Patent Grant
10196920
This application claims priority under the following U.S. Patent Applications, the entire contents of each of which is incorporated by reference herein:
“TURBINE ABRADABLE LAYER WITH PROGRESSIVE WEAR ZONE HAVING A FRANGIBLE OR PIXELATED NIB SURFACE”, filed Feb. 25, 2014, and assigned U.S. Ser. No. 14/188,941; and
“TURBINE ABRADABLE LAYER WITH PROGRESSIVE WEAR ZONE MULTI LEVEL RIDGE ARRAYS”, filed Feb. 25, 2014, and assigned U.S. Ser. No. 14/188,958.
A concurrently filed International Patent Application entitled “TURBINE ABRADABLE LAYER WITH AIRFLOW DIRECTING PIXELATED SURFACE FEATURE PATTERNS”, PCT/US15/16271, and assigned serial number (unknown) is identified as a related application and is incorporated by reference herein.
The following United States Patent Applications are identified as related applications for purposes of examining the presently filed application, the entire contents of each of which is incorporated by reference herein:
“TURBINE ABRADABLE LAYER WITH PROGRESSIVE WEAR ZONE TERRACED RIDGES”, filed Feb. 25, 2014 and assigned U.S. Ser. No. 14/188,992;
“TURBINE ABRADABLE LAYER WITH PROGRESSIVE WEAR ZONE MULTI DEPTH GROOVES”, filed Feb. 25, 2014 and assigned U.S. Ser. No. 14/188,813;
“TURBINE ABRADABLE LAYER WITH ASYMMETRIC RIDGES OR GROOVES”, filed Feb. 25, 2014 and assigned Ser. No. 14/189,035;
“TURBINE ABRADABLE LAYER WITH ZIG-ZAG GROOVE PATTERN”, filed Feb. 25, 2014 and assigned Ser. No. 14/189,081; and
“TURBINE ABRADABLE LAYER WITH NESTED LOOP GROOVE PATTERN”, filed Feb. 25, 2014 and assigned Ser. No. 14/189,011.
The invention relates to combustion or steam turbine engines having thermal barrier coating (TBC) layers on its component surfaces that are exposed to heated working fluids, such as combustion gasses or high-pressure steam, including individual sub-components that incorporate such thermal barrier coatings. The invention also relates to methods for reducing crack propagation or spallation damage to such turbine engine component TBC layers that are often caused by engine thermal cycling or foreign object damage (FOD). More particularly, various embodiments described herein relate to formation of engineered groove features (EGFs) within the thermal barrier coating (TBC). The EGFs are advantageously aligned with likely stress zones within the TBC or randomly aligned in a convenient two-dimensional or polygonal planform pattern on the TBC surface and into the TBC layer. The EGFs localize thermal stress- or foreign object damage (FOD)-induced crack propagation within the TBC that might otherwise allow excessive TBC spallation and subsequent thermal exposure damage to the turbine component underlying substrate.
Known turbine engines, including gas/combustion turbine engines and steam turbine engines, incorporate shaft-mounted turbine blades circumferentially circumscribed by a turbine casing or housing. The remainder of this description focuses on applications within combustion or gas turbine technical application and environment, though exemplary embodiments described herein are applicable to steam turbine engines. In a gas/combustion turbine engine hot combustion gasses flow in a combustion path that initiates within a combustor and are directed through a generally tubular transition into a turbine section. A forward or Row 1 vane directs the combustion gasses past successive alternating rows of turbine blades and vanes. Hot combustion gas striking the turbine blades cause blade rotation, thereby converting thermal energy within the hot gasses to mechanical work, which is available for powering rotating machinery, such as an electrical generator.
Engine internal components within the hot combustion gas path are exposed to combustion temperatures approximately 900 degrees Celsius (1600 degrees Fahrenheit). The engine internal components within the combustion path, such as for example combustion section transitions, vanes and blades are often constructed of high temperature resistant superalloys. Blades and vanes often include cooling passages terminating in cooling holes on component outer surface, for passage of coolant fluid into the combustion path.
Turbine engine internal components often incorporate a thermal barrier coat or 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 (BC) 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 passages in the component further lowers the substrate temperature.
