Ceramic matrix composite vane assembly with shaped load transfer features

A turbine vane assembly adapted for use in a gas turbine engine includes a spar, a turbine vane, and load transfer pins. The spar comprises metallic materials and is configured to support other components of the turbine vane assembly relative to an associated turbine case. The turbine vane comprises ceramic matrix composite materials and is shaped to include an airfoil configured to direct the flow of hot gasses through a primary gas path of the turbine vane assembly.

FIELD OF THE DISCLOSURE

The present disclosure relates generally to vane assemblies for gas turbine engines, and more specifically to vane assemblies that comprise ceramic matrix composite components.

BACKGROUND

Products of the combustion reaction directed into the turbine flow over airfoils included in stationary vanes and rotating blades of the turbine. The interaction of combustion products with the airfoils heats the airfoils to temperatures that require the airfoils to be made from high-temperature resistant materials and/or to be actively cooled by supplying relatively cool air to the vanes and blades. 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 required for the parts.

SUMMARY

A turbine vane assembly adapted for use in a gas turbine engine may include a metallic spar, a turbine vane, and load transfer pins. The metallic spar may extend radially inwardly relative to an associated turbine case arranged around a central reference axis. The metallic spar may be configured to support other components of the turbine vane assembly relative to the associated turbine case. The turbine vane comprising ceramic matrix composite materials may be supported by the metallic spar. The load transfer pins may couple the turbine vane to the metallic spar.

In some embodiments, at least one aperture in the turbine vane may be shaped to extend non-linearly relative the central axis. The shape of the aperture(s) may drive complex movement of the corresponding mount pin relative to the turbine vane due to different amounts of thermal growth caused by material mismatch during operation of the gas turbine engine at various temperatures.

In some designs, the shape of the aperture(s) may drive pre-load being applied to the turbine vane by the metallic support structure when the turbine vane assembly is at ambient temperatures. These preloads may be selected so as to manage or minimize the loading and deflection of the airfoil caused by aerodynamic loading of the turbine vane during use in the gas turbine engine.

In some embodiments, the aperture(s) may be shaped to have a radial and a circumferential component. In some designs, the aperture(s) may have an arcuate shape when viewed in the circumferential direction about the central reference axis.

DETAILED DESCRIPTION OF THE DRAWINGS

A turbine vane assembly10according to the present disclosure is adapted for use in a gas turbine engine as shown inFIGS.1-5. The turbine vane assembly includes a spar12, a turbine vane14, and load transfer pins16as shown inFIG.1. The spar12extends radially inwardly relative to an associated turbine case arranged around a central reference axis. The turbine vane14is shaped to define a passageway18extending radially therethrough and is supported by the metallic spar12that extends through the passageway18. The load transfer pins16couple the turbine vane14to the metallic spar12via apertures30,32that may be shaped to induce pre-load into the turbine vane14under certain conditions so as to address aero loads applied to the turbine vane14during use in the gas turbine engine.

The spar12comprises metallic materials while the turbine vane14comprises ceramic matrix composite materials. The ceramic matrix composite vane14is adapted to withstand high temperatures, but may have a relatively low strength compared the metallic spar12. The spar12is therefore configured to support other components of the turbine vane assembly10, like the ceramic matrix composite turbine vane14, relative to the associated turbine case and carry any aerodynamic loads from the turbine vane14.

Metallic materials of the spar12and ceramic matrix composite materials of the turbine vane14grow and shrink at different rates when exposed to high/low temperatures. More specifically, metallic materials have coefficients of thermal expansion much higher than those of ceramic matrix composites. To accommodate this difference, the turbine vane14is shaped to allow radial movement (sliding) of the corresponding load transfer pins16that couple the turbine vane14to the spar12during operation of the turbine vane assembly10at various temperatures.

The turbine vane14includes an airfoil20, panels22, and vane mounts24as shown inFIGS.1and4. The airfoil20is configured to direct the flow of hot gasses through a primary gas path25of the turbine vane assembly10. The panels22each extend circumferentially from the airfoil20away from the passageway18through the turbine vane14to define the primary gas path25across the turbine vane assembly10. The vane mounts24are each arranged radially outward/inward of the primary gas path25.

