COMPOSITE LAYER SYSTEM HAVING AN ADDITIVELY MANUFACTURED SUBSTRATE AND A CERAMIC THERMAL PROTECTION SYSTEM
A composite layer system is presented. The composite layer system includes a metallic substrate, a structured surface, and a thermal protection system. The structured surface may be additively manufactured onto the metallic substrate and includes structured surface features formed to project above the metallic substrate. Each of the structured surface features are separated from adjacent structured surface features by grooves. The thermal protection coating may be thermally sprayed onto the structured surface and is bonded to each of the structured surface features.
Internal components of gas turbine engines, especially those in the hot combustion gas path, are exposed to temperatures of approximately 900 degrees Celsius or hotter. In order to make engines more efficient, the trend is to produce engines operating under higher firing temperatures. The engine internal components within the combustion path are often constructed of high temperature superalloys. These superalloy components often include cooling passages terminating on the component outer surface for passage of coolant fluid to cool the surfaces exposed to the hot combustion gases.
BRIEF SUMMARYIn one embodiment, a composite layer system is presented. The composite layer system includes a metallic substrate, a structured surface, and a thermal protection system. The structured surface may be additively manufactured onto the metallic substrate and includes structured surface features formed to project above the metallic substrate. Each of the structured surface features are separated from adjacent structured surface features by grooves. The thermal protection coating may be thermally sprayed onto the structured surface and is bonded to each of the structured surface features.
In another embodiment, a method of making a combustion turbine component within a gas turbine engine having a heat insulating outer surface for exposure to a combustion gas is presented. The method includes additively manufacturing a metallic substrate with a structure surface, additively forming the structured surface to include a plurality of structure surface features projection above an intermediate horizonal surface and separated by groove, thermally spraying a thermal barrier coating onto the structured surface, and segmenting the thermal barrier coating.
To easily identify the discussion of any particular element or act, the most significant digit or digits in a reference number refer to the figure number in which that element is first introduced.
The compressor section 102 is in fluid communication with an inlet section 122 to allow the gas turbine engine 100 to draw atmospheric air into the compressor section 102. During operation of the gas turbine engine 100, the compressor section 102 draws in atmospheric air and compresses that air for delivery to the combustor section 106. The illustrated compressor section 102 is an example of one compressor section 102 with other arrangements and designs being possible.
In the illustrated construction, the combustor section 106 includes a plurality of separate combustors that each operate to mix a flow of fuel with the compressed air from the compressor section 102 and to combust that air-fuel mixture to produce a flow of high temperature, high pressure combustion gas. The combustion gas (shown by the arrow) is conveyed by a transition 104 to a turbine section 108 of the engine, where thermal energy is converted to mechanical energy. Of course, many other arrangements of the combustor section 106 are possible.
The turbine section 108 includes a plurality of turbine stages 126 with each turbine stage 126 including a number of rotating turbine blades 128 and a number of stationary turbine vanes 118. The turbine stages 126 are arranged to receive the combustion gas from the transition 104 of combustor section 106 at a turbine inlet 130 and expand that gas to convert thermal and pressure energy into rotating or mechanical work. The turbine section 108 is connected to the compressor section 102 to drive the compressor section 102. For gas turbine engines, used for power generation or as prime movers, the turbine section 108 is also connected to a generator, pump, or other device to be driven. As with the compressor section 102, other designs and arrangements of the turbine section 108 are possible.
An exhaust section 132 is positioned downstream of the turbine section 108 and is arranged to receive the expanded flow of combustion gas from the final turbine stage 126 in the turbine section 108. The exhaust section 132 is arranged to efficiently direct the combustion gas away from the turbine section 108 to assure efficient operation of the turbine section 108. Many variations and design differences are possible in the exhaust section 132. As such, the illustrated exhaust section 132 is but one example of those variations.
Turbine engine internal components often incorporate a thermal barrier coating (TBC) of metal-ceramic material that is applied directly to the external surface of the component substrate surface or over an intermediate metallic bond coat that was previously applied to the substrate surface. The TBC provides an insulating layer over the component substrate, which reduces the substrate temperature. Combination of TBC application along with cooling channels in the component further lowers the substrate temperature.
Thermal Barrier Coatings, TBCs, are prone to cracking due to the extreme heat and exposure to combustion gases. The cracking, at times, can lead to spallation, i.e., the separation of the insulative material from the underlying substrate. Such cracking and spallation can increase the temperatures of the underlying substrate significantly resulting in premature damage and ultimately failure of the component.
