Method for cooling turbine blades

Abstract

A gas turbine engine defines an axial direction and a radial direction and includes: a main cooling circuit, the main cooling circuit being configured to receive a first air flow; and a turbine rotor, the turbine rotor including rotor blades, the rotor blades at least partially defining a front impeller space located in front of the rotor blades, the front impeller space being configured to receive a second air flow, the rotor blades further defining: a first cooling circuit, the first cooling circuit being inside the rotor blades and in fluid communication with the main cooling circuit, for receiving the first air flow from the main cooling circuit; a second cooling circuit, the second cooling circuit being inside the rotor blades and in fluid communication with the front impeller space, for receiving a portion of the second air flow from the front impeller space; and a device for drawing a portion of the second air flow into the second cooling circuit.

Description

Technical Field

[0001] The present disclosure relates to a gas turbine engine and, in particular, to a method for cooling blades in a gas turbine engine. Background Art

[0002] A gas turbine engine generally comprises a turbine and a rotor assembly. Gas turbine engines, such as turbofan engines, can be used for aircraft propulsion. In the case of a turbofan engine, the rotor assembly can be configured as a fan assembly.

[0003] A turbomachine may include a turbine having a plurality of turbine rotor blades. The turbine rotor blades may be exposed to relatively high temperatures during operation. To maintain the turbine rotor blades within an operating temperature range, it may be necessary to provide a cooling air flow to one or more internal cavities defined within the turbine rotor blades. Improvements in this process would be welcome in the art. BRIEF DESCRIPTION OF THE DRAWINGS

[0004] A full and effective disclosure of the present disclosure, including the best mode thereof, to one of ordinary skill in the art is set forth in the specification with reference to the accompanying drawings, in which:

[0005] Figure 1 is a cross-sectional view of a gas turbine engine according to an exemplary aspect of the present disclosure.

[0006] Figure 2 is a cross-sectional view of a portion of a gas turbine engine according to an exemplary aspect of the present disclosure and illustrates a cooling assembly for a turbine blade.

[0007] Figure 3 is a simplified isolated view of a portion of a turbine blade according to an exemplary aspect of the present disclosure and showing cooling passages through the blade.

[0008] Figure 4 is an enlarged isolated view of a root of a turbine blade according to an exemplary aspect of the present disclosure and illustrates the alignment of cooling passages through the root.

[0009] Figure 5 is an isolated view of a drain hole with a curved scoop according to an exemplary aspect of the present disclosure.

[0010] Figure 6 is an isolated view of a drain hole having a recessed opening according to an exemplary aspect of the present disclosure.

[0011] Figure 7 is a flow chart of a method of cooling blades of an engine according to an exemplary aspect of the present disclosure. DETAILED DESCRIPTION

[0012] Reference will now be made in detail to the present embodiments of the present disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. Like or similar designations in the drawings and the description have been used to refer to like or similar parts of the disclosure.

[0013] The word "exemplary" is used herein to mean "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other implementations. Furthermore, unless expressly stated otherwise, all embodiments described herein are to be considered exemplary.

[0014] As used herein, the terms “first,” “second,” and “third” may be used interchangeably to distinguish one component from another and are not intended to indicate the position or importance of each component.

[0015] The terms "fore" and "aft" refer to relative positions within a gas turbine engine or vehicle and refer to the normal operating attitude of the gas turbine engine or vehicle. For example, for a gas turbine engine, the front position refers to the position closer to the engine inlet, and the aft position refers to the position closer to the engine nozzle or exhaust.

[0016] The terms "upstream" and "downstream" refer to relative directions relative to the flow of a fluid in a fluid path. For example, "upstream" refers to the direction from which the fluid is flowing, and "downstream" refers to the direction to which the fluid is flowing.

[0017] The terms “coupled,” “fixed,” “attached,” and the like refer to both direct coupling, fixing, or attachment, as well as indirect coupling, fixing, or attachment through one or more intermediate components or features, unless otherwise indicated herein.

[0018] The singular forms "a," "an," and "the" include plural references unless the context clearly dictates otherwise.

[0019] Approximate language used throughout the specification and claims is applied to modify any quantitative expression that can be allowed to vary without causing a change in its related basic function. Therefore, the values ​​modified by one or more terms such as "approximately", "approximately" and "substantially" are not limited to the specified precise values. In at least some cases, approximate language can correspond to the accuracy of the instrument used to measure the value, or the accuracy of the method or machine used to construct or manufacture parts and / or systems. For example, approximate language can refer to within 1%, 2%, 4%, 10%, 15% or 20% margin. These approximate margins can be applied to a single value, any endpoint or two endpoints of a defined numerical range, and / or the margin of the range between the endpoints.

[0020] Here and throughout the specification and claims, range limitations are combined and interchanged, and unless context or language indicates otherwise, such ranges are identified and include all sub-ranges contained therein. For example, all ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other.

[0021] The terms "low" and "high," or their respective comparative degrees (e.g., higher or lower, as applicable), when used with reference to compressor, turbine, shaft or spool components, etc., each refer to relative speeds within the engine, unless otherwise specified. For example, a "low turbine" or "low-speed turbine" defines a component that is configured to operate at a lower speed (e.g., maximum allowable speed) than a "high turbine" or "high-speed turbine" at the engine.

[0022] The terms “turbomachine” or “turbomachinery” refer to a machine that includes one or more compressors, a heat-generating section (eg, a combustion section), and one or more turbines that together produce a torque output.

[0023] The term "gas turbine engine" refers to an engine having a turbine as all or part of its power source. Example gas turbine engines include turbofan engines, turboprop engines, turbojet engines, turboshaft engines, and the like.

[0024] The term "combustion section" refers to any heat addition system for a turbomachine. For example, the term combustion section may refer to a section that includes one or more of a deflagration combustion assembly, a rotating detonation combustion assembly, a pulse detonation combustion assembly, or other suitable heat addition assemblies. In certain exemplary embodiments, the combustion section may include an annular combustor, a can combustor, a tubular combustor, a trapped vortex combustor (TVC), or other suitable combustion systems, or combinations thereof.

[0025] The present disclosure relates to cooling assemblies for rotor blades used in gas turbine engines. In certain engine designs, the cooling air flow supplied to the turbine blades can be derived from a single, dedicated delivery circuit. This configuration can eliminate the need for cooling air flows from other sources, which are ultimately dumped as purge flow or otherwise do not perform useful power work or thermal management in the engine.

