US20200263557A1

US20200263557A1 - Turbine vane assembly with cooling feature

Info

Publication number: US20200263557A1
Application number: US16/279,469
Authority: US
Inventors: Michael J. WHITTLE; Steven HILLIER
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2019-02-19
Filing date: 2019-02-19
Publication date: 2020-08-20

Abstract

A turbine vane assembly adapted for use in a gas turbine engine includes a support and a turbine vane arranged around the support. The support is made of metallic materials. The turbine vane is made of ceramic matrix composite materials to insulate the metallic materials of the support.

Description

FIELD OF THE DISCLOSURE

The present disclosure relates generally to vane assemblies for gas turbine engines, and more specifically to vanes that comprise ceramic-containing materials.

BACKGROUND

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 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

The present disclosure may comprise one or more of the following features and combinations thereof.
A turbine vane assembly adapted for use in a gas turbine engine may include a vane made of ceramic matrix composite materials and a spar may be made of metallic materials. The spar may be spaced from the airfoil of the vane at all radial locations across the primary gas path such that a gap is maintained between the vane and the spar across the primary gas path.
In some embodiments, the vane may include an outer end wall, an inner end wall, and an airfoil. The inner end wall may be spaced radially inward of the outer end wall relative to a central reference axis to define a primary gas path therebetween. The airfoil may extend from the outer end wall to the inner end wall across the primary gas path.
In some embodiments, the spar may include a mount panel and a rod. The mount panel may be engaged with the vane at at least one location radially spaced from the primary gas path to receive aerodynamic loads from the vane. The rod may extend radially from the mount panel through a radially-extending passageway formed by the airfoil of the vane across the primary gas path.
In some embodiments, the spar may further include a plurality of heat transfer augmentation features. The plurality of heat transfer augmentation features may be arranged at radial locations along the primary gas path. The plurality of heat augmentation features may be configured to induce turbulence in cooling air supplied to the gap between the vane and the spar across the primary gas path during use of the turbine vane such that heat is more effectively transferred from the spar to the cooling air while avoiding conductive heat transfer from the ceramic matrix composite materials of the vane to the metallic materials of the spar that would be caused by contact between the vane and the spar across the primary gas path.
In some embodiments, the plurality of heat transfer augmentation features may include a plurality of protrusions. The plurality of protrusions may extend from the rod toward the airfoil of the vane.
In some embodiments, the plurality of protrusions may be located only along a side of the rod facing a pressure side of the airfoil. In some embodiments, the plurality of protrusions may be located only along a side of the rod facing a suction side of the airfoil.
In some embodiments, the plurality of heat transfer augmentation features may further include a plurality of through holes. The plurality of through holes may extend in a general circumferential direction, tangent to a circumference of the gas turbine engine with respect to the central reference axis through the rod in a portion of the rod, which is a third of a chord length of the rod.
In some embodiments, the plurality of heat transfer augmentation features may include a plurality of depressions. The plurality of depressions may extend inwardly into the rod of the spar.