For gas turbine components, the metallic substrate 202 may include a superalloy material such as Alloy 247 (CM 247 LC®), for example. Other superalloy materials that the metallic substrate 202 may comprise include IN718, IN738, IN939, Rene80, PWA1484, PWA1483, AMH282, and Stall5DE. These materials are especially prone to oxidation because typically these alloys include chromium which is known to oxidize. Typical materials for thermal barrier coatings may include a variety of ceramic materials such as 8YSZ (8 wt. % yttria-stabilized zirconia), 30-50 wt. % yttria stabilized zirconia, pyrochlores (such as Gd2Zr2O7), and bilayer 8YSZ/pyrochlore systems.
A bond coat 206 may be included as an intermediate layer between the metallic substrate 202 and the thermal barrier coating 204 as part of the thermal protection coating, as shown, in order to improve thermal barrier coating 204 adhesion to the metallic substrate 202. In the shown embodiment, the thermal barrier coating 204 includes a bilayer having two different ceramic materials. A top layer may include an abradable material such as 30-50 wt. % yttria stabilized zirconia or Ytterbia stabilized zirconia. The thermal barrier coating 204 generally protects the underlying metallic substrate 202 from extreme heat, such as a component in a gas turbine engine that is exposed to combustion gases. In the illustrated component, the exposure to the extreme heat has resulted in thermal barrier coating spall 208. In some instances, the spallations can increase the substrate temperature high enough to risk full wall oxidation.
The structured surface features 302 may help anchor the TBC layer and/or function as a barrier or wall in order to localize thermal stress or foreign object damage induced crack propagation within the TBC that might otherwise allow excessive TBC spallation and subsequent thermal exposure damage to the turbine component's underlying metallic substrate 202. For example, as seen, in
Additive manufacturing also enables structured surface features 302 having different cross-sectional profiles. For example,
In some embodiments of the composite layer system 300, the metallic substrate 202 includes a thermally spayed overlying thermal protection coating having an overall thermal protection coating thickness 402. While the disclosure references the thermal protection coating applied as thermally sprayed, in alternate embodiments, the protection coating may be vapor deposited or solution/suspension plasma sprayed or may also be a ceramic matrix composite layer. The thermal protection coating may be sprayed so that at least the grooves 308 include the coating. In other embodiments, the thermal protection coating may include a thickness above the tips of the structured surface features 302. Thus, a ratio of the structured surface feature height 404 to the overall thermal protection coating thickness 402 may vary. In certain embodiments, the structured surface features 302 have a projection height greater than 75 percent of the overall thermal protection coating thickness 402. For example, in
As another example, the structured surface features 302, exemplified as rods in
In an embodiment, the grid thickness, i.e., the spacing between the thermal barrier coating 204 segments, may be around 1000-6000 μmeters. A first overall height 606 of the individual segment of thermal barrier coating 204 may lie in a range of 0.5 to 3.5 mm where a second height 604 of the TBC segments above the structured surface features 302 may lie in a range of 0.25-1.5 mm The segmented TBC reduces tensile surface strain which increases the reliability of the TBC and minimizes spallation from the metallic substrate.
Vertical segmentation cracking may be introduced during the thermal spraying process when the ceramic protection coating system is sprayed onto the metallic substrate 202.
As seen in the embodiment shown in
In an alternate embodiment, vertical segmentation cracking may occur during a thermal cycling operation, i.e., during engine operation, which may subsequently occur after the thermal spraying process. A gas turbine component including the composite layer system 300 in the hot gas path i.e., the combustor section 106, and at least first two rows of blades and vanes in the turbine section 108 where the temperatures may reach 1200° C. or higher, may experience adaptive segmentation due to the exposure to hot combustion gas.
An additively manufactured third layer comprising a lattice structure 1002 may be built together with the metallic substrate 202 to abut a second surface 1006 of the composite layer system 300. In an embodiment, the second surface 1006 is a surface opposite the structured surface 304 of the metallic substrate layer. The additively manufactured third layer may comprise a lattice structure 1002. The lattice structure 1002 provides improved structural integrity so that the third layer is more strain compliant than a typical bulk, fully solid metallic substrate. Lattice structures have a higher strain to failure in compression compared to bulk metallic solids and also have a distributed damage propagation under tensile loads. These unique damage progression behaviors allow for tailoring the lattice structure designs to meet local strain requirements. In an embodiment, the lattice structure 1002 comprises a Gyroid lattice structure 1002. A gyroid lattice structure is an example of a continuous surface that is printable in a LPBF process independent of orientation. Other embodiments of the lattice structure may include a truss-based lattice structure.
