阅读说明：本技术 具有后缘的涡轮发动机翼型件 (Turbine engine airfoil having trailing edge ) 是由 庞廷范 海伦·奥格巴辛·加布里乔尔格斯 扎卡里·丹尼尔·韦伯斯特 格里高利·特伦斯·加莱 史 于 2020-04-16 设计创作，主要内容包括：一种用于涡轮发动机的翼型件,可以包括外壁,该外壁界定内部并限定压力侧和吸力侧,该外壁在前缘和后缘之间延伸以限定弦向方向并在根部和末端之间延伸以限定跨度方向。多个出口和多个圆齿状部分可以延伸至后缘附近。(An airfoil for a turbine engine may include an outer wall bounding an interior and defining a pressure side and a suction side, the outer wall extending between a leading edge and a trailing edge to define a chordwise direction and between a root and a tip to define a span direction. The plurality of outlets and the plurality of scalloped portions may extend to near the trailing edge.)

1. An airfoil for a turbine engine, comprising:

an outer wall bounding an interior and defining a pressure side and a suction side, the outer wall extending between a leading edge and a trailing edge to define a chordwise direction and between a root and a tip to define a span direction;

a plurality of outlets extending to proximate the trailing edge; and

a plurality of non-uniform scalloped portions extending along the trailing edge, wherein at least some of the non-uniform scalloped portions are interposed between adjacent outlets.

2. The airfoil of claim 1, wherein the plurality of non-uniform scalloped portions comprise at least one of a non-uniform length, a non-uniform width, a non-uniform centerline, or a non-uniform geometry.

3. The airfoil of claim 1, wherein the plurality of non-uniform scalloped portions comprises a long scalloped portion having a first length and a short scalloped portion having a second length less than the first length.

4. The airfoil of claim 1, wherein the plurality of non-uniform scalloped portions comprises a wide scalloped portion having a first width and a narrow scalloped portion having a second width that is less than the first width.

5. The airfoil of claim 1, wherein the plurality of non-uniform scalloped portions comprises a variable width scalloped portion having a first width at a first location and a second width at a second location.

6. The airfoil of claim 1, wherein the plurality of non-uniform scalloped portions comprises a transition surface located upstream of the trailing edge and joining two adjacent scalloped portions.

7. The airfoil as claimed in any one of claims 1-6 wherein the plurality of non-uniform scalloped portions comprises a first inclined scalloped portion having a first centerline extending in the span direction and the chord direction.

8. The airfoil of claim 7, wherein the plurality of non-uniform scalloped portions comprises a second inclined scalloped portion having a second centerline that is not aligned with the first centerline.

9. The airfoil of claim 1, wherein the plurality of non-uniform scalloped portions comprises smooth scalloped portions having rounded corners and sharp scalloped portions having sharp edges.

10. The airfoil of claim 1, wherein the outlet extends along the trailing edge and the plurality of non-uniform scalloped portions extend along at least one of the pressure side or the suction side.

Technical Field

The present disclosure relates generally to airfoil cooling devices, and more particularly, to cooling outlets along an airfoil trailing edge.

Background

Turbine engines, particularly gas turbine engines or combustion turbine engines, are rotary engines that extract energy from a pressurized flow of combustion gases passing through the engine and transfer the energy to rotating turbine blades.

Turbine engines are typically designed to operate at high temperatures to increase engine efficiency. It would be beneficial to provide cooling measures for engine components (e.g., airfoils) in high temperature environments that can reduce material wear on these components and provide greater structural stability during engine operation.

Disclosure of Invention

In one aspect, the present disclosure is directed to an airfoil of a turbine engine, comprising: an outer wall bounding an interior and defining a pressure side and a suction side, the outer wall extending between a leading edge and a trailing edge to define a chordwise direction and between a root and a tip to define a span direction; a plurality of outlets extending to near the trailing edge; and a plurality of non-uniform scalloped portions extending along the trailing edge, wherein at least some of the non-uniform scalloped portions are between adjacent outlets.