[0026] Aspects of the present disclosure propose extracting air from the forward wheelspace and mixing it with fresh cooling air delivered to, for example, a trailing edge circuit. For example, forward wheelspace cavity air is drawn into the shank of a turbine rotor blade and delivered to the turbine blade's trailing edge cooling circuit, where it mixes with fresh cooling air. This mixed air is then delivered to the trailing edge cooling circuit. By using excess air from the forward wheelspace cavity to cool the trailing edge, the amount of purge air required for cooling may be reduced, potentially improving the engine's specific fuel consumption.

[0027] Referring now to the drawings, in which like numerals refer to like elements throughout, Figure 1 is a schematic cross-sectional view of a propulsion system 10 according to an exemplary embodiment of the present disclosure. More specifically, Figure 1 In an embodiment of the present invention, the propulsion system 10 includes a gas turbine engine, referred to herein as a "turbofan engine 12." In one example, the turbofan engine 12 may be a high-bypass turbofan jet engine. Figure 1 As shown, the turbofan engine 12 defines an axial direction A (extending parallel to a longitudinal centerline 14 for reference) and a radial direction R. Generally, the turbofan engine 12 includes a fan section 16 and a turbine 18 disposed downstream of the fan section 16 .

[0028] The depicted exemplary turbomachine 18 generally includes a substantially tubular outer casing 20 defining an annular inlet 22. The outer casing 20 encloses, in a series flow sequence / relationship, a compressor section including a supercharger or low-pressure compressor 24 ("LP compressor 24") and a high-pressure compressor 26 ("HP compressor 26"); a combustion section 28; a turbine section including a high-pressure turbine 30 ("HP turbine 30") and a low-pressure turbine 32 ("LP turbine 32"); and the combustion section 28. A high-pressure shaft or spool 34 ("HP spool 34") drivingly connects the HP turbine 30 to the HP compressor 26. A low-pressure shaft or spool 36 ("LP spool 36") drivingly connects the LP turbine 32 to the low-pressure compressor 24.

[0029] For the depicted embodiment, fan section 16 includes a variable pitch fan 38 having a plurality of fan blades 40 coupled to a disk 42 in a spaced-apart manner. As depicted, fan blades 40 generally extend outwardly from disk 42 in a radial direction R. Each fan blade 40 can be rotated relative to disk 42 about a pitch axis P by virtue of fan blades 40 being operably coupled to a suitable actuation member 44 that is configured to collectively (e.g., in unison) vary the pitch of fan blades 40. Fan blades 40, disk 42, and actuation members 44 can be rotated together about longitudinal centerline 14 via LP spool 36 across a power gearbox 46. Power gearbox 46 includes a plurality of gears for reducing the rotational speed of LP spool 36 to a more efficient fan speed.

[0030] Still refer to Figure 1In an exemplary embodiment of the present invention, the disk 42 is covered by a rotatable front hub 48 having an aerodynamic profile to facilitate air flow through the plurality of fan blades 40. In addition, the fan section 16 includes an annular fan casing or outer nacelle 50 that circumferentially surrounds the variable pitch fan 38 and / or at least a portion of the turbine 18. It should be understood that in some embodiments, the nacelle 50 is configured to be supported relative to the turbine 18 by a plurality of circumferentially spaced outlet guide vanes 52. Furthermore, a downstream section 54 of the nacelle 50 extends over an outer portion of the turbine 18 to define a bypass airflow passage 56 therebetween.

[0031] During operation of turbofan engine 12, a volume of air 58 enters turbofan engine 12 through nacelle 50 and / or associated inlet 60 of fan section 16. As volume of air 58 passes through fan blades 40, a first portion of air 58, indicated by arrow 62, is directed or channeled into bypass airflow passage 56, and a second portion of air 58, indicated by arrow 64, is directed or channeled into LP compressor 24. The ratio between first portion 62 of air and second portion 64 of air is generally referred to as the bypass ratio. The pressure of second portion 64 of air is then increased as it is directed through high-pressure (HP) compressor 26 and into combustion section 28, where it is mixed with fuel and combusted to provide combustion gases 66. Combustion gases 66 are then directed through HP turbine 30 and LP turbine 32, where a portion of the thermal and / or kinetic energy from combustion gases 66 is extracted.

[0032] The combustion gases 66 are then directed through the combustion section 28 of the turbine 18 to provide propulsive thrust. Simultaneously, the pressure of the first portion of air 62 is significantly increased as it is directed through the bypass airflow passage 56 before being discharged from the fan nozzle exhaust section 68 of the turbofan engine 12, also providing propulsive thrust.

[0033] However, it should be understood that Figure 1 The turbofan engine 12 depicted in FIG. 1 is merely an example, and in other exemplary embodiments, aspects of the present disclosure may additionally or alternatively be applied to any other suitable gas turbine engine. For example, in other exemplary embodiments, the turbofan engine 12 may alternatively be any other suitable aviation gas turbine engine, such as a turbojet engine, a turboshaft engine, a turboprop engine, etc. Additionally, in other exemplary embodiments, the turbofan engine 12 may include any other suitable number and / or configuration of shafts, spools, compressors, turbines, etc.; may be configured as a direct drive engine (e.g., without including a power gearbox 46); may be a fixed pitch fan; etc.

[0034] Now refer to Figure 2 , Figure 2 is a cross-sectional view of a portion of propulsion system 10 and illustrates HP turbine 30 and cooling components.

[0035] The HP turbine 30 includes a Figure 2 Inlet guide vanes 70 are shown at the left side of the figure. In the exemplary embodiment, inlet guide vanes 70 are disposed downstream of combustion section 28 along the axial direction A of propulsion system 10.

[0036] The propulsion system 10 also includes an inducer 72. In an exemplary embodiment, the turbofan engine 12 defines an inducer 72. The inducer 72 is a flow channel that is configured to direct the diversion of the airflow therethrough in a circumferential direction (i.e., a direction extending around the axial direction A). In certain exemplary embodiments, the inducer 72 may include an array of circumferentially spaced blades. The inducer 72 is connected to the primary cooling circuit (e.g., the inducer 72 and the turbine cooling cavity 86) and is disposed upstream of the primary cooling circuit. The inducer 72 is in fluid communication with the primary cooling circuit (e.g., the inducer 72 and the turbine cooling cavity 86).

