Shark-bite tip shelf cooling configuration

An airfoil includes a leading edge, a trailing edge, a suction surface, a pressure surface, and a tip shelf. The suction surface and the pressure surface both extend axially to connect the leading edge to the trailing edge. The suction surface and the pressure surface both extend radially from a root section of the airfoil a tip section of the airfoil. The tip shelf is formed along the tip section, and includes a triangular pocket.

BACKGROUND

This invention relates generally to turbomachinery, and specifically to turbine flow path components for gas turbine engines. In particular, the invention relates to cooling techniques for airfoils and other gas turbine engine components exposed to hot working fluid flow, including, but not limited to, rotor blades and stator vane airfoils, endwall surfaces including platforms, shrouds and compressor and turbine casings, combustor liners, turbine exhaust assemblies, thrust augmentors and exhaust nozzles.

Gas turbine engines are rotary-type combustion turbine engines built around a power core made up of a compressor, combustor and turbine, arranged in flow series with an upstream inlet and downstream exhaust. The compressor section compresses air from the inlet, which is mixed with fuel in the combustor and ignited to generate hot combustion gas. The turbine section extracts energy from the expanding combustion gas, and drives the compressor section via a common shaft. Expanded combustion products are exhausted downstream, and energy is delivered in the form of rotational energy in the shaft, reactive thrust from the exhaust, or both.

Gas turbine engines provide efficient, reliable power for a wide range of applications in aviation, transportation and industrial power generation. Small-scale gas turbine engines typically utilize a one-spool design, with co-rotating compressor and turbine sections. Larger-scale combustion turbines including jet engines and industrial gas turbines (IGTs) are generally arranged into a number of coaxially nested spools. The spools operate at different pressures, temperatures and spool speeds, and may rotate in different directions.

Individual compressor and turbine sections in each spool may also be subdivided into a number of stages, formed of alternating rows of rotor blade and stator vane airfoils. The airfoils are shaped to turn, accelerate and compress the working fluid flow, or to generate lift for conversion to rotational energy in the turbine.

Industrial gas turbines often utilize complex nested spool configurations, and deliver power via an output shaft coupled to an electrical generator or other load, typically using an external gearbox. In combined cycle gas turbines (CCGTs), a steam turbine or other secondary system is used to extract additional energy from the exhaust, improving thermodynamic efficiency. Gas turbine engines are also used in marine and land-based applications, including naval vessels, trains and armored vehicles, and in smaller-scale applications such as auxiliary power units.

Aviation applications include turbojet, turbofan, turboprop and turboshaft engine designs. In turbojet engines, thrust is generated primarily from the exhaust. Modern fixed-wing aircraft generally employ turbofan and turboprop configurations, in which the low pressure spool is coupled to a propulsion fan or propeller. Turboshaft engines are employed on rotary-wing aircraft, including helicopters, typically using a reduction gearbox to control blade speed. Unducted (open rotor) turbofans and ducted propeller engines also known, in a variety of single-rotor and contra-rotating designs with both forward and aft mounting configurations.

Aviation turbines generally utilize two and three-spool configurations, with a corresponding number of coaxially rotating turbine and compressor sections. In two-spool designs, the high pressure turbine drives a high pressure compressor, forming the high pressure spool or high spool. The low-pressure turbine drives the low spool and fan section, or a shaft for a rotor or propeller. In three-spool engines, there is also an intermediate pressure spool. Aviation turbines are also used to power auxiliary devices including electrical generators, hydraulic pumps and elements of the environmental control system, for example using bleed air from the compressor or via an accessory gearbox.

Turbofan engines are commonly divided into high and low bypass configurations. High bypass turbofans generate thrust primarily from the fan, which accelerates airflow through a bypass duct oriented around the engine core. This design is common on commercial aircraft and transports, where noise and fuel efficiency are primary concerns. The fan rotor may also operate as a first stage compressor, or as a pre-compressor stage for the low-pressure compressor or booster module. Variable-area nozzle surfaces can also be deployed to regulate the bypass pressure and improve fan performance, for example during takeoff and landing. Advanced turbofan engines may also utilize a geared fan drive mechanism to provide greater speed control, reducing noise and increasing engine efficiency, or to increase or decrease specific thrust.

