Patent Application
20210108524
The present disclosure relates generally to gas turbine engines, and more specifically to turbine vane assemblies for use with gas turbine engines.
Gas turbine engines are used to power aircraft, watercraft, power generators, and the like. Gas turbine engines typically include a compressor, a combustor, and a turbine. The compressor compresses air drawn into the engine and delivers high pressure air to the combustor. In the combustor, fuel is mixed with the high pressure air and is ignited. Products of the combustion reaction in the combustor are directed into the turbine where work is extracted to drive the compressor and, sometimes, an output shaft. Left-over products of the combustion are exhausted out of the turbine and may provide thrust in some applications.
Products of the combustion reaction directed into the turbine are conducted toward airfoils included in stationary vanes and rotating blades of the turbine. The airfoils are often made from high-temperature resistant materials and/or are actively cooled by supplying relatively cool air to the vanes and blades due to the high temperatures of the combustion products. To this end, some airfoils for vanes and blades are incorporating composite materials adapted to withstand very high temperatures. Design and manufacture of vanes and blades from composite materials presents challenges because of the geometry and strength desired for the parts.
The present disclosure may comprise one or more of the following features and combinations thereof.
A turbine vane assembly for use in a gas turbine engine may include a plurality of ceramic matrix composite turbine vanes and a metallic outer vane support. The plurality of ceramic matrix composite turbine vanes may be adapted to interact with hot gases flowing through a gas path of the gas turbine engine during use of the turbine vane assembly. The metallic outer vane support may be configured to receive force loads applied to the plurality of ceramic matrix composite turbine vanes by the hot gases during use of the turbine vane assembly in the gas turbine engine.
In some embodiments, the plurality of ceramic matrix composite turbine vanes may include a first turbine vane and a second turbine vane. The second turbine vane may be spaced apart circumferentially from the first turbine vane relative to an axis.
In some embodiments, the metallic outer vane support may include an outer mount, a first support spar, and a second support spar. The outer mount may be located radially outward of the plurality of ceramic matrix composite turbine vanes and may extend at least partway circumferentially about the axis. The second support spar may be spaced apart circumferentially from the first support spar relative to the axis.
In some embodiments, the first support spar may extend radially inward from the outer mount through an interior cavity of the first turbine vane. The second support spar may extend radially inward from the outer mount through an interior cavity of the second turbine vane. In some embodiments, the first and second support spars are integrally formed with the outer mount to form a single-piece component.
In some embodiments, the turbine vane assembly may include a metallic inner vane support. The inner vane support may be spaced apart radially from the outer mount relative to the axis to locate the plurality of turbine vanes radially between.
In some embodiments, the metallic inner vane support may include an inner mount and at least two fasteners. The inner mount may extend at least partway circumferentially about the axis. The at least two fasteners may be configured to couple the first and second support spars of the metallic outer vane support to the inner mount to provide a mechanical linkage between the first turbine vane and the second turbine vane. In some embodiments, the mechanical linkage formed between the outer vane support and the inner vane support may reduce twisting of the turbine vane assembly and adjacent turbine vane assemblies relative to one another during use of the turbine vane assembly in the gas turbine engine.
In some embodiments, the inner mount may include an inner mount platform and raised interface surfaces. The inner mount platform may extend at least circumferentially partway about the axis between the plurality of ceramic matrix composite turbine vanes. The raised interface surfaces may be spaced circumferentially apart from one another. In some embodiments, each of the raise interface surfaces may extend radially outward from the inner mount platform and engage one of the first support spar and the second support spar to block radial movement of the inner mount relative to the outer vane support.
In some embodiments, the inner mount may further include anti-rotation pegs. The anti-rotation pegs may each extend radially outward from one of the raised interface surfaces and into a corresponding support spar to block twisting of the inner mount relative to the outer vane support.
In some embodiments, the metallic inner vane support may further include a first nozzle and a second nozzle. The first nozzle may be arranged radially inward from the inner mount platform and configured to receive an inner end of the first support spar. The second nozzle may be arranged radially inward from the inner mount platform and configured to receive an inner end of the second support spar.
In some embodiments, the inner end of each of the first and second support spars may be threaded and the at least two fasteners may be nuts. The nuts may be configured to mate with threads on the inner end of one of the first and second support spars and engage one of the first nozzle and the second nozzle to maintain engagement of the raised interface surfaces and the anti-rotation pegs with the corresponding support spar of the first support spar and the second support spar.
In some embodiments, the inner mount, the first nozzle, and the second nozzle of the inner vane support may be integrally formed. The inner mount, the first nozzle, and the second nozzle of the inner vane support may be integrally formed such that the inner mount, the first nozzle, and the second nozzle are a one-piece, integral component.
In some embodiments, the first nozzle and the second nozzle may each include a cylindrical tube, an anti-rotation notch, and a spout. The cylindrical tube may be configured to receive the inner end of one of the first support spar and the second support spar. The anti-rotation notch may extend into the cylindrical tube and may be configured to receive an anti-rotation tab extending radially inward from the inner mount platform. The spout may extend circumferentially from the cylindrical tube and may be configured to discharge a flow of cooling air.
In some embodiments, the inner end of each of the first and second support spars may be threaded and the at least two fasteners each include a first nut and a second nut. Each of the first nuts may be configured to mate with threads on the inner end of one of the first and second support spars and engage the inner mount platform to maintain engagement of the raised interface surfaces and the anti-rotation pegs with the corresponding support spar of the first support spar and the second support spar. Each of the second nuts may be spaced radially inward of the first nut to locate one of the first nozzle and the second nozzle therebetween.
In some embodiments, each of the second nuts may be configured to mate threads on the inner end of one of the first support spar and the second support spar. The second nuts may be configured to engage one of the first nozzle and the second nozzle to block removal of the one of the first nozzle and the second nozzle off the inner end of the one of the first support spar and the second support spar.
In some embodiments, the inner vane support may further include a first nozzle and a second nozzle. The first nozzle may be arranged radially inward from the inner mount platform and may be configured to receive an inner end of the first support spar. The second nozzle may be arranged radially inward from the inner mount platform and may be configured to receive an inner end of the second support spar.
In some embodiments, the at least two fasteners may include plurality of bolts. The plurality of bolts may each extend through one of the first nozzle and the second nozzle and the inner mount platform into one of the first support spar and the second support spar. The plurality of bolts may be configured to couple each of the first nozzle and the second nozzle to the inner mount platform and block twisting of the inner vane support relative to the outer vane support.
