US20190106989A1

US20190106989A1 - Gas turbine engine airfoil

Info

Publication number: US20190106989A1
Application number: US15/727,683
Authority: US
Inventors: Timothy Charles Nash
Original assignee: United Technologies Corp
Current assignee: RTX Corp
Priority date: 2017-10-09
Filing date: 2017-10-09
Publication date: 2019-04-11
Also published as: EP3467260A1

Abstract

A component for a gas turbine engine includes a platform that has a radially inner side and a radially outer side. A root portion extends from the radially inner portion of the platform. An airfoil extends from the radially outer side of the platform. The airfoil includes a pressure side that extends between a leading edge and a trailing edge. A suction side extends between the leading edge and the trailing edge. A bowed tip portion extends perpendicular to a mid-camber line of the airfoil.

Description

BACKGROUND

This disclosure relates to a gas turbine engine airfoil. More particularly, this disclosure relates to a gas turbine engine airfoil having an improved design.
A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustor section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines.
The shape of an airfoil designed for turbomachinery applications is an important characteristic. It is often a result of multidisciplinary considerations including aerodynamics, durability, structures and manufacturability. However, recent advances in the design of aerodynamically high-performing, high-pressure turbine blades, particularly at the tip, have caused increased difficulties in the design of blades.

SUMMARY

In one exemplary embodiment, a component for a gas turbine engine includes a platform that has a radially inner side and a radially outer side. A root portion extends from the radially inner portion of the platform. An airfoil extends from the radially outer side of the platform. The airfoil includes a pressure side that extends between a leading edge and a trailing edge. A suction side extends between the leading edge and the trailing edge. A bowed tip portion extends perpendicular to a mid-camber line of the airfoil.
In a further embodiment of any of the above, the bowed tip portion extends in a circumferential direction and an axial direction.
In a further embodiment of any of the above, the bowed tip portion beings at between 75% and 85% of a span of the airfoil.
In a further embodiment of any of the above, the bowed tip portion begins at 80% of the span of the airfoil.
In a further embodiment of any of the above, the bowed tip portion is bowed between 3 and 19.6 degrees perpendicular to the mid-camber line in a circumferential direction.
In a further embodiment of any of the above, the bowed surface is bowed between 3 and 19.6 degrees perpendicular to the mid-camber line in an axial direction.
In a further embodiment of any of the above, the bowed surface is bowed by the same degree in both the circumferential direction and the axial direction.
In a further embodiment of any of the above, the bowed tip portion follows a curvilinear profile beginning at 3 degrees and increasing to 19.6 degrees.
In a further embodiment of any of the above, a leading edge of the bowed tip portion includes a rounded contour to reduce vibration and oxidation.
In another exemplary embodiment, a gas turbine engine includes a compressor section and a turbine section. A circumferential array of airfoils are located in one of the compressor section and the turbine section. Each of the airfoils include a pressure side that extends between a leading edge and a trailing edge. A suction side extends between the leading edge and the trailing edge. A bowed tip portion extends perpendicular to a mid-camber line of the airfoil.
In a further embodiment of any of the above, the bowed tip portion extends in a circumferential direction and an axial direction.
In a further embodiment of any of the above, the bowed tip portion beings at between 75% and 85% of a span of the airfoil.
In a further embodiment of any of the above, the bowed tip portion begins at 80% of the span of the airfoil.
In a further embodiment of any of the above, the bowed tip portion is bowed between 2 and 15 degrees perpendicular to the mid-camber line in a circumferential direction.
In a further embodiment of any of the above, the bowed surface is bowed between 3 and 19.6 degrees perpendicular to the mid-camber line in an axial direction.
In a further embodiment of any of the above, the bowed surface is bowed by the same degree in both the circumferential direction and the axial direction.
In a further embodiment of any of the above, the bowed tip portion follows a curvilinear profile beginning at 3 degrees and increasing to 19.6 degrees.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic view of an example gas turbine engine according to a first non-limiting embodiment.

FIG. 2 is a schematic view of a section of the gas turbine engine of FIG. 1, such as a turbine section.

FIG. 3 is a schematic view of a turbine blade.

FIG. 4 is a cross-sectional view through an airfoil according to this disclosure.

FIG. 5 is an aft perspective view of the airfoil according to this disclosure.

FIG. 6 is a forward perspective view of the airfoil according to this disclosure.