In an embodiment, the hybrid three-layer system 1000 includes tunable mechanical properties so that each layer may be modified in thickness and density. By tuning, or selecting, one property such as thickness and density, another property of the layer may be optimized. For example, in an embodiment, a thickness of each layer may be modified. By modifying the thickness of the third layer, for example, the load carrying capability of the hybrid three-layer system 1000 may be tailored for local strain levels. By modifying the thickness of the metallic substrate layer 202, the in-wall cooling effectiveness may be modified. By modifying the height of the structured surface feature 302, and the overall thermal protection coating thickness 402, the overall thermal resistance of the hybrid three-layer system 1000 may be modified. In another embodiment, the metallic substrate layer 202 and the lattice layer 1002, which may be additively manufactured, may include depth varying material properties including, for example, density. Additionally, the shape of the lattice structure 1002 may be varied. Exemplary ways to modify mechanical properties include application of a plurality of layers, enabled by additive manufacturing, that vary slightly in composition. In this way, the hybrid three-layer system 1000 may be adapted to the varying temperatures, strain, and pressure loads placed upon the it locally for any specific component design.
Typically, the transition exit mouth 1102 is machined out of a superalloy plate, subtractively, resulting in a significant waste of materials. This exit mouth 1102 is then welded to the main body 1104 of the transition 104 and to a bracket for mounting to an inner casing of the gas turbine engine 100.
The inventors thus propose an advanced additively manufactured exit mouth 1102 design comprising the hybrid three-layer system 1000 as seen in
The metallic substrate 202, positioned in between the thermal protection coating and the lattice structure 1002, may include cooling channels 1008 for carrying cooling air.
In order to mitigate the effects of the hot combustion gas exposure to turbine vanes, currently different cooling schemes are utilized including impingement cooling, using vane inserts with internal ribs, and connecting the suction and pressure sides for structural integrity. The inventors thus propose using the composite layer system 300 with in-wall and transpiration cooling to cool the turbine vane 118.
In the illustrated embodiment of
In an embodiment of the proposed turbine airfoil, the structured surface features 302 include a height of between 1 and 3 mm enabling a thick, between 3-4 mm, thermal barrier coating 204 capable of high temperatures of approximately 1700-1800° C. As shown in
In an embodiment, as shown in
In an embodiment, the inventors propose an edge portion 1500 having a thinner transition exit mouth wall. In contrast to the conventional side wall thickness (t1), an AM exit mouth side wall comprising the composite layer system 300, shown on the right side of
In an alternate embodiment, the edge portion 1500 is a trailing edge of a turbine vane 118. The metallic substrate 202 may include structured surface features 302 in a range, for example, of 0.5 mm-1 mm, shorter than those in the pressure and suctions sides of the vane airfoil. The trailing edge may be tapered to a unique composite architecture. In order to maintain a thin geometry, for example, the composite layer system may include a total thickness of approximately 0.5-1 mm which may minimize aerodynamic impact on turbine efficiency. Due to the higher thermal resistance of the composite trailing edge, the local substrate temperatures can be lowered. While a couple of embodiments of the edge portion have been discussed in the disclosure, other embodiments of a composite component edge portion may also be possible wherein the increased effective TBC thickness of a composite layer system may result in different cooling channel designs resulting from topology optimization.
Cooling components in increasingly high temperatures in a gas turbine engine has proven to be challenging. Thermal barrier coatings (TBCs) are commonly used, however, they are prone to cracking which then may cause spallations of the coating. The proposed composite layer system including structured surface features projecting from the metallic substrate into the TBC provides a significant increase in reliability over conventional TBC coated metallic substrates without structured surface features, especially upon local TBC spallation.
Additive manufacturing enables the manufacturing of components that were difficult to manufacture using conventional manufacturing techniques. In many of the additive manufacturing processes, the components are manufactured layer by layer. With additive manufacturing, the available design space is significantly larger allowing for unique in-wall cooling channel designs, strain compliant lattice structures and high temperature capable protection systems. The hybrid three-layer system, where two of the three layers may be additively manufactured may also be tunable in thickness and density due to the large design space enabled by additive manufacturing, thus adapting to the varying temperature, strain, and pressure loads of all cooled gas turbine components.
For large gas turbines, size limitations of Laser Powder Bed Fusion (LPBF) machines can be a hurdle in realizing the full AM benefit, however. In these situations, integration of hybrid manufacturing process such as AM and joining, will maximize the AM benefit. For example, a transition, as proposed, having an AM transition exit mouth including the hybrid three-layer system, may be joined to a conventionally manufactured bonded panel transition main body.
Although an exemplary embodiment of the present disclosure has been described in detail, those skilled in the art will understand that various changes, substitutions, variations, and improvements disclosed herein may be made without departing from the spirit and scope of the disclosure in its broadest form.
None of the description in the present application should be read as implying that any particular element, step, act, or function is an essential element, which must be included in the claim scope: the scope of patented subject matter is defined only by the allowed claims. Moreover, none of these claims are intended to invoke a means plus function claim construction unless the exact words “means for” are followed by a participle.