In some embodiments, the plurality of depressions may be located only along a side of the rod facing a pressure side of the airfoil. In some embodiments, the plurality of depressions may be located only along a side of the rod facing a suction side of the airfoil.
In some embodiments, the spar may be formed to include a cooling air conduit. The cooling air conduit may extend from outside the primary gas path into the rod, which may be formed to include cooling air holes sized to discharge cooling air from the cooling air conduit into the gap between the vane and the spar along the primary gas path. In some embodiments, the cooling air holes may be arranged to discharge cooling air toward a leading edge of the airfoil included in the vane to provide some cooling to the vane.
According to an aspect of the present disclosure, a turbine vane assembly adapted for use in a gas turbine engine may include a vane and a spar. The vane may be made of ceramic matrix composite materials and may include an airfoil sized to extend in a radial direction relative to a central reference axis across a primary gas path of the gas turbine engine. The spar may be made of metallic materials and may be spaced from the airfoil of the vane at all radial locations across the primary gas path such that a gap is maintained between the vane and the spar across the primary gas path.
In some embodiments, the spar may include a mount panel, a rod, and a plurality of heat transfer augmentation features. The mount panel may be engaged with the vane at a location radially spaced from the primary gas path. The rod may extend radially from the mount panel through a radially-extending passageway formed by the airfoil of the vane across the primary gas path. The plurality of heat transfer augmentation features may be arranged at radial locations along the primary gas path and may be configured to induce turbulence in cooling air supplied to the gap.
In some embodiments, the plurality of heat transfer augmentation features include a plurality of protrusions. The plurality of protrusions may extend from the rod toward the airfoil of the vane.
In some embodiments, the plurality of protrusions may be located only along a side of the rod facing a pressure side of the airfoil. In some embodiments, the plurality of protrusions may be located only along a side of the rod facing a suction side of the airfoil.
In some embodiments, the plurality of heat transfer augmentation features may include a plurality of depressions. The plurality of depressions may extend inwardly into the rod of the spar.
In some embodiments, the plurality of depressions may located only along a side of the rod facing a pressure side of the airfoil. In some embodiments, the plurality of depressions may located only along a side of the rod facing a suction side of the airfoil.
In some embodiments, the spar may be formed to include a cooling air conduit. The cooling air conduit may extend from outside the primary gas path into the rod.
In some embodiments, the rod may be formed to include cooling air holes. The cooling air holes may be sized to discharge cooling air from the cooling air conduit into the gap between the vane and the spar along the primary gas path.
These and other features of the present disclosure will become more apparent from the following description of the illustrative embodiments.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a perspective view of a turbine vane assembly in accordance with the present disclosure showing the turbine vane assembly includes a metallic spar having a plurality of heat transfer augmentation features provided to encourage active cooling of the metallic spar and a ceramic matrix composite vane arranged around the spar such that the heat transfer augmentation features are located in a gap between the spar and the vane;