Low bypass turbofans produce proportionally more thrust from the exhaust flow, generating greater specific thrust for use in high-performance applications including supersonic jet aircraft. Low bypass turbofan engines may also include variable-area exhaust nozzles and afterburner or augmentor assemblies for flow regulation and short-term thrust enhancement. Specialized high-speed applications include continuously afterburning engines and hybrid turbojet/ramjet configurations.

Across these applications, turbine performance depends on the balance between higher pressure ratios and core gas path temperatures, which tend to increase efficiency, and the related effects on service life and reliability due to increased stress and wear. This balance is particularly relevant to gas turbine engine components in the hot sections of the compressor, combustor, turbine and exhaust sections, where active cooling is required to prevent damage due to high gas path temperatures and pressures.

SUMMARY

An airfoil includes a leading edge, a trailing edge, a suction surface, a pressure surface, and a tip shelf. The suction surface and the pressure surface both extend axially to connect the leading edge to the trailing edge. The suction surface and the pressure surface both extend radially from a root section of the airfoil to a tip section of the airfoil. The tip shelf is formed along the tip section, and includes a triangular cooling pocket.

A rotor blade for a gas turbine engine includes an airfoil and a tip shelf. The airfoil includes a convex surface and a concave surface. Both the convex surface and the concave surface extend radially from a root section to a tip section and axially from a leading edge to a trailing edge. A mean camber line is defined midway between the convex surface and the concave surface. The tip shelf extends axially along the tip section between the mean camber line and the concave surface. The tip shelf includes at least one wall positioned perpendicular to fluid flow streamlines.

An airfoil for a gas turbine engine includes a leading edge, a trailing edge, a pressure surface, a suction surface, and a tip shelf. The leading edge and the trailing edge both extend from a root section to a tip section. The pressure surface and the suction surface both extend between the leading edge and the trailing edge. The pressure surface and the suction surface define a mean camber line therebetween. The tip shelf is located at the tip section, and includes a plurality of triangular recesses. Each triangular recess has three sides, wherein two of the sides are defined by perimeter walls.

DETAILED DESCRIPTION

FIG. 1is a cross-sectional view of gas turbine engine10. Gas turbine engine (or turbine engine)10includes a power core with compressor section12, combustor14and turbine section16arranged in flow series between upstream inlet18and downstream exhaust20. Compressor section12and turbine section16are arranged into a number of alternating stages of rotor airfoils (or blades)22and stator airfoils (or vanes)24.

In the turbofan configuration ofFIG. 1, propulsion fan26is positioned in bypass duct28, which is coaxially oriented about the engine core along centerline (or turbine axis) CL. An open-rotor propulsion stage26may also provided, with turbine engine10operating as a turboprop or unducted turbofan engine. Alternatively, fan rotor26and bypass duct28may be absent, with turbine engine10configured as a turbojet or turboshaft engine, or an industrial gas turbine.

For improved service life and reliability, components of gas turbine engine10are provided with an improved cooling configuration, as described below. Suitable components for the cooling configuration include rotor airfoils22, stator airfoils24and other gas turbine engine components exposed to hot gas flow, including, but not limited to, platforms, shrouds, casings and other endwall surfaces in hot sections of compressor12and turbine16, and liners, nozzles, afterburners, augmentors and other gas wall components in combustor14and exhaust section20.

In the two-spool, high bypass configuration ofFIG. 1, compressor section12includes low pressure compressor (LPC)30and high pressure compressor (HPC)32, and turbine section16includes high pressure turbine (HPT)34and low pressure turbine (LPT)36. Low pressure compressor30is rotationally coupled to low pressure turbine36via low pressure (LP) shaft38, forming the LP spool or low spool. High pressure compressor32is rotationally coupled to high pressure turbine34via high pressure (HP) shaft40, forming the HP spool or high spool.

