Patent Application
20170089259
Field
The present disclosure relates generally to composite thermal barriers for combustion chamber component surfaces in an internal combustion engine.
Technical Background
The efficiency of internal combustion engines may be improved by retaining heat from ignited fuel in the combustion chamber. This can be accomplished by minimizing heat loss to the surrounding engine. One solution has been to insulate parts of the combustion chamber. A problem with insulating the combustion chamber from the surrounding engine may be creating a reliable bond between the thermal barrier and combustion chamber component surfaces.
Accordingly, a need exists for improved thermal barriers within internal combustion engines.
According to one embodiment of the present disclosure, a composite thermal barrier is disclosed. In embodiments, the composite thermal barrier comprises a metallic web and an insulation material. In embodiments, the metallic web includes a metal and a void space. In embodiments, the metallic web also includes a volume is defined by a length, a width, and a thickness. In embodiments, the metallic web is connected to a metallic surface within a combustion chamber of an engine. In embodiments, the insulation material is contained within the void space of the metallic web.
According to another embodiment of the present disclosure, a composite thermal barrier is disclosed. In embodiments, the composite thermal barrier comprises a metallic web and an insulation material. In embodiments, the metallic web includes a metal and a plurality of voids and is defined by a length, a width, and a thickness. In embodiments, the metallic web is connected to a metallic component within the combustion chamber of an engine. In embodiments, the insulation material is contained within the plurality of voids of the metallic web.
According to yet another embodiment of the present disclosure, a method of applying a composite thermal barrier is disclosed. In embodiments, the method includes preparing a metallic surface within a combustion chamber of an engine for application of the composite thermal barrier. In embodiments, the method includes applying the metallic web including void space to the metallic surface. In embodiments, the method includes inserting the insulation material within the void space of the metallic web.
Before turning to the following Detailed Description and Figures, which illustrate exemplary embodiments in detail, it should be understood that the present inventive technology is not limited to the details or methodology set forth in the Detailed Description or illustrated in the Figures. For example, as will be understood by those of ordinary skill in the art, features and attributes associated with embodiments shown in one of the Figures or described in the text relating to one of the embodiments may well be applied to other embodiments shown in another of the Figures or described elsewhere in the text.
The disclosure will be better understood, and features, aspects and advantages other than those set forth above will become apparent when consideration is given to the following detailed description thereof. Such detailed description makes reference to the following drawings.
Unless defined otherwise, all technical and scientific terms used herein have the same meaning as commonly understood by one of ordinary skill in the art to which the disclosure belongs. Although any methods and materials similar to or equivalent to those described herein can be used in the practice or testing of the present disclosure, the exemplary methods and materials are described below.
Engine fuel efficiency is affected by the thermal conductivity of the materials used to make the various components of an engine. This is particularly true for components within the combustion chamber of an engine (e.g., wall of the combustion chamber, pistons, valves, exhaust ports, manifolds, etc.). The higher the thermal conductivity of materials used in the combustion chamber, the more combustion energy lost to heat energy. By lowering the thermal conductivity of materials directly exposed to the combustion reaction, more energy of combustion is available for performing work and powering the engine (i.e., to drive the piston). That is, heat of combustion that is not lost to heat energy can be used to drive a turbocharger in the exhaust manifold and/or more effectively light off the catalytic converter during a cold-start of the engine. Accordingly, the overall efficiency of the engine (including fuel efficiency) may be improved with thermally resistant materials.
Conventional methods for lowering the thermal conductivity of materials within the combustion chamber have included the use of thermal barriers. Conventional thermal barriers for combustion chambers of internal combustion engines may have one or more of several problems. One major shortcoming for conventional thermal barriers may be that the thermal barrier spalls or separates from the surface within the combustion chamber when exposed to the violent combustion kinetics, high pressures (e.g., 10 bars-200 bars), and high temperatures (e.g., 1000° C.-3000° C.) therein. Spalling of thermal barriers including brittle ceramic materials into the combustion chamber can cause damage (e.g., gouge, plug, etc.) to other engine components and the catalytic convertor. Another shortcoming of conventional thermal barriers may be insufficient thermal resistivity properties or a different coefficient of thermal expansion (CTE) than the combustion chamber surface which may lead to separation at high temperatures. Yet another shortcoming may be non-uniform thicknesses of conventional thermal barriers on engine component surfaces.
The present application is directed to a composite thermal barrier 200 on any metallic surface within an internal combustion engine 100.