[0037] The propulsion system 10 further defines a forward wheel space cavity 84. However, it should be understood that in other exemplary embodiments, the forward wheel space cavity may be referred to as a forward wheel space. The forward wheel space cavity 84 is located or disposed forward of the rotor blades 92 in the axial direction A (e.g., at Figure 2 ) and is configured to receive a second air flow (eg, compressed air flow 104 ).

[0038] The propulsion system 10 additionally defines a turbine cooling cavity 86. The turbine cooling cavity 86 is a space or chamber. The turbine cooling cavity 86 is in fluid communication with the inducer 72. In an exemplary embodiment, the inducer 72 and the turbine cooling cavity 86 can define (and be referred to as) a primary cooling circuit. In certain exemplary embodiments, the turbofan engine 12 includes a primary cooling circuit (e.g., the inducer 72 and the turbine cooling cavity 86) and a turbine rotor 88. The primary cooling circuit (e.g., the inducer 72 and the turbine cooling cavity 86) is configured to receive a first air flow (e.g., the cooled cooling air flow 106). Additionally, the primary cooling circuit (e.g., the inducer 72 and the turbine cooling cavity 86) is configured to rotate about the longitudinal centerline 14 of the turbofan engine 12.

[0039] Propulsion system 10 further includes a turbine rotor 88. As will be appreciated, turbine rotor 88 at least partially defines forward wheelspace cavity 84. Turbine rotor 88 is a rotating assembly configured to generate torque for propulsion system 10 in response to airflow through turbine rotor 88. Turbine rotor 88 includes a disk 90. ​​Disk 90 is a hub configured to rotate about longitudinal centerline 14. Turbine rotor 88 also includes rotor blades 92. Rotor blades 92 are airfoils attached to disk 90. ​​Rotor blades 92 are configured to rotate with disk 90 about longitudinal centerline 14.

[0040] The rotor blade 92 includes a shank 94 disposed at an inward portion of the rotor blade 92 in a radial direction, such as in the illustrated embodiment in the radial direction R. The shank 94 is defined by the inward portion of the rotor blade 92 in the radial direction R.

[0041] The rotor blade 92 also includes an airfoil portion 95 connected to the shank 94 and disposed radially outward from the shank 94. In certain exemplary embodiments, the shank 94 may be coupled to or integrally formed with the airfoil portion 95. The airfoil portion 95 of the rotor blade 92 defines a leading edge 96 and a trailing edge 98. The leading edge 96 defines the upstream edge (upstream direction, e.g., the upstream direction) of the rotor blade 92. Figure 2 The leading edge 96 is disposed outwardly from the shank 94 in the radial direction R. The trailing edge 98 defines the downstream edge (downstream direction, as shown) of the rotor blade 92. Figure 2 A trailing edge 98 is provided on an end of the rotor blade 92 opposite the leading edge 96 in the axial direction A.

[0042] The rotor blade 92 further defines an internal passage 100 through a portion of the shank 94. The internal passage 100 is configured to convey a fluid flow in the axial direction A through a portion of the shank 94.

[0043] Turbine rotor 88 defines a gap 102 extending between shanks 94 and outer portions of rotor blades 92 in axial direction A of disk 90 . Gap 102 is fluidly connected to turbine cooling cavity 86 and the cooling circuit of rotor blades 92 .

[0044] Figure 2Also shown are a series of air flows through various portions of the propulsion system 10. In certain exemplary embodiments, the compressed air flow 104 may include, for example, a compressed air flow from a compressor discharge cavity (the cavity between the compressor sections (e.g., the LP compressor 24 and the HP compressor 26) and the combustors of the combustion section 28). During operation, the compressed air flow 104 passes through portions of the propulsion system 10 and portions of the HP turbine 30. For example, the compressed air flow 104 passes through the leading wheel space cavity 84 to the rotor blades 92. As the compressed air flow 104 approaches the rotor blades 92, a portion of the compressed air flow 104 continues in the radial direction R until it reaches the leading edge 96. This portion of the compressed air flow 104 can prevent or minimize the flow of combustion gases 66 into the working gas flow path of the propulsion system 10 between the turbine rotor 88 and the inlet guide vanes 70. Another portion of the internal compressed air flow 104 flows in the axial direction A and enters the internal passage 100 of the shank 94 to similarly purge the aft cavity.

[0045] The series of air flows also includes a cooled cooling air flow 106. In the exemplary embodiment, the cooled cooling air flow 106 is an air flow from a heat exchanger of the propulsion system 10. More specifically, in at least some exemplary aspects, the heat exchanger of the propulsion system 10 may include a cooled cooling air heat exchanger 107. In the exemplary embodiment, the turbofan engine 12 further includes a cooled cooling air heat exchanger 107 as a source of a first air flow (e.g., cooled cooling air flow 106). During operation of the propulsion system 10, the cooled cooling air flow 106 from the cooled cooling air heat exchanger 107 enters and passes through the inducer 72 to the turbine cooling cavity 86. The cooled cooling air flow 106 flows from the turbine cooling cavity 86 toward and into the gap 102. From the gap 102, the cooled cooling air flow 106 is separated into a tip flow 108, a leading edge flow 110 ("LE flow 110"), and a trailing edge flow 112 ("TE flow 112").

[0046] Tip flow 108 is a portion of the cooled cooling air flow 106 that is diverted through the inner passages (shown for clarity from the inside) inside the rotor blade 92. Figure 2 ) and is delivered to the tip 97 of the rotor blade 92. During operation of the propulsion system 10, the LE flow 110 is directed to the leading edge 96 so as to cool the leading edge 96.

[0047] As the TE flow 112 flows outwardly in the radial direction R over the shank 94, the compressed air flow 104 from the internal passage 100 combines with the TE flow 112 to form a mixed flow 116. In this way, air from the forward wheelspace cavity 84 (e.g., the compressed air flow 104) serves as a coolant for the trailing edge 98. For example, the compressed air flow 104 may be at a suitable temperature for cooling the rotor blades 92, and using the compressed air flow 104 to cool the trailing edge 98 reduces excess purge flow (e.g., the purge air 114).

