Engine component with non-diffusing section

An airfoil for a turbine engine can include an outer wall bounding an interior and extending in a chord-wise direction between a leading edge and a trailing edge, and also extending radially between a root and a tip. A tip rail including at least one tip cooling cavity can project from the tip, and cooling holes can be provided in the tip rail.

BACKGROUND

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

Turbine engines are often designed to operate at high temperatures to maximize engine efficiency. It is beneficial to provide cooling measures for components such as airfoils in the high-temperature environment, where such cooling measures can reduce material wear on these components and provide for increased structural stability during engine operation.

BRIEF DESCRIPTION

In one aspect, an airfoil for a turbine engine includes an outer wall bounding an interior and extending axially between a leading edge and a trailing edge to define a chord-wise direction with a chord line, and also extending along a radial line between a root and a tip to define a span-wise direction. A tip rail can project from the tip in the span-wise direction, the tip rail including an exterior surface spaced from an interior surface, at least one cooling passage formed in the interior of the airfoil, at least one cooling cavity provided within the tip rail between the exterior and interior surfaces and fluidly coupled to the at least one cooling passage, and a cooling trench extending in the chord-wise direction in the exterior surface of the tip rail and at least partially defined by first and second cooling holes provided in the tip rail. Each of the first and second cooling holes can include an inlet on the interior surface fluidly coupled to the at least one cooling cavity and an outlet on the exterior surface, where adjacent outlets define at least a portion of the cooling trench.

In another aspect, a turbine engine includes a fan section, compressor section, a combustion section, and a turbine section in axial flow arrangement. At least one of the compressor section and turbine section can have an airfoil including an outer wall bounding an interior and extending axially between a leading edge and a trailing edge to define a chord-wise direction with a chord line, and also extending along a radial line between a root and a tip to define a span-wise direction. A tip rail can project from the tip in the span-wise direction, the tip rail including an exterior surface spaced from an interior surface, at least one cooling passage formed in the interior of the airfoil, at least one cooling cavity provided within the tip rail between the exterior and interior surfaces; and a cooling trench extending in the chord-wise direction in the exterior surface of the tip rail and at least partially defined by first and second cooling holes provided in the tip rail. Each of the first and second cooling holes includes an inlet on the interior surface fluidly coupled to the at least one cooling cavity and an outlet on the exterior surface, where adjacent outlets define at least a portion of the cooling trench.

In yet another aspect, a method of cooling a tip wall of an airfoil in a turbine engine includes emitting cooling air into a trench extending in a chord-wise direction along the tip wall. The trench can be at least partially defined by cooling holes within the tip wall, where the cooling holes include overlapping outlets.

DESCRIPTION OF EMBODIMENTS

The described embodiments of the present disclosure are directed to cooling holes for turbine engine airfoils. For purposes of illustration, the present disclosure will be described with respect to the turbine for an aircraft turbine engine. It will be understood, however, that the disclosure is not so limited and may have general applicability within an engine, including compressors, as well as in 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 closer to the engine inlet as compared to another component. The term “aft” or “downstream” used in conjunction with “forward” or “upstream” refers to a direction toward the rear or outlet of the engine or being relatively closer to the engine outlet as compared to another component.

As used herein, “a set” can include any number of the respectively described elements, including only one element. Additionally, the terms “radial” or “radially” as used herein refer to a dimension extending between a center longitudinal axis of the engine and an outer engine circumference.

FIG. 1is a schematic cross-sectional diagram of a gas turbine engine10for an aircraft. The engine10has a generally longitudinally extending axis or centerline12extending forward14to aft16. The engine10includes, in downstream serial flow relationship, a fan section18including a fan20, a compressor section22including a booster or low pressure (LP) compressor24and a high pressure (HP) compressor26, a combustion section28including a combustor30, a turbine section32including a HP turbine34, and a LP turbine36, and an exhaust section38.

The fan section18includes a fan casing40surrounding the fan20. The fan20includes a plurality of fan blades42disposed radially about the centerline12. The HP compressor26, the combustor30, and the HP turbine34form a core44of the engine10, which generates combustion gases. The core44is surrounded by core casing46, which can be coupled with the fan casing40.

A HP shaft or spool48disposed coaxially about the centerline12of the engine10drivingly connects the HP turbine34to the HP compressor26. ALP shaft or spool50, which is disposed coaxially about the centerline12of the engine10within the larger diameter annular HP spool48, drivingly connects the LP turbine36to the LP compressor24and fan20. The spools48,50are rotatable about the engine centerline and couple to a plurality of rotatable elements, which can collectively define a rotor51.

