US20120070297A1

US20120070297A1 - Aft loaded airfoil

Info

Publication number: US20120070297A1
Application number: US12/886,693
Authority: US
Inventors: Matthew B. Estes; Renee J. Jurek; Shankar S. Magge
Original assignee: Individual
Current assignee: MTU Aero Engines AG
Priority date: 2010-09-21
Filing date: 2010-09-21
Publication date: 2012-03-22
Also published as: EP2441919A1

Abstract

An example airfoil array includes an airfoil array configured to rotate within a turbomachine at a different rotational speed than a fan within the turbomachine. The airfoil array includes a first airfoil that is circumferentially spaced a distance from an adjacent second airfoil. A ratio of the distance to an axial chord length of the first airfoil is less than 1.3. An uncovered turning angle of the first airfoil is greater than 15 degrees.

Description

BACKGROUND

This disclosure relates generally to a gas turbine engine and, more particularly, to loading an aft end of a low pressure turbine airfoil.
As known, gas turbine engines include multiple sections, such as a fan section, a compression section, a combustor sections, a turbine section, and an exhaust nozzle section. The compression section and the turbine section include airfoil arrays mounted for rotation about an engine axis. The airfoil arrays include multiple individual airfoils (or blades) that extend radially from a mounting platform to a tip.
Air moves into the engine though the fan section. Rotating the combustion section's airfoil arrays compresses the air. The compressed air is then mixed with fuel and combusted in the combustor section. The products of combustion expand to rotatably drive the airfoil arrays in the turbine section. Rotating the airfoil arrays in the turbine section drives rotation of the fan section.
The turbine section often includes a low pressure turbine and a high pressure turbine. Each airfoil within the low pressure turbine's airfoil array is designed with the goal to position the laminar to turbulent boundary layer transition to reduce the likelihood of boundary layer separation and severe performance degradation. This typically results in a design with more forward loading, moving the transition location forward and a reducing laminar flow length.
As can be appreciated, reducing the length of the laminar flow increases the length of the turbulent boundary layer. The turbulent boundary layer performs more poorly than laminar flow provided the boundary layer has not separated.

SUMMARY

An example airfoil array includes an airfoil array configured to rotate within a turbomachine at a different rotational speed than a fan within the turbomachine. The airfoils in the array are arranged such that the pitch-to-chord ratio measured at midspan is less than 1.3, wherein said pitch is a distance between the trailing edges of adjacent airfoils and said chord is the axial chord length of said airfoils. An uncovered turning angle of the airfoil is greater than 15 degrees.
An example turbomachine arrangement includes a fan rotatable about an axis, and a turbine section that rotates at a first speed. The turbine section is configured to rotatably drive the fan at second, different speed. The turbine section includes an airfoil array having a first airfoil that is circumferentially spaced a pitch from an adjacent second airfoil. A ratio of the pitch to an axial chord length of the first airfoil is less than 1.3. An uncovered turning angle of the first airfoil is greater than 15 degrees.
An example method of establishing a transition location for a turbomachine airfoil includes positioning a first airfoil relative to a second airfoil within a turbomachine such that a ratio of a circumferential distance between the first airfoil and the second airfoil to an axial chord length of the first airfoil is less than 1.3. The first airfoil has an uncovered turning angle that is greater than 15 degrees and is configured to be rotated about an axis of the turbomachine at a different rotational speed than a fan of the turbomachine.
These and other features of the disclosed examples can be best understood from the following specification and drawings, the following of which is a brief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 shows a schematic view of an example gas turbine engine.

FIG. 2 shows a perspective view of an example airfoil assembly from a low pressure turbine section of the FIG. 1 engine.

FIG. 3 shows a section view at line 3-3 in FIG. 2.