Due to differences in thermal expansion, fracture toughness and elastic modulus—among other things—between typical metal-ceramic TBC materials and typical superalloy materials used to manufacture the aforementioned exemplary turbine components, there is potential risk of cracking the TBC layer as well as TBC/turbine component adhesion loss at the interface of the dissimilar materials. The cracks and/or adhesion loss/delamination negatively affect the TBC layer structural integrity and potentially lead to its spallation, i.e., separation of the insulative material from the turbine component. For example, vertical cracks developing within the TBC layer can propagate to the TBC/substrate interface, and then spread horizontally. Similarly, horizontally oriented cracks can originate within the TBC layer or proximal the TBC/substrate interface. Such cracking loss of TBC structural integrity can lead to further, premature damage to the underlying component substrate. When the TBC layer breaks away from underlying substrate the latter loses its protective thermal layer coating. During continued operation of the turbine engine, it is possible over time that the hot combustion gasses will erode or otherwise damage the exposed component substrate surface, potentially reducing engine operation service life. Potential spallation risk increases with successive powering on/off cycles as the engine is brought on line to generate electrical power in response to electric grid increased load demands and idling down as grid load demand decreases. In order to manage the TBC spallation risk and other engine operational maintenance needs, combustion turbine engines are often taken out of service for inspection and maintenance after a defined number of powering on/off thermal cycles.
In addition to thermal or vibration stress crack susceptibility, the TBC layer on engine components is also susceptible to foreign object damage (FOD) as contaminant particles within the hot combustion gasses strike the relatively brittle TBC material. A foreign object impact can crack the TBC surface, ultimately causing spallation loss of surface integrity that is analogous to a road pothole. Once foreign object impact spalls of a portion of the TBC layer, the remaining TBC material is susceptible to structural crack propagation and/or further spalling of the insulative layer. In addition to environmental damage of the TBC layer by foreign objects, contaminants in the combustion gasses, such as calcium, magnesium, aluminum, and silicon (often referred to as “CMAS”) can adhere to or react with the TBC layer, increasing the probability of TBC spallation and exposing the underlying bond coat.
Past attempts to enhance TBC layer structural integrity and affixation to underlying turbine component substrates have included development of stronger TBC materials better able to resist thermal cracking or FOD, but with tradeoffs in reduced thermal resistivity or increased material cost. Generally, the relatively stronger, less brittle potential materials for TBC application have had lower thermal resistivity. Alternatively, as a compromise separately applied multiple layers of TBC materials having different advantageous properties have been applied to turbine component substrates—for example a more brittle or softer TBC material having better insulative properties that is in turn covered by a stronger, lower insulative value TBC material as a tougher “armor” outer coating better able to resist FOD and/or CMAS contaminant adhesion. In order to improve TBC adhesion to the underlying substrate, intermediate metallic bond coat (BC) layers have been applied directly over the substrate. Structural surface properties and/or profile of the substrate or BC interface to the TBC have also been modified from a flat, bare surface. Some known substrate and/or BC surface modifications (e.g., so-called “rough bond coats” or RBCs) have included roughening the surface by ablation or other blasting, thermal spray deposit or the like. In some instances, the BC or substrate surface has been photoresist or laser etched to include surface features approximately a few microns (μm) height and spacing width across the surface planform. Features have been formed directly on the substrate surface of turbine blade tips to mitigate stress experienced in blade tip coatings. Rough bond coats have been thermally sprayed to leave porous surfaces of a few micron-sized features. TBC layers have been applied by locally varying homogeneity of the applied ceramic-metallic material to create pre-weakened zones for attracting crack propagation in controlled directions. For example a weakened zone has been created in the TBC layer corresponding to a known or likely stress concentration zone, so that any cracks developing therein are propagated in a desired direction to minimize overall structural damage to the TBC layer.
Various embodiments of turbine component construction and methods for making turbine components that are described herein help preserve turbine component thermal barrier coating (TBC) layer structural integrity during turbine engine operation. In some embodiments engineered surface features (ESFs) formed directly in the component substrate or in, intermediate layers applied over the substrate enhance TBC layer adhesion thereto. In some embodiments, the ESFs function as walls or barriers that contain or isolate cracks in the TBC layer, inhibiting additional crack propagation within that layer or delamination from adjoining coupled layers.
In some embodiments engineered groove features (EGFs) are formed in the TBC layer through the outer surface thereof, such as by laser or water jet ablation or mechanical cutting into a previously formed TBC layer. The EGFs—functioning as the equivalent of a fire line that prevents a fire from spreading across a void or gap in combustible material—stop further crack propagation in the TBC layer across the groove to other zones in the TBC layer. EGFs in some embodiments are aligned with stress zones that are susceptible to development of cracks during engine operation. In such embodiments, formation of a groove in the stress zone removes material that possibly or likely will form a stress crack during engine operation. In other embodiments, EGFs are formed in convenient two dimensional or polygonal planform patterns into the TBC layer. The EGFs localize thermal stress- or foreign object damage (FOD)-induced crack propagation within the TBC that might otherwise allow excessive TBC spallation and subsequent thermal exposure damage to the turbine component underlying substrate. A given TBC surface area that has developed one or more stress cracks is isolated from non-cracked portions that are outside of the EGFs. Therefore, if the cracked portion isolated by one or more EGFs spalls from the component the remaining TBC surface outside the crack containing grooves will not spall off because of the contained crack(s).