FIG. 2 is a cross sectional view of the turbine vane assembly of FIG. 1 taken along line 2-2 showing that the plurality of heat transfer augmentation features are arranged on both sides of the spar facing a suction side and pressure side of the vane;

FIG. 3 is a detail view of the heat transfer augmentation features included in the turbine vane assembly of FIG. 1 showing that the exemplary heat transfer augmentation features can include protrusions and holes;

FIG. 4 is an exploded view of the turbine vane assembly of FIG. 1 showing that the spar includes an outer mount panel and a rod and showing that the heat transfer augmentation features are arranged radially along the length of the rod;

FIG. 5 is a cross sectional view of the turbine vane assembly of FIG. 1 taken along line 5-5 showing that the heat transfer augmentation features extend radially along the length of the rod from an outer end wall to an inner end wall of the vane and suggesting the heat transfer augmentation features may also include flow separators to separate a flow of cooling air in the gap between the spar and the vane;

FIG. 6 is a detail end view of the heat transfer augmentation features of FIG. 4 showing that the protrusions can have an long ovular cross sectional shape;

FIG. 7 is a detail end view of the heat transfer augmentation features of FIG. 4 showing that the protrusions can have a circular cross sectional shape.

DETAILED DESCRIPTION OF THE DRAWINGS

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 illustrative turbine vane assembly 10 extends partway about a central axis for use in a gas turbine engine is shown in FIG. 1. The turbine vane assembly 10 includes a vane 12 and a spar 14 as shown in FIGS. 1-5. The vane 12 is made of ceramic matrix composite materials, while the spar 14 is made of metallic materials. The spar 14 provides structural support for the turbine vane assembly 10 and may be adapted for mounting in a ring or to a turbine case. In the illustrative embodiment, the spar 14 includes heat transfer augmentation features 40 that encourage cooling of the metallic spar 14 components that are adjacent to hot components of the vane 12. These heat transfer augmentation features 40 can include protrusions 50 (pins/fins), depressions 54, flow separators 53, holes 52, and other features that drive turbulence in cooling air moving along the spar 14 so that more heat can be withdrawn by the air as it moves along the spar 14.
The vane 12 defines a primary gas path 16 adapted to conduct hot gases during use of the turbine vane assembly 10 in a gas turbine engine. The vane 12 is arranged around at least a portion of the spar 14 so that the vane 12 insulates the metallic materials of the spar 14 from high temperatures in the primary gas path 16 defined through the turbine vane assembly 10. The spar 14 is spaced from an airfoil 24 of the vane 12 at all radial locations across the primary gas path 16 such that a gap 18 is maintained between the vane 12 and the spar 14 across the primary gas path 16.
The vane 12 includes an outer end wall 20, an inner end wall 22, and an airfoil 24 as shown in FIGS. 1, 4, and 5. The outer end wall 20 defines a radially outer boundary of the primary gas path 16 and the inner end wall 22 defines a radially inner boundary of the primary gas path 16. The inner end wall 22 is spaced radially inward of the outer end wall 20 relative to a central reference axis 11 to define the primary gas path 16 therebetween The outer end wall 20 shields an outer mount panel 34 of the spar 14 from the primary gas path 16 and the inner end wall 22 shields an inner mount panel 38 of the spar 14 from the primary gas path 16. The airfoil 24 is shaped to redirect air flowing through the gas turbine engine and shield a rod 36 of the spar 14 from the primary gas path 16. The airfoil 24 is also formed to define a radially-extending passageway 25.
In the illustrative embodiment, the outer end wall 20, inner end wall 22, and the airfoil 24 of the vane 12 are integrally formed from ceramic matrix composite materials such that the outer end wall 20, the inner end wall 22, and the airfoil 24 are included in a one-piece vane component 12 as shown in FIGS. 3-9. In other embodiments, the outer end wall 20, inner end wall 22, and the airfoil 24 may be formed as separate components.
The airfoil 24 includes an outer surface 26 and an interior surface 28 as shown in FIG. 2. The outer surface 26 faces the primary gas path 16 and extends between the outer end wall 20 and the inner end wall 22. The interior surface 28 is spaced apart from the outer surface 26 and defines the radially-extending passageway 25 that extends radially through the airfoil 24.
The outer surface 26 of the airfoil 24 defines a leading edge 30, a trailing edge 31, a pressure side 32, and a suction side 33 as shown in FIG. 2. The trailing edge 31 is axially spaced apart from the leading edge 30. The suction side 33 is circumferentially spaced apart from the pressure side 32. The pressure side 32 and the suction side 33 extend between and interconnect the leading edge 30 and the trailing edge 31.