In some embodiments, the metallic outer vane support may include an outer mount platform and a plurality of reinforcement extensions. The outer mount platform may extend circumferentially at least partway about the axis. The outer mount platform may be configured to be coupled to a turbine case of the gas turbine engine. The plurality of reinforcement extensions may extend radially outward from an outer surface of the outer mount platform relative to the axis. The reinforcement extensions may be configured to minimize resulting stresses in the outer mount platform due to the twisting of the turbine vane assembly.
In some embodiments, the plurality of reinforcement extensions may include a plurality of axially extending reinforcement ribs and a plurality of circumferentially extending reinforcement ribs. The axially extending reinforcement ribs may extend radially outward from and axially along the outer surface of the outer mount platform relative to the axis. The circumferentially extending reinforcement ribs may extend radially outward from and circumferentially along the outer surface of the outer mount platform relative to the axis.
In some embodiments, the turbine vane assembly may further comprise a metallic inner vane support spaced apart radially from the outer mount relative to the axis to locate the plurality of ceramic matrix composite turbine vanes radially between. The metallic inner vane support may include an inner mount platform, a first mating feature, and a second mating feature. The inner mount platform may extend at least partway circumferentially about the axis. The first mating feature may engage an inner end of the first support spar to block rotation of the metallic outer vane support about a spar axis relative to the metallic inner vane support. The second mating feature may couple to an inner end of the second support spar to block radial movement of the metallic outer vane support relative to the metallic inner vane support.
In some embodiments, the metallic inner vane support may further include a locking pin. The locking pin may extend through the inner mount platform and into the first support spar to block circumferential rotation of the metallic outer vane support about the axis relative to the metallic inner vane support.
In some embodiments, the first mating feature may be a rotational stop. The rotational stop may extend radially outward from the inner mount and engage the inner end of the first support spar. The rotational stop may provide load transfer from the inner mount platform to the first support spar of the outer vane support.
In some embodiments, the second mating feature may be at least one locking notch formed in the inner mount platform. The second support spar may include at least one locking tab that extends circumferentially from the inner end of the second support spar and into the notch to provide a bayonet fitting therebetween. The bayonet fitting may block radial movement of the outer vane support relative to the inner vane support.
In some embodiments, the turbine vane assembly may further comprise a metallic inner vane support spaced apart radially from the outer mount relative to the axis to locate the plurality of ceramic matrix composite turbine vanes radially between. The metallic inner vane support may include an inner mount and a retainer plate. The inner mount may extend at least partway circumferentially about the axis. The retainer plate may be located radially inward of the inner mount.
In some embodiments, the retainer plate may couple to an inner end of the second support spar to block radial movement of the metallic outer vane support relative to the metallic inner vane support. The retainer plate may engage an inner end of the first support spar to block rotation of the metallic outer vane support about a spar axis relative to the metallic inner vane support.
In some embodiments, the turbine vane assembly may further comprise a metallic inner vane support spaced apart radially from the outer mount relative to the axis to locate the plurality of ceramic matrix composite turbine vanes radially between. The metallic inner vane support may include an inner mount platform and at least one locking notch formed in the inner mount platform. The inner mount platform may extend at least partway circumferentially about the axis. The at least one locking notch may receive at least one locking tab formed on an inner end of the second support spar to block radial movement of the metallic outer vane support relative to the metallic inner vane support.
In some embodiments, the metallic inner vane support may further include a locking pin. The locking pin may extend through the inner mount platform and into the first support spar to block circumferential rotation of the metallic outer vane support about the axis relative to the metallic inner vane support.
In some embodiments, the metallic inner vane support may further include a rotational stop. The rotational stop may engage an inner end of the first support spar to block rotation of the metallic outer vane support about a spar axis relative to the metallic inner vane support.
According to another aspect of the present disclosure, a turbine vane assembly for use in a gas turbine engine may include a plurality of turbine vanes and an outer vane support. In some embodiments, the outer vane support may at least one outer mount and a plurality of support spars.
In some embodiments, the outer mount may be located radially outward of the plurality of ceramic matrix composite turbine vanes and may extend circumferentially at least partway about an axis. The plurality of support spars may each extend radially inward from the at least one outer mount through an interior cavity of one turbine vane of the plurality of turbine vanes. In some embodiments, wherein the plurality of support spars may be integrally formed with the at least one outer mount to form a single-piece component.
In some embodiments, the turbine vane assembly further may include an inner vane support. The inner vane support may be spaced apart radially from the at least one outer mount relative to the axis to locate the plurality of turbine vanes radially between.
In some embodiments, the inner vane support may include at least one inner mount and a plurality of fasteners. The at least one inner mount may extend circumferentially at least partway about the axis. The plurality of fasteners may each be configured to couple a corresponding support spar of the plurality of support spars of the outer vane support to the at least one inner mount.
In some embodiments, the plurality of turbine vanes may include at least two turbine vanes. In some embodiments, the plurality of support spars may include at least two support spars.
In some embodiments, the plurality of turbine vanes may include at least three turbine vanes. In some embodiments, the plurality of support spars may include at least three support spars.
In some embodiments, the outer vane support may include at least two outer mounts. The at least two outer vane mounts may have a second outer mount spaced apart circumferentially from a first outer mount.
In some embodiments, the plurality of support spars includes a first support spar, a second support spar, a third support spar, and a fourth support spar. The first support spar may extend radially inward from the first outer mount through a first turbine vane of the plurality of turbine vanes. The second support spar may be spaced apart circumferentially from the first support spar relative to the axis and may extend radially inward from the first outer mount through a second turbine vane of the plurality of turbine vanes. The third support spar may extend radially inward from the second outer mount through a third turbine vane of the plurality of turbine vanes. The fourth support spar may be spaced apart circumferentially from the third support spar relative to the axis and may extend radially inward from the second outer mount through a fourth turbine vane of the plurality of turbine vanes.
In some embodiments, the first outer mount and the second outer mount may each include an outer mount platform and a plurality of reinforcement extensions. Each outer mount platform may extend at least partway about the axis and may be configured to be coupled to a turbine case. The plurality of reinforcement extensions may extend radially outward from an outer surface of the outer mount platform relative to the axis.