[0048] It is worth noting that while internal passage 100 is shown as providing a portion of compressed air flow 104 to combine with TE flow 112 to provide cooling to trailing edge 98 , in alternative embodiments, shank 94 may define a plurality of internal passages fluidly coupled to forward wheelspace cavity 84 , with some of these internal passages 100 providing only a discharge air flow cavity behind turbine rotor 88 , e.g., for purge, while others of these internal passages 100 provide only a compressed air flow 104 to an internal cooling circuit of turbine rotor 88 .

[0049] Here, the benefit of the propulsion system 10 using the compressed air flow 104 (from the forward wheelspace cavity 84) mixed with the TE flow 112 to cool the trailing edge 98 is to offset or partially offset the use of the TE flow 112. This offset or reduction in the use of the TE flow 112 results in a reduction in the use of the cooling air flow 106 and improves the specific fuel consumption of the propulsion system 10 because the net amount of air drawn from the compressor section for cooling the rotor blades 92 has been reduced.

[0050] Now refer to Figure 3 , Figure 3 is a simplified isolated view of rotor blade 92 according to an exemplary aspect of the present disclosure. Figure 3 The examples provided in Figure 2 The configuration is similar to the one described above. Figure 3 Including the above Figure 2 The same or similar parts as described herein.

[0051] In addition to the above Figure 2 In addition to the elements discussed above, rotor blade 92 additionally includes a tang 118. Tang 118 is a curved protrusion of solid material that extends in a direction transverse to axial direction A and perpendicular to both axial direction A and radial direction R. Tang 118 projects outward from the center of rotor blade 92. Shank 94 defines tang 118. In this particular exemplary embodiment, one tang 118 is shown from a perspective cut through a cross-section of rotor blade 92. However, as will be appreciated, in other exemplary embodiments, rotor blade 92 may include two or more tangs 118 along various portions of shank 94.

[0052] In certain exemplary embodiments, tangs 118 may be configured to engage correspondingly shaped slots of disk 90 to mechanically attach rotor blade 92 to disk 90. ​​In certain exemplary embodiments, tangs 118 define inlets (e.g., openings 128A and 128B) for a second cooling circuit (e.g., internal passage 100), which are disposed at an end face 119 of shank 94.

[0053] Also in Figure 3 , rotor blade 92 is shown as including a cooling circuit having a tip channel 120, a leading edge channel 122 ("LE channel 122"), and a trailing edge channel 124 ("TE channel 124"). In certain exemplary embodiments, tip channel 120, LE channel 122, and TE channel 124 may be collectively referred to as a first cooling circuit. Tip channel 120, LE channel 122, and TE channel 124 are each internal cooling circuits configured to individually convey one or more fluids therethrough. In the exemplary embodiment, tip channel 120, LE channel 122, and TE channel 124 are each in contact with gap 102 (see, e.g., FIG. 1 ). Figure 2 In this manner, each of the tip channel 120 , the LE channel 122 , and the TE channel 124 is in fluid communication with the turbine cooling cavity 86 to receive a portion of the cooled cooling air flow 106 .

[0054] During operation of the propulsion system 10, the tip channel 120 provides a tip flow 108 to the tip 97 of the rotor blade 92, the LE channel 122 provides an LE flow 110 to the leading edge 96, and the TE channel 124 provides a TE flow 112 to the trailing edge 98. In certain exemplary embodiments, the turbine rotor 88 includes the rotor blade 92 and at least partially defines the forward wheel space cavity 84, a first cooling circuit (e.g., the tip channel 120, the LE channel 122, and the TE channel 124), and a second cooling circuit (e.g., the internal passage 100). The first cooling circuit (e.g., the tip channel 120, the LE channel 122, and the TE channel 124) is internal to the rotor blade 92 and is configured to be in fluid communication with the primary cooling circuit (e.g., the inducer 72 and the turbine cooling cavity 86) for receiving a first air flow (e.g., the cooled cooling air flow 106) from the primary cooling circuit (e.g., the inducer 72 and the turbine cooling cavity 86) when installed in the turbofan engine 12 and during operation of the turbofan engine 12. A second cooling circuit (e.g., internal passage 100 ) is internal to rotor blade 92 and is configured to be in fluid communication with forward wheelspace cavity 84 for receiving a portion of a second air flow (e.g., compressed air flow 104 ) from forward wheelspace cavity 84 when installed in turbofan engine 12 and during operation of turbofan engine 12 .

[0055] In certain exemplary embodiments, rotor blade 92 further defines means for drawing a portion of the second air flow (e.g., compressed air flow 104) into the second cooling circuit (e.g., internal passage 100). More specifically, in at least certain exemplary aspects, the means for drawing a portion of the second air flow (e.g., compressed air flow 104) into the second cooling circuit (e.g., internal passage 100) includes a nozzle 127 defined by a jet pump 126 (e.g., an ejector).

[0056] In the exemplary embodiment depicted, propulsion system 10 further includes a means for drawing a portion of the second air flow into the second cooling circuit, or more specifically, for the depicted embodiment, propulsion system 10 further includes a means for drawing a portion of the exhaust air flow into internal passage 100. More specifically, for the depicted embodiment, the means for drawing a portion of the exhaust air flow into internal passage 100 includes a jet pump 126. Jet pump 126 is a pump with no moving parts that is configured to deliver a first, high-pressure fluid to draw a second, low-pressure fluid into the combined flow of the first and second fluids. In an exemplary embodiment, jet pump 126 may include an ejector. In such an exemplary embodiment, the means for drawing a portion of the exhaust air flow into internal passage 100 includes an ejector in fluid communication with trailing edge cooling circuit 125 and the second cooling circuit (e.g., internal passage 100). However, as will be appreciated, in other exemplary embodiments, jet pump 126 may include an ejector, syringe, or other type of jet pump. Jet pump 126 is in fluid communication with internal passage 100 and TE channel 124.

[0057] Rotor blade 92 includes a trailing edge 98 and further defines a trailing edge cooling circuit 125 configured to deliver air to trailing edge 98 of rotor blade 92. In certain exemplary embodiments, a first cooling circuit (e.g., tip channels 120, LE channels 122, and TE channels 124) includes a trailing edge cooling circuit 125 configured to deliver air to trailing edge 98 of rotor blade 92.

