Patent Application
20210025282
The present disclosure relates generally to airfoil assemblies for gas turbine engines, and more specifically to airfoils that comprise ceramic materials.
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.
An airfoil assembly for a gas turbine engine may include a ceramic matrix composite vane, a metallic support strut, and a metallic inner carrier. The ceramic matrix composite vane is adapted to interact with hot gases flowing around the airfoil assembly during use of the airfoil assembly. The ceramic matrix composite vane may include an outer platform that defines an outer boundary of a gas path, an inner platform spaced apart radially from the outer platform relative to an axis to define an inner boundary of the gas path, and an airfoil that extends radially between and interconnects the outer platform and the inner platform. The airfoil may be formed to define an interior cavity that extends radially through the airfoil. The metallic support strut may be located in the interior cavity formed in the airfoil. The metallic support strut may be configured to receive force loads applied to the ceramic matrix composite vane by the hot gases during use of the airfoil assembly.
The metallic inner carrier may be coupled with the metallic support strut and adapted to block the hot gases from flowing radially inward toward the axis. The metallic inner carrier may include a body, a tab, and a void. The body extends axially and circumferentially relative to the axis. The tab extends circumferentially away from a first end of the body and may be configured to be received in an adjacent metallic inner carrier and the void extends circumferentially into a second end of the body and may be configured to receive a portion of another adjacent metallic inner carrier to interlock the airfoil assembly with the adjacent metallic inner carriers to reduce twisting of the airfoil assembly during use of the airfoil assembly.
In some embodiments, the body has an outer surface and an inner surface spaced apart radially from the outer surface. The tab may be spaced apart radially from at least one of the outer surface and the inner surface.
In some embodiments, the body may have a fore surface and an aft surface spaced apart axially from the fore surface. The tab may be spaced apart axially from the aft surface.
According to another aspect of the present disclosure, an airfoil assembly for a gas turbine engine may include an airfoil, a support strut, and an inner carrier. The support strut may extend radially into the airfoil relative to an axis. The inner carrier may be coupled with the support strut to locate the inner carrier radially between the airfoil and the axis. The inner carrier may include a body, a tab that extends circumferentially away from the body, and a void that extends circumferentially into the body.
In some embodiments, the body has an outer surface and an inner surface spaced apart radially form the outer surface. The tab may be spaced apart radially from at least one of the outer surface and the inner surface. In some embodiments, the tab may be spaced apart radially from both the outer surface and the inner surface.
In some embodiments, the body has an outer surface and an inner surface spaced apart radially from the outer surface. The tab may be flush radially with at least one of the outer surface and the inner surface.
In some embodiments, the tab may be flush with one of the outer surface and the inner surface. The tab may be spaced apart radially from the other of the outer surface and the inner surface.
In some embodiments, the body includes a fore surface and an aft surface spaced apart axially from the fore surface. A surface of the tab may be flush with the fore surface. The void may extend axially into the fore surface partway into the body.
In some embodiments, the body may be formed to further include a seal slot that extends circumferentially into the body. In some embodiments, the tab includes a curvilinear surface.
According to another aspect of the present disclosure, an airfoil assembly for a gas turbine engine includes a first airfoil unit and a second airfoil unit. The first airfoil unit may include includes a first airfoil, a first support strut that extends radially into the first airfoil relative to an axis, and a first inner carrier coupled with the first support strut to locate the first inner carrier radially inward of the first airfoil. The second airfoil unit may be located circumferentially adjacent the first airfoil unit relative to the axis. The second airfoil unit may include a second airfoil, a second support strut that extends radially into the second airfoil, and a second inner carrier coupled with the second support strut to locate the second inner carrier radially inward of the second airfoil. The first inner carrier may be interlocked with the second inner carrier.
In some embodiments, the first inner carrier may include a first body, a first tab, and a first void. The first tab may extend circumferentially away from the first body. The first void may extend circumferentially into the first body.
In some embodiments, the second inner carrier includes a second body, a second tab, and a second void. The second tab may extend circumferentially away from the second body and into the first void. The second void may extend circumferentially into the second body. In some embodiments, at least one surface of the first tab may be curvilinear.