In some embodiments, the at least one inner mount may include an inner mount platform raising interface surfaces, and anti-rotation pegs. The inner mount platform may extend at least circumferentially partway about the axis between the plurality of turbine vanes. The raised interface surfaces may be spaced circumferentially apart from one another and each extend radially outward from the inner mount platform. The anti-rotation pegs may each extend radially outward from one of the raised interface surfaces.
In some embodiments, the raised interface surfaces may engage one of the plurality of support spars to block radial movement of the at least one inner mount relative to the outer vane support. The anti-rotation pegs may each extend radially outward into one support spar of the plurality of support spars to block twisting of the at least one inner mount relative to the outer vane support.
In some embodiments, the inner vane support may include at least two inner mounts. The at least two inner mounts may have a second inner mount spaced apart circumferentially from a first inner mount.
In some embodiments, the plurality of fasteners may include a first fastener, a second fastener, a third fastener, and a fourth fastener. The first fastener may be configured to couple a first support spar of the plurality of support spars to the first inner mount. The second fastener may be configured to couple a second support spar of the plurality of support spars to the first inner mount. The third fastener may be configured to couple a third support spar of the plurality of support spars to the second inner mount. The fourth fastener may be configured to couple a fourth support spar of the plurality of support spars to the second inner mount.
In some embodiments, the first inner mount and the second inner mount may each include an inner mount platform raised interface surfaces, and anti-rotation pegs. The inner mount platform may extend at least circumferentially partway about the axis between at least two turbine vanes of the plurality of turbine vanes. The raised interface surfaces may be spaced circumferentially apart from one another and each extend radially outward from the inner mount platform. Each of the raised interface surfaces may engage one of the plurality of support spars to block radial movement of the at least two inner mounts relative to the outer vane support. The anti-rotation pegs may each extend radially outward from one of the raised interface surfaces and into one support spar of the plurality of support spars.
In some embodiments, the inner vane support may further include a plurality of nozzles. Each nozzle of the plurality of nozzles may be configured to receive an inner end of one support spar of the plurality of support spars.
In some embodiments, the turbine vane assembly may further comprise an inner vane support spaced apart radially from an outer mount included in the outer vane support relative to the axis to locate the plurality of turbine vanes radially between. The inner vane support may include an inner mount platform, a first mating feature, and a second mating feature. The inner mount platform may extend at least partway circumferentially about the axis. The first mating feature may engage an inner end of a first support spar included in the plurality of support spars to block rotation of the outer vane support about a spar axis relative to the metallic inner vane support. The second mating feature may couple to an inner end of a second support spar included in the plurality of support spars to block radial movement of the outer vane support relative to the inner vane support.
In some embodiments, the inner vane support may further include a locking pin. The locking pin may extend through the inner mount platform and into the first support spar to block circumferential rotation of the outer vane support about the axis relative to the inner vane support.
In some embodiments, the first mating feature may be a rotational stop. The rotational stop may extend radially outward from the inner mount and engage the inner end of the first support spar. The rotational stop may provide load transfer from the inner mount platform to the first support spar of the outer vane support.
In some embodiments, the second mating feature may be at least one locking notch formed in the inner mount platform. The second support spar may include at least one locking tab that extends circumferentially from the inner end of the second support spar and into the notch to provide a bayonet fitting therebetween. The bayonet fitting may block radial movement of the outer vane support relative to the inner vane support.
In some embodiments, the turbine vane assembly may further comprise an inner vane support spaced apart radially from an outer mount included in the outer vane support relative to the axis to locate the plurality of turbine vanes radially between. The inner vane support may include an inner mount and a retainer plate. The inner mount may extend at least partway circumferentially about the axis. The retainer plate may be located radially inward of the inner mount.
In some embodiments, the retainer plate may couple to an inner end of a first support spar included in the plurality of support spars to block radial movement of the outer vane support relative to the inner vane support. The retainer plate may engage an inner end of a second support spar included in the plurality of support spars to block rotation of the outer vane support about a spar axis relative to the inner vane support.
In some embodiments, the turbine vane assembly may further comprise an inner vane support spaced apart radially from an outer mount included in the outer vane support relative to the axis to locate the plurality of turbine vanes radially between. The inner vane support may include an inner mount platform and a locking pin. The inner mount platform may extend at least partway circumferentially about the axis. The locking pin may extend through the inner mount platform and into a first support spar included in the plurality of support spars to block circumferential rotation of the outer vane support about the axis relative to the inner vane support.
In some embodiments, the inner vane support may further include at least one locking notch formed in the inner mount platform. The at least one locking notch may receive at least one locking tab formed on an inner end of a second support spar included in the plurality of support spars to block radial movement of the outer vane support relative to the inner vane support.
In some embodiments, the metallic inner vane support may further include a rotational stop. The rotational stop may engage an inner end of the first support spar to block rotation of the outer vane support about a spar axis relative to the inner vane support.
These and other features of the present disclosure will become more apparent from the following description of the illustrative embodiments.
For the purposes of promoting an understanding of the principles of the disclosure, reference will now be made to a number of illustrative embodiments illustrated in the drawings and specific language will be used to describe the same.
A turbine vane assembly 10 for use in a gas turbine engine 110 is shown in
The vanes 12 comprise ceramic materials, while the outer and inner vane supports 14, 16 comprise metallic materials in the illustrative embodiment. As such, the ceramic matrix composite vanes 12 are adapted to withstand high temperatures, but may have relatively low strength compared to the metallic vane supports 14, 16. The vane supports 14, 16 provide structural strength to the turbine vane assembly 10 by receiving the force loads applied to the vanes 12 and transferring them to a casing 20 that surrounds the turbine vane assembly 10.
The turbine vane assembly 10 is adapted for use in the gas turbine engine 110, which includes a fan 112, a compressor 114, a combustor 116, and a turbine 118 as shown in
The turbine 118 includes a turbine case 20, a plurality of static turbine vane rings 22 that are fixed relative to the axis 19, and a plurality of bladed rotating wheel assemblies 24 as suggested in
The force loads received by the outer and inner vane supports 14, 16 from the turbine vanes 12 and/or other components 80 of the gas turbine engine 110 may impart a rotation on each turbine vane assembly 10 included in the turbine vane ring 22. The resulting rotation may result in increased leakage between the vane assemblies 10. To minimize twisting between assemblies 10, the outer and inner vane supports 14, 16 are arranged to extend partway about the axis 19 and provide a mechanical linkage between circumferentially adjacent turbine vanes 12. The increased surface area and structural reinforcement at both radially outer and inner ends of the turbine vane 12 reduces the non-trivial rotation between adjacent turbine vane assemblies 10 and therefore reduces the leakage and increases engine 110 performance.