In another aspect, the present disclosure is directed to an airfoil of a turbine engine, comprising: an outer wall bounding an interior and defining a pressure side and a suction side, the outer wall extending between a leading edge and a trailing edge to define a chordwise direction and between a root and a tip to define a span direction; a plurality of outlets extending to near the trailing edge; a plurality of scalloped portions extending along the trailing edge between adjacent outlets; and at least one contact point between two of the plurality of scalloped portions.

In yet another aspect, the present disclosure is directed to a method of cooling an airfoil in a turbine engine. The method includes supplying cooling air to an interior of an airfoil having a plurality of non-uniform scalloped portions extending along a trailing edge of the airfoil and a plurality of outlets, wherein at least some of the non-uniform scalloped portions are interposed between adjacent outlets; and discharging the cooling air through a plurality of outlets.

Drawings

In the drawings:

FIG. 1 is a cross-sectional schematic view of a turbine engine for an aircraft.

FIG. 2 is a perspective view of an airfoil that may be used in the turbine engine of FIG. 1.

FIG. 3 is a perspective view of a first region of the airfoil of FIG. 2 illustrating a trailing edge having a plurality of scalloped portions in accordance with aspects described herein.

FIG. 4 is a perspective view of a second region of the airfoil of FIG. 2, illustrating a trailing edge having an additional plurality of scalloped portions in accordance with aspects described herein.

Detailed Description

Various aspects of the present disclosure are directed to airfoils. For ease of description, the airfoil will be described in the context of a turbine engine. It will be appreciated that the present disclosure may have general applicability to any airfoil, including rotating or non-rotating airfoils, as well as to airfoils located anywhere within a turbine engine, including in a turbine section or a compressor section. The invention may also have general applicability in non-airfoil engine components as well as non-aircraft applications, such as other mobile applications and non-mobile industrial, commercial, and residential applications.

As used herein, the term "forward" or "upstream" refers to moving in a direction toward the engine inlet, or a component being relatively close to the engine inlet as compared to another component. The terms "rearward" or "downstream" are used with "forward" or "upstream" and refer to a direction toward the rear or outlet of the engine, or relatively closer to the engine outlet than another component.

As used herein, a "set" may include any number of the separately described elements, including only one element. Further, the terms "radial" or "radially" as used herein refer to a dimension extending between a central longitudinal axis of the engine and an outer engine circumference.

All directional references (e.g., radial, axial, proximal, distal, upper, lower, upward, downward, left, right, lateral, front, rear, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, forward, rearward, etc.) are only used for identification purposes to aid the reader's understanding of the present disclosure, and do not create limitations, particularly as to the position, orientation, or use of the disclosure. Unless otherwise specified, connection references (e.g., attached, coupled, connected, and engaged) are to be construed broadly and may include intermediate members between a collection of elements and relative movement between members. Thus, joinder references do not necessarily infer that two elements are directly connected and in fixed relation to each other. The exemplary drawings are for illustrative purposes only and the dimensions, locations, order and relative sizes reflected in the accompanying drawings may vary.

FIG. 1 is a schematic cross-sectional view of a gas turbine engine 10 for an aircraft. The engine 10 has a generally longitudinally extending shaft or centerline 12 extending from a forward portion 14 to an aft portion 16. Engine 10 includes a fan section 18 (including a fan 20), a compressor section 22 (including a booster or Low Pressure (LP) compressor 24 and a High Pressure (HP) compressor 26), a combustion section 28 (including a combustor 30), a turbine section 32 (including an HP turbine 34 and an LP turbine 36), and an exhaust section 38 in downstream serial flow relationship.

The fan section 18 includes a fan housing 40 surrounding the fan 20. The fan 20 includes a plurality of fan blades 42 arranged radially about the centerline 12. The HP compressor 26, combustor 30, and HP turbine 34 constitute a core 44 of the engine 10, which generates combustion gases. The core 44 is surrounded by a core housing 46, and the core housing 46 may be coupled with the fan housing 40.

An HP shaft or spool 48, coaxially disposed about the centerline 12 of the engine 10, drivingly connects the HP turbine 34 to the HP compressor 26. An LP shaft or spool 50, disposed coaxially about the centerline 12 of the engine 10 within the larger diameter annular HP spool 48, drivingly connects the LP turbine 36 to the LP compressor 24 and fan 20. The spools 48, 50 are rotatable about the engine centerline and are coupled to a plurality of rotatable elements, which may collectively define a rotor 51.

