Patent Application
20200080427
This disclosure relates to cooling for a component of a gas turbine engine.
Gas turbine engines can include a fan for propulsion air and to cool components. The fan also delivers air into a core engine where it is compressed. The compressed air is then delivered into a combustion section, where it is mixed with fuel and ignited. The combustion gas expands downstream over and drives turbine blades. Static vanes are positioned adjacent to the turbine blades to control the flow of the products of combustion. The blades and vanes are subject to extreme heat, and thus cooling schemes are utilized for each.
An airfoil for a gas turbine engine according to an example of the present disclosure includes an airfoil section that extends in a radial direction from a platform to a tip, and extends in a thickness direction between a pressure side and a suction side that meet together at both a leading edge and a trailing edge. The tip defines a tip pocket that extends inwardly from the tip to a floor. The airfoil section includes a purge partition that extends in a chordwise direction between the leading and trailing edges and extends in the thickness direction between the pressure and suction sides to divide the airfoil section into a first region and a second region. The first region defines at least one internal cooling circuit. The second region defines a purge cavity. The purge partition defines a first set of apertures that fluidly couple the at least one internal cooling circuit and the purge cavity, and the purge cavity extends along the floor between the leading and trailing edges.
In a further embodiment of any of the foregoing embodiments, the at least one internal cooling circuit extends between the leading and trailing edges.
In a further embodiment of any of the foregoing embodiments, the airfoil section extends in the radial direction between 0% span at the platform and 100% span at the tip, and the purge partition is defined at a radial position between 75% and 98% span.
In a further embodiment of any of the foregoing embodiments, the airfoil section defines a second set of apertures in the floor that fluidly couple the purge cavity and the tip pocket, the airfoil section defines a purge slot along an external wall that bounds the purge cavity. A total cross-sectional area of all of the second set of apertures is greater than a combined sum, the combined sum defined as a total cross-sectional area of all of the first set of apertures and a cross-sectional area of the purge slot, and the cross-sectional area of the purge slot is greater than the cross-sectional area of each of the first set of apertures.
In a further embodiment of any of the foregoing embodiments, the airfoil section includes a rim wall feature that follows a perimeter of the airfoil section along the tip to bound the tip pocket.
In a further embodiment of any of the foregoing embodiments, the airfoil section defines a purge slot along an external wall that bounds the purge cavity.
In a further embodiment of any of the foregoing embodiments, the purge slot is defined adjacent to the trailing edge.
In a further embodiment of any of the foregoing embodiments, a cross-sectional area of the purge slot is greater than a cross-sectional area of each of the first set of apertures.
In a further embodiment of any of the foregoing embodiments, the airfoil section includes a plurality of ribs arranged to define a serpentine flow path in the at least one internal cooling circuit.
In a further embodiment of any of the foregoing embodiments, an external wall of the airfoil section defines a plurality of film cooling holes extending from the purge cavity.
In a further embodiment of any of the foregoing embodiments, the airfoil is a turbine blade.
A gas turbine engine according to an example of the present disclosure includes an array of blades rotatable about an engine axis, and an array of blade outer air seals (BOAS) arranged about the array of blades to bound a gas path. At least one of the array of blades includes an airfoil section that extends in a radial direction from a platform to a tip, and extends in a thickness direction between a pressure side and a suction side that meet together at both a leading edge and a trailing edge. A purge partition radially divides the airfoil section. A perimeter of the purge partition follows along external walls of the airfoil section. The tip defines a tip pocket that extends radially inwardly from the tip to a floor. The airfoil section includes at least one internal cooling circuit and a purge cavity opposed to the at least one internal cooling circuit along the purge partition, with the purge partition bounded by the floor. The purge partition defines a first set of apertures that fluidly couple the at least one internal cooling circuit and the purge partition.
In a further embodiment of any of the foregoing embodiments, the purge partition extends in the thickness direction between the pressure side and the suction side.
In a further embodiment of any of the foregoing embodiments, the floor defines a second set of apertures that fluidly couple the purge cavity and the tip pocket.
