Gas turbine engine mid turbine frame with flow turning features

• US 8,424,313 B1
• Filed: 02/09/2012
• Issued: 04/23/2013
• Est. Priority Date: 01/31/2012
• Status: Active Grant

• Current Assignee: Rtx Corporation
• Original Assignee: United Technologies Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Inventors: Praisner, Thomas J., Magge, Shankar S., Estes, Matthew B.
• Primary Examiner(s): Wongwian, Phutthiwat
• Application Number:
  US13/369,366
• Time in Patent Office: 439 Days
• Field of Search: 607/96, 608/05, 415/191, 415/211.2, 416/241.R
• US Class Current: 60/796
• CPC Class Codes: F01D 25/28 Supporting or mounting arrangements, e.g. for turbine casing
• 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: F01D 9/04 forming ring or sector
• CPC Class Codes: F01D 9/041 using blades F01D5/148 takes precedence
• CPC Class Codes: F01D 9/065 Fluid supply or removal conduits traversing the working fluid flow, e.g. for lubrication-, cooling-, or sealing fluids see also F01D25/16, F01D25/24 and F01D25/26
• CPC Class Codes: F02K 3/06 with front fan
• CPC Class Codes: F05D 2240/12 Fluid guiding means, e.g. vanes
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

A gas turbine engine includes first and second stages having a rotational axis. A mid turbine frame is arranged axially between the first and second stages. The mid turbine frame includes a circumferential array of airfoils, and the airfoils each 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 plane and the curvature at airfoil leading and trailing edges. The angle is equal to or greater than about 10°, for example. In one example, an airfoil aspect ratio is less than 1.5.

Claims

1. A gas turbine engine comprising:

a first stage and a second stage having a rotational axis;

a frame arranged axially between the first stage and the second stage, the frame including a circumferential array of airfoils, 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 intersection of the midspan plane and the curvature at the airfoil leading and trailing edges, the angle being equal to or greater than about 10°

.

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 in a range of 20°

-60°

.

3. The gas turbine engine according to claim 1, wherein the frame includes an inner 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 a leading edge span and a trailing edge span, and the airfoil extends 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 first stage and the second stage are configured to rotate in opposite directions.

5. The gas turbine engine according to claim 1, wherein a rotational axis plane extends through the rotational axis and intersects the trailing edge and the curvature, a first angle provided between the rotational axis plane and the second line, the angle providing a second angle, and wherein the first angle is greater than 20°

.

6. The gas turbine engine according to claim 1, wherein the array includes twenty or fewer airfoils.

7. The gas turbine engine according to claim 1, comprising:

a compressor section comprising a high pressure compressor and a low pressure compressor;

a combustor fluidly connected to the compressor section;

a turbine section fluidly connected to the combustor, the turbine section comprising;

the first stage is a high pressure turbine;

the second stage is a low pressure turbine; and

wherein the frame is provided by a mid-turbine frame positioned between the high pressure turbine and the low pressure turbine.

8. The gas turbine engine according to claim 7, further comprising a fan fluidly connected to the compressor section.

9. The gas turbine engine according to claim 8, comprising a geared architecture is interconnected between the fan and the low pressure turbine.

10. The gas turbine engine according to claim 8, wherein the gas turbine engine is a high bypass geared aircraft engine having a bypass ratio of greater than about six (6).

11. The gas turbine engine according to claim 8, wherein the gas turbine engine includes a Fan Pressure Ratio of less than about 1.45.

12. The gas turbine engine according to claim 8, wherein the low pressure turbine has a pressure ratio that is greater than about 5.

13. A gas turbine engine comprising:

a first stage and a second stage having a rotational axis;

a frame arranged axially between the first stage and the second stage, the frame including a circumferential array of airfoils, the frame includes inner and outer cases joined by the airfoils, leading and trailing edges respectively extending in a generally radial direction from the inner and outer cases a leading edge span and a trailing edge span, and the airfoil extends in an axial direction an axial chord length between the leading and trailing edges, the airfoils each having an aspect ratio range having a lower limit of greater than 1.0 to an upper limit of about 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.

14. The gas turbine engine according to claim 13, the airfoils each having a curvature provided equidistantly between pressure and suction sides and extending from the leading edge to the trailing edge at a midspan plane along the airfoil, a rotational axis plane extending through the rotational axis and intersecting the trailing edge and the curvature, first and second lines respectively tangent to the curvature at the leading and trailing edges, a first angle provided between the rotational axis plane and the second line and a second angle provided between the second and first lines, wherein the second angle is greater than 10°

.

15. The gas turbine engine according to claim 14, wherein the midspan plane is oriented at a flow path angle relative to the rotational axis in a range of 20°

-60°

.

16. The gas turbine engine according to claim 13, wherein the first stage and the second stage are configured to rotate in opposite directions.

17. The gas turbine engine according to claim 14, wherein the first angle is greater than 20°

.

18. The gas turbine engine according to claim 13, wherein the array includes twenty or fewer airfoils.

19. The gas turbine engine according to claim 13, comprising:

a compressor section comprising a high pressure compressor and a low pressure compressor;

a combustor fluidly connected to the compressor section;

a turbine section fluidly connected to the combustor, the turbine section comprising;

the first stage is a high pressure turbine;

the second stage is a low pressure turbine; and

wherein the frame is provided by a mid-turbine frame positioned between the high pressure turbine and the low pressure turbine.

20. The gas turbine engine according to claim 19, wherein the gas turbine engine includes a fan and is a high bypass geared aircraft engine having a bypass ratio of greater than about six (6).

21. The gas turbine engine according to claim 20, wherein the gas turbine engine includes a Fan Pressure Ratio of less than about 1.45.

22. The gas turbine engine according to claim 19, wherein the low pressure turbine has a pressure ratio that is greater than about 5.

23. A mid turbine frame for a gas turbine engine, comprising:

a frame of the mid turbine frame including a circumferential array of airfoils, 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 intersection of the midspan plane and the curvature at the airfoil leading and trailing edges, the angle being equal to or greater than about 10°

.

24. A turbine section for a gas turbine engine, comprising:

a first turbine and a second turbine having a rotational axis;

a mid turbine frame arranged axially between the first turbine and the second turbine, a mid turbine frame including a circumferential array of airfoils, 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 intersection of the midspan plane and the curvature at the airfoil leading and trailing edges, the angle being equal to or greater than about 10°

.

25. The turbine section according to claim 24, wherein one of the first and second turbines is a low pressure turbine, the low pressure turbine has a pressure ratio that is greater than about 5.

26. The turbine section according to claim 24, wherein the array includes twenty or fewer airfoils.