Gas turbine engine component having suction side cutback opening

An airfoil for a gas turbine engine, according to an exemplary aspect of the present disclosure includes, among other things, a pressure side wall and a suction side wall spaced apart from the pressure side wall and each extending between a leading edge portion and a trailing edge portion. A plurality of cutback openings are spaced along a radial axis of the suction side wall.

BACKGROUND

This disclosure relates to a gas turbine engine, and more particularly to a gas turbine engine component that includes a suction side cutback opening.

Gas turbine engines typically include a compressor section, a combustor section and a turbine section. In general, during operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases flow through the turbine section which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads.

Due to exposure to hot combustion gases, numerous components of the gas turbine engine may include cooling circuits that circulate cooling airflow throughout various internal and external surfaces of the components during engine operation. Certain portions of components may be difficult to cool notwithstanding such internal cooling circuits. In addition, cooling airflow must typically be extracted from the gas path of upstream sections of the gas turbine engine which can result in efficiency losses.

SUMMARY

An airfoil for a gas turbine engine, according to an exemplary aspect of the present disclosure includes, among other things, a pressure side wall and a suction side wall spaced apart from the pressure side wall and each extending between a leading edge portion and a trailing edge portion. A plurality of cutback openings are spaced along a radial axis of the suction side wall.

In a further non-limiting embodiment of the foregoing airfoil, the plurality of cutback openings are positioned at the trailing edge portion of the suction side wall.

In a further non-limiting embodiment of either of the foregoing airfoils, at least a portion of the plurality of cutback openings are slots.

In a further non-limiting embodiment of any of the foregoing airfoils, the plurality of cutback openings are positioned along an entire radial span of the suction side wall.

In a further non-limiting embodiment of any of the foregoing airfoils, a rib extends between adjacent cutback openings of the plurality of cutback openings.

In a further non-limiting embodiment of any of the foregoing airfoils, the airfoil includes at least one film cooling hole at the trailing edge portion of the pressure side wall.

In a further non-limiting embodiment of any of the foregoing airfoils, the airfoil includes at least one film cooling hole in a gas path accelerating region of the suction side wall.

In a further non-limiting embodiment of any of the foregoing airfoils, a gas path decelerating region of the suction side wall does not include film cooling holes.

In a further non-limiting embodiment of any of the foregoing airfoils, cooling airflow is communicated through each of the plurality of cutback openings at a surface angle that is less than about 10 degrees relative to an exterior surface of the suction side wall.

A component for a gas turbine engine, according to an exemplary aspect of the present disclosure includes, among other things, a body portion that includes a pressure side wall and a suction side wall spaced apart from the pressure side wall. A plurality of cutback openings are spaced along a radial axis of the suction side wall.

In a further non-limiting embodiment of the foregoing component, the body portion is an airfoil of a blade.

In a further non-limiting embodiment of either of the foregoing components, the body portion is part of a blade outer air seal (BOAS).

In a further non-limiting embodiment of any of the foregoing components, the plurality of cutback openings are part of a cooling circuit disposed inside the body portion that also includes at least a first cavity and a second cavity in fluid communication with the first cavity.

In a further non-limiting embodiment of any of the foregoing components, the plurality of cutback openings are positioned at a trailing edge portion of the body portion.

In a further non-limiting embodiment of any of the foregoing components, the component comprises at least one film cooling hole at a trailing edge portion of the pressure side wall.

In a further non-limiting embodiment of any of the foregoing components, a gas path decelerating region of the suction side wall does not include film cooling holes.

A gas turbine engine, according to an exemplary aspect of the present disclosure includes, among other things, a compressor section, a combustor section in fluid communication with the compressor section, and a turbine section in fluid communication with the combustor section. A first component is disposed in at least one of the compressor section and the turbine section. The first component includes a body portion having a pressure side wall and a suction side wall spaced apart from the pressure side wall and that each extend between a leading edge portion and a trailing edge portion. A plurality of cutback openings are spaced along a radial axis of the suction side wall.

In a further non-limiting embodiment of the foregoing gas turbine engine, a second component is circumferentially adjacent to the first component. A gauge area extends between a trailing edge portion of the second component and the suction side wall of the first component.

In a further non-limiting embodiment of either of the foregoing gas turbine engines, the suction side wall includes a gas path accelerating region upstream from the gauge area and a gas path decelerating region downstream from the gauge area.