The plurality of turbine vanes 12 includes a first turbine vane 26 and a second turbine vane 28 as shown in
The outer vane support 14 includes an outer mount 32, a first support spar 34, and a second support spar 36 as shown in
The outer mount 32 includes an outer mount platform 40 and a plurality of reinforcement extensions 42, 44 as shown in
In some embodiments, the reinforcement ribs 42, 44 may be configured to help minimize the axial deflection of the turbine vane assembly 10. The reinforcement ribs 42, 44 may also be configured to help minimize resulting stresses in the outer mount platform 40 due to twisting of the turbine vane assembly 10.
Each of the support spars 34, 36 include an outer end 46, an inner end 48, and a strut 50 as shown in
Each of the support spars 34, 36 are also shaped to include a cooling channel 52 as shown in
The inner end 48 of each support spar 34, 36 is shaped to include a cooling air exit hole 54 as shown in
The strut 50 of each support spar 34, 36 is shaped to include inner interface surface 60 and an anti-rotation notch 62 as shown in
The inner vane support 16 includes an inner mount 64, a plurality of fasteners 66, 68, and a plurality of nozzles 70, 72 as shown in
In the illustrative embodiment, the inner mount 64 is configured to be coupled to an inter-stage seal 80 included in the turbine section 118 as shown in
In the illustrative embodiment, the plurality of fasteners 66, 68 includes a first fastener 66 and a second fastener 68 as shown in
In the illustrative embodiment, the plurality of nozzles 70, 72 includes first nozzle 70 and a second nozzle 72 as shown in
In the illustrative embodiment, the nozzles 70, 72 are integrally formed with the inner mount 64 such that the inner mount 64, the first nozzle 70, and the second nozzle 72 are a one-piece, integral component. In other embodiments, the nozzles 70, 72 may be separate pieces from the inner mount 64.
The inner mount 64 includes an inner mount platform 74, raised interface surfaces 76, and anti-rotation pegs 78 as shown in
In the illustrative embodiment, the inner mount platform 74 is machined to form the raised interface surfaces 76. In other embodiments, the inner mount platform 74 may be machined so that the interface surfaces 76 extend radially into the inner mount platform 74.
In the illustrative embodiment, the anti-rotation peg 78 extends radially outward from the inner mount platform 74. In other embodiments, the anti-rotation feature arrangement may be reversed so that the anti-rotation notch 62 is machined into the inner mount platform 74 and the strut 50 of the support spar 34, 36 includes the anti-rotation peg 78.
Each nozzle 70, 72 includes a cylindrical tube 86 and a spout 88 as shown in
Each fastener 66, 68 includes a nut 90 and a pin 92 as shown in
In other embodiments, the fastener may only include the pin 92. In such embodiments, the pin 92 may extend through a portion of the nozzle 70, 72 to block removal of the nozzle 70, 72 from the inner end 48 of the support spar 34, 36. In other embodiments, the pin 92 may extend through a portion of the inner mount platform 74 into the inner end 48 of the support spar 34, 36 to couple the support spar 34, 36 to the inner mount platform 74. In other embodiments, the fastener may be another suitable nut-locking feature or joint coupling.
In the illustrative embodiment, each nut 90 engages the corresponding nozzle 70, 72 to cause the raised interface surface 76 of the inner mount platform 74 to engage the inner interface surface 60 of the strut 50. The maintained engagement of the inner mount 64 with the strut 50 maintains the anti-rotation features and minimizes twisting of the vane supports 14, 16.
Turning again to the turbine vanes 26, 28, each turbine vane 26, 28 is shaped to include an outer platform 94, an inner platform 96, and an airfoil 98 as shown in
The airfoil 98 is also formed to define the interior cavity 30 that extends radially into the airfoil 98 as shown in
In the illustrative embodiment, the outer platform 94, the inner platform 96, and the airfoil 98 of the vane 26, 28 are integrally formed from ceramic matrix composite materials. As such, the outer platform 94, the inner platform 96, and the airfoil 98 provide a single, integral, one-piece vane 26, 28 as shown in
A method of assembling the turbine vane assembly 10 may include several steps. The method may include arranging the first support spar 34 through the first turbine vane 26, arranging the second support spar 36 through the second turbine vane 28, and coupling the inner mount 64 to the inner ends 48 of the first support spar 34 and the second support spar 36.
In the illustrative embodiment, the coupling step includes arranging the inner end 48 of each support spar 34, 36 through corresponding apertures in the inner mount platform 74 and into the corresponding nozzles 70, 72, fixing the first fastener 66 to the inner end 48 of the first support spar 34, and fixing the second fastener 68 to the inner end 48 of the second support spar 36.
In the illustrative embodiment, the arranging step of the inner end 48 of the support spars 34, 36 includes engaging the interface surface 60 of each support spar 34, 36 with the raised interface surface 76 on the inner mount 64. The arranging step may also include engaging the anti-rotation peg 78 extends into the anti-rotation notch 62 in the strut 50 so as to align the exit holes 54 with the corresponding spout 88 of the nozzles 70, 72. In the illustrative embodiment, the fixing step of the fasteners 66, 68 includes mating the nut 90 with the threads of the inner end 48 of the support spar 34, 36 and arranging the pin 92 in the pin hole 93 in the inner end 48 of the support spar 34, 36.
Another embodiment of a turbine vane assembly 210 in accordance with the present disclosure is shown in
The turbine vane assembly 210 includes a plurality of turbine vanes 12, an outer vane support 214, and an inner vane support 216 as shown in
The inner vane support 216 includes an inner mount 264, a plurality of fasteners 266, 268, and a plurality of nozzles 270, 272 as shown in
The inner mount 264 includes an inner mount platform 274 and anti-rotation tabs 281 as shown in
In the illustrative embodiment, the plurality of fasteners 266, 268 includes a first nut 266 and a second nut 268 as shown in
Each nozzle 270, 272 includes a cylindrical tube 286, an anti-rotation notch 287, and a spout 288 as shown in
A method of assembling the turbine vane assembly 210 may include several steps. The method may include arranging the first support spar 234 through one of the turbine vanes 12, arranging the second support spar 236 through another turbine vane 12, and coupling the inner mount 264 to the inner ends 248 of the first support spar 234 and the second support spar 236.