[0058] Returning to the jet pump 126 , in at least some exemplary aspects, the device for drawing a portion of the second air flow (e.g., the compressed air flow 104 ) into the second cooling circuit (e.g., the internal passage 100 ) includes a jet pump 126 (e.g., an ejector) that is in fluid communication with the trailing edge cooling circuit 125 and the second cooling circuit (e.g., the internal passage 100 ).

[0059] For example, as the TE flow 112 passes through the converging portion or nozzle 127 of the jet pump 126, the TE flow 112 is accelerated into the trailing edge cooling circuit 125 (e.g., a portion of the TE channel 124 downstream of the jet pump 126, wherein the downstream direction is oriented as shown in FIG. Figure 3 125 ). As a result, the static pressure at the nozzle 127 of the jet pump 126 decreases in response to the TE flow 112 accelerating into the trailing edge cooling circuit 125. In response to the decreased static pressure at the nozzle 127 of the jet pump 126, the compressed air flow 104 is drawn into the trailing edge circuit (e.g., into the TE channel 124) at the jet pump 126. In this manner, as the compressed air flow 104 is drawn into the trailing edge cooling circuit, it mixes with the TE flow 112 to form a mixed flow 116. In the exemplary embodiment depicted, the means for drawing a portion of the exhaust air flow into the internal passage 100 includes a nozzle 127 defined by the jet pump 126 (e.g., an ejector).

[0060] like Figure 3 As shown, the jet pump 126 can be integrated into the shank 94 of the rotor blade 92 so that the TE flow 112 (e.g., fresh coolant) exits the nozzle 127 of the jet pump 126 and transfers momentum to the compressed air flow 104 (e.g., incoming front cavity air), thereby drawing more of the compressed air flow 104 into the TE channel 124.

[0061] Now refer to Figure 4 , Figure 4 is an enlarged isolated view of the handle 94 according to an exemplary aspect of the present disclosure.

[0062] Shank 94 defines internal passage 100A and internal passage 100B. In certain exemplary embodiments, shank 94 defines tangs 118A and 118B that project outward from the center of rotor blade 92. Tangs 118A and 118B define an inlet (e.g., openings 128A and 128B) for a second cooling circuit (e.g., internal passage 100), which is disposed at an end face 119 of shank 94.

[0063] In certain exemplary embodiments, the internal channel 100A can be disposed in a portion of the tang 118A and the internal channel 100B can be disposed in a portion of the other tang 118B. In such an example, since the internal channels 100A and 100B are positioned in the tangs 118A and 118B, respectively, and toward the edge of the shank 94, the internal channels 100A and 100B are aligned with the tip channel 120 and the LE channel 122 (removed for clarity). Figure 4In this way, the compressed air streams 104A and 104B can bypass and avoid the tip channel 120 and the LE channel 122 because the inner channel 100A and the inner channel 100B are spaced apart at the jet pump 126 (see, for example, Figure 3 ) delivers the compressed air streams 104A and 104B to the TE channel 124. For example, the internal passage 100A and the internal passage 100B may converge at the TE channel 124 at the jet pump 126, such that the internal passage 100A and the internal passage 100B deliver the compressed air streams 104A and 104B to the TE channel 124 at a single location (e.g., at the jet pump 126).

[0064] Opening 128A and opening 128B may be further defined by shank 94 of rotor blade 92. Opening 128A and opening 128B are holes or inlets 94 disposed in and defined by shank 94 of rotor blade 92 along end surface 119 of the shank. Opening 128A is in fluid communication with interior passage 100A and opening 128B is in fluid communication with interior passage 100B.

[0065] In this manner, the compressed air flows 104A and 104B are prevented from communicating or mixing with the tip flow 108 and the LE flow 110 , thereby preserving the thermal and pressure characteristics of the respective air flows.

[0066] However, as will be appreciated, in other embodiments, propulsion system 10 may further include any additional or alternative means for drawing a portion of the second air flow into the second cooling circuit, or more specifically, into internal passage 100. For example, in certain exemplary embodiments, the means for drawing a portion of the second air flow into internal passage 100 may include an inlet airflow feature. An inlet airflow feature may be any feature at the inlet into the internal passage that is configured to increase the amount of airflow entering internal passage 100.

[0067] For example, now refer to Figure 5 , Figure 5 is an isolated view of the internal passage 100 and illustrates a first exemplary embodiment of an inlet airflow feature. For the illustrated embodiment, the inlet airflow feature is a curved scoop at the opening 128, in accordance with exemplary aspects of the present disclosure.

[0068] like Figure 5 As depicted in FIG, the direction of rotation 130 represents the direction of rotation of the rotor blades 92 (via rotation of the disk 90 about the longitudinal centerline 14, see e.g. Figure 1-4During operation of the propulsion system 10, as the rotor blades 92 rotate at a certain speed, the compressed air flow 104 (e.g., within the forward wheel space cavity 84, see e.g., Figure 2 ) moves relative to rotor blade 92 through opening 128 of internal passage 100 .

[0069] In certain exemplary embodiments, the inlet airflow feature is more specifically a scoop 132 located at the opening 128 of the internal passage 100. In other words, the rotor blade 92 includes a scoop 132 located at the inlet (e.g., the opening 128) of the second cooling circuit (e.g., the internal passage 100). More specifically, in at least certain exemplary aspects, the scoop 132 may include a shrouded scoop or a curved surface such as a spherical wedge (e.g., a spherical wedge having a dihedral angle equal to or less than π / 2 radians, such as equal to or less than π / 4 radians, such as equal to or less than π / 8 radians, such as greater than 0 radians) or an ellipsoidal wedge. In at least certain exemplary aspects, the scoop 132 protrudes outwardly and away from the body of the shank 94 of the rotor blade 92 (see, e.g., FIG. 2 ). Figure 2-4 ).

[0070] The scoop 132 is shaped to protrude from the handle 94 and into a portion of the forward wheelspace cavity 84. Because the scoop 132 protrudes into a portion of the forward wheelspace cavity 84, the compressed air flow 104 passes over the rotor blades 92 at a certain speed as the rotor blades 92 rotate relative to the forward wheelspace cavity 84. As the compressed air flow 104 contacts the scoop 132, the scoop 132 captures some of the dynamic head of the compressed air flow 104 and redirects a portion of the compressed air flow 104 into the opening 128. As the portion of the compressed air flow 104 is redirected by the scoop 132 into the opening 128, a portion of the compressed air flow 104 is drawn by the scoop 132 and through the internal passage 100.