In some embodiments, the first body may include a fore face and an aft face spaced apart axially from the first face. The first tab may extend flush with one of the fore face and the aft face. The first tab may be spaced apart axially from the other of the fore face and the aft face.
In some embodiments, the first airfoil unit may include a first vane that includes the first airfoil, a first outer platform coupled with the first airfoil, and a first inner platform coupled with the first airfoil. The first inner platform may be located radially between the first outer platform and the first inner carrier.
In some embodiments, the second airfoil unit includes a second vane that includes the second airfoil, a second outer platform coupled with the second airfoil, and a second inner platform coupled with the second airfoil. The second inner platform may be located radially between the second outer platform and the second inner carrier. The first inner platform may extend circumferentially into the second inner platform to interlock the first inner platform with the second inner platform.
In some embodiments, the second airfoil unit may include a second vane that includes the second airfoil, a second outer platform coupled with the second airfoil, and a second inner platform coupled with the second airfoil. The second inner platform may be located radially between the second outer platform and the second inner carrier. The first outer platform may extend circumferentially into the second outer platform to interlock the first outer platform with the second outer platform. In some embodiments, the first inner carrier and the second inner carrier may overlap when viewed axially.
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.
An airfoil assembly 10 is shown in
The ceramic matrix composite vane 12 interacts with hot gases conducted through a gas path 24 of the gas turbine engine 110 and conducts the hot gases around the airfoil assembly 10 toward a rotating wheel assembly 22 located downstream of the airfoil assembly 10 as suggested in
The metallic inner carrier 16 of each airfoil unit 18 interlocks with circumferentially adjacent airfoil units 18 to reduce the rotation of the metallic inner carriers 16 caused by aerodynamic loading. Reducing the rotation of the inner carriers 16 may reduce secondary air system leakages and improve engine performance. By each inner carrier 16 engaging the inner carrier 16 of adjacent airfoil units 18, the reaction between the airfoil units 18 acts in the opposite direction to the circumferential aerodynamic load and reduces the net loading and net deflection of each of the inner carriers 16.
The airfoil 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 plurality of the static turbine vane rings 20 that are fixed relative to the axis 11 and a plurality of the bladed wheel assemblies 22 as suggested in
The ceramic matrix composite vane 12 is adapted to withstand high temperatures, but may have relatively low strength compared to the metallic support strut 14. The support strut 14 provides structural strength to the airfoil assembly 10 by receiving the force loads applied to the vane 12 and transferring them to a casing that surrounds the airfoil assembly 10. The support strut 14 and inner carrier 16 may not be capable of withstanding directly the high temperatures experienced by the vane 12.
In illustrative embodiments, the vane 12 comprises ceramic materials while the support strut 14 and inner carrier 16 comprise metallic materials. Illustratively, the vane 12 comprises ceramic matrix composite materials. In other embodiments, each of the vane 12, support strut 14, and inner carrier 16 may comprise any suitable materials including ceramics, ceramic matrix composites, metals, alloys, super alloys, etc.
The vane 12 of each airfoil assembly 10 includes an outer platform 26, an inner platform 28, and an airfoil 30 as shown in
The airfoil 30 is also formed to define an interior cavity 36 that extends radially into the airfoil 30 as shown in
In the illustrative embodiment, the outer platform 26, the inner platform 28, and the airfoil 30 of the vane 12 are integrally formed from ceramic matrix composite materials. As such, the outer platform 26, the inner platform 28, and the airfoil 30 provide a single, integral, one-piece vane 12 as shown in
The support strut 14 extends at least partway into the airfoil 30 as shown by the broken away portion of the vane 12 in
The support strut 14 extends radially between a first end 38 and a second end 40 as shown in
The support strut 14 is hollow in the illustrative embodiment. In some embodiments, the support strut 14 includes holes that extend through the support strut 14 to allow cooling air to pass through the hollow support strut 14 and flow into the interior cavity 36. In other embodiments, the support strut 14 is solid material between the first end 38 to the second end 40.