In some embodiments, spallation of cracked TBC material that is constrained within ESFs and/or EGFs leaves a partial underlying TBC layer that is analogous to a road pothole. The underlying TBC material that forms the floor or base of the “pot hole” provides continuing thermal protection for the turbine engine component underlying substrate.
In some embodiments a turbine component has a thermally sprayed overlying thermal barrier coating (TBC) with depth-varying material properties. Exemplary depth-varying material properties include elastic modulus, fracture toughness, and thermal conductivity that vary from the TBC layer inner to outer surface. Exemplary ways to modify physical properties include application of plural separate overlying layers of different material composition or by varying the applied material composition during the thermal spray application of the TBC layer.
Some embodiments also apply a calcium-magnesium-aluminum-silicon (CMAS)-retardant material over the TBC layer to retard reaction with or adhesion of CMAS containing combustion particulates to the TBC layer. When CMAS-retardant layers are applied over EGFs, they inhibit accumulation of foreign material within the grooves and provide smoother boundary layer surfaces to enhance combustion gas flow aerodynamic efficiency.
More particularly, embodiments of the invention described herein feature a combustion turbine component having a heat insulating outer surface for exposure to combustion gas, which includes a metallic substrate having a substrate surface and an anchoring layer built upon the substrate surface. A planform pattern of engineered surface features (ESFs) is formed in and projects from the anchoring layer. A thermally sprayed or vapor deposited or solution/suspension plasma sprayed outer thermal barrier coat (OTBC) having an OTBC inner surface is applied over and coupled to the anchoring layer and an OTBC outer surface for exposure to combustion gas. Engineered groove features (EGFs) are formed into and penetrating the previously applied OTBC layer through the OTBC outer surface, having a groove depth.
Other embodiments of the invention described herein feature a method for controlling crack propagation in a thermal barrier coating (TBC) outer layer of combustion turbine engine component, by providing a combustion turbine engine that includes a component having a heat insulating outer surface for exposure to combustion gas, which include a metallic substrate having a substrate surface; an anchoring layer built upon the substrate surface; and a thermally sprayed or vapor deposited or solution/suspension plasma sprayed outer thermal barrier coat (OTBC) having an OTBC inner surface applied over and coupled to the anchoring layer and an OTBC outer surface for exposure to combustion gas. The provided component also has a planform pattern of engineered groove features (EGFs) formed into and penetrating the previously applied OTBC layer through the OTBC outer surface, having a groove depth. The method is practiced by operating the engine, inducing thermal or mechanical stress in the OTBC during engine thermal cycling, or inducing mechanical stress in the OTBC by foreign object impact, where any of the induced stresses generates a crack in the OTBC. Crack propagation in the OTBC is arrested when the crack intersects one or more of the EGFs.
Yet other embodiments of the invention described herein feature a method for controlling crack propagation in a thermal barrier coating (TBC) outer layer of combustion turbine engine component, by providing a combustion turbine engine that includes a component having a heat insulating outer surface for exposure to combustion gas, which include a metallic substrate having a substrate surface; an anchoring layer built upon the substrate surface; and a planform pattern of engineered surface features (ESFs) projecting from the anchoring layer that are in contact with the OTBC layer. A thermally sprayed or vapor deposited or solution/suspension plasma sprayed outer thermal barrier coat (OTBC) having an OTBC inner surface is applied over and coupled to the anchoring layer and an OTBC outer surface for exposure to combustion gas. A planform pattern of engineered groove features (EGFs), having a groove depth, is formed into the previously applied OTBC layer, and penetrates through the OTBC outer surface. The method is practiced by operating the engine, inducing thermal or mechanical stress in the OTBC during engine thermal cycling, or inducing mechanical stress in the OTBC by foreign object impact, where any of the induced stresses generates a crack in the OTBC. Crack propagation in the OTBC is arrested upon intersection of the crack with one or more of the EGFs or ESFs.
The respective features of the various embodiments described herein invention may be applied jointly or severally in any combination or sub-combination.