Flow F at inlet18divides into primary (core) flow FPand secondary (bypass) flow Fsdownstream of fan rotor26. Fan rotor26accelerates secondary flow FSthrough bypass duct28, with fan exit guide vanes (FEGVs)42to reduce swirl and improve thrust performance. In some designs, structural guide vanes (SGVs)42are used, providing combined flow turning and load bearing capabilities.

Primary flow FPis compressed in low pressure compressor30and high pressure compressor32, then mixed with fuel in combustor14and ignited to generate hot combustion gas. The combustion gas expands to provide rotational energy in high pressure turbine34and low pressure turbine36, driving high pressure compressor32and low pressure compressor30, respectively. Expanded combustion gases exit through exhaust section (or exhaust nozzle)20, which can be shaped or actuated to regulate the exhaust flow and improve thrust performance.

Low pressure shaft38and high pressure shaft40are mounted coaxially about centerline CL, and rotate at different speeds. Fan rotor (or other propulsion stage)26is rotationally coupled to low pressure shaft38. In advanced designs, fan drive gear system44is provided for additional fan speed control, improving thrust performance and efficiency with reduced noise output.

Fan rotor26may also function as a first-stage compressor for gas turbine engine10, and LPC30may be configured as an intermediate compressor or booster. Alternatively, propulsion stage26has an open rotor design, or is absent, as described above. Gas turbine engine10thus encompasses a wide range of different shaft, spool and turbine engine configurations, including one, two and three-spool turboprop and (high or low bypass) turbofan engines, turboshaft engines, turbojet engines, and multi-spool industrial gas turbines.

In each of these applications, turbine efficiency and performance depend on the overall pressure ratio, defined by the total pressure at inlet18as compared to the exit pressure of compressor section12, for example at the outlet of high pressure compressor32, entering combustor14. Higher pressure ratios, however, also result in greater gas path temperatures, increasing the cooling loads on rotor airfoils22, stator airfoils24and other components of gas turbine engine10. To reduce operating temperatures, increase service life and maintain engine efficiency, these components are provided with improved cooling configurations, as described below. Suitable components include, but not limited to, cooled gas turbine engine components in compressor sections30and32, combustor14, turbine sections34and36, and exhaust section20of gas turbine engine10.

FIG. 2Ais a perspective view of rotor airfoil (or blade)22for gas turbine engine10, as shown inFIG. 1, or for another turbomachine. Rotor airfoil22extends axially from leading edge51to trailing edge52, defining pressure surface53(front) and suction surface54(back) therebetween.

Pressure and suction surfaces53and54form the major opposing surfaces or walls of airfoil22, extending axially between leading edge51and trailing edge52, and radially from root section55, adjacent inner diameter (ID) platform56, to tip section57, opposite ID platform56. In some designs, tip section57is shrouded.

Cooling holes or outlets60are provided on one or more surfaces of airfoil22, for example along leading edge51, trailing edge52, pressure (or concave) surface53, or suction (or convex) surface54, or a combination thereof. Cooling holes or passages60may also be provided on the endwall surfaces of airfoil22, for example along ID platform56, or on a shroud or engine casing adjacent tip section57.

Cooling holes or outlets60are provided along one or more surfaces of airfoil24, for example leading or trailing edge61or62, pressure (concave) or suction (convex) surface63or64, or a combination thereof. Cooling holes or passages60may also be provided on the endwall surfaces of airfoil24, for example along ID platform66and OD platform68.

Rotor airfoils22(FIG. 2A) and stator airfoils24(FIG. 2B) are formed of high strength, heat resistant materials such as high temperature alloys and superalloys, and are provided with thermal and erosion-resistant coatings. Airfoils22and24are also provided with internal cooling passages and cooling holes60to reduce thermal fatigue and wear, and to prevent melting when exposed to hot gas flow in the higher temperature regions of a gas turbine engine or other turbomachine. Cooling holes60deliver cooling fluid (e.g., steam or air from a compressor) through the outer walls and platform structures of airfoils22and24, creating a thin layer (or film) of cooling fluid to protect the outer (gas path) surfaces from high temperature flow.