The LP compressor24and the HP compressor26respectively include a plurality of compressor stages52,54, in which a set of compressor blades56,58rotate relative to a corresponding set of static compressor vanes60,62to compress or pressurize the stream of fluid passing through the stage. In a single compressor stage52,54, multiple compressor blades56,58can be provided in a ring and can extend radially outwardly relative to the centerline12, from a blade platform to a blade tip, while the corresponding static compressor vanes60,62are positioned upstream of and adjacent to the rotating blades56,58. It is noted that the number of blades, vanes, and compressor stages shown inFIG. 1were selected for illustrative purposes only, and that other numbers are possible.

The blades56,58for a stage of the compressor can be mounted to (or integral to) a disk61, which is mounted to the corresponding one of the HP and LP spools48,50. The vanes60,62for a stage of the compressor can be mounted to the core casing46in a circumferential arrangement.

The HP turbine34and the LP turbine36respectively include a plurality of turbine stages64,66, in which a set of turbine blades68,70are rotated relative to a corresponding set of static turbine vanes72,74(also called a nozzle) to extract energy from the stream of fluid passing through the stage. In a single turbine stage64,66, multiple turbine blades68,70can be provided in a ring and can extend radially outwardly relative to the centerline12while the corresponding static turbine vanes72,74are positioned upstream of and adjacent to the rotating blades68,70. It is noted that the number of blades, vanes, and turbine stages shown inFIG. 1were selected for illustrative purposes only, and that other numbers are possible.

The blades68,70for a stage of the turbine can be mounted to a disk71, which is mounted to the corresponding one of the HP and LP spools48,50. The vanes72,74for a stage of the compressor can be mounted to the core casing46in a circumferential arrangement.

Complementary to the rotor portion, the stationary portions of the engine10, such as the static vanes60,62,72,74among the compressor and turbine section22,32are also referred to individually or collectively as a stator63. As such, the stator63can refer to the combination of non-rotating elements throughout the engine10.

In operation, the airflow exiting the fan section18is split such that a portion of the airflow is channeled into the LP compressor24, which then supplies pressurized air76to the HP compressor26, which further pressurizes the air. The pressurized air76from the HP compressor26is mixed with fuel in the combustor30and ignited, thereby generating combustion gases. Some work is extracted from these gases by the HP turbine34, which drives the HP compressor26. The combustion gases are discharged into the LP turbine36, which extracts additional work to drive the LP compressor24, and the exhaust gas is ultimately discharged from the engine10via the exhaust section38. The driving of the LP turbine36drives the LP spool50to rotate the fan20and the LP compressor24.

A portion of the pressurized airflow76can be drawn from the compressor section22as bleed air77. The bleed air77can be drawn from the pressurized airflow76and provided to engine components requiring cooling. The temperature of pressurized airflow76entering the combustor30is significantly increased. As such, cooling provided by the bleed air77is necessary for operating of such engine components in the heightened temperature environments.

A remaining portion of the airflow78bypasses the LP compressor24and engine core44and exits the engine assembly10through a stationary vane row, and more particularly an outlet guide vane assembly80, comprising a plurality of airfoil guide vanes82, at the fan exhaust side84. More specifically, a circumferential row of radially extending airfoil guide vanes82are utilized adjacent the fan section18to exert some directional control of the airflow78.

Some of the air supplied by the fan20can bypass the engine core44and be used for cooling of portions, especially hot portions, of the engine10, and/or used to cool or power other aspects of the aircraft. In the context of a turbine engine, the hot portions of the engine are normally downstream of the combustor30, especially the turbine section32, with the HP turbine34being the hottest portion as it is directly downstream of the combustion section28. Other sources of cooling fluid can be, but are not limited to, fluid discharged from the LP compressor24or the HP compressor26.

Referring now toFIG. 2, an engine component in the form of one of the turbine blades68includes an airfoil100extending from a dovetail115. The airfoil100includes an outer wall101that bounds an interior102and includes a pressure side103and a suction side104. The outer wall101extends axially between a leading edge105and a trailing edge106to define a chord-wise direction indicated by a chord line L, and also extends along a radial line R between a root107and a tip108to define a span-wise direction. It will be understood that the airfoil100can be any rotating or non-rotating airfoil within the turbine engine10, including in the compressor section22or turbine section32.