The spar 14 includes an outer mount panel 34, a rod 36, and an inner mount panel 38 as shown in FIGS. 1, 4 and 5. The outer mount panel 34 is engaged with the vane 12 at least one location radially spaced from the primary gas path 16 to receive aerodynamic loads from the vane 12. The inner mount panel 38 is spaced radially inward from the outer mount panel 34 relative to the axis 11. The rod 36 extends radially inward from the outer mount panel 34 through the radially-extending passageway 25 formed by the airfoil 24 of the vane 12 across the primary gas path 16 and couples to the inner mount panel 38.
The spar 14 further includes a plurality of heat transfer augmentation features 40 and a cooling air conduit 42 as shown in FIGS. 1-5. The plurality of heat transfer augmentation features 40 are arranged at radial locations along the primary gas path 16. The heat transfer augmentation features 40 are configured to induce turbulence in cooling air supplied to the gap 18 between the vane 12 and the spar 14 across the primary gas path 16 during use of the turbine vane 10 such that heat is more effectively transferred from the spar 14 to the cooling air while avoiding conductive heat transfer from the ceramic matrix composite materials of the vane 12 to the metallic materials of the spar 14 that would be caused by contact between the vane 12 and the spar 14 across the primary gas path 16. The cooling air conduit 42 extends from outside the primary gas path into the outer mount panel 34 and the rod 36 and receives cooling air from a cooling air source.
The rod 36 includes an outermost surface 44 and cooling air holes 45 as shown in FIGS. 2-4. The outermost surface 44 faces the interior surface 28 of the airfoil 24 and is spaced apart from the airfoil 24 at all locations radially between the outer end wall 20 and the inner end wall 22 to define the gap 18 between the metallic spar 14 and the airfoil 24. The cooling air holes 45 are sized to discharge cooling air from the cooling air conduit 42 into the gap 18 between the vane 12 and the spar 14 along the primary gas path 16. The cooling air holes 45 are fluidly connected to the cooling air conduit 42 and are arranged to discharge cooling air toward the leading edge 30 of the airfoil 24 included in the vane 12 to provide some cooling to the vane 12.
In some embodiments, the rod 36 may further be shaped to include a radial rib 43 as suggested in FIG. 2. The radial rib 43 extends circumferentially outward from the outermost surface 44 toward the interior surface 28 of the airfoil 24. The rib 43 extends along the radial length of the rod 36 to form radial zones in the gap 18 between the airfoil 24 and the rod 36. The rib 43 is arranged in the gap 18 along the radial length of the spar 14 to bias the cooling air around either a suction side 48 or a pressure side 49 of the spar 14 after the cooling air has exited the cooling air holes 45. In some embodiments, the rib 43 may extend to, but not contact, the interior surface 28 of the airfoil 24 to form a small gap between the rib 43 and the interior surface 28 of the airfoil 24. In other embodiments, the rib 43 may contact the interior surface 28 of the airfoil 24.
The outermost surface 44 of the rod 36 is shaped to form a leading edge 46, a trailing edge 47, a suction side 48, and a pressure side 49 as shown in FIG. 2. The trailing edge 47 of the rod 36 is axially spaced apart from the leading edge 46 of the rod 36. The suction side 48 of the rod 36 is circumferentially spaced apart from the pressure side 49 of the rod. The suction side 48 and the pressure side 49 of the rod 36 extend between and interconnect the leading edge 46 and the trailing edge 47 of the rod 36. In the illustrative embodiment, the outermost surface 44 of the rod 36 is airfoil shaped and the rod 36 has a chord length. Additionally, the cooling air conduit 42 is formed in the rod 36 toward the leading edge 46 of the rod 36.
In the illustrative embodiment, the rib 43 extends from the outermost surface 44 of the rod 36 at a point spaced apart from the leading edge 46 of the rod 36. The rib 43 is located on the suction side 48 of the rod 36 in the illustrative embodiment. In other embodiments, the rib 43 is located on the pressure side 49 of the rod 36. In some embodiments, the rib 43 may be located at the leading edge 46 of the rod 36.