The LP and HP compressors 24, 26 each include a plurality of compressor stages 52, 54 with a set of compressor blades 56, 58 rotating relative to a corresponding set of stationary compressor vanes 60, 62 to compress or pressurize the fluid flow through the stages. In a single compressor stage 52, 54, a plurality of compressor blades 56, 58 may be arranged in a ring and may extend radially outward relative to the centerline 12, from the blade platform to the blade tip, while respective static compressor vanes 60, 62 are positioned upstream and adjacent to the rotating blades 56, 58. It is noted that the number of blades, vanes, and compressor stages shown in FIG. 1 is chosen for illustration purposes only, and that other numbers may be chosen.

The blades 56, 58 of the compressor stages may be mounted to (or integrated into) a disc 61, the disc 61 being mounted to a respective one of the HP and LP spools 48, 50. The buckets 60, 62 of the compressor stages may be mounted to the core housing 46 in a circumferential arrangement.

The HP and LP turbines 34, 36 each include a plurality of turbine stages 64, 66, with a set of turbine blades 68, 70 rotating relative to a corresponding set of stationary turbine vanes 72, 74 (also referred to as nozzles) to extract energy from the fluid flow through the stages. In a single turbine stage 64, 66, a plurality of turbine blades 68, 70 may be arranged in a ring and may extend radially outward relative to the centerline 12, while respective static turbine vanes 72, 74 are positioned upstream and adjacent to the rotating blades 68, 70. It is noted that the number of blades, buckets, and turbine stages shown in FIG. 1 is chosen for illustration purposes only, and that other numbers may be chosen.

The blades 68, 70 of the turbine stages may be mounted to a disc 71, the disc 71 being mounted on a respective one of the HP and LP spools 48, 50. The vanes 72, 74 of the compressor stages may be mounted to the core housing 46 in a circumferential arrangement.

In addition to the rotor portion, stationary portions of the engine 10, such as the stationary vanes 60, 62, 72, 74 between the compressor and turbine sections 22, 32, are also referred to individually or collectively as the stator 63. Thus, the stator 63 may refer to a combination of non-rotating elements throughout the engine 10.

In operation, the airflow flowing from fan section 18 is split such that a portion of the airflow is channeled to LP compressor 24, LP compressor 24 subsequently supplies pressurized air 76 to HP compressor 26, and HP compressor 26 further pressurizes the air. Pressurized air 76 from the HP compressor 26 mixes with fuel in the combustor 30 and is ignited, generating combustion gases. Some work is extracted from these gases by the HP turbine 34, which drives the HP compressor 26. The combustion gases are exhausted into the LP turbine 36, the LP turbine 36 extracts additional work to drive the LP compressor 24, and the exhaust gases are ultimately exhausted from the engine 10 via an exhaust section 38. The drive of the LP turbine 36 drives the LP spool 50 to rotate the fan 20 and the LP compressor 24.

A portion of pressurized gas stream 76 may be withdrawn from compressor section 22 as a bleed gas (bleedair) 77. Exhaust gases 77 may be drawn from pressurized gas stream 76 and provided to engine components that require cooling. The temperature of the pressurized gas stream 76 entering the combustor 30 increases significantly. Thus, the cooling provided by the exhaust gases 77 is necessary to operate such engine components in a high temperature environment.

The remainder of airflow 78 bypasses LP compressor 24 and engine core 44 and exits engine assembly 10 at fan exhaust side 84 through a stationary vane row (more specifically, an outlet guide vane assembly 80 comprising a plurality of airfoil guide vanes 82). More specifically, a circumferential row of radially extending airfoil guide vanes 82 is utilized adjacent fan section 18 for some directional control of airflow 78.

Some of the air supplied by fan 20 may bypass engine core 44 and be used to cool portions of engine 10 (particularly hot portions) and/or to cool or power other aspects of the aircraft. In the case of a turbine engine, the hot portion of the engine is typically located downstream of the combustor 30, particularly the turbine section 32, with the HP turbine 34 being the hottest portion as it is located directly downstream of the combustion section 28. Other sources of cooling liquid may be, but are not limited to, fluid discharged from the LP compressor 24 or the HP compressor 26.