In the illustrative embodiment, the coupling step includes arranging the inner end 248 of each support spar 234, 236 through corresponding apertures in the inner mount platform 274, arranging the first nozzle 270 on the inner end 248 of the first support spar 234, arranging the second nozzle 272 on the inner end 248 of the second support spar 236, fixing the first fastener 266 to the inner end 248 of the first support spar 234, and fixing the second fastener 268 to the inner end 248 of the second support spar 236.
In the illustrative embodiment, the arranging step of the nozzles 270, 272 on the support spars 234, 236 includes placing the cylindrical tube 286 over the inner end 248 of the support spar 234, 236 and aligning the anti-rotation notch 287 with the anti-rotation tab 281 of the inner mount platform 274.
Another embodiment of a turbine vane assembly 310 in accordance with the present disclosure is shown in
The turbine vane assembly 310 includes a plurality of turbine vanes 12, an outer vane support 314, and an inner vane support 316 as shown in
The outer vane support 314 includes a first support spar 334 and second support spar 336 as shown in
Each of the support spars 334, 336 include an inner end 348 and a strut 350 as shown in
The inner end 348 of each support spar 334, 336 is shaped to include a cooling air exit hole 354 as shown in
The strut 350 of each support spar 334, 336 is shaped to include inner interface surface 360 and bolt holes 362 as shown in
The inner vane support 316 includes an inner mount 364, a plurality of fasteners 366, 368, 369, 371, and a plurality of nozzles 370, 372 as shown in
The inner mount 364 includes an inner mount platform 374, raised interface surfaces 376, and bolt holes 378 as shown in
Each nozzle 370, 372 includes an attachment plate 377, a cylindrical tube 386 and a spout 388 as shown in
In the illustrative embodiment, each of the fasteners 366, 368, 369, 371 extend through the attachment plate 377 and the inner mount platform 374 of the inner mount 364 into the strut 350 of the corresponding support spar 334, 336 to couple the nozzle 370, 372 to the inner mount platform 374. The fasteners 366, 368, 369, 371 are also configured to act as anti-rotation features and block twisting of the inner vane support 316 relative to the outer vane support 314.
Another embodiment of a turbine vane assembly 410 in accordance with the present disclosure is shown in
The turbine vane assembly 410 includes a plurality of turbine vanes 412, an outer vane support 414, and an inner vane support 416 as shown in
The plurality of turbine vanes 412 includes a first turbine vane 426 and a second turbine vane 428 as shown in
Each turbine vane 426, 428 is shaped to include an outer platform 494, an inner platform 496, and an airfoil 498 as shown in
The outer vane support 414 includes an outer mount 432, a first support spar 434, and a second support spar 436 as shown in
The outer mount 432 includes an outer mount platform 440 and a plurality of reinforcement extensions 442, 444 as shown in
In the illustrative embodiment, the reinforcement collars 442, 444 may stiffen the outer mount platform 440 and minimize the compliance of the outer mount 432 and resulting deflections. In some embodiments, the reinforcement collars 442, 444 may help minimize the axial deflection of the turbine vane assembly 410. The reinforcement collars 442, 444 may also help minimize resulting stresses in the outer mount platform 440 due to the twisting of the turbine vane assembly 410.
Each of the support spars 434, 436 include an outer end 446, an inner end 448, and a strut 450 as shown in
Each of the support spars 434, 436 are also shaped to include a cooling channel 452 and an impingement channel 453 as shown in
Each of the impingement channels 453 is configured to supply a flow of cooling air to the vanes 426, 428 in the interior cavity 430 through impingement holes (not shown) in the support spar 434, 436. In some embodiments, the support spars 434, 436 may also be shaped to include impingement holes that extend from the cooling channel 452 and supply the flow of cooling air to the vanes 426, 428 in the interior cavity 430.
The inner end 448 of each support spar 434, 436 is shaped to include a cooling air exit hole 454 as shown in
The inner end 448 of each support spar 434, 436 is shaped to include a plurality of threads 455, 457 as shown in
The inner vane support 416 includes an inner mount 464, a plurality of fasteners 466, 468, 469, 471, and a plurality of nozzles 470, 472 as shown in
In the illustrative embodiment, the plurality of fasteners 466, 468, 469, 471 includes a first fastener 66, a second fastener 68, a third fastener 469, and a fourth fastener 471 as shown in
In the illustrative embodiment, the plurality of nozzles 470, 472 includes first nozzle 470 and a second nozzle 472 as shown in
In the illustrative embodiment, each of the fasteners 466, 468, 469, 471 are nuts as shown in
The inner mount 64 includes an inner mount platform 474, a first inner load transfer collar 473, and a second inner load transfer collar 475 as shown in
Another embodiment of a turbine vane assembly 510 in accordance with the present disclosure is shown in
The turbine vane assembly 510 includes a plurality of turbine vanes 12, an outer vane support 514, and an inner vane support 516 as shown in
The outer vane support 514 includes an outer mount 532, a first support spar 534, a second support spar 536, and a third support spar 538 as shown in
In the illustrative embodiment, the support spars 534, 536, 538 are integrally formed with the outer mount 532. The support spars 534, 536, 538 are integrally formed with the outer mount 532 to reduce the number of gaps.
The inner vane support 516 includes an inner mount 564 and a plurality of couplings 566, 568, 569 as shown in
Another embodiment of a turbine vane assembly 610 in accordance with the present disclosure is shown in
The turbine vane assembly 610 includes a plurality of turbine vanes 12, an outer vane support 614, and an inner vane support 616 as shown in
The outer vane support 614 includes at least two outer mounts 632, 633 and a plurality of support spars 634, 635, 636, 638 as shown in
The plurality of support spars 634, 635, 636, 638 includes a first support spar 634, a second support spar 635, a third support spar 636, and a fourth support spar 638 as shown in
In the illustrative embodiment, the first support spar 636 extends radially inward from the first outer mount 632 through one of the plurality of turbine vanes 12. The second support spar 635 extends radially inward from the first outer mount 632 through another one of the plurality of turbine vanes 12. The third support spar 636 extends radially inward from the second outer mount 633 through another one of the plurality of turbine vanes 12. The fourth support spar 638 extends radially inward from the second outer mount 633 through another one of the plurality of turbine vanes 12.