In
Composite thermal barrier 200 of the present disclosure may be on any metallic surface within engine 100. In an exemplary embodiment, composite thermal barrier 200 is on a metallic surface 101 within combustion chamber 102. Metallic surface 101 may be surfaces defining compression exhaust volume 121, surfaces defining chamber exhaust volume 122, or surfaces defining chamber intake volume 123. In one embodiment, surface 101 may not be wall surfaces of combustion chamber 102 contacted by piston 104. That is, composite thermal barrier 200 may be excluded from surfaces in chamber 102 subjected to mechanical friction from piston 104 or areas along the crevice quench that may wear or separate composite thermal barrier 200 from that surface. In another exemplary embodiment, metallic surface 101 is piston top surface 120, wall surfaces and end surfaces of combustion chamber 102, a surface of intake valve 106, a surface of exhaust valve 114, walls of exhaust duct 118, or walls of intake duct 119.
Composite thermal barrier 200 of the present disclosure includes a metallic web 202 and an insulation material 204. Metallic web 202 includes a metal 203 and a void space 205. Metallic web 202 also includes a volume defined by a length, a width, and a thickness T1. That is, metal 203 and void space 205 of metallic web 202 together may delineate the volume of metallic web 202 (referred to herein as metallic web 202 volume). Void space 205 also has a smaller volume within metallic web 202 volume (referred to herein as void space 205 volume). In one embodiment, metallic web 202 volume includes from about 1% to 95% metal 203. In alternative embodiments, metallic web 202 volume may be from about 1% to about 90% metal 203, or from about 2% to about 80% metal 203, or from about 3% to about 70% metal 203, or even from about 4% to about 60% metal 203. Accordingly, metallic web 202 volume may be from 5% to about 99% void space 205, or about 10% to about 99% void space 205, or from about 20% to about 98% void space 205, or even from about 30% to about 97% void space 205. Void space 205 may extend across the entire thickness T1 of metallic web 202. In alternative embodiments, void space 205 may extend across at least 50% of the thickness T1 of metallic web 202. Void space 205 within metallic web 202 may be a singular void space or a plurality of discrete and/or interconnected voids. Plurality of voids of void space 205 may have a diameter ranging from about 0.01 mm to about 4 mm, or about 50 microns to about 5000 microns. Plurality of voids of void space 205 may have a median diameter D50 from about 0.02 mm to about 4 mm, or about 200 microns to about 4000 microns. In an exemplary embodiment, the diameters of the plurality of voids of void space 205 are larger than particle sizes of insulation material 204 so the insulation material may be inserted into void space 205. Void space 205 of metallic web 202 may be a plurality of voids and extend at least 50% of the metallic web length, up to 100% of the metallic web length.
The structure of metal 203 within metallic web 202 is webbed.
The structure of metal 203 in metallic web 202 is capable of retaining its shape on surface 101 and around void space 205. The structure of metal 203 is also capable of containing insulation material 204 within the void space. The structure of metal 203 acts as an anchor capable of interlocking insulation material 204 within the void space. The structure of metallic web 202 may be sufficiently rigid and has thermo mechanical fatigue resistance so as to withstand the combustion temperatures and pressures within combustion chamber 102 during operation of engine 100. The volume of metallic web 202 may be in any shape including, but not limited to, rectangular, cubic, annular, hemispherical, or cylindrical. The shape of the volume of metallic web 202 may also conform to the rounded or non-uniform shapes of surface 101 to which it is connected, including a curved piston top surface 120.
The length and width of metallic web 202 (including metal 203 and void space 205) can have any suitable lateral dimensions (e.g., from about 0.1 mm to about 100 cm), including equal dimensions. Thickness T1 of metallic web 202 may be from about 0.01 mm to about 10 mm, or from about 0.1 mm to about 2 mm, or from about 0.4 mm to about 2 mm, or even from about 0.5 mm to about 1 mm. In exemplary embodiments, thickness T1 is uniform across the length and the width of metallic web 202. Thickness T1 of metallic web 202 may be measured from surface 101 to a termination point of metallic web 202 away from surface 101. Thickness T1 may also be measured from surface 101 to an average thickness of metallic web 202 away from surface 101. Surface 101 within combustion chamber 102 may be identified from metallic web 202 by a lack of void space. That is, thickness T1 of metallic web 202 may be distinct from a thickness of material comprising surface 101 by the presence of void space 205 (or insulation material 204 filled therein) within thickness T1. Alternatively, metallic web 202 thickness T1 may be identified from surface 101 by a distinct interface of the connection caused by the application method.