In a further embodiment of any of the foregoing embodiments, each of the second set of apertures defines a passage axis oriented to eject flow into a clearance gap defined in the radial direction between the tip and an adjacent one of the BOAS.
A further embodiment of any of the foregoing embodiments includes a turbine section driving a compressor section. The airfoil is a turbine blade of the turbine section.
A method of cooling a gas turbine engine component according to an example of the present disclosure includes communicating cooling flow to an internal cooling circuit of an airfoil section. A purge partition radially divides the airfoil section. The airfoil section defines a purge cavity opposed to the internal cooling circuit along the purge partition. The method includes communicating the cooling flow through a first set of apertures in the purge partition and then to the purge cavity to purge debris in the internal cooling circuit, and communicating a portion of the cooling flow from the purge cavity to a second set of apertures defined along a floor of a tip pocket. The tip pocket extends inwardly from a tip of the airfoil section and ejects the portion of the cooling flow from the second set of apertures to a clearance gap defined between the tip and a blade outer air seal (BOAS).
In a further embodiment of any of the foregoing embodiments, the gas turbine engine component is a turbine blade, and the airfoil section extends radially from a platform to the tip.
In a further embodiment of any of the foregoing embodiments, the tip pocket is bounded by a rim wall feature that follows a perimeter of the airfoil section along a tip portion to define the tip, and the method further includes contacting the rim wall feature with the BOAS in response to rotating the airfoil section.
A further embodiment of any of the foregoing embodiments includes removing the tip portion that has walls of the airfoil section bounding the purge cavity, and attaching a second tip portion to a remainder of the airfoil section to define another instance of the purge cavity and another instance of the tip pocket.
The various features and advantages of this invention will become apparent to those skilled in the art from the following detailed description of an embodiment. The drawings that accompany the detailed description can be briefly described as follows.
The exemplary engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of bearing systems 38 may be varied as appropriate to the application.
The low speed spool 30 generally includes an inner shaft 40 that interconnects, a first (or low) pressure compressor 44 and a first (or low) pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive a fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a second (or high) pressure compressor 52 and a second (or high) pressure turbine 54. A combustor 56 is arranged in exemplary gas turbine 20 between the high pressure compressor 52 and the high pressure turbine 54. A mid-turbine frame 57 of the engine static structure 36 may be arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The mid-turbine frame 57 further supports bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.
The core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 57 includes airfoils 59 which are in the core airflow path C. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion. It will be appreciated that each of the positions of the fan section 22, compressor section 24, combustor section 26, turbine section 28, and fan drive gear system 48 may be varied. For example, gear system 48 may be located aft of the low pressure compressor, or aft of the combustor section 26 or even aft of turbine section 28, and fan 42 may be positioned forward or aft of the location of gear system 48.
The engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about five 5:1. Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. The geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1 and less than about 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 invention is applicable to other gas turbine engines including direct drive turbofans.
A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of the engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and 35,000 ft (10,668 meters), with the engine at its best fuel consumption —also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R)/(518.7°R)]0.5. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.5 meters/second).
A vane 66 is positioned along the engine axis A and adjacent to the airfoil 62. The vane 66 is spaced axially from the adjacent airfoil 62. The vane 66 includes an airfoil section 66A extending between an inner platform 66B and an outer platform 66C to define a portion of the core flow path C. The turbine section 28 includes an array of airfoils 62, BOAS 64, and vanes 66 arranged circumferentially about the engine axis A.