While surface cooling extends service life and increases reliability, injecting cooling fluid into the gas path also reduces engine efficiency, and the cost in efficiency increases with the required cooling flow. Cooling holes60are thus provided with improved metering and inlet geometry to reduce jets and blow off, and improved diffusion and exit geometry to reduce flow separation. Cooling holes60reduce flow requirements and improve the spread of cooling fluid across the hot outer surfaces of airfoils22and24, and other gas turbine engine components, so that less flow is needed for cooling and efficiency is maintained or increased.

FIG. 3Ais a perspective view of airfoil102in a rotor blade configuration, andFIG. 3Bis an enlarged perspective view of tip section100for airfoil102. Airfoil102includes root section98, tip section100, leading edge104, trailing edge106, pressure or concave surface108, suction or convex surface110, tip surface112, squealer pocket114having floor116and wall118, tip shelf120, and cooling outlets122. Tip shelf120includes shelf pockets124, each shelf pocket124having three sides (first side126, second side128, third side130) and three points (first point132, second point134, and third point136). Also shown inFIG. 3Bare chord line CH and camber line CA. Shelf pockets124forming tip shelf120are triangular in shape to reduce fluid flow leakage past tip section100, improve temperature distribution within tip section100, and facilitate filling of cooling air for tip section100.

Airfoil102is similar to rotor blade airfoil22described above with reference toFIG. 2A. As shown inFIG. 3A, airfoil102extends axially from leading edge104to trailing edge106, with trailing edge106located downstream of leading edge104. Pressure or concave surface108forms front of airfoil102, and suction or convex surface110forms back of airfoil102, such that pressure surface108and suction surface110define a profile of airfoil102. Pressure surface108and suction surface110form the major opposing surfaces or walls of airfoil102. Pressure surface108and suction surface110extend axially between leading edge104and trailing edge106. Pressure surface108and suction surface110extend radially from root section98to tip section100, thereby defining a span height of airfoil102.

Tip section100extends axially from leading edge104to trailing edge106, and transversely or circumferentially from pressure surface108to suction surface110. Tip section100includes a substantially planar tip surface112having two major recesses, pockets, or cut-outs: squealer pocket114and tip shelf120. As shown inFIG. 3B, at least one cooling outlet122is present in each of squealer pocket114and tip shelf120, although more and less cooling outlets are equally contemplated.

Squealer pocket114is relatively small, located near leading edge104, and is at least partially upstream of tip shelf120. In the depicted embodiment, squealer pocket114extends for about 10% of chord line CH (a straight line connecting leading edge104to trailing edge106), although squealer pocket114could be designed as a larger or smaller recess. As shown, squealer pocket114extends across camber line CA (a curved line midway between pressure surface108and suction surface110), although other configurations are contemplated. Squealer pocket114has floor116, which is recessed in comparison to, and substantially parallel with, tip surface112. Extending radially upwards from floor116is substantially perpendicular wall118. Wall118forms a continuous perimeter for squealer pocket114, which is closed on all sides (i.e. fully enclosed axially, radially, and transversely). In the depicted embodiment, squealer pocket114is tear drop shaped with an enlarged portion positioned near leading edge104, although other suitable shapes are within the scope of this disclosure. Likewise, one cooling outlet122(e.g. cooling hole) is located centrally within squealer pocket114, but more, less, or different cooling outlet features are possible.

Tip shelf120is relatively large and extends axially from a location near leading edge104(but downstream of a least a portion squealer pocket114), along the intersection of tip section100and pressure surface108, to a location near trailing edge106. In the depicted embodiment, tip shelf120extends for about 75% of chord line CH, although tip shelf120could be designed as a larger or smaller axial recess. As shown, tip shelf120is located wholly on pressure side108of camber line CA, although other configurations are contemplated (i.e. a larger tip shelf could beyond camber line CA toward suction surface110). Tip shelf120is formed from a plurality of triangular shaped shelf pockets124. In the depicted embodiment, each shelf pocket includes at least two cooling outlets122, but more, less, or different cooling outlet features are possible. Cooling outlets122as used herein includes cooling outlet features of any shape (e.g. circular, rectangular slot, etc.). Cooling outlets122are also located on other surfaces of airfoil102, as shown and described with reference to cooling outlets60on rotor blade22inFIG. 2A.