The inner carrier 16 is coupled with the support strut 14 to locate the inner platform 28 radially between the inner carrier 16 and the outer platform 26 as shown in
During operation of the gas turbine engine, force loads are applied to the airfoil assembly 10 as suggested in
The force loads applied to the components of the airfoil units 18 are applied axially and circumferentially and urge the airfoil units 18 to twist as suggested by arrow 34 in
The inner carrier 16 includes a body 50, a tab 52, and a void 54 as shown in
The body 50 includes an outer surface 56, an inner surface 58 spaced apart radially from the outer surface, a fore surface 60, an aft surface 62 spaced apart axially from the fore surface 60, a first circumferential surface 64, and a second circumferential surface 66 spaced apart circumferentially from the first circumferential surface 64 as shown in
The tab 52 of each inner carrier 16 overlaps with the body 50 of a neighboring inner carrier 16. The tab 52 overlaps with the body 50 of a neighboring inner carrier 16 when viewed axially such that an imaginary axially extending pin 44 would intersect neighboring inner carriers 16 as shown in
Illustratively, the tab 52 extends circumferentially away from the first circumferential surface 64 as shown in
The tab 52 is rectangular in the embodiment shown in
The tab 52 of the embodiment of
The tab 52 of the embodiment of
In some embodiments, the tab 52 overlaps with the body 50 of a neighboring inner carrier 16 such that the imaginary radially extending pin 46 would intersect both neighboring carriers 16 as shown in
As shown in
The body 50 is further formed to include a first seal slot 80 and a second seal slot 82 as shown in
In some embodiments, the tab 52 overlaps with the body 50 of a neighboring inner carrier 16 both radially and axially as suggested in
In some embodiments, the outer platform 26 and/or the inner platform 28 of the vane 12 interlock according to the features of the present disclosure as suggested in
The present disclosure provides features to reduce the rotation of ceramic matrix composite airfoil units 18 caused by aerodynamic loading. The reduction in rotation may be leveraged to reduce secondary air system leakages and improve engine performance.
The load from high pressure stage two nozzle guide vanes 12 is transmitted outboard to the high-pressure turbine casing via the support strut 14. The ceramic matrix composite airfoil 30 and/or support structure 14 is cantilevered from a radially outboard attachment 42 in the illustrative embodiment. In some embodiments, the strut 14 is used to transmit the axial load from the inter-stage seal to the high-pressure turbine casing 42. In addition, the strut 14 may be used to support the ceramic matrix composite airfoil 30. The inter-stage seal induces an axial load onto the strut 14, while the airfoil 30 induces an axial component in addition to a circumferential component due to the hot gases applying force load to the airfoil 30.
As a result of the circumferential loading, the structure may be urged to rotate. This may be pronounced if the metallic inner platform carrier 16 structure is segmented, which may occur to balance thermal growth mismatches if the strut 14 is rigidly connected to the high-pressure turbine casing 42. Such rotation could induce relative movement and increase leakages through seals at the inner end of the structure. Ceramic matrix composite materials, such as the ceramic matrix composite airfoil 30, are intended to reduce cooling flow requirements, but if the leakages are greater than an equivalent metallic design then this rotation may erode the fuel burn benefit of ceramic matrix composite parts.
The present disclosure introduces a mechanical linkage to reduce rotation of the structure. By engaging the adjacent part, the reaction acts in the opposite direction to the circumferential aerodynamic load and reduces the net loading and net deflection of every part.
The mechanical linkage may be sensitive to a number of criteria. For example, the load path created when adjacent parts engage. i.e. face contact on the full axial extent of the feature or just on the protruding circumferential faces. The shape and size of these features may be optimized to reduce overall movement, but avoid over-stressing the components, see
The linkage features may be applied to any ceramic matrix composite vane, blade, airfoil, seal segment, or other structure with loads that may result in twisting. Linkage features may be applied to the ceramic matrix composite platforms in addition to the linkage features in the inner carrier or instead of the linkage features in the inner carrier. The linkage features could be pre-formed as ceramic preforms of approximately the final size or pre-formed larger than desired and then machined to size after the infiltration process.
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.