FIG. 7 is a suction side perspective view of the airfoil according to this disclosure.

FIG. 8 is a pressure side perspective view of the airfoil according to this disclosure.

The embodiments, examples and alternatives of the preceding paragraphs, the claims, or the following description and drawings, including any of their various aspects or respective individual features, may be taken independently or in any combination. Features described in connection with one embodiment are applicable to all embodiments, unless such features are incompatible.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28. Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section 22 drives air along a bypass flow path B in a bypass duct defined within a nacelle 15, and also drives air along a core flow path C for compression and communication into the combustor section 26 then expansion through the turbine section 28. Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures.
The exemplary engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of bearing systems 38 may be varied as appropriate to the application.
The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a first (or low) pressure compressor 44 and a first (or low) pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a second (or high) pressure compressor 52 and a second (or high) pressure turbine 54. A combustor 56 is arranged in exemplary gas turbine 20 between the high pressure compressor 52 and the high pressure turbine 54. A mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The mid-turbine frame 57 further supports bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.
The core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 57 includes airfoils 59 which are in the core airflow path C. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion. It will be appreciated that each of the positions of the fan section 22, compressor section 24, combustor section 26, turbine section 28, and fan drive gear system 48 may be varied. For example, gear system 48 may be located aft of combustor section 26 or even aft of turbine section 28, and fan section 22 may be positioned forward or aft of the location of gear system 48.
The engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about five 5:1. Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. The geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.
A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of the engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and 35,000 ft (10,668 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)]^0.5. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.5 meters/second).
Referring to FIG. 2, a cross-sectional view through a high pressure turbine section 54 is illustrated. In the example high pressure turbine section 54, first and second arrays of circumferentially spaced fixed vanes 60, 62 are axially spaced apart from one another. A first stage array of circumferentially spaced turbine blades 64, mounted to a rotor disk 68, is arranged axially between the first and second fixed vane arrays. A second stage array of circumferentially spaced turbine blades 66 is arranged aft of the second array of fixed vanes 62. It should be understood that any number of stages may be used. Moreover, the disclosed airfoil may be used in a compressor section, turbine section and/or fixed or rotating stages.
The turbine blades each include a tip 80 adjacent to a blade outer air seal 70 of a case structure 72, which provides an outer flow path. The first and second stage arrays of turbine vanes and first and second stage arrays of turbine blades are arranged within a core flow path C and are operatively connected to a spool 32, for example.
Each blade 64 includes an inner platform 76 respectively defining inner flow path. The platform inner platform 76 supports an airfoil 78 that extends in a radial direction R, as shown in FIG. 3. It should be understood that the turbine vanes 60, 62 may be discrete from one another or arranged in integrated clusters. The airfoil 78 includes a leading edge 82 and a trailing edge 84.
The airfoil 78 is provided between pressure side 94 (predominantly concave) and suction side (predominantly convex) 96 in an airfoil thickness direction (FIG. 4), which is generally perpendicular to a chord-wise direction provided between the leading and trailing edges 82, 84. Multiple turbine blades 64 are arranged in a circumferentially spaced apart manner in a circumferential direction Y (FIG. 4). The airfoil 78 includes multiple film cooling holes 90, 92 respectively schematically illustrated on the leading edge 82 and the pressure side 94 (FIG. 4).
The turbine blades 64 are constructed from a high strength, heat resistant material such as a nickel-based or cobalt-based superalloy, or of a high temperature, stress resistant ceramic or composite material. In cooled configurations, internal fluid passages and external cooling apertures provide for a combination of impingement and film cooling. Other cooling approaches may be used such as trip strips, pedestals or other convective cooling techniques. In addition, one or more thermal barrier coatings, abrasion-resistant coatings or other protective coatings may be applied to the turbine vanes 62.
FIGS. 3 and 4 schematically illustrate an airfoil including pressure and suction sides joined at leading and trailing edges 82, 84. An attachment or root 74 supports the platform 76. The root 74 may include a fir tree that is received in a correspondingly shaped slot in the rotor disk 68, as is known. The airfoil 78 extends a span from a support, such as an inner platform 76 to an end, such as a tip 80 in a radial direction R from a radially outer side of the platform 76. The 0% span and the 100% span positions, respectively, correspond to the radial airfoil positions at the support and the end. The leading and trailing edges 82, 84 are spaced apart from one another and an axial chord b_xlength (FIG. 4) extends in the axial direction X.
As shown in FIG. 3, a radially outer end of the air foil 78 includes a bowed tip portion 100. The bowed tip portion 100 is bowed in a direction perpendicular to a mid-camber line 98 (FIG. 4) of the airfoil 78 such that the bowed tip portion 100 extends towards the suction side 96 in a circumferential or tangential direction and towards the trailing edge 82 in an axial downstream direction. The geometry of the bowed tip portion 100 results in a non-pointed or rounded bowed tip leading edge 102 that reduces vulnerable of a leading edge of the bowed tip portion 100 to damaging vibrations and oxidation.
As shown in FIG. 3, the bowed tip portion 100 beings at 80% of the span of the airfoil 78 and continues to the tip 80. In another example, the bowed tip portion 100 begins between 75% and 85% of the bowed tip portion.
As shown in FIGS. 5 and 6, the bowed tip portion 100 bends in a circumferential direction towards the suction side 96 at an angle α. In the illustrated embodiment, the angle α is between 2 and 15 degrees. Similarly, as shown in FIGS. 7 and 8, the bowed tip portion 100 bends in an axially downstream direction towards at an angle β. In the illustrated embodiment, the angle β is between 3 and 19.6 degrees. Moreover, in the illustrated embodiment, the bowed tip portion 100 is bowed by the same degree in both the circumferential direction and the axially downstream direction. In another example embodiment, the bowed tip portion 100 follows a curvilinear profile beginning at 3 degrees and increasing to 19.6 degrees at the tip 80.
It should also be understood that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit herefrom. Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present invention.
Although the different examples have specific components shown in the illustrations, embodiments of this invention are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples.
Although an example embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of the claims. For that reason, the following claims should be studied to determine their true scope and content.