In the depicted embodiment, tip shelf120is formed from four triangular shaped shelf pockets124, although more or less shelf pockets124are possible. The triangular design of tip shelf120resembles a shark-bite or a zig-zag. Each shelf pocket124is defined by three sides (first side126, second side128, third side130), which are joined to one another at three points (first point132, second point134, and third point136). One or more of the triangular shaped shelf pockets124can be a right-angle triangle. In the embodiment depicted, a single shelf pocket124extends for about 20% of the chord line CH, but smaller and larger shelf pockets124are within the scope of this disclosure.

The following description refers to one of the shelf pockets124, but is equally applicable to each of the shelf pockets124. First side126extends axially along the junction of tip section100and pressure surface108, such that first side126is open to and continuous with pressure surface108. Second side128is attached to first side126at first point132, which is located at the junction of tip section100and pressure surface108. Second side128extends transversely across tip surface122from first point132at an acute angle to first side126. Second side128extends angularly toward suction surface110to second point134, but does not extend beyond camber line CA. In other embodiments, second side128does extend beyond camber line CA closer to suction side110. Third side130is attached to second side128at second point134, which is located at or near camber line CA. Third side130extends transversely across tip surface122from second point134at an approximately right angle to second side128. Third side130extends angularly toward pressure surface112to third point136located at a junction of tip section and pressure surface108. Third side130is attached to first side at third point136to complete a perimeter of shelf pocket124. While the points or corners and sides or walls are described as angular (e.g. acute or right), it should be appreciated that, as a result of manufacturing, these junctions may not be perfectly sharp and may be slightly rounded.

In between the perimeter formed by first side126, second side128, and third side130is a floor of tip pocket124, which is recessed from, and substantially parallel to, tip surface112. While first side126is open, second side128and third side130are closed. More specifically, a wall extends axially upward and substantially perpendicularly from the floor along both second side128and third side130. Second side128and third side130include these vertical walls to collectively form a pocket around cooling holes(s)122and/or cooling slots124for pooling cooling air at tip shelf120.

When airfoil102is exposed to high temperature fluid flow, for example in the turbine and high pressure compressor sections of gas turbine engine, tip section100experiences oxidation, erosion, burn-through, thermal mechanical fatigue, and other high temperature effects. To address this problem, tip section100of airfoil102is formed with squealer pocket118and tip shelf120, which aid in cooling of tip section100by pooling or collecting cooling fluid. Cooling fluid (e.g. air) is introduced into squealer pocket118and tip shelf120by cooling outlets122. Squealer pocket118then maintains a region or pocket of cooling fluid along tip section100of airfoil102, between pressure surface108and suction surface110. Similarly, tip shelf120maintains a region or pocket of cooling fluid along pressure surface120, between leading edge104and trailing edge106of tip section100. These pockets of pooled cooling fluid provide a more uniform cooling temperature along tip section100. Since the overall area of tip shelf120is broken up into discrete shelf pockets124each having a reduced volume, shelf pockets124will fill more quickly and fully with cooling fluid, as well as require less cooling fluid.

The triangular design of tip shelf120also decreases the heat transfer coefficient across tip section100, improving the performance and service life of airfoil21. The triangular structure of tip shelf120reduces mass flow across the tip, so the heat transfer coefficient is also reduced in tip section100. That means there is less heat transfer from the working fluid into airfoil tip section100, resulting in decreased thermal effects and improved service life for airfoil102. Additional benefits of the present disclosure are discussed below with reference toFIG. 4.