The plurality of heat transfer augmentation features 40 include a plurality of protrusions 50 and a plurality of through holes 52 as shown in FIGS. 1-5. The plurality of protrusions 50 extend from the outermost surface 44 of the rod 36 toward the interior surface 28 of the airfoil 24 of the vane 12. The plurality of protrusions 50 are spaced apart from the interior surface 28 of the airfoil 24 in the illustrative embodiment. The protrusions 50 are spaced apart radially along the radial length of the rod 36 in the primary gas path 16 and along the chord length of the rod 36 starting at a point spaced apart from the leading edge 46 of the rod 36. The plurality of through holes 52 extend through the rod 36 in a general circumferential direction, tangent to a circumference of the gas turbine engine. The holes 52 are formed in a portion of the rod 36 toward the trailing edge 47 of the rod 36. In the illustrative embodiment, the portion of the rod 36 including the holes 52 is about a third the chord length of the rod 36. In other embodiments, the portion of the rod 36 including the holes 52 is about half the chord length of the rod 36.
In the illustrative embodiment, the plurality of protrusions 50 are located along the sides 48, 49 of the rod 36 facing the pressure side 32 and the suction side 33 of the airfoil 24. In other embodiment, the protrusions 50 are located only along the side 49 of the rod 36 facing the pressure side 32 of the airfoil 24. In other embodiments, the protrusions 50 are located only along the side 48 of the rod 36 facing the suction side 33 of the airfoil 24.
In the illustrative embodiment, the spacing between each of the plurality of protrusions 50 exponentially decreases moving along the chord length of the rod 36 from the leading edge 46 to the trailing edge 47 of the rod 36. In other embodiments, the protrusions are located in discreet spaced-apart, increased-frequency patches over the rod 36 to increase heat transfer at predetermined locations associated with expected hot spots on either the airfoil 24 or the spar 14.
In the illustrative embodiment, the protrusions 50 have an oblong cross-sectional shape as shown in FIG. 6. In other embodiments, the protrusions 50 have a circular cross-sectional shape as shown in FIG. 7. In other embodiments, the protrusions 50 may have any other suitable shape.
In some embodiments, the plurality of heat transfer augmentation features 40 may further include a plurality of flow separators 53 as suggested in FIG. 5. The flow separators 53 are located in the gap 18 between the rod 36 and the airfoil 24. The flow separators 53 extend circumferential outward from the outermost surface of the rod 36 toward the interior surface 28 of the airfoil 24 and extend axially along the chord length of the rod 36 from the leading edge 46 to the trailing edge 47 of the rod 36 on the suction side 48 and/or the pressure side 49 of the rod 36. The flow separators 53 may be radially spaced apart from one another along the radial length of the rod 36 to measure and segregate the flow of cooling air at multiple radial heights along the radial length of the rod 36. In other embodiments, only one flow separator 53 may be located in the gap 18.
In some embodiments, the plurality of heat transfer augmentation features 40 may further include a plurality of depressions 54 as suggested in FIGS. 2 and 5. The plurality of depressions 54 may extend inwardly into the rod 36 of the spar 14. As suggested in FIG. 2, the plurality of depressions 54 may be located along the sides 48, 49 of the rod 36 facing both the pressure side 32 and the suction side 33 of the airfoil 24. In other embodiments, the depressions 54 may be located only along the side 49 of the rod 36 facing the pressure side 32 of the airfoil 24. In other embodiments, the depressions 54 may be located only along the side 48 of the rod 36 facing the suction side 33 of the airfoil 24.
Similar to the protrusions, the spacing between each of the plurality of depressions 54 may exponentially decrease moving along the chord length of the rod 36 from the leading edge 46 to the trailing edge 47 of the rod 36. In other embodiments, the depressions 54 may be located in discreet spaced-apart, increased-frequency patches over the rod 36 to increase heat transfer at predetermined locations associated with expected hot spots.
Turning again to the airfoil 24, the airfoil 24 includes a radial outer end 60, a radial inner end 62, and a body 64 as shown in FIG. 4. The radial outer end 60 extends radially-outwardly past the outer end wall 20 outside the primary gas path 16. The radial inner end 62 is spaced apart from the radial outer end 60 relative to the axis 11 and extends radially-inwardly past the inner end wall 22 outside the primary gas path 16. The body 64 extends radially entirely between and interconnects the radial outer end 60 and the radial inner end 62.
The radial outer end 60 of the airfoil 24 is shaped to form a load transfer tab 66 as shown in FIG. 5. The load transfer tab 66 is located outside of the primary gas path 16 and extends from the vane 12 to the spar 14. The load transfer tab 66 is contacted by a load transfer rib 68 of the spar 14 to transfer loads applied to the vane 12 to the spar 14 at the radial outer end 60 of the airfoil 24 outside of the primary gas path 16.