[0071] In this manner, the effective pressure of the compressed air flow 104 delivered to the trailing edge cooling circuit 125 is increased, resulting in a greater amount of compressed air flow 104 being delivered to the TE channels 124. Since more of the compressed air flow 104 is used to cool the trailing edge 98, a smaller amount of TE flow 112 is required, and thus a smaller amount of cooled cooling air flow 106 is required to cool the trailing edge 98. As a result, the overall cost of cooling the rotor blades 92 is reduced, thereby improving the specific fuel consumption of the propulsion system 10.

[0072] Now refer to Figure 6 , Figure 6 is an isolated view of the internal passage 100 and illustrates a second exemplary embodiment of an inlet airflow feature. For the embodiment shown, the inlet airflow feature is a recessed opening at the opening 128 of the internal passage 100.

[0073] like Figure 6 As shown, opening 128 defines a scoop 132'. In certain exemplary embodiments, scoop 132' defines a recessed opening (e.g., opening 128) and comprises a recessed scoop or scoop recessed into a portion of handle 94. Scoop 132' incorporates a low-resistance air inlet design that is configured to draw in a portion 104' of compressed air flow 104 with minimal disturbance of compressed air flow 104 flowing through opening 128.

[0074] Scoop 132' is configured to draw a portion of a second air flow (e.g., compressed air flow 104) into a second cooling circuit (e.g., internal passage 100) using a recessed opening (e.g., opening 128) of scoop 132'. For example, during operation, as rotor blade 92 moves in rotational direction 130, compressed air flow 104 moves through opening 128. As compressed air flow 104 passes through opening 128, portion 104' of compressed air flow 104 is drawn into internal passage 100 via scoop 132' at opening 128. As a result, portion 104' of compressed air flow 104 entering internal passage 100 has the effect of increasing the overall pressure of combined flow 116 used by trailing edge cooling circuit 125 to cool trailing edge 98 of rotor blade 92.

[0075] However, as will be appreciated, in other exemplary embodiments, turbine rotor 88 or shank 94 of turbine rotor 88 may define or include any other suitable airflow feature as a means for drawing a portion of the second airflow into the second cooling circuit, or more specifically, into internal passage 100. For example, the airflow feature may be a scoop (e.g., protruding from shank 94), a recess (e.g., extending into shank 94), or a combination thereof, of any other geometric shape. Additionally, in certain exemplary embodiments, the airflow feature may be used in conjunction with other means for drawing a portion of the second airflow into the second cooling circuit, or may be used as the sole means for drawing a portion of the second airflow into the second cooling circuit, or more specifically, into internal passage 100.

[0076] Now refer to Figure 7 , a method 200 for cooling a blade of an engine according to an exemplary aspect of the present disclosure is provided. The method can be used in combination with one or more of the above exemplary embodiments, or can be used with any other suitable structure.

[0077] The method 200 includes, at 202, cooling the first airflow (e.g., the cooled cooling airflow 106) using a cooled cooling air heat exchanger 107 before providing the first airflow (e.g., the cooled cooling airflow 106) to a primary cooling circuit (e.g., the inducer 72 and the turbine cooling cavity 86). The method 200 includes, at 204, passing the first airflow (e.g., the cooled cooling airflow 106) through the inducer 72 in a radial direction R of the turbofan engine 12 before providing the first airflow (e.g., the cooled cooling airflow 106) to the primary cooling circuit (e.g., the inducer 72 and the turbine cooling cavity 86). The method 200 includes, at 206, providing the first airflow (e.g., the cooled cooling airflow 106) through the primary cooling circuit (e.g., the inducer 72 and the turbine cooling cavity 86) in fluid communication with a first cooling circuit (e.g., the tip channel 120, the LE channel 122, and the TE channel 124) defined within a blade (e.g., the turbine rotor blade 92). Method 200 includes, at 208, providing a second air flow (e.g., compressed air flow 104) into a forward wheelspace cavity 84 of the turbofan engine 12, which is located forward of a shank 94 of a blade (e.g., turbine rotor blade 92) and is in fluid communication with a second cooling circuit (e.g., internal passage 100) defined within the blade (e.g., turbine rotor blade 92).

[0078] Method 200 includes drawing a portion of a second air flow (e.g., compressed air flow 104) into a second cooling circuit (e.g., internal passage 100) at 210. In certain exemplary embodiments, step 210 of method 200 may include drawing a portion of the second air flow (e.g., compressed air flow 104) into the second cooling circuit (e.g., internal passage 100) at least partially along with the first air flow (e.g., cooled cooling air flow 106) at 212. In certain exemplary embodiments, step 210 of method 200 may also include providing a portion of the first air flow (e.g., cooled cooling air flow 106) through an ejector (e.g., ejector pump 126) defined by blades (e.g., turbine rotor blades 92) at 214.

[0079] The method 200 includes, at 216, accelerating, using an ejector (e.g., the jet pump 126), a first air flow (e.g., the cooled cooling air flow 106) flowing through the ejector (e.g., the jet pump 126) into a trailing edge cooling circuit 125 defined within a blade (e.g., the turbine rotor blade 92). The method 200 includes, at 218, reducing a static pressure at a nozzle 127 of the ejector (e.g., the jet pump 126) in response to accelerating the first air flow (e.g., the cooled cooling air flow 106) into the trailing edge cooling circuit 125. The method 200 includes, at 220, drawing a portion of a second air flow (e.g., the compressed air flow 104) into the trailing edge cooling circuit 125 in response to reducing the static pressure at the nozzle 127 of the ejector (e.g., the jet pump 126).

[0080] The method 200 includes, at 222, delivering a portion of a first air flow (e.g., the cooled cooling air flow 106) to a leading edge 96 of a blade (e.g., a turbine rotor blade 92), a tip 97 of a blade (e.g., a turbine rotor blade 92), or both using a primary cooling circuit (e.g., the inducer 72 and the turbine cooling cavity 86). The method 200 includes, at 224, delivering a portion of a second air flow (e.g., the compressed air flow 104) drawn into a second cooling circuit (e.g., the inner passage 100) to a trailing edge 98 of the blade (e.g., the turbine rotor blade 92) using the first cooling circuit (e.g., the tip channel 120, the LE channel 122, and the TE channel 124).