In a further non-limiting embodiment of any of the foregoing gas turbine engines, the gas path accelerating region includes at least one film cooling hole and the gas path decelerating region does not include film cooling holes.

DETAILED DESCRIPTION

FIG. 1schematically illustrates a gas turbine engine20. The exemplary gas turbine engine20is a two-spool turbofan engine that generally incorporates a fan section22, a compressor section24, a combustor section26and a turbine section28. Alternative engines might include an augmenter section (not shown) among other systems for features. The fan section22drives air along a bypass flow path B, while the compressor section24drives air along a core flow path C for compression and communication into the combustor section26. The hot combustion gases generated in the combustor section26are expanded through the turbine section28. Although depicted as a turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to turbofan engines and these teachings could extend to other types of engines, including but not limited to, three-spool engine architectures.

The gas turbine engine20generally includes a low speed spool30and a high speed spool32mounted for rotation about an engine centerline longitudinal axis A. The low speed spool30and the high speed spool32may be mounted relative to an engine static structure33via several bearing systems31. It should be understood that other bearing systems31may alternatively or additionally be provided.

The low speed spool30generally includes an inner shaft34that interconnects a fan36, a low pressure compressor38and a low pressure turbine39. The inner shaft34can be connected to the fan36through a geared architecture45to drive the fan36at a lower speed than the low speed spool30. The high speed spool32includes an outer shaft35that interconnects a high pressure compressor37and a high pressure turbine40. In this embodiment, the inner shaft34and the outer shaft35are supported at various axial locations by bearing systems31positioned within the engine static structure33.

A combustor42is arranged between the high pressure compressor37and the high pressure turbine40. A mid-turbine frame44may be arranged generally between the high pressure turbine40and the low pressure turbine39. The mid-turbine frame44can support one or more bearing systems31of the turbine section28. The mid-turbine frame44may include one or more airfoils46that extend within the core flow path C.

The inner shaft34and the outer shaft35are concentric and rotate via the bearing systems31about the engine centerline longitudinal axis A, which is co-linear with their longitudinal axes. The core airflow is compressed by the low pressure compressor38and the high pressure compressor37, is mixed with fuel and burned in the combustor42, and is then expanded over the high pressure turbine40and the low pressure turbine39. The high pressure turbine40and the low pressure turbine39rotationally drive the respective high speed spool32and the low speed spool30in response to the expansion.

The pressure ratio of the low pressure turbine39can be pressure measured prior to the inlet of the low pressure turbine39as related to the pressure at the outlet of the low pressure turbine39and prior to an exhaust nozzle of the gas turbine engine20. In one non-limiting embodiment, the bypass ratio of the gas turbine engine20is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor38, and the low pressure turbine39has a pressure ratio that is greater than about five (5:1). It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines, including direct drive turbofans.

In this embodiment of the exemplary gas turbine engine20, a significant amount of thrust is provided by the bypass flow path B due to the high bypass ratio. The fan section22of the gas turbine engine20is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. This flight condition, with the gas turbine engine20at its best fuel consumption, is also known as bucket cruise Thrust Specific Fuel Consumption (TSFC). TSFC is an industry standard parameter of fuel consumption per unit of thrust.

Fan Pressure Ratio is the pressure ratio across a blade of the fan section22without the use of a Fan Exit Guide Vane system. The low Fan Pressure Ratio according to one non-limiting embodiment of the example gas turbine engine20is less than 1.45. Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of [(Tram° R)/(518.7° R)]0.5, where T represents the ambient temperature in degrees Rankine. The Low Corrected Fan Tip Speed according to one non-limiting embodiment of the example gas turbine engine20is less than about 1150 fps (351 m/s).

Each of the compressor section24and the turbine section28may include alternating rows of rotor assemblies and vane assemblies (shown schematically) that carry airfoils that extend into the core flow path C. For example, the rotor assemblies can carry a plurality of rotating blades25, while each vane assembly can carry a plurality of vanes27that extend into the core flow path C. The blades25of the rotor assemblies create or extract energy (in the form of pressure) from the core airflow that is communicated through the gas turbine engine20along the core flow path C. The vanes27of the vane assemblies direct the core airflow to the blades25to either add or extract energy.