The outer mount panel 34 of the spar 14 is shaped to include a load transfer rib 68 as shown in FIG. 5. The load transfer rib 68 is located on a radial inward side of the outer mount panel 34 and outside of the primary gas path 16. The load transfer rib 68 is configured to carry circumferential and axial loads applied to the vane 12 to the spar 14 at the radial outer end 60 of the vane 12 outside of the primary gas path 16 where the temperature of the turbine vane assembly 10 is lower.
The present disclosure relates to a turbine vane assembly 10 with increased heat transfer coefficient within the ceramic matrix composite (CMC) internal cavity 25 of the airfoil 24. The increased heat transfer coefficient also increases the cooling effectiveness and reduces the CMC temperature without consuming additional cooling flows.
In many metallic vanes designs, the metallic vanes do not need sparred supports, and therefore do not require CMC cooling. However, the CMC cooling requirements will depend on their material temperature capability and engine cycle design. In some embodiments, the spar 14 may be coated in a low conductivity thermal barrier coating to reduce heat transfer. In other embodiments, the spar 14 may be made of a capable material on the external surface.
The turbine vane 10 may be configured to support other gas turbine engine components, such as an inter-stage seal. Accordingly, an application of cooling flows may be used to maintain an acceptable temperature between the turbine vane assembly 10 components 12, 14 so that the structural strength of the materials is maintained and may support the other gas turbine engine components, such as the inter-stage seal.
The present disclosure relates to the use of augmentation features 40 applied to the spar 14 to increase the heat transfer coefficient of the coolant. The potential application zones are illustrated in FIGS. 1-4.
The spar 14 is likely to be a cast feature therefore, the augmentation 40 may take several forms. In some embodiments, a pedestal/pin fin banks may be used for the augmentation features. In other embodiments, turbulators/ribs or chordal ribs/fins may be used as shown in FIGS. 6 and 7. Alternatively, augmentation could be machined into the spar 14 structure.
The augmentation features 40 may be applied generally to the pressure and suction sides of the metallic spar 14. Or alternatively, the augmentation features 40 may be applied to discrete regions that require an increased level of cooling relative to the surrounding material. In some embodiments, a leading edge flow separation may be arranged at the leading edge of the spar 14.
CMC materials are relatively low strength and are sensitive to thermal gradients. Rather than minimizing the reduction in CMC temperature, the augmentation features 40 may be applied to manage the thermal gradients i.e. increase the heat transfer coefficient (HTC) of the coolant locally in line with the external heat transfer coefficient distribution to drive the CMC material to become more isothermal.
Alternatively, the HTCs may be increased by application of increased roughness of the CMC on the internal structure of the CMC vane. The roughness may be increased by application of discrete ribs and/or features or specific CMC fabric architectures that offer increased roughness. A surface roughness above 10 Micron Ra can provide augmentation features 40 and may result in a suitable increase in heat transfer coefficient and any value above this will further increase the HTC.
The augmentation features 40 may be applied as a cast feature. In some designs, the augmentation features 40 may be machined in the surface and/or applied to the spar 14 by other machining methods, such as welding, brazing, bonding, or by another means of additive layer manufacture. In other embodiments, no additional coolant would be required to further reduce wall temperatures. In each embodiment, the application of the features may be tailored to influence local hot spots or overall surfaces. In other embodiments, the features may also be applied at CMC vane platform interfaces 20, 22.
In some embodiments, the cooling augmentation features 43 may be arranged to segregate the CMC-spar cavity. The ribs 43 may form radial ‘zones’ in the cavity and/or radial ribs 43 to bias cooling air after the cooling air has exited the leading edge impingement holes 45 to move around the pressure or suction side cavities. The ribs 3 may form a very small gap (or no gap) between the metal rib and the CMC wall to introduce a large pressure drop if the fluid were to cross the rib and therefore the ribs 43 would bias the flows in other directions.
The turbine vane assembly 10 may include a number of different features in combination to achieve the desired cooling scheme and/or material temperature profile. As suggested in FIGS. 2 and 4, the augmentation features 40 also include more exampled of cooling augmentation features 40, such as ribs 43, axial flow separators 53, depressions 54, and etc.
While the disclosure has been illustrated and described in detail end 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.