Metal 203 within metallic web 202 may be an element or an alloy and may include metals and metal alloys commonly used in combustion chamber 102 manufacturing. Metal 203 may include carbon steel, stainless steel, alloy aluminum, aluminum, nickel plated aluminum, titanium, hastelloy, and combinations thereof for example. Metal 203 may also be other super alloys including nickel, chromium, cobalt, or combinations thereof. Metal 203 may have the same coefficient of thermal expansion (CTE) as the material encompassing surface 101 (assuming similar operating temperature ranges) to minimize thermal expansion stresses and failures at their connection. In an exemplary embodiment, the CTE of metal 203 may be within 150% of the CTE as the material encompassing surface 101 (assuming similar operating temperature ranges).
Composite thermal barrier 200 also includes insulation material 204. Insulation material 204 is contained with void space 205 of metallic web 202. In one embodiment, insulation material 204 is contained with the plurality of voids of void space 205. The presence of insulation material 204 within void space 205 of metallic web 202 inherently eliminates void space 205 within metallic web 202. Insulation material 204 may fill from 5% to 100% of void space 205. Referring back to
In an exemplary embodiment, insulation material 204 is interlocked within the webbed structure of metal 203 such that it does not escape, spall, or flake out from metallic web 202 into combustion chamber 102 during operation of engine 100.
In an embodiment where insulation material 204 includes ceramic material, the ceramic material may have a porosity from about 10% to about 90%, or from about 30% to about 70%. The pores of the ceramic material may include air, argon, nitrogen, helium, and combinations thereof. Alternatively, the pores of the ceramic material may have a vacuum pressure less than atmospheric pressure. Example ceramic materials include, but are not limited to, yttria stabilized zirconia (YSZ), zirconium dioxide, lanthanum zirconate, gadolinium zirconate, lanthanum magnesium hexaaluminate, gadolinium magnesium hexaaluminate, lanthanum-lithium hexaaluminate, barium zirconate, strontium zirconate, calcium zirconate, sodium zirconium phosphate, mullite, aluminum oxide, cerium oxide, and combinations thereof. The ceramic material of exemplary embodiments may be ceramic foam. The ceramic material of exemplary embodiments may also be formed from aluminates, zirconates, silicates, titanates, and combinations thereof.
Composite thermal barrier 200 may also include a metallic skin 206. Metallic skin 206 has a length, a width, and a thickness T2. In one embodiment, metallic skin 206 is adjacent metallic web 202. Referring again to
Metallic skin 206 may be an element or an alloy and may include metals and metal alloys (e.gs., aluminum, carbon steel, Inconel, etc.) commonly used in combustion chamber 102 manufacturing. Metallic skin 206 may be the same metal as metal 203, or different. In one embodiment, metallic skin 206 is the same as the material encompassing surface 101. In an exemplary embodiment, the CTE of metallic skin 206 is the same as the material encompassing surface 101 (assuming similar operating temperature ranges) such that they expand and contract at relatively the same rate. Alternatively, the CTE of metallic skin 206 may be within 150% of the CTE of the material encompassing surface 101. The CTE of metallic skin 206 may also be within 150% of the CTE of metal 203 (assuming similar operating temperature ranges) so as to minimize thermal expansion stresses and failures at their connection.
The length and width of metallic skin 206 can have any suitable lateral dimensions (e.g., from about 0.1 mm to about 100 cm), including equal dimensions. Metallic skin 206 lateral dimensions may extend across at least 50% of the length or the width of metallic web 202, up to 100%. Accordingly, metallic skin 206 may include plurality of discrete lengths and widths.