Claims
1. A composite layer system, comprising:
- a metallic substrate;
- a structured surface additively manufactured onto the metallic substrate and including structured surface features formed to project above the metallic substrate, each of the structured surface features separated from adjacent structured surface features by grooves; and
- a thermal protection coating thermally sprayed onto the structured surface and bonded to each of the structured surface features.
2. The composite layer system of claim 1 wherein the plurality of structured surface features includes a profile shape selected from the group consisting of rounded, triangular, and rectangular.
3. The composite layer system of claim 1, wherein a projection height of each of the structured surface features is greater than 75 percent of an overall thermal protection coating thickness.
4. The composite layer system of claim 3, wherein a tip of each of the structured surface features extends to an outer surface of the composite layer system so that the projection height of each of the plurality of structured surface features is 100 percent of the overall thermal protection coating thickness.
5. The composite layer system of claim 1, wherein the structured surface includes a repeating three-dimensional arrayed pattern.
6. The composite layer system of claim 5, wherein the three-dimensional arrayed pattern is a hexagonal pattern.
7. The composite layer system of claim 1, wherein the thermal protection coating includes a thermal barrier coating (TBC) layer and a bond coat (BC) layer, the bond coat layer coupled to the metallic substrate.
8. The composite layer system of claim 7, wherein the thermal barrier coating is a bi-layer thermal barrier coating, wherein an inner TBC layer is applied over the bond coat layer and an outer TBC layer is applied over the inner TBC layer for exposure to a combustion gas.
9. The composite layer system of claim 8, wherein the inner TBC layer is an 8YSZ layer and the outer TBC layer is an abradable 48YSZ layer.
10. The composite layer system of claim 7, wherein the thermal barrier coating is segmented.
11. The composite layer system of claim 10, wherein the segmented thermal barrier coating comprises vertical segmentation cracking so that a crack aligns with a tip of each of the plurality of structured surface features and extends through the second layer to the outer surface of the composite layer system.
12. The composite layer system of claim 10, wherein the segmented thermal barrier coating comprises vertical segmentation gaps in the thermal barrier coating that segments the thermal barrier coating.
13. The composite layer system of claim 11, wherein the composite layer system comprises at least a portion of a gas turbine component having a heat insulating outer layer for exposure to a combustion gas.
14. The composite layer system of claim 13, wherein the structured surface includes a repeating three-dimensional arrayed pattern and wherein the arrayed pattern is a hexagonal arrayed pattern.
15. The composite layer system of claim 14, wherein the gas turbine component is a turbine vane insert.
16. The composite layer system of claim 15, wherein the composite layer system comprises a strip positioned on a leading edge of the turbine vane insert.
17. The composite layer system of claim 13, wherein the composite layer system comprises an edge portion of a gas turbine component having at least two outer walls that intersect in an edge.
18. The composite layer system of claim 17,
- wherein the gas turbine component is a turbine vane and the edge portion is the trailing edge of the turbine vane, and
- wherein the thickness of the edge is in a range of 0.5 mm to 1 mm
19. The composite layer system of claim 17,
- wherein the gas turbine component is a transition and the edge portion is a side wall of the exit mouth of a transition, and
- wherein a thickness of the edge is in a range of 4 mm to 5 mm.
20. A method for making a combustion turbine component within a gas turbine engine having a heat insulating outer surface for exposure to a combustion gas, comprising:
- additively manufacturing a metallic substrate with a structured surface;
- additively forming the structured surface to include a plurality of structured surface features projecting above an intermediate horizontal surface and separated by grooves;
- thermally spraying a thermal barrier coating onto the structured surface; and
- segmenting the thermal barrier coating.
21. The method of claim 20, the segmenting occurring during the thermal spraying such that vertical cracks align with a tip of each of the structured surface features extending through the thermal barrier coating to the outer surface of the combustion turbine component.
22. The method of claim 20, the segmenting occurring during a thermal cycling procedure such that vertical cracks align with a tip of each of the structured surface features extending through the thermal barrier coating to the outer surface of the combustion turbine component.
23. The method of claim 20, the segmenting occurring during a thermal cycling procedure on the gas turbine component by exposing the plurality of structured surface features to the hot combustion gas in excess of 1200° C. wherein the plurality of structured surface features oxidize and recede into the thermal barrier coating leaving vertical segmentation gaps.
Type: Application
Filed: May 29, 2020
Publication Date: Dec 1, 2022
Inventors: Ramesh Subramanian (Oviedo, FL), Lieke Wang (Oviedo, FL), Philip Clissold Howell (Ottobrunn)
Application Number: 17/754,356