The tip108can include a tip wall109at least partially bounding the airfoil interior102. A tip rail110can project from the tip wall109in the span-wise direction as shown; the tip rail110can include an exterior surface111spaced from an interior surface112, where an outer edge113can connect the exterior and interior surfaces111,112. It is contemplated that the tip rail110can circumscribe the airfoil100, and a plurality of cooling holes120can be provided in the tip rail110.

The airfoil100can be affixed to the dovetail115at the root107by way of a platform114; in one non-limiting example, the dovetail115can be configured to mount to the turbine rotor disk71on the engine10to drive the blade68. The dovetail115is also illustrated herein as having three inlet passages116extending through the dovetail115and the platform114. The inlet passages116can fluidly couple to the airfoil interior102via passage outlets117. In addition, the airfoil100can also include at least one cooling passage118formed in the interior102; the cooling passage118can fluidly couple to the passage outlets117, one or more other cooling passages118, or other features formed within the airfoil100or dovetail115as desired.

A tip cooling cavity119can be formed within the tip rail110between the exterior and interior surfaces111,112and can also be fluidly coupled to the cooling passage118. In addition, at least one cooling hole120can be provided in the tip rail110adjacent the pressure side103or the suction side104.

Referring now toFIG. 3, the cooling hole120can include an inlet121fluidly coupled to the tip cooling cavity119and an outlet122in the exterior surface111. It is also contemplated that the outlet122can be formed on the outer edge113or interior surface112of the tip rail110, and it should be understood that the cooling hole120can also be provided in the tip rail110adjacent the suction side104as well. Furthermore, the inlet121and outlet122can have any desired shape including round, square, oval, rectangular, rounded rectangular, or an irregular shape as desired.

In one example, a plurality of cooling holes120can be formed in the tip rail110wherein a first cooling hole131is formed adjacent a second cooling hole132. The first cooling hole131has a first inlet141in the tip cooling cavity119and a first outlet151in the exterior surface111; similarly, the second cooling hole132can have a second inlet142in the tip cooling cavity119and a second outlet152in the exterior surface111.

The first and second cooling holes131,132can define respective first and second centerlines161,162as shown. It is contemplated that either or both of the first and second centerlines161,162can be non-orthogonal to the chord line L, including 50 degrees or smaller in a non-limiting example. In the example ofFIG. 3, the first cooling hole131forms a first angle171with the chord line L and the second cooling hole132forms a second angle172with the chord line L. The first angle171can differ from the second angle172, and in a non-limiting example both of the first and second angles171,172can be non-orthogonal to the chord line L, for example smaller than 50 degrees. Furthermore, the centerline angle can differ for cooling holes120proximate the leading edge105as compared to those proximate the trailing edge106; in one example, the first angle171can be greater than the second angle172, where the first cooling hole120is positioned closer to the leading edge105while the second cooling hole132is positioned closer to the trailing edge106.

It is further contemplated that a plurality of cooling holes120can each have an outlet122on the exterior surface111, where each cooling hole120includes a centerline160which is non-orthogonal to the chord line L. In one example the centerlines160can form angles with the chord line L which generally decrease toward the trailing edge106, such as that illustrated by the first and second angles171,172. In another example the centerlines160can form angles with the chord line L which continuously increase at a constant rate toward the trailing edge106. In still another example, the centerlines160can form angles with the chord line L which are non-orthogonal and alternately increase and decrease moving toward the leading edge105.

Additionally, adjacent outlets122of the cooling holes120can be positioned axially along the exterior surface111and share at least one contact point125or at least partially overlap. In this manner the outlets122can at least partially define a cooling trench180extending in the chord-wise direction in the exterior surface111. The cooling trench180can include a trench width181and trench length182as shown. It is contemplated that the trench width181can be greater than or equal to a maximum cross dimension126of the cooling hole outlet122. In non-limiting examples, the trench width181can be smaller than 0.04 cm while the maximum cross dimension126can be smaller than 0.005 cm; or the maximum cross dimension126of the outlet122can be larger than a cross dimension of the inlet121(not shown). It is also contemplated that the outlets122and cooling trench180can also be formed axially along the interior surface112or along the outer edge113as desired.

In operation, cooling air C (FIG. 2) from within the airfoil interior102can flow through the cooling passage118and tip cooling cavity119, entering through the inlet121and flowing to the external surface111via the cooling trench180. In another non-limiting example where the cooling trench180is positioned on the interior surface112, cooling air can also flow from the tip cooling cavity119to the interior surface112via the cooling holes120and cooling trench180. In another non-limiting example the cooling air can flow from the cooling passage118directly to the cooling holes120and cooling trench180.