Referring now to FIG. 2, an airfoil assembly 95 that may be used in the turbine engine 10 of FIG. 1 is shown. Airfoil assembly 95 includes an airfoil 100, which airfoil 100 may be, in non-limiting examples, any rotating or non-rotating airfoil, such as a blade or vane in fan section 18, compressor section 22, or turbine section 32.

The airfoil 100 includes an outer wall 103, the outer wall 103 bounding an interior 104 and defining a pressure side 106 and a suction side 108. Outer wall 103 also extends axially between leading edge 110 and trailing edge 112 to define a chordwise direction C, and also extends radially between a root 114 and a tip 116 to define a span direction S.

The airfoil assembly 95 may also include a platform 118, the platform 118 being coupled to the airfoil 100 at the root 114. In one example, the airfoil 100 is in the form of a blade extending from a dovetail 119, such as the HP turbine blade 68 of FIG. 1. In this case, platform 118 may form at least a portion of dovetail 119. In another example, the airfoil 100 may be in the form of a bucket, such as the LP turbine bucket 72, and the platform 118 may form at least a portion of an inner or outer band (not shown) coupled to the root 114.

Dovetail 119 may be configured to be mounted to turbine rotor disk 71 on engine 10. The dovetail 119 may include a set of inlet passages 120, illustratively three inlet passages, extending through the dovetail 119 to provide fluid communication with the interior of the airfoil 100. It should be understood that the dovetail 119 is shown in cross-section, with the inlet passage 120 located within the body of the dovetail 119.

The plurality of outlets 130 may extend to near the trailing edge 112. The outlet 130 is shown as extending along the trailing edge 112, and it is also contemplated that the outlet 130 may extend upstream of the trailing edge 112 or be located upstream of the trailing edge 112, such as adjacent the trailing edge 112 on the pressure side 106 or the suction side 108. Further, in a non-limiting example, the outlet 130 is shown as a circular injection hole, and the outlet 130 may also include a series diffuser, diffusion slot, exhaust slot, membrane hole, injection hole, or passage. Although illustrated as circular, in non-limiting examples, the outlet 130 may also have any suitable geometric profile, including oval, square with rounded corners, or asymmetric/irregular.

In another non-limiting example, the outlet 130 may be centered along the trailing edge 112. In yet another example, the outlet 130 may extend along the trailing edge 112 in a direction that is not aligned with the span direction S. In another example, the outlet 130 may extend in the span direction S along the trailing edge 112 and be positioned or biased to a position closer to the pressure side 106 or the suction side 108.

The airfoil 100 may further include a cooling air circuit 125, the cooling air circuit 125 fluidly coupling the set of inlet passages 120 to the outlet 130. For example, the cooling air circuit 125 may include an inlet passage 120, at least one internal cooling passage 128 within the airfoil 100, and an outlet 130. The cooling air 126 supplied through the at least one inlet passage 120 may flow through a cooling air circuit 125 (e.g., at least one interior cooling passage 128) and form an outlet airflow 127 through an outlet 130. It will be appreciated that the interior cooling passage 128 is shown in simplified view or schematic form, and that in a non-limiting example, the interior cooling passage 128 may have any suitable geometric profile (including linear or curvilinear), and constant or varying cross-sectional area, or at least one bifurcated portion. Further, while a single internal cooling passage 128 is shown, multiple internal cooling passages may be provided, including multiple fluidly coupled internal cooling passages within the airfoil 100. In this manner, any or all of the inlet channels 120 may be fluidly coupled to any or all of the outlets 130.

A plurality of scalloped portions 140 may be provided in outer wall 103 and extend to near trailing edge 112. At least some of the scalloped portions 140 may be interposed between adjacent outlets 130. In the example shown, each scalloped portion 140 is positioned between adjacent outlets 130. In another example (not shown), a plurality of scalloped portions may be positioned between adjacent outlets.