FIG. 4is a perspective view of tip section100fromFIGS. 3A-3Bwith fluid flow streamlines138shown. Airfoil102includes tip section100, leading edge104, trailing edge106, pressure surface108, suction surface110, tip surface112, squealer pocket114, and tip shelf120. Tip shelf120includes shelf pockets124, each shelf pocket124having three sides (first side126, second side128, third side130). Second side128includes second rib or wall140and third side130includes third rib or wall142. Third wall142is oriented substantially perpendicular to fluid flow streamlines138in order to reduce fluid flow past tip section100.

As described above with reference toFIGS. 3A-3B, tip section100extends axially from leading edge104to trailing edge106, and transversely or circumferentially from pressure surface108to suction surface110. Tip section100includes a substantially planar tip surface112having two major cavities, recesses, pockets, or cut-outs: squealer pocket114and tip shelf120. Squealer pocket114extends between pressure surface108and suction surface110near leading edge104. Tip shelf120extends adjacent pressure surface108, from a location near leading edge104to a location near trailing edge106.

Tip shelf120defines an open (or discontinuous) perimeter radial recess in tip section120. That is, the tip shelf120is open along pressure surface108. Tip shelf120is formed from a plurality of tip pockets124, each of which individually forms an open perimeter radial recess. Each shelf pocket124is triangular and includes a first side126open along and continuous with pressure surface108. Second side128is closed by second wall140extending radially or vertically upwards from a floor of tip pocket124. Similarly, third side130is closed by third wall142extending radially or vertically upwards from a floor of tip pocket124. Second wall140and third wall142close second side128and third side30to retain a pocket of cooling air in tip pocket124as described above with reference toFIG. 3. Further, third wall142is oriented substantially perpendicular to fluid flow streamlines138in order to block fluid flow across tip section. More specifically, third wall142for each tip pocket124is oriented perpendicular to fluid flow streamlines138, which requires each third wall142to have a slightly different angle along tip shelf120thereby maximizing fluid flow resistance.

During use, working fluid (hot air or gas) flows along pressure side108of airfoil102. Some of the working fluid leaks or migrates across tip section100, as shown by fluid flow streamlines138. This leakage across tip section100is undesirable as it is unused “work” that reduces engine efficiency. Tip shelf120is formed from triangular tip pockets124that reduce fluid leakage or maximize flow resistance across tip section100. More specifically, third wall142of each tip pocket124is oriented substantially perpendicular to fluid flow streamlines138in order to directly block fluid from flowing across tip section100. Since third wall142extends to the junction of tip shelf120and suction surface108, fluid flow is also blocked from entering a downstream tip pocket124. The angular nature of tip shelf120provides a more effective blockade to fluid flow than a conventional tip shelf.

The gas turbine engine components, gas path walls and cooling features (cooling holes, cooling slots, squealer pocket, tip shelf, etc.) described herein can thus be manufactured using one or more of a variety of different processes. These techniques provide each cooling feature with its own particular configuration, including, but not limited to, inlet, metering, transition, diffusion, outlet, upstream wall, downstream wall, lateral wall, longitudinal, lobe and downstream edge features. In some cases, multiple techniques can be combined to improve overall cooling performance or reproducibility, or to reduce manufacturing costs.

Suitable manufacturing techniques for forming the cooling configurations described here include, but are not limited to, electrical discharge machining (EDM), laser drilling, laser machining, electrical chemical machining (ECM), water jet machining, casting, conventional machining and combinations thereof. Electrical discharge machining includes both machining using a shaped electrode as well as multiple pass methods using a hollow spindle or similar electrode component. Laser machining methods include, but are not limited to, material removal by ablation, trepanning and percussion laser machining. Conventional machining methods include, but are not limited to, milling, drilling and grinding.

The gas flow path walls and outer surfaces of some gas turbine engine components include one or more coatings, such as bond coats, thermal barrier coatings, abrasive coatings, abradable coatings and erosion or erosion-resistant coatings. For components having a coating, the inlet, metering portion, transition, diffusion portion and outlet cooling features may be formed prior to coating application, after a first coating (e.g., a bond coat) is applied, or after a second or third (e.g., interlayer) coating process, or a final coating (e.g., environmental or thermal barrier) coating process. Depending on component type, cooling feature location, repair requirements and other considerations, the diffusion portion and outlet features may be located within a wall or substrate, within a thermal barrier coating or other coating layer applied to a wall or substrate, or based on combinations thereof. The cooling geometry and other features may remain as described above, regardless of position relative to the wall and coating materials or airfoil materials.