The turbine section 28 includes a cooling arrangement 70. The cooling arrangement 70 includes one or more cooling cavities or plenums 68, 69 defined by an engine static structure such as the engine case 37 or another portion of the engine static structure 36 (
Referring to
Referring to
The airfoil 162 defines a cooling arrangement 170 for transporting cooling air flow F through portions of the airfoil 162. The cooling arrangement 170 includes at least one internal cooling cavity or circuit 174 in a thickness of the airfoil 162. The cooling circuit 174 can include one or more portions defined at various locations of the airfoil 162 and at various orientations. In the illustrated example of
The cooling circuit 174 includes one or more segments or cooling passages 174-2 that extend through the airfoil section 162B for cooling portions of the airfoil 162. The cooling passages 174-2 can be interconnected to circulate the cooling flow F in the airfoil section 162B. For example, the airfoil 162 can include a plurality of ribs 162R defined by the internal walls 162N that are arranged to define a multi-pass serpentine flow path in the cooling circuit 174. The internal cooling circuit 174 extends between the leading edge 162LE and the trailing edge 162TE. The airfoil 162 can define one or more apertures 176 for ejecting cooling flow F from the cooling circuit 174 to provide convective and/or film cooling to adjacent portions of the airfoil 162. The apertures 176 can be defined adjacent the leading and/or trailing edges 162LE, 162TE, for example.
Referring to
The tip pocket 178 extends predominately radially inwardly from the tip 162C to a floor 180, which forms the outer tip wall, such that the tip pocket 178 is exposed to the gas path GP along the tip 162C. The squealer pocket and/or tip floor 180 is a wall 162E of the airfoil 162 that extends in the chord-wise and thickness directions C, T. The arrangement of the tip pocket 178 and rim wall feature 182 may be referred to as a “tip squealer” arrangement. In other examples, the tip pocket 178 is omitted such that the tip 162C is relatively planar and flush and of equivalent radial height between the leading and trailing edges 162LE, 162TE and the pressure and suction sides 162P, 162S.
During operation, dirt or sand particles and other debris may be communicated in the internal cooling circuit 174. The debris may deposit or collect on internal surfaces of the internal cooling circuit 174, which may block or otherwise impede flow of the cooling flow F through portions of the cooling arrangement 170. Accumulation of debris can act as an insulator which reduces the thermal cooling effectiveness and convective cooling of adjacent portions of the airfoil 162. Reductions in local thermal cooling effectiveness due to dirt particulate accumulation may cause increases in the floor 180 wall metal temperatures, and in local airfoil tip 162C thermal strains, resulting in premature and accelerated durability failure modes related to local airfoil tip 162C oxidation, and thermal mechanical fatigue (TMF) cracking in the airfoil 162. Thru wall oxidation and TMF cracking in the airfoil tip 162C, may adversely impact the functional thermal cooling performance associated with the internal convective heat transfer augmentation and film cooling characteristics which would have otherwise been provided by a non-compromised internal cooling circuit 174. The redirecting of cooling air flow F, thru local distressed regions of the airfoil tip 162C, reduces the overall cooling efficiency and durability capability of the airfoil 162. Accumulation of debris can also cause an increase in turbine temperature and an increase in the clearance gap G, which can reduce efficiency of the airfoil 162, as well as stage and engine performance.
Again with reference to
In the illustrated example of
In examples, the purge partition 184 is defined at a radial position between 75% and 98% span, or more narrowly at a radial position greater than or equal to 85% span, such that the first region R1 has a volume that is less than a volume of the second region R2. The radial position of the purge partition 184 can be defined with respect to one or more predetermined vibratory modes of the airfoil 162. In examples, the volume of the first region R1 is less than 25% of the volume of the second region R2, or more narrowly less than 5% or 10%. The volumes of the first and second regions R1, R2 are defined with respect to a spatial boundary of the airfoil section 162B.
The first region R1 defines cooling arrangement 170 to include a purge cavity 186 that is configured to receive cooling flow F and debris from the internal cooling circuit(s) 174. The purge cavity 186 extends between and is bounded in the radial direction R between the squealer pocket and/or tip floor 180 and the purge partition 184. The purge cavity 186 is opposed to the internal cooling circuit(s) 174 along the purge partition 184.