In other embodiments, the first, second, and third support spars 634, 635, 636 may extend radially inward from the first outer mount 632, while the fourth support spar 638 extends radially inward from the second outer mount 633. Similarly, the first support spar 634 may extend radially inward from the first outer mount 632, while the second, third, and fourth support spars 635, 636, 638 extend radially inward from the second outer mount 633.
The inner vane support 616 includes an inner mount 664 and a plurality of couplings 666, 668, 669, 671 as shown in
Another embodiment of a turbine vane assembly 710 in accordance with the present disclosure is shown in
The turbine vane assembly 710 includes a plurality of turbine vanes 12, an outer vane support 714, and an inner vane support 716 as shown in
The outer vane support 714 includes an outer mount 732 and a plurality of support spars 734, 735, 736, 738 as shown in
The plurality of support spars 734, 735, 736, 738 includes a first support spar 734, a second support spar 735, a third support spar 736, and a fourth support spar 738 as shown in
The inner vane support 716 includes at least two inner mounts 764, 765 and a plurality of couplings 766, 768, 769, 771 as shown in
In the illustrative embodiment, the first support spar 734 and the second support spar 735 are coupled to the first inner mount 764, while the third support spar 736 and the fourth support spar 738 are coupled to the second inner mount 765. In other embodiments, the first, second, and third support spars 734, 735, 736 may be coupled to the first inner mount 764, while the four support spar 738 is coupled to the second inner mount 765. Similarly, in other embodiments, the first support spar 734 may be coupled to the first inner mount 764, while the second, third, and fourth support spars 735, 736, 738 are coupled to the second inner mount 765.
Another embodiment of a turbine vane assembly 810 in accordance with the present disclosure is shown in
The turbine vane assembly 810 includes an outer vane support 814 and an inner vane support 816 as shown in
In the illustrative embodiments, the inner vane support 816 includes an inner mount platform 874 and mating features 876, 878 as shown in
The mating features 876, 878 mate with the corresponding support spar 834, 836 to block radial and circumferential movement of the metallic outer vane support 814 relative to the metallic inner vane support 816. In other embodiments, plastic deformation may be induced on the support spars, i.e. bending/distorting the support spars during assembly of the support spars with the inner vane support 816. To avoid distortion of the support spars 834, 836, the mating features 876, 878 engage the support spars 834, 836 to independently control the radial and circumferential locations i.e. radial on one spar 836, circumferential on the other spar 834.
In some embodiments, the mating features 876, 878 may be configured to plastically deform upon engagement with the support spars 834, 836. The mating features 876, 878 may plastically deform to lock the inner vane support 816 and the outer vane support 814 together.
In the illustrative embodiments, the inner vane support 816 includes the inner mount platform 874, the first mating feature 876, 877, the second mating feature 878, and a locking pin 879 as shown in
In the illustrative embodiments, the first mating feature 876, 877 includes a rotational stop 876 and a radial locator 877 as shown in
In the illustrative embodiment, a radially-inwardly facing surface 855 of the inner end 848 of the first support spar 834 abuts the outer surface 865 of the inner mount platform 874, while the inner end 848 of the second support spar 836 extends through the inner mount platform 874. The locking pin 879 extends through the inner mount platform 874 and into the surface 855 of the first support spar 834, blocking rotation of the inner vane support 816 relative to the outer vane support 814. The inner end 848 of the second support spar 836 extends through a hole 884 formed in the inner mount platform 874.
In the illustrative embodiment, the second mating feature 878 includes a plurality of bayonet notches 878 as shown in
In some embodiments, the second mating feature 878 may include a single notch 878 that receives a single locking tab 860. In other embodiments, the second mating feature 878 may include a different number of notches 878 with the same number of locking tabs 860.
In the illustrative embodiment, the notches 878 extend into the inner mount platform 874 so that the locking tabs 860 extend into the inner mount platform 874. In other words, the notches 878 extend partway into the outer surface 865 so that the locking tabs 860 are engage a surface located radially between the outer surface 865 and the inner surface 863.
In other embodiments, the notches 878 extend through both surfaces 865, 863 of the inner mount platform 874. In such embodiments, the inner end of the second support spar 836 extends through the inner mount platform 874 so that the bayonet notches 878 are exposed and open radially inward as shown. The locking tabs 860 may then engage an inner surface 863 of the inner mount platform 874 in the respective bayonet notches 878 to radially retain the outer vane support 814.
Turning again to the outer vane support 814, the outer vane support 814 includes an outer mount 832 and the plurality of support spars 834, 836 as shown in
Each of the support spars 834, 836 include an outer end 846, an inner end 848, and a strut 850 as shown in
The inner end 848 of the first support spar 834 is shaped to include a groove 861 as shown in
The inner end 848 of the second support spar 836 is shaped to include the locking tabs 860 as shown in
Once the turbine vanes 12 are assembled on the support spars 834, 836, the inner vane support 816 is assembled with the support spars 834, 836 of the outer vane support 814. To assemble the inner vane support 816 with the support spars 834, 836, the inner end 848 of second support spar 836 is inserted into a corresponding hole 884 formed in the inner mount platform 874. As the inner end 848 of the second spar 836 is inserted into the hole 884, the locking tabs 860 are aligned with the corresponding notches 878 in the inner mount platform 874.
A method of assembling the turbine vane assembly 10 may include several steps. Once the inner end 848 is inserted through the hole 884 so that the locking tabs 860 extend into the corresponding notches 878, the outer vane support 814 is rotated about a spar axis 831 of the second support spar 836. The outer vane support 814 is rotated until the rotational stop 876 engages the strut 850 of the first support spar 834 and the locking tabs 860 engage the inner surface 863 of the inner mount platform 874.
The locking tabs 860 engage with the inner mount platform 874 to form the bayonet fitting and block radial movement, while the rotational stop 876 engages the strut 850 of the first support spar 834 to block circumferential movement. In the illustrative embodiment, the rotational stop 876 engages a leading edge of the strut 850.
