Gas turbine engine airfoil

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 side 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 curvature inflection point is located between 30% and 70% of an axial chord length of the airfoil.

Claims

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 side of the platform; and an airfoil is a blade and extends from the radially outer side of the platform to an unshrouded end, 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 plurality of curvature inflection points located on the pressure side of the airfoil between 30% and 70% of an axial chord length of the airfoil with a curvature inflection point occurring at a span location with a smallest radius of curvature for the span location, the smallest radius of curvature is measured in an X-Y plane about an axis parallel to a radial direction of the airfoil, and the plurality of curvature inflection point includes a first curvature inflection point located at 50% span of the airfoil and between 40% and 50% of the axial chord length.

2. The component of claim 1, wherein the plurality of curvature inflection points includes another curvature inflection point on the pressure side of the airfoil located at 0% span of the airfoil and between 60% and 70% of the axial chord length.

3. The component of claim 1, wherein the plurality of curvature inflection points includes another curvature inflection point on the pressure side of the airfoil located at 25% span of the airfoil and between 50% and 60% of the axial chord length.

4. The component of claim 1, wherein the plurality of curvature inflection points includes another curvature inflection point on the pressure side of the airfoil located at 75% span of the airfoil and between 30% and 40% of the axial chord length.

5. The component of claim 1, wherein the plurality of curvature inflection points includes another curvature inflection point on the pressure side of the airfoil located at 90% span of the airfoil and between 30% and 40% of the axial chord length.

6. The component of claim 1, wherein the plurality of curvature inflection points on the pressure side of the airfoil includes a second curvature inflection point located at 0% span of the airfoil and between 60% and 70% of the axial chord length, a third curvature inflection point located at 25% span of the airfoil and between 50% and 60% of the axial chord length, a fourth curvature inflection point located at 75% span of the airfoil and between 30% and 40% of the axial chord length, and a fifth curvature inflection point located at 90% span of the airfoil and between 30% and 40% of the axial chord length.

7. The component of claim 6, wherein the second curvature inflection point located at 0% span of the airfoil is axially downstream of the first curvature inflection point located at 50% span of the airfoil.

8. The component of claim 1, wherein the plurality of curvature inflection points includes another curvature inflection point located at 0% span of the airfoil that is axially downstream of the first curvature inflection point located at 50% span of the airfoil.

9. 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 the array of airfoils are blades and extends from a platform to an unshrouded end and each airfoil includes: 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 plurality of curvature inflection points located on the pressure side of the airfoil between 30% and 70% of an axial chord length of the airfoil, a curvature inflection point occurring at a span location with a smallest radius of curvature for the span location, the smallest radius of curvature is measured in an X-Y plane about an axis parallel to a radial direction of the airfoil, and the plurality of curvature inflection points includes a first curvature inflection point located at 75% span of the airfoil and between 30% and 40% of the axial chord length.

10. The gas turbine engine of claim 9, wherein the plurality of curvature inflection points includes another curvature inflection point on the pressure side of the airfoil located at 0% span of the airfoil and between 60% and 70% of the axial chord length.

11. The gas turbine engine of claim 9, wherein the plurality of curvature inflection points includes another curvature inflection point on the pressure side of the airfoil located at 25% span of the airfoil and between 50% and 60% of the axial chord length.

12. The gas turbine engine of claim 9, wherein the plurality of curvature inflection points includes another curvature inflection point on the pressure side of the airfoil located at 50% span of the airfoil and between 40% and 50% of the axial chord length.

13. The gas turbine engine of claim 9, wherein the plurality of curvature inflection points includes another curvature inflection point on the pressure side of the airfoil located at 90% span of the airfoil and between 30% and 40% of the axial chord length.

14. The gas turbine engine of claim 9, wherein the plurality of curvature inflection points on the pressure side of the airfoil includes a second curvature inflection point located at 0% span of the airfoil and between 60% and 70% of the axial chord length, a third curvature inflection point located at 25% span of the airfoil and between 50% and 60% of the axial chord length, a fourth curvature inflection point located at 50% span of the airfoil and between 40% and 50% of the axial chord length, and a fifth curvature inflection point located at 90% span of the airfoil and between 30% and 40% of the axial chord length.

15. The gas turbine engine of claim 9, wherein the plurality of curvature inflection points includes a second curvature inflection point located at 0% span of the airfoil that is axially downstream of a fourth curvature inflection point located at 50% span of the airfoil.