The purge cavity 186 extends along the floor 180 between the leading and trailing edges 162LE, 162TE. In the illustrated example of
The purge partition 184 defines a first set of apertures 188 spaced apart from the floor 180. The apertures 188 interconnect or otherwise fluidly couple the internal cooling circuit(s) 174 and the purge cavity 186. In examples, the apertures 188 provide the sole flow path between each and every cooling flow path or cavity defined in the second region R2, including internal cooling circuit(s) 174, and each and every cooling flow path or cavity defined in the first region R1, including the purge cavity 186 and tip pocket 178. Cooling flow F communicated from the apertures 188 can provide convective cooling to adjacent portions of the purge cavity 186. The apertures 188 can be oriented to eject cooling flow F to impinge on the backside surfaces, defined by the inboard radius of the squealer pocket and/or tip floor 180 bounding the purge cavity 186.
The airfoil section 162B can define one or more purge slots 190 (one shown for illustrative purposes) along an external wall 162E that bounds the purge cavity 186. In the illustrated example of
The purge slot 190 can be dimensioned relative to the apertures 188 to establish a predetermined pressure in the purge cavity 186 and across the purge slot 190. In examples, a cross-sectional area of the purge slot 190 is greater than a cross-sectional area of each of the individual apertures 188. The purge slot 190 can be dimensioned to provide a low pressure sink for relatively fine dirt/sand particulate or debris D to evacuate the purge cavity 186. In examples, the cross-sectional area of each of the individual apertures 188 taken along the purge cavity 186 is less than 5%, or more narrowly less than 3%, of the total cross-sectional area of the purge partition 184 taken along the purge cavity 186.
One or more of the external walls 162E of the airfoil section 162B can define one or more film cooling holes 192 (shown in dashed lines in
The airfoil section 162B defines a second set of apertures 194 in the squealer pocket and/or tip floor 180. The apertures 194 interconnect or otherwise fluidly couple the purge cavity 186 and the tip pocket 178. Each of the apertures 194 defines a passage axis PA. The passage axis PA is oriented to eject cooling flow F into clearance gap G defined in the radial direction R between the airfoil tip 162C and adjacent Blade Outer Air Seal (BOAS) 164 (shown schematically in dashed lines in
Continuing reference to
The cooling flow F carries, transports or otherwise moves the debris D towards the purge cavity 186 and/or second set of apertures 194. The debris D is ejected outwardly through the purge slot 190 and/or apertures 194 and into the gas path GP (
During operation, the radial positions of the airfoil tip 162C and/or adjacent BOAS 164 (shown in dashed lines) can change to cause a dimension of the clearance gap G to vary. The rim wall feature 182 can rub against or otherwise directly contact an inner diameter of the BOAS 164 in response to rotating the airfoil 162 including airfoil section 162B about the axis X (
As illustrated by
To repair the airfoil 162, material of the tip portion 162TP that extends outwardly from the purge partition 184 can be removed, including walls of the airfoil section 162B bounding the purge cavity 186. Referring to
The components 161/261 disclosed herein can be made of metal, ceramic or composite materials, for example. Example metals include high temperature nickel and cobalt alloys. Various manufacturing techniques can be utilized to fabricate the purge partition 184/284 and apertures 188/288. In examples, the purge partition 184/284 is a separate and distinct component that is mechanically attached or otherwise secured to an interior of the airfoil 162/262. In other examples, the purge partition 184/284 is integrally formed with walls of the airfoil section 162B/262B. Example manufacturing techniques include additive manufacturing such as laser sintering power bed methods and investment casting utilizing a frangible core and lost wax. Another example manufacturing technique includes a Tomo-Lithographic Molding (TOMO) process, which combines lithography, molding and casting manufacturing methods to manufacture an integrated free form core.
It should be understood that relative positional terms such as “forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like are with reference to the normal operational attitude of the vehicle and should not be considered otherwise limiting.
Although the different examples have the specific components shown in the illustrations, embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples.
Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present disclosure.
The foregoing description is exemplary rather than defined by the limitations within. Various non-limiting embodiments are disclosed herein, however, one of ordinary skill in the art would recognize that various modifications and variations in light of the above teachings will fall within the scope of the appended claims. It is therefore to be understood that within the scope of the appended claims, the disclosure may be practiced other than as specifically described. For that reason the appended claims should be studied to determine true scope and content.