The embodiments shown and described herein can be understood by considering the following detailed description in conjunction with the accompanying drawings, in which:
To facilitate understanding, identical reference numerals have been used, where possible, to designate identical elements that are common to the figures. The figures are not drawn to scale. The following common designators for dimensions, cross sections, fluid flow, axial or radial orientation and turbine blade rotation have been utilized throughout the various invention embodiments described herein:
C-C cross section;
DG groove depth;
F flow direction through turbine engine;
G turbine blade tip to abradable surface gap;
H height of a surface feature;
HR ridge height;
L length of a surface feature;
R turbine blade rotational direction;
R1 Row 1 of the turbine engine turbine section;
R2 Row 2 of the turbine engine turbine section;
SR ridge centerline spacing;
SG groove spacing;
T thermal barrier coat (TBC) layer thickness;
W width of a surface feature;
WG groove width;
WR abradable ridge width;
Δ groove skew angle relative to abradable ridge longitudinal/axial axis; and
σ stress concentration in a thermal barrier coating (TBC).
Exemplary embodiments of the present invention enhance performance of the thermal barrier coatings (TBCs) that are applied to surfaces of turbine engine components, including combustion or gas turbine engines, as well as steam turbine engines. In exemplary embodiments of the invention that are described in detail herein, engineered groove features (EGFs) are formed within the thermal barrier coating (TBC). The EGFs are advantageously aligned with likely stress zones within the TBC or randomly aligned in a convenient two-dimensional or polygonal planform pattern on the TBC surface and into the TBC layer. The EGFs isolate and localize thermal stress- or foreign object damage (FOD)-induced crack propagation within the TBC layer—by isolating the damage to one side of the groove that faces the damage and preventing it from jumping across the groove to otherwise undamaged portions of the TBC layer—that might otherwise allow excessive TBC spallation and subsequent thermal exposure damage to the turbine component underlying substrate.
General Summary of Thermally Sprayed TBC
Application in Combustion Turbine Engine Components
Referring to
As previously noted, turbine component surfaces that are exposed to combustion gasses are often constructed with a thermal barrier coating (TBC) layer for insulation of their underlying substrates. Typical TBC coated surfaces include the turbine blades 92, the vanes 104, 106 and related turbine vane carrier surfaces and combustion section transitions 85. The TBC layer for blade 92, vane 104, 106 and transition 85 exposed surfaces are often applied by thermal sprayed or vapor deposition or solution/suspension plasma spray methods, with a total TBC layer thickness of 300-2000 microns (μm).
Turbine Blade Tip Abradable Component TBC Application
Insulative layers of greater thickness than 1000 microns are often applied to sector shaped turbine blade tip abradable components 110 (hereafter referred to generally as an “abradable component”) that line the turbine engine 80 turbine casing 100 in opposed relationship with the blade tips 94. The abradable components 110 having a support surface 112 retained within and coupled to the casing and an insulative abradable substrate 120 that is in opposed, spaced relationship with the blade tip by a blade tip gap G. The abradable substrate is often constructed of a metallic/ceramic material, similar to the TBC coating materials that are applied to blade 92, vane 104, 106 and transition 85 combustion gas exposed surfaces. Those abradable substrate materials have high thermal and thermal erosion resistance and maintain structural integrity at high combustion temperatures. Generally, it should be understood that some form of TBC layer is formed over the blade tip abradable component 110 bare underlying metallic support surface substrate 112 for insulative protection, plus the insulative substrate thickness that projects at additional height over the TBC. Thus it should be understood that abradable components 110 have a functionally equivalent TBC layer to the TBC layer applied over the turbine transition 85, blade 92 and vane 102/104, The abradable surface 120 function is analogous to a shoe sole or heel that protects the abradable component support surface substrate 112 from wear and provides an additional layer of thermal protection. Exemplary materials used for blade tip abradable surface ridges/grooves include pyrochlore, cubic or partially stabilized yttria stabilized zirconia. As the abradable surface 120 metallic ceramic materials is often more abrasive than the turbine blade tip 94 material a blade tip gap G is maintained to avoid contact between the two opposed components that might at best cause premature blade tip wear and in worse case circumstances might cause engine damage.