[0081] This written description uses examples to disclose the present disclosure, including the best mode, and also to enable any person skilled in the art to practice the present disclosure, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the present disclosure is defined by the claims, and may include other examples that occur to those skilled in the art. If such other examples include 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 language of the claims, they are intended to be within the scope of the claims.

[0082] Further aspects are provided by the subject matter of the following clauses:

[0083] 19. The gas turbine engine of claim 18, wherein the cooling fan is configured to cool the at least one air flow to the exhaust gas fan and the cooling fan is configured to cool the at least one air flow to the exhaust gas fan.

[0084] The gas turbine engine according to one or more of these clauses, wherein the rotor blade includes a trailing edge, wherein the first cooling circuit includes a trailing edge cooling circuit configured to deliver air to the trailing edge of the rotor blade.

[0085] The gas turbine engine according to one or more of these clauses, wherein the first cooling circuit comprises a trailing edge cooling circuit, wherein the means for drawing a portion of the second air flow into the second cooling circuit comprises an ejector in fluid communication with the trailing edge cooling circuit, the first cooling circuit and the second cooling circuit.

[0086] The gas turbine engine according to one or more of these clauses, wherein the means for drawing a portion of the second air flow into the second cooling circuit comprises a nozzle defined by the ejector.

[0087] A gas turbine engine according to one or more of these clauses, wherein the rotor blade further includes: a shank defined by an inward portion of the rotor blade along the radial direction; a leading edge disposed outwardly from the shank along the radial direction; and a trailing edge disposed at an end of the rotor blade opposite the leading edge along the axial direction, wherein the first cooling circuit includes a trailing edge cooling circuit configured to deliver air to the trailing edge of the rotor blade.

[0088] A gas turbine engine according to one or more of these clauses, wherein the rotor blade further includes: a shank defined by an inward portion of the rotor blade along the radial direction; a leading edge disposed outwardly from the shank along the radial direction; and a trailing edge disposed at an end of the rotor blade opposite the leading edge along the axial direction, wherein the shank defines a tang protruding outwardly from a center of the rotor blade, wherein the tang defines an inlet of the second cooling circuit, wherein the inlet is disposed at an end face of the shank.

[0089] The gas turbine engine according to one or more of these clauses, wherein the gas turbine engine defines an inducer, wherein the inducer is connected to the primary cooling circuit and disposed upstream of the primary cooling circuit, wherein the inducer is in fluid communication with the primary cooling circuit.

[0090] The gas turbine engine according to one or more of these clauses, further comprising a source of the first air flow, wherein the source of the first air flow comprises a cooled cooling air heat exchanger.

[0091] The gas turbine engine according to one or more of these clauses, wherein the primary cooling circuit is configured to rotate about a longitudinal centerline of the gas turbine engine.

[0092] A gas turbine engine according to one or more of these clauses, wherein the rotor blade further comprises: a shank defined by an inward portion of the rotor blade in the radial direction; a leading edge disposed outwardly from the shank in the radial direction; and a trailing edge disposed at an end of the rotor blade opposite the leading edge in the axial direction, wherein the means for drawing a portion of the second air flow into the second cooling circuit comprises a scoop at an inlet of the second cooling circuit, wherein the scoop protrudes outwardly and away from the shank of the rotor blade.

[0093] A gas turbine engine according to one or more of these clauses, wherein the means for drawing a portion of the second air flow into the second cooling circuit comprises a scoop located at an inlet of the second cooling circuit, wherein the scoop defines a recessed opening, wherein the scoop is configured to draw a portion of the second air flow into the second cooling circuit using the recessed opening of the scoop.

[0094] 19. The air filter assembly of claim 18, wherein the cooling fan is configured to cool the exhaust gas fan to a temperature of the exhaust gas fan and to cool the exhaust gas fan to a temperature of the exhaust gas fan. The cooling fan is configured to cool the exhaust gas fan to a temperature of the exhaust gas fan. The cooling fan is configured to cool the exhaust gas fan to a temperature of the exhaust gas fan.

[0095] A method for cooling a blade of an engine, the engine defining an axial direction and a radial direction, the blade including a shank, the method comprising: providing a first air flow through a main cooling circuit, the main cooling circuit being in fluid communication with a first cooling circuit defined within the blade; providing a second air flow into a forward impeller space of the engine, the forward impeller space being forward of the shank along the axial direction, the forward impeller space being in fluid communication with a second cooling circuit defined within the blade; and drawing a portion of the second air flow into the second cooling circuit.

[0096] The method of one or more of these clauses, wherein drawing a portion of the second air flow into the second cooling circuit comprises drawing a portion of the second air flow into the second cooling circuit at least partially together with the first air flow.

[0097] The method of one or more of these clauses, wherein the blade defines an ejector, wherein drawing a portion of the second air flow at least partially with the first air flow into the second cooling circuit includes providing a portion of the first air flow through the ejector.

[0098] The method of one or more of these clauses, further comprising: accelerating, with the ejector, a first airflow flowing through the ejector into a trailing edge cooling circuit defined within the blade; reducing a static pressure at a nozzle of the ejector in response to accelerating the first airflow into the trailing edge cooling circuit; and drawing a portion of the second airflow into the trailing edge cooling circuit in response to reducing the static pressure at the nozzle of the ejector.

[0099] The method of one or more of these clauses further comprises delivering a portion of the first air flow to any cooling circuit defined within the blade using the primary cooling circuit.

[0100] The method of one or more of these clauses further comprising delivering, using the first cooling circuit, a portion of the second air flow drawn into the second cooling circuit to a trailing edge of the blade.

[0101] The method of one or more of these clauses, further comprising cooling the first air flow using a cooled cooling air heat exchanger before providing the first air flow through the main cooling circuit.

[0102] The method of one or more of these clauses, further comprising passing the first air flow through a flow inducer before providing the first air flow to the primary cooling circuit.