Various components of a gas turbine engine20, including but not limited to the airfoils of the blades25and the vanes27of the compressor section24and the turbine section28, may be subjected to repetitive thermal cycling under widely ranging temperatures and pressures. The hardware of the turbine section28is particularly subjected to relatively extreme operating conditions. Therefore, some components may require internal cooling circuits for cooling the parts during engine operation. Example cooling circuits that include features such as suction side cutback openings are discussed below.

FIGS. 2, 3 and 4illustrate a component50that can be incorporated into a gas turbine engine, such as the gas turbine engine20ofFIG. 1. The component50includes a body portion52that axially extends between a leading edge portion54and a trailing edge portion56. The body portion52also includes a pressure side wall58(seeFIG. 2) and a suction side wall60(seeFIG. 3) that are spaced apart from one another and axially extend between the leading edge portion54and the trailing edge portion56.

In this embodiment, the body portion52is representative of an airfoil. For example, the body portion52could be an airfoil that extends from a platform portion51that is connected to a root portion53, or could alternatively extend between inner and outer platforms where the component50is a vane (not shown). In yet another embodiment, the component50could be a non-airfoil component, including but not limited to, a blade outer air seal (BOAS), a combustor liner, a turbine exhaust case liner, or any other part that may require dedicated cooling.

A gas path62is communicated axially downstream through the gas turbine engine20along the core flow path C in a direction that extends from the leading edge portion54toward the trailing edge portion56of the body portion52. The gas path62represents the communication of core airflow along the core flow path C (seeFIG. 1). The body portion52may extend radially across a span S.

A cooling circuit64(best seen inFIG. 4) may be disposed inside of the body portion52for cooling the internal and external surfaces of the component50. For example, the cooling circuit64can include one or more cavities that may radially, axially and/or circumferentially extend inside of the body portion52to establish cooling passages for receiving a cooling airflow68to cool the component50. The cooling airflow68may be communicated into one or more of the cavities from an airflow source70that is external to the component50. The cooling airflow68can be communicated into an inlet at either the body portion52or the root portion53, for example.

The cooling airflow68is generally of a lower temperature than the airflow of the gas path62that is communicated across the body portion52. In one particular embodiment, the cooling airflow68is a bleed airflow that can be sourced from the compressor section24or any other portion of the gas turbine engine20that is upstream from the component50. The cooling airflow68can be circulated through the cooling circuit64to transfer thermal energy from the component50to the cooling airflow68thereby cooling the internal and external surfaces of the component50.

As best illustrated inFIG. 3, the component50can include a plurality of cutback openings80that are part of the cooling circuit64. The cutback openings80are removed sections in the suction side wall60of the component50that establish passages for expelling the cooling airflow68from the cooling circuit64. As removed sections, a chord C1of the body portion52that extends through a centerline of each cutback opening80extends a shorter distance than the chord C2of the entire axial distance of the suction side wall60. In other words, a distal most portion77of suction side wall60is offset (by an axial distance D1) from a distal most portion79of the pressure side wall58at each cutback opening80(seeFIG. 5).

The cutback openings80can be positioned to extend along the trailing edge portion56of the component50. In this manner, the cooling circuit64can adequately cool the trailing edge portion56of the suction side wall60in a manner that requires the communication of a relatively small amount of cooling airflow68to the suction side wall60. The cutback openings80could be alternatively positioned at other locations along the suction side wall60.

A plurality of cutback openings80can be spaced along a radial axis RA of the suction side wall60. The radial axis RA is generally parallel to the span S of the body portion52. A rib82can extend between adjacent cutback openings80. The chord C2extends through each rib82, in this embodiment. In one embodiment, the plurality of cutback openings80extend along the entire span S of the suction side wall60. The number of cutback openings80that are formed into the suction side wall60can vary depending upon design specific parameters including but not limited to the cooling requirements of the component50.

In one embodiment, the plurality of cutback openings80are slots. In another embodiment, the cutback openings80are thumbnail shaped. However, the plurality of cutback openings80could embody other shapes within the scope of this disclosure.

The cutback openings80may be cast features of the component50. However, other techniques can also be utilized to manufacture the cutback openings80into the component50.

FIG. 4illustrates additional features of the cooling circuit64that can be disposed inside of the component50. The pressure side wall58and the suction side wall60include interior surfaces55as well as exterior surfaces57(i.e., gas path surfaces). The interior surfaces55are remote from the gas path62and establish portions of the cooling circuit64, whereas the exterior surfaces57are positioned within the gas path62. The cooling circuit64can communicate cooling airflow68to cool both the interior surfaces55and the exterior surfaces57.