Blade tip abradable components 110 are often constructed with a metallic base layer support surface 112, to which is applied a thermally sprayed ceramic/metallic abradable substrate layer 120 of many thousands of microns thickness, i.e., multiples of the typical transition 85 blade 92 or vane 104/106 TBC layer thickness. As will be described in greater detail herein, the abradable layer of exemplary turbine blade tip opposing abradable surface planform and projection profile invention embodiments described in the related patent applications for which priority is claimed herein include grooves, depressions or ridges in the abradable substrate layer 120 to reduce abradable surface material cross section for potential blade tip 94 wear reduction and for directing combustion airflow in the gap region G. Commercial desire to enhance engine efficiency for fuel conservation has driven smaller blade tip gap G specifications: preferably no more than 2 millimeters and desirably approaching 1 millimeter (100 μm).
With respect to the
With the progressive wear zones, construction of some blade tip abradable wear surface 120 embodiments of the prior applications for which priority is claimed herein, blade tip gap G can be reduced from previously acceptable known dimensions. For example, if a known acceptable blade gap G design specification is 1 mm the higher ridges in wear zone I can be increased in height so that the blade tip gap is reduced to 0.5 mm. The lower ridges that establish the boundary for wear zone II are set at a height so that their distal tip portions are spaced 1 mm from the blade tip. In this manner a 50% tighter blade tip gap G is established for routine turbine operation, with acceptance of some potential wear caused by blade contact with the upper ridges in zone I. Continued localized progressive blade wearing in zone II will only be initiated if the blade tip encroaches into the lower zone, but in any event, the blade tip gap G of 1 mm is no worse than known blade tip gap specifications. In some exemplary embodiments the upper zone I height is approximately ⅓ to ⅔ of the lower zone II height. If the blade tip gap G becomes reduced for any one or more blades due to turbine casing 100 distortion, fast engine startup mode or other reason initial contact between the blade tip 94 and the abradable component 10 will occur at the higher ridge tips forming Zone I. While still in zone I the blade tips 94, only rub the alternate staggered higher ridges. If the blade gap G progressively becomes smaller, the higher ridges will be abraded until they are worn all the way through zone I and start to contact the lower ridge tips in zone II. Once in Zone II the turbine blade tip 94 rubs all of the remaining ridges at the localized wear zone, but in other localized portions of the turbine casing there may be no reduction in the blade tip gap G and the upper ridges may be intact at their full height. Thus the alternating height rib construction of some of the abradable component 110 embodiments accommodates localized wear within zones I and II, but preserve the blade tip gap G and the aerodynamic control of blade tip leakage in those localized areas where there is no turbine casing 100 or blade 92 distortion.
Multi-height wear zone constructions in abradable components are also beneficial for so-called “fast start” mode engines that require faster full power ramp up (order of 40-50 Mw/minute). Aggressive ramp-up rates exacerbate potential higher incursion of blade tips into ring segment abradable coating 120, resulting from quicker thermal and mechanical growth and higher distortion and greater mismatch in growth rates between rotating and stationary components. When either standard or fast start or both engine operation modes are desired the taller ridges Zone I form the primary layer of clearance, with the smallest blade tip gap G, providing the best energy efficiency clearance for machines that typically utilize lower ramp rates or that do not perform warm starts. Generally the ridge height for the lower ridge tips in Zone II is between 25%-75% of the higher ridge tip height of those forming Zone I.
More particularly,
Progressive wear zones in abradable component surfaces 120 of the embodiments of
In the turbine blade tip abradable component 230 embodiment of
With thermally sprayed blade tip abradable component construction, the cross sections and heights of upper wear zone I thermally sprayed abradable material can be configured to conform to different degrees of blade tip intrusion by defining arrays of micro ribs or nibs, as shown in
In the alternative embodiment of
Pixelated nib 242A and groove 248A/C dimensional boundaries are identified in
Generally, the upper wear zone I ridge height in the abradable component can be chosen so that the ideal blade tip gap is 0.25 mm. The 3:00 and 9:00 turbine casing circumferential wear zones are likely to maintain the desired 0.25 mm blade tip gap throughout the engine operational cycles, but there is greater likelihood of turbine casing/abradable component distortion at other circumferential positions. The lower ridge height may be selected to set its ridge tip at an idealized blade tip gap of 1.0 mm so that in the higher wear zones the blade tip only wears deeper into the wear zone I and never contacts the lower ridge tip that sets the boundary for the lower wear zone II. If despite best calculations the blade tip continues to wear into the wear zone II, the resultant blade tip wear operational conditions are no worse than in previously known abradable layer constructions. However in the remainder of the localized circumferential positions about the abradable layer the turbine is successfully operating with a lower blade tip gap G and thus at higher operational efficiency, with little or no adverse increased wear on the blade tips.