Claims

1. A gas turbine engine, the gas turbine engine defining an axial direction and a radial direction, characterized in that: The gas turbine engine comprises: a primary cooling circuit configured to receive a first air flow; and a turbine rotor including rotor blades at least partially defining a forward wheel space forward of the rotor blades, the forward wheel space being configured to receive a second air flow, the rotor blades further defining: a first cooling circuit inside the rotor blade and in fluid communication with the main cooling circuit for receiving the first air flow from the main cooling circuit; a second cooling circuit internal to the rotor blade and in fluid communication with the forward wheel space for receiving a portion of the second air flow from the forward wheel space; and means for drawing a portion of the second air flow into the second cooling circuit; Wherein, drawing the portion of the second air flow into the second cooling circuit includes drawing the portion of the second air flow at least partially together with the first air flow into the second cooling circuit.

2. The gas turbine engine according to claim 1, wherein: in, The rotor blade includes a trailing edge, wherein the first cooling circuit includes a trailing edge cooling circuit configured to deliver air to the trailing edge of the rotor blade.

3. The gas turbine engine according to claim 1, wherein: in, The first cooling circuit comprises a trailing edge cooling circuit, wherein the means for drawing the portion of the second air flow into the second cooling circuit comprises an ejector in fluid communication with the trailing edge cooling circuit, the first cooling circuit, and the second cooling circuit.

4. The gas turbine engine according to claim 3, characterized in that in, The means for drawing the portion of the second air flow into the second cooling circuit includes a nozzle defined by the ejector.

5. The gas turbine engine according to claim 1, wherein in, The rotor blade further comprises: a shank defined by an inward portion of the rotor blade along the radial direction; a leading edge disposed outwardly from the shank along the radial direction; and a trailing edge, the trailing edge being provided at an end of the rotor blade opposite to the leading edge in the axial direction, The first cooling circuit includes a trailing edge cooling circuit configured to deliver air to the trailing edge of the rotor blade.

6. The gas turbine engine according to claim 1, wherein: in, The rotor blade further comprises: a shank defined by an inward portion of the rotor blade along the radial direction; a leading edge disposed outwardly from the shank along the radial direction; and a trailing edge, the trailing edge being provided at an end of the rotor blade opposite to the leading edge in the axial direction, wherein the shank defines a tang protruding outwardly from a center of the rotor blade, wherein the tang defines an inlet of the second cooling circuit, wherein the inlet is disposed at an end surface of the shank.

7. The gas turbine engine according to claim 1, wherein: in, The gas turbine engine defines an inducer, wherein the inducer is connected to the primary cooling circuit and disposed upstream of the primary cooling circuit, wherein the inducer is in fluid communication with the primary cooling circuit.

8. The gas turbine engine according to claim 1, wherein: Further included is a source of the first air stream, wherein the source of the first air stream comprises a cooled cooling air heat exchanger.

9. The gas turbine engine according to claim 1, wherein: in, The primary cooling circuit is configured to rotate about a longitudinal centerline of the gas turbine engine.

10. The gas turbine engine according to claim 1, wherein: in, The rotor blade further comprises: a shank defined by an inward portion of the rotor blade along the radial direction; a leading edge disposed outwardly from the shank along the radial direction; a trailing edge disposed at an end of the rotor blade opposite to the leading edge along the axial direction; and wherein the means for drawing the portion of the second air flow into the second cooling circuit comprises a scoop at an inlet of the second cooling circuit, wherein the scoop projects outwardly and away from the shank of the rotor blade.

11. The gas turbine engine according to claim 1, wherein: in, The means for drawing the portion of the second air flow into the second cooling circuit comprises a scoop at an inlet of the second cooling circuit, wherein the scoop defines a concave opening, wherein the scoop is configured to draw the portion of the second air flow into the second cooling circuit using the concave opening of the scoop.

12. A rotor blade for a turbine of a gas turbine engine, characterized in that: The gas turbine engine defines a primary cooling circuit configured to receive a first air flow and defines a forward wheel space configured to receive a second air flow, the rotor blades defining: a first cooling circuit internal to the rotor blade and configured to be in fluid communication with the primary cooling circuit for receiving the first air flow from the primary cooling circuit when installed in the gas turbine engine and during operation of the gas turbine engine; a second cooling circuit internal to the rotor blade and configured to be in fluid communication with a forward wheelspace of the gas turbine engine for receiving a portion of a second air flow from the forward wheelspace when installed in the gas turbine engine and during operation of the gas turbine engine; and means for drawing a portion of the second air flow into the second cooling circuit; Wherein, drawing the portion of the second air flow into the second cooling circuit includes drawing the portion of the second air flow at least partially together with the first air flow into the second cooling circuit.

13. A method of cooling a blade of an engine, the engine defining an axial direction and a radial direction, the blade including a shank, characterized in that: The method comprises: providing a first air flow through a primary cooling circuit in fluid communication with a first cooling circuit defined within the blade; providing a second air flow into a forward wheel space of the engine, the forward wheel space being located forward of the shank in the axial direction, the forward wheel space being in fluid communication with a second cooling circuit defined within the blade; and drawing a portion of the second air flow into the second cooling circuit; Wherein, drawing the portion of the second air flow into the second cooling circuit includes drawing the portion of the second air flow at least partially together with the first air flow into the second cooling circuit.

14. The method according to claim 13, wherein: in, The vane defines an ejector, wherein drawing the portion of the second air flow at least partially into the second cooling circuit along with the first air flow includes providing a portion of the first air flow through the ejector.

15. The method according to claim 14, characterized in that Further including: utilizing the ejector to accelerate the first air flow flowing through the ejector into a trailing edge cooling circuit defined within the blade; reducing a static pressure at a nozzle of the ejector in response to accelerating the first airflow into the trailing edge cooling circuit; as well as In response to decreasing the static pressure at the nozzle of the ejector, the portion of the second air flow is drawn into the trailing edge cooling circuit.

16. The method according to claim 13, characterized in that Further included delivering a portion of the first air flow to any cooling circuits defined within the blade using the primary cooling circuit.

17. The method according to claim 13, characterized in that Further comprising delivering the portion of the second airflow drawn into the second cooling circuit to the trailing edge of the blade using the first cooling circuit.

18. The method according to claim 13, characterized in that Further included cooling the first air flow using a cooled cooling air heat exchanger before providing the first air flow through the primary cooling circuit.

19. The method according to claim 13, wherein Further included passing the first air flow through a flow inducer before providing the first air flow to the primary cooling circuit.

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