The scalloped portion 140 may extend along one or both of the pressure side 106 and the suction side 108. Further, as shown, a plurality of scalloped portions 140 may extend from trailing edge 112 at least partially in chordwise direction C. In another example (not shown), the first span region may include scalloped portions along the trailing edge between adjacent outlets and the second span region may include outlets along the trailing edge, the scalloped portions being located along the pressure side 106 or the suction side 108 and not extending to the trailing edge 112. It should be understood that "between" as used herein may refer to any portion of the scalloped portion 140 disposed between adjacent outlets 130. In other words, the scalloped portion 140 may extend upstream of the trailing edge 112 such that the first region is between adjacent outlets 130. In another example, the entire scalloped portion 140 may be positioned at or between adjacent outlets 130.

Referring now to FIG. 3, the first portion 101 of the airfoil 100 illustrates a plurality of outlets 130 and a scalloped portion 140. The scalloped portions 140 may be formed as discrete portions with spaces between adjacent scalloped portions 140.

The plurality of scalloped portions 140 may further include a plurality of non-uniform scalloped portions 142. For example, the plurality of non-uniform scalloped portions 142 may include at least one of a non-uniform length, a non-uniform width, a non-uniform centerline, or a non-uniform geometric profile. In other words, in non-limiting examples, at least one of the size, overall size, orientation, shape, or surface features (e.g., smooth versus rough, or sharp versus rounded) of the at least two non-uniform scalloped portions 142 may be different.

In the example of fig. 3, the plurality of non-uniform scalloped portions 142 include, but are not limited to, a long scalloped portion 150, a short scalloped portion 151, a first oblique scalloped portion 160, a second oblique scalloped portion 161, a smooth scalloped portion 170, and a sharp scalloped portion 171, a wide scalloped portion 180, a narrow scalloped portion 181, and a variable width scalloped portion 190.

The oblong tooth portion 150 may have a first length 155. In the example shown, first length 155 extends in chordwise direction C, and other directions are also contemplated, including at least partially extending in span direction S. The short scalloped portion 151 may have a second length 157 that is less than the first length 155. For example, in a non-limiting example, the second length 157 may be a predetermined fraction of the first length 155, such as half the first length 155, or 75% of the first length 155.

The first inclined scalloped portion 160 may define a first centerline 165. In the example shown, the first centerline 165 extends in the combined chordwise direction C and span direction S. The second beveled scalloped portion 161 may define a second centerline 167, the second centerline 167 being misaligned or different from the first centerline 165 of the first beveled scalloped portion 160. It is contemplated that first centerline 165 and second centerline 167 may differ by a predetermined amount, such as, in a non-limiting example, by 5-60 degrees.

It is further contemplated that the scalloped portion 140 may form the outlet airflow 127 (fig. 2) through the outlet 130. As shown, the first outlet airflow direction 127A through the first outlet 131 is in the chordwise direction C, and also along the camber line 107 of the airfoil 100. Due to the asymmetric positioning, shaping, or orientation of the first and second scalloped portions 141 and 142, the second outlet airflow direction 127B through the second outlet 132 is not aligned with the arc 107. It is also contemplated that the two outlet airflows 127 may be misaligned with each other by a predetermined angular difference, such as between 2 and 50 degrees in a non-limiting example. In this manner, the outlet airflow 127 (FIG. 2) may be customized to have various directions along the trailing edge 112 due to the shaping or positioning of the scalloped portion 140.

The smooth scalloped portion 170 may include rounded corners 175, the rounded corners 175 providing a rounded, curved, or otherwise smooth transition between the smooth scalloped portion 170 and the outer wall 103. The sharp scalloped portion 172 may include a sharp boundary or sharp edge 177 relative to the outer wall 103. Some non-limiting examples of sharp edges 177 include beveled, chamfered, pointed, or truncated edges.

Turning to FIG. 4, the second portion 102 of the airfoil 100 is shown with an outlet 130 and additional scalloped portions 140, wherein at least some of the scalloped portions may be non-uniform scalloped portions 142.

The wide scalloped portion 180 is shown as having a first width 185, such as a width in span. The narrow scalloped portion 181 is shown as having a second width 187 that is less than the first width 185. In the example shown, the wide scalloped portion 180 is located on the suction side 106 and the narrow scalloped portion 181 is located on the pressure side 106.