In the blade tip abradable embodiments of
The MSFs in the PMPPs of some embodiments are generated from a cast in or an engineered surface feature formed directly in the substrate material. In other embodiments, the MSFs in the PMPPs are generated in the substrate or in an overlying bond coat (BC) layer by an ablative or additive surface modification technique such as water jet or electron beam or laser cutting or by laser sintering methods. The engineered surface features are subsequently coated with high temperature abradable thermal barrier coating (TBC), with, or without an intermediate bond coat layer applied on the engineered MSF features in the PMPP, to produce a discontinuous surface that will abrade more efficiently than a current state of the art coating. Once contacted (by a passing blade tip), released (abraded) particles are removed via a tortuous, convoluted (above or subsurface) path in gaps between the MSFs or additional slots formed within the abradable surface between the MSFs. Optional continuous slots and/or gaps are oriented to provide a tortuous path for hot gas ejection, thereby maintaining the sealing efficiency of the primary (contact) surface. The surface configuration, which reduces potential rubbing contact surface area between the blade tips and the discontinuous MSFs, reduces frictional heat generated in the blade tip. Reduced frictional heat in the blade tip potentially reduces worn blade tip material loss attributable to tip over heating and metal smear/transfer onto the surface of the abradable. Further benefits include the ability to deposit thicker, more robust thermal barrier coatings over the MSFs than normally possible with known continuous abradable rib designs, thereby imparting potentially extended design life for ring segments.
The micro surface feature (MSF) in its simplest form can be basic shape geometry, repeated in unit cells across the surface of the ring segment with gaps between respective cells. The unit cell MSFs are analogous to pixels that in aggregate forms the PMPP's larger pattern. In more optimized forms, the MSF can be modified according to the requirement of the blade tip relationship of the thermal behavior of the component during operation. In such circumstances, feature depth, orientation, angle, and aspect ratio may be modified within the surface to produce optimized abradable performance from beginning to end of blade sweep. Other optimization parameters include ability of thermal spray equipment that forms the TBC to penetrate fully captive areas within the surface and allow for an effective continuous TBC coating across the entire surface.
As previously noted, the abradable component with the PMPPs comprising arrays of MSFs is formed by casting the MSFs directly into the abradable substrate during its manufacture or built up on the substrate (such as by thermal spray or additive manufacturing techniques, e.g., electron beam or laser beam deposition) or by ablation of substrate material. In the first-noted formation process, a surface feature can be formed in a wax pattern, which is then shelled and cast per standardized investment casting procedures. Alternatively, a ceramic shell insert can be used on the outside of the wax pattern to form part of the shell structure. When utilizing a ceramic shell insert the MSFs can be more effectively protected during the abradable component manufacture handing and can more exotic in feature shape and geometry (i.e., can contain undercuts or fragile protruding features that would not survive a normal shelling operation.
MSFs can be staggered (stepped) to accept and specifically deflect plasma splats for optimum TBC penetration. Surface features cast-in and deposited onto the substrate may not necessarily fully translate in form to a fully TBC coated surface. During coating, ceramic deposition will build upon the substrate in a generally transformative nature but will not directly duplicate the original engineered surface feature. The thermal spray thickness can also be a factor in determining final surface form. Generally, the thicker the thermal spray coating, the more dissipated the final surface geometry. This is not necessarily problematical but needs to be taking into consideration when designing the engineered surface feature (both initial size and aspect ratio. For example, a chevron-shaped MSF formed in the substrate, when subsequently coated by an intermediate bond coat layer and a TBC top layer may dissipate as a crescent- or mound-shaped protrusion in the finished abradable surface projecting profile.
Where exemplary MSF unit cells are shown in
Various exemplary embodiments described herein, which incorporate pixelated major planform patterns (PMPP) of discontinuous micro surface features (MSF) jointly or severally in different combinations have at least some of the following features:
Exemplary embodiments of turbine abradable components including pixelated major planform patterns (PMPP) of discontinuous micro surface features (MSF) are shown in
On the uppermost portion of the abradable component 260 a thermal barrier coating (TBC) 266 has been applied directly over the MSFs 263, leaving mound or crescent-shaped profile projections 267 on the abradable component in a PMPP 262 that are arrayed for directing hot gas flow between the abradable component and a rotating turbine blade tip. In the event of contact between the blade tip and the opposing surface of the abradable component 260, the relatively small cross sectional surface area MSFs 263 will rub against and be abraded by the blade tip. The MSF 263 and turbine blade tip contact is less likely to cause blade tip erosion or spallation of the abradable surface 260 from the contact, compared to previously known continuous single height or solid surface abradable components that do not have the benefit of the abradable upper and lower Zones I and II, such as those shown in
On the lowermost portion of the abradable component 260 a metallic bond coat (BC) 264 is applied to the naked metallic substrate 261 and the chevron-shaped MSFs 265 are formed in the BC by additive or ablative manufacturing processes. The BC 264 and the MSFs 265, arrayed in the PMPP 262, are then covered with a TBC 266 leaving generally chevron-shaped MSFs 268 that project from the substrate 260 surface.