In addition, the order in which cooling features are formed and coatings are applied may affect selection of manufacturing techniques, including techniques used in forming the inlet, metering portion, transition, outlet, diffusion portion and other cooling features. For example, when a thermal barrier coat or other coating is applied to the outer surface of a gas path wall before the cooling hole or passage is produced, laser ablation or laser drilling may be used. Alternatively, either laser drilling or water jet machining may be used on a surface without a thermal barrier coat. Additionally, different machining methods may be more or less suitable for forming different features of the cooling hole or cooling passage, for example, different EDM, laser machining and other machining techniques may be used for forming the outlet and diffusion features, and for forming the transition, metering and inlet features.

Discussion of Possible Embodiments

An airfoil includes a leading edge, a trailing edge, a suction surface, a pressure surface, and a tip shelf. The suction surface and the pressure surface both extend axially to connect the leading edge to the trailing edge. The suction surface and the pressure surface both extend radially from a root section of the airfoil to a tip section of the airfoil. The tip shelf is formed along the tip section, and includes a triangular cooling pocket.

the triangular pocket can include a first side that is open to and continuous with the pressure surface;

the triangular pocket can include a second side attached to the first side at a first point located near the pressure surface, the second side having a second side wall extending from the first point across the tip section toward the suction surface at an acute angle;

the triangular pocket can include a third side attached to the second side at a second point located between the pressure surface and the suction surface, the third side having a third side wall extending from the second point across the tip section toward the pressure surface;

the third side can be attached to the first side at a third point located near the pressure surface, the third side wall extending from the third point across the tip section to the second point at an acute angle;

a cooling hole can be located inside of the triangular pocket; and/or

a cooling slot can be located inside of the triangular pocket.

A rotor blade for a gas turbine engine includes an airfoil and a tip shelf. The airfoil includes a convex surface and a concave surface. Both the convex surface and the concave surface extend radially from a root section to a tip section and axially from a leading edge to a trailing edge. A mean camber line is defined midway between the convex surface and the concave surface. The tip shelf extends axially along the tip section between the mean camber line and the concave surface. The tip shelf includes at least one wall positioned perpendicular to fluid flow streamlines.

The tip shelf can include a plurality of shelf sections and each self section includes a wall positioned perpendicular to fluid flow streamlines;

the walls can be positioned perpendicular to fluid flow streamlines are connected by a plurality of opposing walls to form a zig-zag perimeter for the tip shelf;the tip shelf can be radially open to the concave surface;a squealer pocket can be located on the tip section;at least a portion of the squealer pocket can be located axially upstream from the tip shelf; and/ora plurality of cooling outlets can be formed in the tip shelf to maintain a pocket of cooling fluid along the tip section.

An airfoil for a gas turbine engine includes a leading edge, a trailing edge, a pressure surface, a suction surface, and a tip shelf. The leading edge and the trailing edge both extend from a root section to a tip section. The pressure surface and the suction surface both extend between the leading edge and the trailing edge. The pressure surface and the suction surface define a mean camber line therebetween. The tip shelf is located at the tip section, and includes a plurality of triangular recesses. Each triangular recess has three sides, wherein two of the sides are defined by perimeter walls.

The airfoil of the preceding paragraph can optionally include, additionally and/or alternatively any, one or more of the following features, configurations and/or additional components:one side of each triangular recess can be open to the pressure surface;the tip shelf can extend along more than 50% of a chord line of the airfoil;the triangular recesses can be located between the mean camber line and the pressure surface;at least one cooling hole or cooling slot can be located in each triangular recess; and/orat least one perimeter wall of each triangular recess can be oriented perpendicular to streamlines of fluid flow.