The vane mounts24include an outer vane mount26and an inner vane mount28as shown inFIG.4. The outer vane mount26is arranged radially outward of the primary gas path25and the inner vane mount28is arranged radially inward of the primary gas path25.

Each of the vane mounts26,28includes apertures30,32that receive the load transfer pins16as shown inFIGS.1-2and4-5. The outer vane mount26includes outer apertures30and the inner vane mount28includes inner apertures32. In the illustrative embodiments, one of the outer apertures30and the inner apertures32are sized to allow radial movement of the corresponding load transfer pins16to create the sliding support that transfers loading of the turbine vane14to the metallic spar12while accounting for materials mismatch of the metallic spar12and the turbine vane14during operation of the turbine vane assembly10at various temperatures.

The load transfer pins16include outer mount pins34and inner mount pins36as shown inFIGS.1-5. The outer mount pin34extends away from the spar12and through the outer aperture30in the outer vane mount26of the turbine vane14. The inner mount pin36that extends away from the spar12and through the inner aperture32in the inner vane mount28of the turbine vane14. In other embodiments, the load transfer pins16may also be used to couple the turbine vane14to other types of metallic support structures (inner collars, turbine case, flanges, etc.).

In the illustrative embodiment, the inner apertures32are shaped to allow radial movement of the corresponding inner mount pins36and to induce pre-load into the turbine vane14under certain conditions as suggested inFIGS.1and2. The outer apertures30are sized to block radial movement of the corresponding outer mount pins34so as to fix the turbine vane14in place radially relative to the metallic spar12at the outer mount pins34. In other embodiments, the outer apertures30may be sized to allow radial movement of the corresponding outer mount pins34and/or to contribute to pre-loading.

In the illustrative embodiment ofFIG.2, each of the inner apertures32sized to allow radial movement of the corresponding inner mount pins36has an enclosed arcuate shape. More generally, the inner apertures32have radial and circumferential components to their shape. In some embodiments, the inner apertures32can have axial components as part of the complex shape developed. The illustrated apertures32have an arcuate C-shape when viewed in the circumferential direction about the central reference axis. The enclosed arcuate shape sets limits on relative radial length change of the metallic spar12relative to the turbine vane14. In the illustrative embodiment ofFIG.2A, each of the inner apertures32sized to allow radial movement of the corresponding inner mount pins36is provided by a radially open-ended arcuate slot32.

The shape of the apertures32extends non-linearly relative the central axis. The shape of the apertures32drives complex, or multi-directional, movement of the corresponding mount pin16relative to the turbine vane14due to different amounts of thermal growth caused by material mismatch during operation of the gas turbine engine at various temperatures.

The shape of the apertures30,32may drive pre-load being applied to the turbine vane14by the metallic support structure when the turbine vane assembly is at ambient temperatures. These preloads may be selected so as to manage or minimize the loading and deflection of the airfoil caused by aerodynamic loading of the turbine vane14during use in the gas turbine engine.

The vane mounts24are also shaped to include a pressure side wall38, a suction side wall40, a leading edge42, and a trialing edge44as shown inFIG.3. The suction side40wall is spaced apart circumferentially from the pressure side wall38. The pressure side38wall and the suction side wall40extend axially between and interconnect the leading edge42and the trailing edge44.

In some embodiments, the apertures30,32may instead by blind features or blind holes. The blind holes may extend into the side walls38,40of the outer and inner vane mounts26,28, so that the pins16may engage the blind feature and the side walls38,40, not just the side walls38,40of the vane mounts26,28.

In the illustrative embodiments, the outer apertures30and the inner apertures32extend through one of the pressure side wall38and the suction side wall40and open into the passageway18. Each of the load transfer pins16extend through one of the pressure side wall38and suction side wall40of one of the outer vane mount26and the inner vane mount28and into the spar12.

In the illustrative embodiment, the outer vane mount26includes suction side apertures30and a pressure side aperture31as shown inFIGS.3and4. The suction side apertures30extend through the suction side wall40of the outer vane mount26. The pressure side aperture31extends through the pressure side wall38of the outer vane mount26.

In the illustrative embodiment, the outer vane mount26includes suction side apertures32and a pressure side aperture33as suggested inFIG.4The suction side apertures32extend through the suction side wall40of the inner vane mount28. The pressure side aperture33extends through the pressure side wall38of the inner vane mount28.