Another embodiment of a turbine vane assembly 910 in accordance with the present disclosure is shown in
The turbine vane assembly 910 includes an outer vane support 914 and an inner vane support 916 as shown in
In the illustrative embodiments, the inner vane support 916 includes an inner mount platform 974 that extends at least partway circumferentially about the axis and mating features (not shown) that mate with support spars 934, 936 included in the outer vane support 914 to radially locate the inner mount platform 974 relative to the outer vane support 914 and block rotation of the supports 914, 916 relative to each other.
In the illustrative embodiments, the inner vane support 916 further includes a locking pin 979 as shown in
The locking pin 979 extends into a circumferential side surface 967 of the inner mount platform 974 as shown in
The locking pin 979 is illustrated as a headed pin that may be installed with an interference fit on the head. In other embodiments, the locking pin 979 may be a larger pin. Alternatively, the interference fit may be on the length of the locking pin 979.
Another embodiment of a turbine vane assembly 1010 in accordance with the present disclosure is shown in
The turbine vane assembly 1010 includes an outer vane support 1014 and an inner vane support 1016 as shown in
In the illustrative embodiments, the retainer plate 1065 includes a first mating feature 1076, 1077, the second mating feature 1078, and a locking pin 1079 as shown in
In the illustrative embodiment, the retainer plate 1065 includes an endwall 1075 as shown in
The inner end 1048 of the first support spar 1034 is shaped to include a groove 1061 as shown in
The inner end 1048 of the second support spar 1036 is shaped to include the locking tabs 1060 as shown in
In some embodiments, the groove 1061 and/or the locking tabs 1060 may be discrete features fabricated in or onto the support spars 1034, 1036 before assembly with the inner vane support 1016. In other embodiments, the inner ends 1048 of the support spars 1034, 1036 may be inserted through the inner mount 1064 and the features 1060, 1061 on the spars 1034, 1036 may be fabricated after assembly. Fabricated the features after assembly may allow for easy repair or reuse of the outer vane support 1014.
Once the turbine vanes 12 are assembled on the support spars 1034, 1036, the inner vane support 1016 is assembled with the support spars 1034, 1036 of the outer vane support 1014. To assemble the inner vane support 1016 with the support spars 1034, 1036, the inner end 1048 of second support spar 1036 is inserted into a corresponding hole 1084 formed in the inner mount 1064. As the inner end 1048 of the second spar 1036 is inserted into the hole 1084, the locking tabs 1060 are aligned with the corresponding notches 1078 in the inner mount 1064.
Once the inner end 1048 is inserted through the hole 1084 so that the locking tabs 1060 extend into the corresponding notches 1078, the outer vane support 1014 is rotated about a spar axis 1031 of the second support spar 1036. The outer vane support 1014 is rotated until the rotational stop 1076 engages the strut 1050 of the first support spar 1034 and the locking tabs 1060 engage the inner surface 1063 of the inner mount 1064.
The locking tabs 1060 engage with the inner mount 1064 to form the bayonet fitting and block radial movement, while the rotational stop 1076 engages the strut 1050 of the first support spar 1034 to block circumferential movement. In the illustrative embodiment, the rotational stop 1076 engages a leading edge of the strut 1050.
The present disclosure relates to reducing the rotation of ceramic matrix composite airfoils and metallic support structures by mechanically linking adjacent structures. The reduction in rotation may be leveraged to reduce the secondary air system leakages and improve engine performance.
In some embodiments, a spar may be used to support a turbine vane and inner stage seal. The differing pressures in the cavities on either side of the inner stage seal may may result in an axial load on the spar. Additionally, the force loads applied to the vanes 12 by the hot gases in the gas path 18 may result in an axial component in addition to a circumferential component on the spar also. The present disclosure teaches a turbine vane assembly 10 for minimizing the deflections under these loads, in an effort to maximise sealing performance.
In the illustrative embodiments, the turbine vane assembly 10, 210, 310, 410, 510, 610, 710, 810, 910, 1010 includes discrete load transfer features between the support spars 34, 36, 234, 236, 334, 336, 434, 436, 534, 536, 538, 634, 635, 636, 638, 734, 735, 736, 738, 834, 836, 934, 936, 1034, 1036 and the turbine vanes 12, 412. In some embodiments, the support spars 34, 36, 234, 236, 334, 336, 434, 436, 534, 536, 538, 634, 635, 636, 638, 734, 735, 736, 738, 834, 836, 934, 936, 1034, 1036 may include discrete load transfer features that engage the turbine vane 12, 412 radially inward and/or outerward of the gas path 18.
In some embodiments, to assemble a turbine vane assembly within a gas turbine engine 110, the turbine vane assembly may be fabricated individually then introduced radially to the inner stage seal for fastener to the inner stage seal bird-mouth. Rotation of the turbine vane assembly may be used to properly engage seals. Once the turbine vane assembly is coupled to the inner stage seal, the sub-assembly is lowered into the turbine case 20 and restraint features e.g. hooks into the casing 20 are engaged.
In the illustrative embodiment, the outer mount platform 40 is coupled to the turbine case 20 with a plurality of rails that extend into corresponding features in the case 20. In other embodiments, the outer mount platform 40 may be shaped to include a plurality of hooks that couple the outer mount platform 40 to the case 20. The use of hooks may avoid introducing bending at interface between the outer mount platform 40 and the hook.
In some embodiments, the inner vane support 16 may be segmented resulting in a non-trivial rotation of the assembly. This could induce relative movement and challenge seal clearances. The present disclosure teaches a turbine vane assembly 10 that introduces a mechanical linkage to reduce to rotation of the structure 10, effectively creating a torsion box.
The mechanical linkage is formed by the outer vane support 14 and the inner vane support 16. The inner mount 64 of the inner vane support 16 may span the same number of turbine vanes 12 as the outer mount 32 of the outer vane support structure 14 as shown in
The arrangement of the mechanical linkage may be a balance of change in stiffness and/or deflection as a function of increasing span of turbine vanes 12. The mechanical linkage arrangement may also be a balance of part count in the gas turbine engine 110, the number of gaps between adjacent turbine vane assemblies 10, and the number of seals and amount of leakage between the assemblies 10.
The mechanical linkage arrangement may also be a balance of mechanical stresses as a result of unequal load sharing and relative movements. The arrangement of the mechanical linkage between the outer vane support 14 and the inner vane support 16 may be a balance of thermal stresses as a result of circumferential temperature gradients. In other embodiment, the mechanical linkage arrangement may also be a balance of the redundancy.