Dimensions of an exemplary chevron-shaped MSF 272 are shown in
As with the blade tip abradable components embodiments shown in
As previously discussed, the micro surface features MSFs can be formed in the substrate or in a bond coat of an abradable component. In
Engineered Surface Features (ESFs) Enhance TBC Adhesion and Crack Isolation
Some exemplary turbine component embodiments incorporate an anchoring layer of engineered surface features (ESFs) that aid mechanical interlocking of the TBC layer and aid in isolation of cracks in the TBC layer, so that they do not spread beyond the ESF. In some blade tip abradable applications the solid ridge and groove projecting surface features as well as MSFs function as ESFs, depending upon the former's physical dimensions and relative spacing between them, but they are too large for more general application to turbine components other than blade tip abradable components. For exemplary turbine blade, vane or combustor transition applications the ESFs are formed in an anchoring layer that is coupled to an inner surface layer of the TBC layer and they are sized to anchor the TBC layer coating thickness range of 300-2000 microns (μm) applied to those components without changing an otherwise generally flat outer surface of the TBC layer that is exposed to combustion gas. Generally, the ESFs have heights and three-dimensional planform spacing on the turbine component surface sufficient to provide mechanical anchoring and crack isolation within the total thickness of the TBC layer. Thus, the ESFs will be shorter than the total TBC layer thickness but taller than etched or engraved surface features that are allegedly provided to enhance adhesion bonding between the TBC and the adjoining lower layer (e.g., an underlying naked substrate or intermediate bond coat layer interposed between the naked substrate and the TBC layer). Generally, in exemplary embodiments the ESFs have a projection height between approximately 2-75 percent of the TBC layer's total thickness. In some preferred embodiments, the ESFs have a projection height of at least approximately 33 percent of the TBC layer's total thickness. In some exemplary embodiments, the ESFs define an aggregate surface area at least 20 percent greater than an equivalent flat surface area.
In
In
Engineered surface feature (ESF) cross sectional profiles, their planform array patterns, and their respective dimensions may be varied during design and manufacture of the turbine component to optimize thermal protection by inhibiting crack formation, crack propagation, and TBC layer spallation. Different exemplary permutations of ESF cross sectional profiles their three-dimensional planform array patterns and their respective dimensions are shown in
In exemplary embodiments, the ESFs are selectively arrayed in three-dimensional planform linear or polygonal patterns. For example, the ESF planform pattern of parallel vertical projections shown in
As previously mentioned, in addition to TBC layer anchoring advantages provided by the ESFs described herein, they also localize TBC layer crack propagation. In the turbine component 380 of
Now compare the crack propagation resistant construction of the turbine component 390 shown in
Engineered Groove Features (EGFs) Enhance TBC Crack Isolation
Some exemplary turbine component embodiments incorporate planform arrays of engineered groove features (EGFs), which are formed in the outer surface of the TBC after the TBC layer application. The EGFs groove axes are selectively oriented, at any skew angle relative to the TBC outer surface and extend into the TBC layer. Analogous to a firefighter fire line, the EGFs isolate cracks in the TBC layer, so that they do propagate across the boundary of a groove void into other portions of adjoining TBC material. Generally, if a crack in the TBC ultimately results in spallation of material above the crack the EGF array surrounding the crack forms a localized boundary perimeter of the spall site, leaving TBC material outside the boundary intact. Within the spallation zone bounded by the EGFs, damage will be generally limited to loss of material above the EGF groove depth. Thus in many exemplary embodiments EGF depth is limited to less than the total thickness of all TBC layers, so that a volume and depth of intact TBC material remains to provide thermal protection for the local underlying component metallic substrate. In some embodiments, the EGF arrays are combined with ESF arrays to provide additional TBC integrity than either might provide alone.