In the illustrative embodiment, the outer vane mount26and the inner vane mount28each include only three apertures30,32. In other embodiments, the mounts26,28may each include more than three apertures30,32. In other embodiments, the outer vane mount26may include a different number of apertures30than the inner vane mount28and vice versa.

The panels22include an outer panel46and an inner panel48as shown inFIGS.4and5. Each of the panels46,48extends circumferentially from the airfoil20away from the passageway18through the turbine vane14to define the primary gas path25across the turbine vane assembly10. The outer panel46is arranged radially inward of the outer vane mount26and the inner panel48is arranged radially outward of the inner vane mount28.

In the illustrative embodiment, the turbine vane14is a one-piece component. The airfoil20, the outer vane mount26, the inner vane mount28, the outer panel46, and the inner panel48are integrated into a single piece of ceramic matrix composite material that has been co-infiltrated with ceramic matrix material. In other embodiments, the airfoil20, the panels22, and the vane mounts24are formed as separate components.

Turning again to the spar12, the spar12is shaped to include a cooling air duct62, outer pin receivers64, and inner pin receivers66as shown inFIG.5. The cooling air duct62extends radially through the spar12. Each of the outer pin receivers64extend through the spar12at a radially outer end of the spar12and open into the air duct62. Each of the inner pin receivers66extend through the spar12at a radially inner end of the spar12and open into the air duct62. In other embodiments, the pin receivers64,66may only extend partway into the spar12.

In the illustrative embodiment, the spar12is arranged to extend radially through the passageway18of the vane14such that the outer and inner pin receivers64,66align with the outer and inner apertures30,32of the vane mounts26,28as shown inFIG.4. Each of the outer mount pins34extends through the outer aperture30of the outer vane mount26and into the outer pin receiver64of the spar12. Each of the inner mount pins36extends through the inner aperture32of the inner vane mount28and into the inner pin receiver66of the spar12.

In the illustrative embodiment, the outer and inner pin receivers64,66are sized to receive the outer and inner mount pins34,36so as to fix the pins16in place and couple the pins16to the metallic spar12. The outer mount pins34extend into the outer pin receiver64and inner pins36extend into the inner pin receivers66of the spar12to couple to the spar12.

It is also contemplated that the spar12may be omitted and that other metallic support structures may be used to couple the turbine vane14to the turbine case. In some embodiments where the spar is omitted, a collar support arrangement may be used to couple the turbine vane14to the turbine case. The load transfer pins16may extend through that collar support arrangement and engage the turbine vane mounts26,28of the turbine vane14to couple the vane14to the case and transfer the loads applied to the vane14to the turbine case. In other embodiments, an inner load transfer collar may be used to transmit forces at the radially inner load transfer pins36. The inner load transfer collar may be attached to the spar12or may be supported by interspersed metallic vanes.

In the illustrative embodiment, the load transfer pins16are made from metallic materials and are mechanically fastened or joined to the metallic spar12. In other embodiments, the load transfer pins16may be made of non-metallic materials to better match the ceramic matrix composite materials coefficient of thermal expansion and improve chemically compatibility.

Each of the load transfer pins16includes a through hole68as shown inFIGS.3and4. The through hole68extends through the length of the pin16. In the illustrative embodiment, the through hole68opens into the passageway18of the vane14.

In some embodiments, the through holes68may be used to supply cooling air to the cooling air duct62of the spar12and transit flow to feed other cavities with the cooling air. In other embodiments, the through holes68of the pins16may be used to cool the pins16and reduce the temperature of the contact between the ceramic vane14and the load transfer pins16. In some embodiments, the spar12may be also be formed to include one or more cooling air impingement holes to allow cooling air from a secondary air system to flow through the spar12and cool the ceramic matrix composite material of the vane14.

In the illustrative embodiments, the load transfer pins16are circular is shape when viewed in cross-section. In other embodiments, the pins16may be another suitable shape, such as ovular.

