module for gas turbine engine comprising: a circumferential array of stationary airfoils disposed about a rotational axis with at least one of the airfoils having a curvature provided equidistantly between pressure and suction sides and extending from a leading edge to a trailing edge at a midspan plane along the airfoil, and an angle defined between first and second lines respectively tangent to the curvature at the airfoil leading and trailing edges, the angle being equal to or greater than 10°, wherein a rotational axis plane extends through the rotational axis and intersects the trailing edge and the curvature and a first angle provided between the rotational axis plane and the second line is greater than 20°. 21. The turbine module as recited in claim 20, wherein the midspan plane is oriented at a flow path angle relative to the rotational axis in a range between 20° and 60°. 22. The turbine module as recited in claim 21, comprising an inner case and an outer case joined by the airfoils, the leading and trailing edges respectively extending in a generally radial direction from the inner case and from the outer case, and the airfoils extend in an axial direction an axial chord length between the leading and trailing edges, the at least one of the airfoils having an aspect ratio of less than 1.5, wherein the aspect ratio is an average of the sum of the leading and trailing edge spans divided by the axial chord length. 23. The turbine module as recited in claim 20, wherein the array of stationary airfoils are supported between an inner case and an outer case. 24. A method of designing a gas turbine engine comprising: defining a turbine section about a rotational axis to include a circumferential array of stationary airfoils arranged with at least one of the airfoils having a curvature provided equidistantly between pressure and suction sides and extending from a leading edge to a trailing edge at a midspan plane along the airfoil, and an angle defined between first and second lines respectively tangent to the curvature at the airfoil leading and trailing edges, the angle being equal to or greater than 10°, wherein a rotational axis plane extends through the rotational axis and intersects the trailing edge and curvature with a first angle provided between the rotational axis plane and the second line that is greater than 20°. 25. The method as recited in claim 24, including defining the midspan plane to be oriented at a flow path angle relative to the rotational axis in a range between 20° and 60°. 26. The method as recited in claim 25, including defining an inner case and an outer case joined by the airfoils such that the leading and trailing edges respectively extend in a generally radial direction from the inner case and from the outer case, and the airfoils extend in an axial direction an axial chord length between the leading and trailing edges, and configuring at least one of the airfoils to include an aspect ratio of less than 1.5, wherein the aspect ratio is an average of the sum of the leading and trailing edge spans divided by the axial chord length. 27. The method as recited in claim 24, including defining a compressor section to include a first compressor and a second compressor; configuring a combustor to be in communication with the compressor section and the turbine section to be in communication with the combustor section; and configuring the turbine section to include at least a first turbine and a second turbine. 28. The method as recited in claim 27, including configuring a geared architecture to drive a fan and one of the first turbine and the second turbine to drive the geared architecture.
Turner,Mark Graham; Orkwis,Paul David; Cedar,Richard David, Row of long and short chord length and high and low temperature capability turbine airfoils.
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