The illustrated variable width scalloped portion 190 is shown as having a first width 191 at a first location 193 and a second width 192 at a second location 194 downstream of the first location 193. The second width 192 in the illustrated example is less than the first width 191. In a non-limiting example, it is contemplated that other variable width scalloped portions (not shown) may have a plurality of widths, including alternately increasing and decreasing widths, continuously increasing widths toward trailing edge 112, or continuously decreasing widths toward trailing edge 112.

It is also contemplated that adjacent scalloped portions 140 may share at least one contact point 200. In one example, the contact point 200 may be in the form of a single point contact between adjacent scalloped portions 140 and is located across the span between adjacent outlets 130. In another example, the contact point 200 may form a transition surface 202 between two or more scalloped portions 140. Transition surface 202 may extend between pressure side 106 and suction side 108, between adjacent outlets 130, through trailing edge 112. The transition surface 202 may also be positioned on the pressure side 106 or the suction side 108 upstream of the trailing edge 112 and extend at least in the span direction S. In this manner, the transition surface 202 may incorporate at least two adjacent scalloped portions 140.

In another example (not shown), it is contemplated that the scalloped portion 140 may be used to modify, adjust, or customize an airfoil throat region, also referred to in the art as the minimum distance between circumferentially adjacent airfoils within the turbine engine 10 (FIG. 1), as measured from the first airfoil trailing edge to the outer wall of the second airfoil. It will be appreciated that the scalloped portion 140 may cause a change in airfoil thickness near the trailing edge 112, which may also cause a change in at least the span in the throat region. Such a varying or customized throat region may provide improved overall engine efficiency due to the modified airflow through the variable throat region.

It should be understood that the centerline, contact point, transition surface, width and length of the scalloped portion 140 may vary along the trailing edge 112, including in a repeating pattern or a random pattern as desired. In one example, scalloped portions 140 may have alternating centerlines along trailing edge 112. In another example, scalloped portion 140 may have a repeating pattern of greater and lesser chordal lengths. In another example, the scalloped portion 140 may have a combination of customized or tailored (custom machined) centerlines, contact points, transition surfaces, lengths and widths to form the airflow present at and around the trailing edge 112.

Further, while airfoil 100 is shown with a combination of non-uniform scalloped portion 142 and contact point 200, it is contemplated that an airfoil within turbine engine 10 (FIG. 1) may include one or both of non-uniform scalloped portion 142 and contact point 200. For example, the airfoil may include a plurality of uniformly formed and distributed scalloped portions along a trailing edge thereof, forming contact points or transition surfaces between at least two of the plurality of scalloped portions. In another example, the airfoil may include a plurality of non-uniform scalloped portions along its trailing edge without any contact points or transition surfaces. Any combination of conforming scalloped portions, non-conforming scalloped portions, contact points, or transition surfaces may be utilized, including geometric profiles, widths, lengths, centerlines, and the like.

A method of cooling an airfoil in a turbine engine includes supplying cooling air to an interior of the airfoil having a plurality of non-uniform scalloped portions (fig. 2-4) extending to near a trailing edge of the airfoil and a plurality of outlets (fig. 2), at least some of the scalloped portions being interposed between adjacent outlets. The method further includes exhausting the cooling air through a plurality of outlets. Optionally, the method may include defining an airflow outlet direction through at least one outlet of the plurality of outlets via a plurality of non-uniform scalloped portions, as shown in fig. 3.

Various aspects of the present disclosure provide various benefits, including that the local airflow through the outlet and near the trailing edge may be adjusted using non-uniform scalloped portions to provide airfoil cooling, mix the local airflow near the trailing edge, or purge stagnant air that may be located near the trailing edge. Further, the scalloped portions may provide a thinner trailing edge that is more efficient than conventional airfoil trailing edges, which may reduce the weight of the airfoil, improve the hole cooling of the airfoil, and improve the aerodynamic performance of the airfoil.

It should be understood that the application of the disclosed design is not limited to turbine engines having fan and booster sections, but is also applicable to turbojets and turboshafts.