In this embodiment, the exemplary cooling circuit64includes a first cavity72A (i.e., a leading edge cavity), a second cavity72B (i.e., a first intermediate cavity), a third cavity72C (i.e., a second intermediate cavity), a fourth cavity72D (i.e., a third intermediate cavity), and a plurality of trailing edge cavities72E,72F,72G and72H. However, the cooling circuit64could alternatively include a greater or fewer number of cavities. The cavities72A,72B,72C,72D,72E,72F,72G and72H can communicate the cooling airflow68through the cooling circuit64, including along a serpentine path, to cool the body portion52. In other words, the cavities72A through72H may be in fluid communication with one another in order to circulate the cooling airflow68throughout the cooling circuit64.

Ribs74may extend between the pressure side wall58and the suction side wall60of the body portion52. In this particular embodiment, a first rib74A is positioned between the first cavity72A and the second cavity72B, a second rib74B is positioned between the second cavity72B and the third cavity72C, a third rib74C is positioned between the third cavity72C and the fourth cavity72D and a fourth rib74D is positioned between the fourth cavity72D and the fifth cavity72E. The trailing edge cavities72E,72F,72G and72H may include one or more pedestals73.

Cooling airflow68from the cooling circuit64can be communicated through the plurality of cutback openings80and returned to the gas path62. In one embodiment, the cooling airflow68is injected through the plurality of cutback openings80at a low surface angle. For example, as shown inFIG. 5, the cooling airflow68can be injected through the cutback openings80at a surface angle α that is less than about 10° relative to the exterior surface57of the suction side wall60. In this manner, the cooling airflow68can be introduced in both accelerating and decelerating regions of the component50with relatively minimal mixing losses.

The component50can also include a plurality of film cooling holes90. The film cooling holes90can be disposed on the pressure side wall58and on portions of the suction side wall60. For example, the leading edge portion54of the suction side wall60can include one or more film cooling holes90. In this embodiment, the trailing edge portion56of the suction side wall60is free of film cooling holes.

FIG. 6illustrates portions of an assembly100, such as a rotor assembly, that can be incorporated into either the compressor section24or turbine section28of the gas turbine engine20. This particular assembly100incorporates at least a first component50A and a second component50B that each include cutback openings80. The first and second components50A and50B are circumferentially adjacent to one another and are positioned about the assembly100. Although two components50A,50B are illustrated, the assembly100could include any number of components.

A gauge area92(i.e., a throat area between the first component50A and the second component50B) extends between the trailing edge portion56of the second component50B and the suction side wall60of the first component50A. The gauge area92divides the suction side wall60into a gas path accelerating region94and a gas path decelerating region96. In this example, the gas path accelerating region94of the suction side wall60of the first component50A is upstream of the gauge area92and the gas path decelerating region96is downstream from the gauge area92. Since the entire chord of the pressure side wall58is upstream from the gauge area92, the pressure side wall58includes a gas path accelerating region98only.

In one embodiment, the gas path accelerating region94of the suction side wall60of the first component50A includes one or more film cooling holes90. However, no film cooling holes90are located within the gas path decelerating region96of the suction side wall60. Therefore, the first component50A does not include film cooling holes90in the trailing edge portion56of the suction side wall60. The cutback openings80provide the cooling in this region of the component50A. Film cooling holes90can be located along any portion (including the trailing edge portion56) of the pressure side wall58since it includes a gas path accelerating region98and no decelerating region.

The cutback openings80described in this disclosure may provide more efficient cooling in low loss regions of a component50, including at the trailing edge, suction side wall. Moreover, the cooling airflow68can be injected into gas path accelerating and decelerating regions of the suction side wall60with minimal loss by injecting the airflow at relatively low surface angles. Other features such as film cooling holes90can also be incorporated to effectively cool the trailing edge, pressure side of the component.

Although the different non-limiting embodiments are illustrated as having specific components, the embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from any of the non-limiting embodiments in combination with features or components from any of the other non-limiting embodiments.

It should be understood that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be understood that although a particular component arrangement is disclosed and illustrated in these exemplary embodiments, other arrangements could also benefit from the teachings of this disclosure.