In some embodiments, some of the load transfer pins16may have a smaller diameter than the other load transfer pins16. The larger diameter pins16may be configured to datum the vane14and engage the apertures30,32upon assembly. The smaller diameter pins16may be configured to transfer the loading and engage upon loading of the vane14during use of the assembly10thereby driving orientation and/or stress in the vane14based on both pin16size and aperture30,32shape.

The present disclosure teaches a turbine vane assembly10that manages the coefficient of thermal expansion mismatch between the metallic support structure12and the ceramic matrix composite vane14. The turbine vane assembly10accounts for the relative thermal growth mismatch due to differences in coefficients of thermal expansion (CTE) between the ceramic matrix composite vane14and the metallic support structure12. More specifically, different radial growth of these components is addressed through the application of compliance through appropriately shaped sliding features16. In other ceramic matrix composite vane arrangements, this material mis-match can sometimes accounted for by using a spring or other compliant feature.

Variation in coefficients of thermal expansion between ceramic matrix composites and metallic spars necessitate features that allow for radial movement while still enabling transfer of aerodynamic loads from the ceramic matrix composite to the spar. These aero loads, particularly at higher temperatures, can result in undesirable deformation of the ceramic matrix composite that results in reduced aero performance. Aero loads, temperatures and axial elongation all tend to increase with increasing thrust and thus it is feasible to apply a preload to the ceramic matrix composite by controlling the shape of the slots30,32of the transfer feature16, and thus minimizing the aerodynamic deflection.

This effect can be exploited in isolation to minimise ceramic matrix composite stresses or in combination with other parts in the assembly for example with a Chordal clamp seal whereby, the stress in the metallic part is reduced, minimising creep concerns. Optionally, the sliding features16have apertures30,32shaped to tailor the orientation of ceramic matrix composite aerofoils20at different points across the flight cycle (i.e. during engine startup, takeoff, and cruise modes) to provide for better control of turbine capacity.

In the illustrative embodiments, the load from vanes14transmitted outboard to the high-pressure turbine casing. In metallic embodiments, the turbine vane assembly structure is supported on hooks and/or rails attached to the outer platform. However, the hook/rail design does not work for structures manufactured from the lower strength SiC/SiC ceramic matrix composite materials. Therefore, in some ceramic matrix composite vanes embodiments, a metallic spar is often incorporated in addition to conventional hook/rail joints.

The present disclosure teaches transferring the aerodynamic loading at both the inner and outer extents of the vane14. The load may be transmitted from the vane14through a metallic structure or spar12out into the casing. However, the ceramic matrix composite material has a significantly lower coefficient of thermal expansion compared to the high temperature metals of the spar12. Therefore, the components of the turbine vane assembly10has a significant mismatch in radial height between cold build and the hottest operating conditions.

In the illustrative embodiments, the turbine vane assembly10transfers the aerodynamic loading imparted on the ceramic matrix composite vane14through a series of load pins16arranged in apertures30,32. The apertures30,32may be arranged on the suction surface of the vane14and may sit on either side of the resultant aerodynamic load vector. The locations of the apertures30,32may be optimized to equally share loading at each aperture30,32.

The apertures30,32may also be arranged on pressure side of the vane14. The apertures30,32on the pressure side prevents rotation of the vane14,214,314and maintains stability of the structure12.

The location of the apertures30,32may be chosen such that each aperture30,32is always on the same side of the aerodynamic lift vector throughout the engine running range. In this way, there is no unloading or reversal of loading when the pressure distribution varies. Additionally, the distance from the vector to the apertures30,32may be small, but measureable as the larger it is, the larger the moment becomes and therefore, the larger the stress in the ceramic matrix composite vane14, and pin16.

In the illustrative embodiments, three load transfer pins16are used at each radial end of the vane14. In other embodiments, the vane assembly10may include more than three load pins16, especially if three of the pins16are used to datum the ceramic matrix composite vane14relative to the support structure12and transfer loading. The remaining pins16may be used solely to transfer loading. The sole load transfer pins16may be undersized pins16that would engage upon loading, while the datum pins16engage on assembly.

In the illustrative embodiments, the load transfer pins16may comprise alumina or other non-metallic material. The pins16may provide a friendly ceramic matrix composite load transfer feature.

In the illustrative embodiments, the pins16may also be hollow. The through holes68may be provided cooling flows to reduce the temperature of the contact between the pins16and the vane14.