To the extent not already described, the different features and structures of the various embodiments may be used in combination or substituted for one another as desired. That one feature is not shown in all embodiments, which does not mean that it cannot be so shown, but is done for simplicity of description. Thus, the various features of the different embodiments can be mixed and matched as desired to form new embodiments, whether or not such new embodiments are explicitly described. The present disclosure encompasses all combinations or permutations of features described herein.

This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.

Further aspects of the invention are provided by the subject matter of the following clauses:

1. an airfoil for a turbine engine, comprising: an outer wall bounding an interior and defining a pressure side and a suction side, the outer wall extending between a leading edge and a trailing edge to define a chordwise direction and between a root and a tip to define a span direction; a plurality of outlets extending to near the trailing edge; and a plurality of non-uniform scalloped portions extending along the trailing edge, wherein at least some of the non-uniform scalloped portions are between adjacent outlets.

2. The airfoil of any preceding claim, wherein the plurality of non-uniform scalloped portions comprise at least one of a non-uniform length, a non-uniform width, a non-uniform centerline, or a non-uniform geometry.

3. The airfoil of any preceding claim, wherein the plurality of non-uniform scalloped portions includes a long scalloped portion having a first length and a short scalloped portion having a second length less than the first length.

4. The airfoil of any preceding clause, wherein the plurality of non-uniform scalloped portions comprises a wide scalloped portion having a first width and a narrow scalloped portion having a second width less than the first width.

5. The airfoil of any preceding claim, wherein the plurality of non-uniform scalloped portions comprises a variable width scalloped portion having a first width at a first location and a second width at a second location.

6. The airfoil of any preceding claim, wherein the plurality of non-uniform scalloped portions comprises a transition surface located upstream of the trailing edge and joining two adjacent scalloped portions.

7. The airfoil of any preceding item, wherein the plurality of non-uniform scalloped portions comprises a first inclined scalloped portion having a first centerline extending in a span direction and a chordwise direction.

8. The airfoil of any preceding claim, wherein the plurality of non-uniform scalloped portions comprises a second skewed scalloped portion having a second centerline that is not aligned with the first centerline.

9. The airfoil of any preceding claim, wherein the plurality of non-uniform scalloped portions comprises smooth scalloped portions having rounded corners and sharp scalloped portions having sharp edges.

10. The airfoil of any preceding claim, wherein the outlet extends along the trailing edge and the plurality of non-uniform scalloped portions extend along at least one of the pressure side or the suction side.

11. An airfoil for a turbine engine, comprising: an outer wall bounding an interior and defining a pressure side and a suction side, the outer wall extending between a leading edge and a trailing edge to define a chordwise direction and between a root and a tip to define a span direction; a plurality of outlets extending to near the trailing edge; a plurality of scalloped portions extending along the trailing edge between adjacent outlets; and at least one contact point between two of the plurality of scalloped portions.

12. The airfoil of any preceding claim, wherein the first scalloped portion is located on a pressure side and the second scalloped portion is located on a suction side.

13. The airfoil of any preceding clause, wherein the at least one contact point is located at a trailing edge between the first scalloped portion and the second scalloped portion and is located across a span between adjacent outlets.

14. The airfoil of any preceding clause, wherein the first scalloped portion and the second scalloped portion are both located on one of the pressure side or the suction side, and the at least one point of contact is between the first scalloped portion and the second scalloped portion.

15. The airfoil of any preceding claim, wherein the at least one contact point is upstream of the trailing edge.

16. The airfoil of any preceding claim, wherein the at least one contact point further comprises a transition surface between two of the plurality of scalloped portions.

17. The airfoil of any preceding claim, wherein the transition surface extends across the trailing edge between the pressure side and the suction side.

18. The airfoil of any preceding claim, wherein the transition surface extends at least in a span direction and is located on one of the pressure side or the suction side.

19. A method of cooling an airfoil in a turbine engine, the method comprising supplying cooling air to an interior of the airfoil, the airfoil having a plurality of non-uniform scalloped portions extending along a trailing edge of the airfoil and a plurality of outlets, wherein at least some of the non-uniform scalloped portions are interposed between adjacent outlets; and discharging the cooling air through a plurality of outlets.

20. The method of any preceding clause, further comprising defining an airflow outlet direction through at least one outlet of the plurality of outlets via a plurality of non-uniform scalloped portions.