Claims

What is claimed is:

1. A turbine vane assembly adapted for use in a gas turbine engine, the assembly comprising

a vane made of ceramic matrix composite materials, the vane including an outer end wall, an inner end wall spaced radially inward of the outer end wall relative to a central reference axis to define a primary gas path therebetween, and an airfoil that extends from the outer end wall to the inner end wall across the primary gas path,

a spar made of metallic materials that is spaced from the airfoil of the vane at all radial locations across the primary gas path such that a gap is maintained between the vane and the spar across the primary gas path, the spar including a mount panel engaged with the vane at at least one location radially spaced from the primary gas path to receive aerodynamic loads from the vane and a rod that extends radially from the mount panel through a radially-extending passageway formed by the airfoil of the vane across the primary gas path,

wherein the spar further includes a plurality of heat transfer augmentation features arranged at radial locations along the primary gas path that are configured to induce turbulence in cooling air supplied to the gap between the vane and the spar across the primary gas path during use of the turbine vane such that heat is more effectively transferred from the spar to the cooling air while avoiding conductive heat transfer from the ceramic matrix composite materials of the vane to the metallic materials of the spar that would be caused by contact between the vane and the spar across the primary gas path.

2. The assembly of claim 1, wherein the plurality of heat transfer augmentation features include a plurality of protrusions that extend from the rod toward the airfoil of the vane.

3. The assembly of claim 2, wherein the plurality of protrusions are located only along a side of the rod facing a pressure side of the airfoil.

4. The assembly of claim 2, wherein the plurality of protrusions are located only along a side of the rod facing a suction side of the airfoil.

5. The assembly of claim 2, wherein the plurality of heat transfer augmentation features further include a plurality of through holes that extend in a general circumferential direction, tangent to a circumference of the gas turbine engine with respect to the central reference axis through the rod in a portion of the rod which is a third of a chord length of the rod.

6. The assembly of claim 1, wherein the plurality of heat transfer augmentation features include a plurality of depressions that extend inwardly into the rod of the spar.

7. The assembly of claim 6, wherein the plurality of depressions are located only along a side of the rod facing a pressure side of the airfoil.

8. The assembly of claim 6, wherein the plurality of depressions are located only along a side of the rod facing a suction side of the airfoil.

9. The assembly of claim 6, wherein the plurality of heat transfer augmentation features further include a plurality of through holes that extend in a general circumferential direction, tangent to a circumference of the gas turbine engine with respect to the central reference axis through the rod in a portion of the rod which is a third of a chord length of the rod

10. The assembly of claim 1, wherein the spar is formed to include a cooling air conduit that extends from outside the primary gas path into the rod which is formed to include cooling air holes sized to discharge cooling air from the cooling air conduit into the gap between the vane and the spar along the primary gas path.

11. The assembly of claim 10, wherein the cooling air holes are arranged to discharge cooling air toward a leading edge of the airfoil included in the vane to provide some cooling to the vane.

12. A turbine vane assembly adapted for use in a gas turbine engine, the assembly comprising

a vane made of ceramic matrix composite materials, the vane including an airfoil sized to extend in a radial direction relative to a central reference axis across a primary gas path of the gas turbine engine,

a spar made of metallic materials that is spaced from the airfoil of the vane at all radial locations across the primary gas path such that a gap is maintained between the vane and the spar across the primary gas path, the spar including (i) a mount panel engaged with the vane at a location radially spaced from the primary gas path, (ii) a rod that extends radially from the mount panel through a radially-extending passageway formed by the airfoil of the vane across the primary gas path, and (iii) a plurality of heat transfer augmentation features arranged at radial locations along the primary gas path configured to induce turbulence in cooling air supplied to the gap.

13. The assembly of claim 12, wherein the plurality of heat transfer augmentation features include a plurality of protrusions that extend from the rod toward the airfoil of the vane.

14. The assembly of claim 13, wherein the plurality of protrusions are located only along a side of the rod facing a pressure side of the airfoil.

15. The assembly of claim 13, wherein the plurality of protrusions are located only along a side of the rod facing a suction side of the airfoil.

16. The assembly of claim 12, wherein the plurality of heat transfer augmentation features include a plurality of depressions that extend inwardly into the rod of the spar.

17. The assembly of claim 16, wherein the plurality of depressions are located only along a side of the rod facing a pressure side of the airfoil.

18. The assembly of claim 16, wherein the plurality of depressions are located only along a side of the rod facing a suction side of the airfoil.

20. The assembly of claim 12, wherein the spar is formed to include a cooling air conduit that extends from outside the primary gas path into the rod which is formed to include cooling air holes sized to discharge cooling air from the cooling air conduit into the gap between the vane and the spar along the primary gas path.