16. The gas turbine engine of claim 14, wherein the second curvature inflection point located at 0% span of the airfoil is axially downstream of the fourth curvature inflection point at 50% span of the airfoil.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

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

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

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

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

(5) FIG. 5 is a graphical representation of a curvature inflection point on the airfoil at 0% span.

(6) FIG. 6 is a graphical representation of a curvature inflection point on the airfoil at 25% span.

(7) FIG. 7 is a graphical representation of a curvature inflection point on the airfoil at 50% span.

(8) FIG. 8 is a graphical representation of a curvature inflection point on the airfoil at 75% span.

(9) FIG. 9 is a graphical representation of a curvature inflection point on the airfoil at 90% span.

(10) 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

(11) 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.

(12) 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.

(13) 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.

(14) 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.

(15) 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.

(16) 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 conditiontypically 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 consumptionalso 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)/(518.7R)].sup.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).

(17) 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.

(18) 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. In the illustrated embodiment, the tip 80 is unshrouded. 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.

(19) 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.

(20) 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).

(21) 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.

(22) FIGS. 3 and 4 schematically illustrates an airfoil including pressure and suction sides joined at leading and trailing edges 82, 84. A 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 68, as is known. The airfoil extends a span from a support, such as an inner platform 76 to an end, such as the tip 80 in a radial direction R. 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.sub.x length (FIG. 4) extends in the axial direction X.

(23) FIGS. 5-9 graphically represent changes in curvature at different axial positions of the airfoil 78 along the axial chord b.sub.x. The relative position along the axial chord length b.sub.x is represented by X/b.sub.x such that the graphical representations can apply to airfoils of varying axial chord length b.sub.x.

(24) A cross-section of the airfoil 78, as illustrated in FIG. 4, at a particular span location includes a curvature inflection point 100 that corresponds to the location of the smallest radius of curvature r and the greatest amount of curvature. In general, the curvature of the pressure side 94 increases moving downstream along the pressure side 94 in the axial chord length bx until reaching the curvature inflection point 100. Downstream of the inflection point, the amount of curvature begins to decrease. In the illustrated embodiment, the curvature inflection point 100 occurs between 30% and 70% of the axial chord bx.

(25) FIG. 5 illustrates a non-limiting embodiment of a line graph representing the change in curvature at 0% span for varying percentages of the axial chord length b.sub.x of the airfoil for three different examples. As shown in FIG. 5, the curvature inflection point 100 occurs at or approximately below 70% of the axial chord length b.sub.x.

(26) FIG. 6 illustrates a non-limiting embodiment of a line graph representing the change in curvature at 25% span for varying percentages of the axial chord length b.sub.x of the airfoil for three different examples. As shown in FIG. 6, the curvature inflection point 100 occurs between 50% and 60% of the axial chord length b.sub.x or at approximately 55% of the axial chord length b.sub.x.

(27) FIG. 7 illustrates a non-limiting embodiment of a line graph representing the change in curvature at 50% span for varying percentages of the axial chord length b.sub.x of the airfoil for three different examples. As shown in FIG. 7, the curvature inflection point 100 occurs between 40% and 50% of the axial chord length b.sub.x or at approximately 45% of the axial chord length b.sub.x.

(28) FIG. 8 illustrates a non-limiting embodiment of a line graph representing the change in curvature at 75% span for varying percentages of the axial chord length b.sub.x of the airfoil for three different examples. As shown in FIG. 8, the curvature inflection point 100 occurs between 30% and 40% of the axial chord length b.sub.x or at approximately 35% of the axial chord length b.sub.x.

(29) FIG. 9 illustrates a non-limiting embodiment of a line graph representing the change in curvature at 90% span for varying percentages of the axial chord length b.sub.x of the airfoil for three different examples. As shown in FIG. 9, the curvature inflection point 100 occurs between 30% and 40% of the axial chord length b.sub.x or at approximately 35% of the axial chord length b.sub.x.

(30) As shown in FIGS. 5-9, the curvature inflection point 100 moves axially forward when moving from 0% span radially outward. By locating the curvature inflection point 100 further aft along the axial chord length, a region immediately downstream of the leading edge 82 on the pressure side 94 includes non-stagnated flow. By reducing the amount of stagnated flow on the pressure side 94, more efficient ways of cooling the airfoil 78 can be utilized to reduce the overall cooling load required by the airfoil 78 and improve overall gas turbine engine efficiency.

(31) 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.

(32) 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.

(33) 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.