Exemplary engineered groove feature crack isolation capabilities are shown in
Unlike prior known TBC stress crack relief mechanisms that create voids or discontinuities within the applied thermally sprayed or vapor deposited TBC layer, such as by altering layer application orientation or material porosity, the engineered groove feature (EGF) embodiments herein form cut or ablated grooves or other voids through the previously formed TBC layer outer surface to a desired depth. As shown in
The turbine component embodiments of
In
Engineered Groove Features (EGFs)
Inhibit TBC Delamination Around Cooling Holes
Advantageously, engineered groove features can be formed in the TBC layer around part of or the entire periphery of turbine component cooling holes or other surface discontinuities, in order to limit delamination of the TBC over layer along the cooling hole or other discontinuity margins in the component substrate. The TBC layer at the extreme margin of the cooling hole can initiate separation from the metallic substrate that can spread laterally/horizontally within the TBC layer away from the hole. Creation of an EGF at a laterally spaced distance from the cooling hole margin—such as at a depth that contacts the anchoring layer or the metallic substrate—limits further delamination beyond the groove.
Various cooling hole periphery EGF embodiments are shown in
In
Material Varying Multi-Layer and Graded TBC Construction
As was previously discussed, the aggregate thermally sprayed TBC layer of any turbine component embodiment described herein may have different local material properties laterally across the component surface or within the TBC layer thickness dimension. As one example, one or more separately applied TBC layers closest to the anchoring layer may have greater strength, ductility, toughness and elastic modulus material properties than layers closer to the component outer surface but the higher level layers may have greater thermal resistivity and brittleness material properties. Multi-layer TBC embodiments are shown in
Exemplary material compositions for thermal barrier coat (TBC) layers include yttria stabilized zirconia, rare-earth stabilized zirconia with a pyrochlore structure, rare-earth stabilized fully stabilized cubic structure, or complex oxide crystal structures such as magnetoplumbite or perovskite or defective crystal structures. Other exemplary TBC material compositions include multi-element doped oxides with high defect concentrations. Examples of CMAS retardant compositions include alumina, yttrium aluminum oxide garnet, slurry deposited/infiltrated highly porous TBC materials (the same materials that are utilized for OTBC or LTBC compositions), and porous aluminum oxidized to form porous alumina.
In
In the embodiment of
The continuously-applied, thermally sprayed and graded TBC layer construction turbine component 530 of
In the embodiment of
Segmented TBC Construction
Segmented TBC construction embodiments, which are conceptually analogous to an ear of corn or maize, combine engineered surface features (ESFs) and engineered groove features (EGFs) embodiments along with optional combinations of multi-layer or graded material-varying thermal barrier coat and CMAS-resistant surface coatings. The segmented TBC construction is suitable for curved as well as flat surfaces of turbine engine components, such as combustion section transitions, blades, and vanes. Exemplary segmented TBC protected, curved surface turbine components are shown in
In
The turbine component embodiment 570 of
Although various embodiments that incorporate the teachings of the invention have been shown and described in detail herein, those skilled in the art can readily devise many other varied embodiments that still incorporate these teachings. The invention is not limited in its application to the exemplary embodiment details of construction and the arrangement of components set forth in the description or illustrated in the drawings. The invention is capable of other embodiments and of being practiced or of being carried out in various ways. For example, various ridge and groove profiles may be incorporated in different planform arrays that also may be locally varied about a circumference of a particular engine application. In addition, it is to be understood that the phraseology and terminology used herein is for the purpose of description and should not be regarded as limiting. The use of “including,” “comprising,” or “having” and variations thereof herein is meant to encompass the items listed thereafter and equivalents thereof as well as additional items. Unless specified or limited otherwise, the terms “mounted”, “connected”, “supported”, and “coupled” and variations thereof are used broadly and encompass direct and indirect mountings, connections, supports, and couplings. Further, “connected” and “coupled” are not restricted to physical or mechanical connections or couplings.
| Filing Document | Filing Date | Country | Kind |
|---|---|---|---|
| PCT/US2015/016318 | 2/18/2015 | WO | 00 |
| Publishing Document | Publishing Date | Country | Kind |
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| WO2015/130526 | 9/3/2015 | WO | A |
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| Entry |
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| PCT International Search Report and Written Opinion dated May 6, 2016 corresponding to PCT Application PCT/US2015/016318 filed Feb. 18, 2015. (13 pages). |
| Number | Date | Country | |
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| 20160362989 A1 | Dec 2016 | US |
| Number | Date | Country | |
|---|---|---|---|
| Parent | 14188941 | Feb 2014 | US |
| Child | 15121429 | US | |
| Parent | 14188958 | Feb 2014 | US |
| Child | 14188941 | US |