FIG. 4 shows the position of the FIG. 2 airfoil assembly relative to an adjacent airfoil within the FIG. 1 engine.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates an example geared gas turbine engine 10 including (in serial flow communication) a fan 14, a low pressure compressor 18, a high pressure compressor 22, a combustor 26, a high pressure turbine 30, and a low pressure turbine 34. The gas turbine engine 10 is circumferentially disposed about an engine centerline X.
During operation, air is pulled into the gas turbine engine 10 by the fan 14, pressurized by the compressors 18 and 22, mixed with fuel, and burned in the combustor 26. The turbines 30 and 34 extract energy from the hot combustion gases flowing from the combustor 26. In a two-spool design, the high pressure turbine 30 utilizes the extracted energy from the hot combustion gases to power the high pressure compressor 22 through a high speed shaft 38. The low pressure turbine 34 utilizes the extracted energy from the hot combustion gases to power the low pressure compressor 18 and the fan 14 through a low speed shaft 42.
The examples described in this disclosure are not limited to the two-spool engine architecture described and may be used in other architectures, such as a single spool axial design, a three-spool axial design, and still other architectures. That is, there are various types of engines that can benefit from the examples disclosed herein, which are not limited to the design shown.
In this example, the low pressure turbine 34 drives the fan 14 through a gear system 36. The gear system 36 can be any known suitable gear system, such as a planetary gear system with orbiting planet gears, planetary system with non-orbiting planet gears, or another type of gear system. In the disclosed example, the gear system 36 has a constant gear ratio, and enables the low pressure turbine 34 to rotate at a higher speed than the fan 14. In the prior art, the rotational speed of the low pressure turbine is the same as the fan.
Referring now to FIGS. 2 and 3 with continuing reference to FIG. 1, an airfoil assembly 50 from the low pressure turbine 34 includes an attachment 54 and an airfoil 58. The airfoil 58 extends from a leading edge 62 to a trailing edge 66 relative to flow through the low pressure turbine 34.
The airfoil 58 has a pressure surface 70, a suction surface 74, and a chord length 78. The chord length 78 represents the distance between the leading edge 62 and the trailing edge 66.
The airfoil 58 also has an axial chord 82, which is a projection of the chord length 78 onto a plane containing the axis X of the gas turbine engine 10 (FIG. 1).
During operation of the gas turbine engine 10, fluid exiting the high pressure turbine 30 moves through passages established between adjacent ones of the airfoil 58. Within the passages, the fluid has a laminar flow portion 86 and a turbulent flow portion 90 relative to the suction surface 74 of the airfoil 58. The laminar flow portion 86 transitions to the turbulent flow portion 90 at a transition location 94.
The transition location 94 is closer to the trailing edge 66 of the airfoil 58 than in prior art designs. Accordingly, laminar flow along a suction surface 74 of the airfoil 58 extends further than in the prior art.
Referring to FIG. 4 with continuing reference to FIGS. 2 and 3, the example airfoil 58 is positioned in a specific location relative to an adjacent airfoil 58 a in the low pressure turbine 34. For example, the airfoil 58 is positioned such a ratio of a pitch distance 98 to the axial chord length 82 is less than 1.3. The pitch distance 98 represents the distance, or circumferential spacing, between the trailing edge 66 of the airfoil 58 and the airfoil 58 a.
The narrowest area of the passage 106 between the airfoil 58 and the airfoil 58 a is referred to as a throat 102. A throat location 110 of the airfoil 58 represents a position of the airfoil 58 corresponding to the narrowest area of a passage 106 between the airfoil 58 and the airfoil 58 a.
In this example, the airfoil 58 has an uncovered turning angle aft of the throat location 110 that is greater than 15 degrees. This amount of uncovered turning facilitates locating the transition location 94 for the airfoil 58.
The uncovered turning angle in this example represents the difference between a surface angle of the airfoil 58 at the throat location 110 and a surface angle of the airfoil 58 at the trailing edge 66. A person having skill in this art and the benefit of this disclosure would understand how to determine uncovered turning angle for an airfoil.
In one non-limiting example, an axial chord length is 1.27 inches (32.36 mm) and a pitch distance is 1.29 inches (32.77 mm) for a pitch-to-chord ratio of 1.02. The uncovered turning angle is 17 degrees in this example. The ratio of 1.02 and angle of 17 degrees performs particularly well in the engine 10. Other combinations of ratios and angles may perform well in other engines and turbomachines. That is, the specific ratio and angle can be adjusted to obtain the desired performance characteristics. The adjustments may be required due to the operating environment and other factors, for example.
Features of the disclosed examples include an airfoil curved and positioned such that the transition location is located closer to the aft end of the airfoil than in the prior art. The example low pressure turbine is able to rotate at a different speed than a fan of the fan section, which facilitates designing the airfoil to position the transition location toward the aft end of the airfoil portion. The gearing between the fan and the low pressure turbine allows the low pressure turbine to rotate at a higher speed than the fan.
The preceding description is exemplary rather than limiting in nature. Variations and modifications to the disclosed examples may become apparent to those skilled in the art that do not necessarily depart from the essence of this disclosure. Thus, the scope of legal protection given to this disclosure can only be determined by studying the following claims.

Claims

We claim:

1. An airfoil array comprising:

an airfoil array configured to rotate within a turbomachine at a different rotational speed than a fan within the turbomachine, the airfoil array including a first airfoil that is circumferentially spaced a distance from an adjacent second airfoil, wherein a ratio of the distance to an axial chord length of the first airfoil is less than 1.3, and an uncovered turning angle between the first airfoil and the second airfoil is greater than 15 degrees.

2. The airfoil assembly of claim 1, wherein the turbine airfoil array comprises a low pressure turbine airfoil array.

3. The airfoil assembly of claim 1, wherein the airfoil array is configured to rotate faster than the fan.

4. The airfoil assembly of claim 1, wherein the amount of uncovered turning is the difference between a throat surface angle of the first airfoil and a trailing edge surface angle of the first airfoil.

5. The airfoil assembly of claim 1, wherein the distance is a pitch between the first airfoil and the second airfoil.

6. A turbomachine arrangement comprising:

a fan rotatable about an axis;

a turbine section that rotates at a first speed, the turbine section configured to rotatably drive the fan at second, different speed; and

a airfoil array of the turbine section, the airfoil array including a first airfoil that is circumferentially spaced a distance from an adjacent second airfoil, wherein a ratio of the distance to an axial chord length of the first airfoil is less than 1.3, and an uncovered turning angle of the first airfoil is greater than 15 degrees.

7. The turbomachine arrangement of claim 6, wherein the adjacent airfoil is circumferentially adjacent the first airfoil.

8. The turbomachine arrangement of claim 6, wherein the fan and turbine section are components of gas turbine engine.

9. The turbomachine arrangement of claim 8, wherein the gas turbine engine is a geared gas turbine engine.

10. The turbomachine arrangement of claim 8, wherein the turbine section is a low pressure turbine section.

11. The turbomachine arrangement of claim 6, wherein the first speed is greater than the second speed.

12. A method of establishing a transition location for a turbomachine airfoil comprising:

positioning a first airfoil relative to a second airfoil within a turbomachine such that a ratio of a circumferential distance between the first airfoil and the second airfoil to an axial chord length of the first airfoil is less than 1.3, wherein the first airfoil has an uncovered turning angle that is greater than 15 degrees and is configured to be rotated about an axis of the turbomachine at a different rotational speed than a fan of the turbomachine.

13. The method of claim 12, wherein the turbomachine is a geared gas turbine engine.

14. The method of claim 12, wherein the circumferential distance is a pitch.

15. The method of claim 12, wherein the first airfoil is configured to be rotated about the axis of the turbomachine at a greater rotational speed than the fan of the turbomachine.