The mechanical linkage between the outer vane support 14 and the inner vane support 16 may be fastened or coupled with a range of embodiments. In some embodiments, the inner vane support 16 may be bolted to the outer vane support 14. In other embodiments, the inner vane support 16 and the outer vane support may be clamped together.
In other embodiments, the outer vane support 14 may be, interference fit with the inner vane support 16. In other embodiments, the inner vane support 16 and the outer vane support 14 may be bi-cast, welded, etc. No matter the fastener arrangement between the outer vane support 14 and the inner vane support 16, the fastener arrangement may minimize compliance (increased deflection) while easily permitting assembly/dis-assembly, introducing acceptable stresses and minimising part count/complexity.
In the illustrative embodiments, the inner vane support 16 includes a hollow passage and/or nozzle arrangement to direct cooling flow. In other embodiments, the inner vane support 16 does not include a passage or nozzle to permit a flow of cooling air. In embodiments, without the nozzles, the flow of cooling air may be transmitted from somewhere else in the gas turbine engine 110 to the cavity 56.
The wedge face of the inner mount 464 may be an axial segmentation as shown in
In the illustrative embodiments, of
In some embodiments, plastic deformation may be induced at spar assembly i.e. bending/distorting the two support spars. To avoid this distortion of the support spars, the radial and circumferential locations may be controlled independently i.e. radial on one spar, circumferential on the other.
In the illustrative embodiments, a bayonet fitting may be applied to one of the support spars 836 as shown in
In some embodiments, the turbine vane assembly 810 may include a cam feature. The cam feature may be configured to plastically deform and ‘lock’ the outer vane support 814 with the inner vane support 816. The other spar 834 will rotate and engage a rotational stop 876, which permits load transfer from inner vane support 816 to the first support spar 834.
In some embodiments, the turbine vane assembly 810 may include an engagement, or radial slot 861, with a shape that encourages plastic deformation e.g. angled slot to ‘lock’ the structure together. The bayonet fitting 889 may also reduced the part count and minimize the number of small parts that may likely fall into disc cavity if they become disengaged.
The mating features 876, 877, 878 also provide a fail-safe in that the load applied increases the engagement of the features. The redundant features may also provide added safety as more than one feature would need to fail for radial position of the inner mount platform 874 to be lost e.g. the bayonet fitting 889 on its own retains the inner mount platform 874 and permits load transfer.
To avoid constraining the ceramic vanes 12 to interface at the inner mount 1064, the functionality may be split between the inner mount 1064 and retainer plate 1065 whereby, the assembly of the turbine vane assembly 1010 may including (i) positioning the ceramic turbine vanes 12 and any seals onto spar structure 1014, (ii) loading the spar structure 1014 into assembly fixture to accurately position parts, (iii) dropping the inner mount 1064 on-top of spar structure 1014, (iv) installing the retainer plate 1065 on spar so that the locking tabs 1060 align with the bayonet notches 1078, and (v) rotating the retainer plate 1065 until the radial locator 1077 extends into the groove 1061.
The bayonet fitting may include a cam type feature that would plastically deform and ‘lock’ the structure together. In some embodiments, the radial locator 877 is the cam feature. In other embodiments, the surfaces of the notches 878, 1078 are angles to increase engagement with the locking tabs 860, 1060 and deform the locking tabs 860, 1060 to lock the structure together.
The other spar engages radially locates the assembly. Plastic deformation at interface may also ‘lock’ the structure together.
Clamping the inner mount 1064 radially between the support spars 1034, 1036 and the retainer plate 1065 allows the structure to transmit axial load through the spar interface and is anti-rotated through the pair of spars. To radially clamp the inner mount 1064, ramped radial clamp surfaces may be included in the notches 1078 to increase engagement on rotation of the retainer plate 1065. The locking tabs 1060 may be shaped to plastically deform and prevent relative movement between the retainer plate 1065 and the support spar 1034. In some embodiments, a ramped radial protrusion may extend radially outward from the retainer plate 1065 that plastically deforms against the inner mount 1064 to lock the components together. Furthermore, grooves may be added to prevent relative axial/circumferential movement if necessary.
In some embodiments, the inner mount 1064 may be pre-loaded on the chordal seals. This may be applied by applying a load between the retainer plate 1065 (radially located onto the spar) and the inner mount 1064 (able to slide radially). This feature my be configured to act like an inverted chordal clamp seal and may eliminate the need for a outer mounted sprung seal.
In other embodiments, springs (not shown) may be located in pockets 1069 of the retainer plate 1065 so engage the inner mount 1064 at the interference therebetween. Corresponding pockets (not shown) may be located in the inner mount 1064 to prevent the springs from escaping the assembly 1010. The height of the pockets may be greater than the expected relative thermal expansion mismatch to ensure that the spring engages on both sets of radial walls. Although the walls are illustrated as a simple pocket 1069, they may be aligned with the spar assembly vector to ensure even loading on the inner mount 1064. A large range of high temperature and creep resistant spring are conceivable.
Further retention features such as internal spigots may be added to locate and trap the springs. For example, a pin attached to the retainer plate 1065 may support a stack of belleville washers while the pin may be engaged in a blind hole in the inner mount 1064 that never disengages with thermal expansion. Alternatively, a wave spring may be located with an outer wall.
Advantageously, when the engine heats up, due to the relative thermal growths a gap would form between turbine vane 12 and inner mount 1064, which means that the stress on a spring feature pre-loading the inner mount 1064 into the spars 1034, 1036 may reduce with temperature, this is beneficial to the springs creep capability. In some embodiments, multiple springs may be introduced as a means of reducing the stress in each part and provide redundancy.
While the disclosure has been illustrated and described in detail in the foregoing drawings and description, the same is to be considered as exemplary and not restrictive in character, it being understood that only illustrative embodiments thereof have been shown and described and that all changes and modifications that come within the spirit of the disclosure are desired to be protected.
This application claims priority to and the benefit of U.S. Provisional Patent Application No. 62/912,950, filed 9 Oct. 2019, the disclosure of which is now expressly incorporated herein by reference.
| Number | Date | Country | |
|---|---|---|---|
| 62912950 | Oct 2019 | US |