Claims

What is claimed is:

1. A component for a gas turbine engine comprising:

a platform having a radially inner side and a radially outer side;

a root portion extending from the radially inner portion of the platform; and

an airfoil extending from the radially outer side of the platform, the airfoil including:

a pressure side extending between a leading edge and a trailing edge;

a suction side extending between the leading edge and the trailing edge; and

a bowed tip portion extending perpendicular to a mid-camber line of the airfoil.

2. The component of claim 1, wherein the bowed tip portion extends in a circumferential direction and an axial direction.

3. The component of claim 2, wherein the bowed tip portion beings at between 75% and 85% of a span of the airfoil.

4. The component of claim 3, wherein the bowed tip portion begins at 80% of the span of the airfoil.

5. The component of claim 1, wherein said bowed tip portion is bowed between 3 and 19.6 degrees perpendicular to the mid-camber line in a circumferential direction.

6. The component of claim 5, wherein the bowed surface is bowed between 3 and 19.6 degrees perpendicular to the mid-camber line in an axial direction.

7. The component of claim 6, wherein the bowed surface is bowed by the same degree in both the circumferential direction and the axial direction.

8. The component of claim 6, wherein the bowed tip portion follows a curvilinear profile beginning at 3 degrees and increasing to 19.6 degrees.

9. The component of claim 6, wherein a leading edge of the bowed tip portion includes a rounded contour to reduce vibration and oxidation.

10. A gas turbine engine comprising:

a compressor section and a turbine section; and

a circumferential array of airfoils located in one of the compressor section and the turbine section, wherein each of the airfoils include:

a pressure side extending between a leading edge and a trailing edge;

a suction side extending between the leading edge and the trailing edge; and

11. The gas turbine engine of claim 10, wherein the bowed tip portion extends in a circumferential direction and an axial direction.

12. The gas turbine engine of claim 11, wherein the bowed tip portion beings at between 75% and 85% of a span of the airfoil.

13. The gas turbine engine of claim 12, wherein the bowed tip portion begins at 80% of the span of the airfoil.

14. The gas turbine engine of claim 10, wherein said bowed tip portion is bowed between 2 and 15 degrees perpendicular to the mid-camber line in a circumferential direction.

15. The gas turbine engine of claim 14, wherein the bowed surface is bowed between 3 and 19.6 degrees perpendicular to the mid-camber line in an axial direction.

16. The gas turbine engine of claim 15, wherein the bowed surface is bowed by the same degree in both the circumferential direction and the axial direction.

17. The gas turbine engine of claim 15, wherein the bowed tip portion follows a curvilinear profile beginning at 3 degrees and increasing to 19.6 degrees.