Turning toFIG. 4, another airfoil200is illustrated which can be utilized in the turbine engine10ofFIG. 1. The airfoil200is similar to the airfoil100; therefore, like parts will be identified with like numerals increased by 100, with it being understood that the description of the like parts of the airfoil100applies to the airfoil200, unless otherwise noted.

The airfoil200can include an outer wall201bounding an interior202and having a pressure side203and suction side204. The outer wall can extend along the chord line L between a leading edge205and a trailing edge206to define a chord-wise direction. A tip208of the airfoil200can include a tip wall209; a tip rail210can project from the tip wall and include an exterior surface211spaced from an interior surface212, where an outer edge213can connect the exterior and interior surfaces211,212. The tip rail210can also include a tip cooling cavity219which can fluidly couple to other cooling passages218within the airfoil200(shown inFIG. 5).

A set of cooling holes220can be formed in the tip rail210. At least one cooling hole220can include an inlet221fluidly coupled to the tip cooling cavity219, and the outlet222can include a diffuser290fluidly coupled to the airfoil exterior. Additionally, the diffuser290can have a variable diffuser diameter291as shown. Furthermore, first and second cooling holes231,232can define respective first and second centerlines261,262; either or both of which can be non-orthogonal to the chord line L in a similar manner to that of the first and second centerlines161,162ofFIG. 3.

The example ofFIG. 4also illustrates that a cooling slot295can be provided in the exterior surface211of the tip rail210and extending in the chord-wise direction, where the outlets222of the cooling holes220can be fluidly coupled to the cooling slot295. In this manner the cooling slot295can define a cooling trench280having a trench width281and trench length282. In one example the trench width281can be larger than the diffuser diameter291.

It is also contemplated in another non-limiting example that adjacent diffusers290can abut or overlap to define the cooling trench280in a manner similar to the example ofFIG. 3. In such a case, the cooling trench280can have a trench width281equal to that of the diffuser diameter291, which can form a maximum cross dimension of the outlet222.

FIG. 5illustrates a cross-sectional view of the airfoil200ofFIG. 4, where it can be seen that the cooling slot295includes a depth296less than a thickness297of the tip rail210; in this manner, the cooling slot295can be formed with the inner surface298in the tip rail210as shown. Additionally, at least one cooling hole220can have a curvilinear profile between the inlet221and outlet222. Cooling holes220are also provided in the tip rail210adjacent the suction side204, where inlets221can be fluidly coupled to tip cooling cavities219, and outlets222can be formed on an inner surface298of the cooling slot295, in the interior surface212, or in the outer edge213. It is contemplated that any cooling hole described herein can include a centerline which is non-orthogonal to the chord line L. Furthermore, any cooling hole described herein can include a centerline which is non-aligned with the radial line R (FIG. 2).

A method of cooling the tip wall109of the airfoil100in the turbine engine10can include emitting cooling air into the cooling trench180extending in the chord-wise direction along the tip wall109, the trench180being at least partially defined by the cooling holes120within the tip wall109having abutting or overlapping outlets122. Cooling air can also be supplied via the cooling passage118(FIG. 2) fluidly coupled to the tip cooling cavity119(FIG. 3), where cooling air can flow from the tip cooling cavity119to the cooling holes120and cooling trench180(FIG. 3).

It can be appreciated that the non-orthogonal angles formed by the cooling hole centerlines with the chord line can allow for better cooling air coverage along the airfoil surface in a direction that improves cooling performance. The non-orthogonal angles can result in longer cooling holes which provide for improved bore cooling and improved diffusion of the cooling air. The cooling holes can be formed with chord-wise or radial angles, and can be linear or curvilinear, and the improved positioning can increase durability or performance of the airfoil by more efficient and directional use of the coolant. Additionally, the cooling trench formed by the cooling hole outlets can further improve cooling performance by keeping cooling air closer to the airfoil surface during operation.

It should be understood that application of the disclosed design is not limited to turbine engines with fan and booster sections, but is applicable to turbojets and turboshaft engines as well.

To the extent not already described, the different features and structures of the various embodiments can be used in combination, or in substitution with each other as desired. That one feature is not illustrated in all of the embodiments is not meant to be construed that it cannot be so illustrated, but is done for brevity of description. Thus, the various features of the different embodiments can be mixed and matched as desired to form new embodiments, whether or not the new embodiments are expressly described. All combinations or permutations of features described herein are covered by this disclosure.