Thickness T2 of metallic skin 206 may be from about 0.001 mm to about 5 mm, or from about 0.1 mm to about 2 mm, or even from about 0.1 mm to about 1 mm. In exemplary embodiments, thickness T2 is uniform across the length and the width of metallic web 202. As shown in
Again, composite thermal barrier 200 of the present disclosure includes metallic web 202 and insulation material 204. Composite thermal barrier 200 may also include metallic skin 206 adjacent metallic web 202 that may assist in containing insulation material 204 within metallic web 202 so it does not spall or flake out into combustion chamber 102 during operation of engine 100. In an exemplary embodiment, composite thermal barrier 200 has a thermal conductivity of about 0.1 W/m·K to about 12 W/m·K at 400° C., or about 1 W/m·K to about 5 W/m·K at 400° C. Various embodiments of composite thermal barrier 200 on a surface within engine 100 are provided in
The present disclosure also includes methods of applying composite thermal barrier 200 to metallic surface 101 within combustion chamber 102 of engine 100. The method includes preparing metallic surface 101 for application of metallic web 202. Preparing metallic surface 101 may include roughening, chemical etching, drilling, cleaning, or other processes of readying surface 101 for application of metallic web 202 thereon. It is envisioned that the method of preparation of surface 101 will likely depend on the method of applying metallic web 202 on surface 101.
The method of applying composite thermal barrier 200 to metallic surface 101 includes applying metallic web 202 to surface 101. Applying metallic web 202 to surface 101 may be accomplished with 3-D printing, metallic plating, mechanical fastening or threading, fusion welding, brazing, resistance welding, diffusion bonding, sintering, or other conventional methods of metallically bonding metal 203 to surface 101 via metal-to-metal bonds. Methods of applying metallic web 202 to surface 101 include the formation of void space 205 around metal 203 within metallic web 202. These methods are also applicable to forming metallic skin 206 adjacent to metallic web 202.
The method of applying composite thermal barrier 200 to metallic surface 101 also includes inserting insulation material 204 within void space 205 of the metallic web 202. Methods of inserting insulation material 204 within void space 205 include pressure application, injection, pressing, impregnating, and other conventional methods of inserting a solid or gas insulator in void space 205. It is envisioned that inserting insulation material 204 within void space 205 may be accomplished while applying metallic web 202 to surface 101, when metallic web includes void space 205 with an insulation material (e.g., air, vacuum, etc.) already present.
The method of applying composite thermal barrier 200 to metallic surface 101 may also include forming metallic skin 206 adjacent to metallic web 202. Methods of forming metallic skin 206 adjacent to metallic web 202 may include 3-D printing, metallic plating, welding (arc, laser, plasma, or friction), brazing, plasma spraying, mechanical fastening, dip coating, deposition, and other conventional methods of forming a uniform metallic skin on the periphery of a metallic web. These methods are also applicable to applying metallic web 202 to surface 101.
Alternative methods include forming metal 203 on metallic skin 206 prior to applying metal 203 to surface 101. In this embodiment, metal 203 may be formed on metallic skin 206 by sheet metal fabrication, superplastic forming, hydroforming, chemical etching, electrical discharge machining, mechanical milling, pressing and sintering, and other similar processes. Subsequently, metal 203 is applied or connected to surface 101 by above mentioned methods.
The present disclosure will be further clarified with reference to the following examples. The following examples should be construed as illustrative and in no way limiting as to the present disclosure.
The effective thermal conductivity of theoretical composite thermal barriers where modeled using the following equation:
where is effective thermal conductivity of composite thermal barrier 200 in W/m·K, where ki is the thermal conductivity of a component (e.g., metal 203) in W/m·K, and wherein φi is volume fraction of a component.
Using the above referenced equation, the volume fractions of a metal component (which could be metal 203 with or without metallic skin 206) and an insulation component (insulation material 204) where varied for Examples 1-13 to model a composite thermal barrier (CTB). Tables 1a and 1b provide the relative fractions and materials for each of Examples 1-13 and comparative examples (CE) 1-4. Tables 1a and 1b also provide the modeled thermal properties of including the CTB thermal conductivity, CTB effective density, CTB effective heat capacity, and CTB thermal diffusivity. Also shown is the thickness of the CTB modeled to effectively function when exposed to a combustion environment at 2000° C. when the piston surface is at 400° C., and maximum temperature increase at thickness depth of 250° C.
The 14 modeled composite thermal barriers (CTB) in Table 1a and Table 1b provide that a CTB thickness from about 0.5 mm to about 2 mm on a steel piston have an effective thermal conductivity of the modeled examples of the present disclosure are from about 3.0 W/m·K to about 12 W/m·K. These modeled example CTBs also have an effective thermal diffusivity from about 8.0E-7 to about 7.0E-6. Comparative example 4 provides an example prior art CTB where the insulation material (YSZ) is not held within a metallic web according to the present disclosure.
Five composite thermal barriers (Examples 2-6) in accordance with the present disclosure were prepared on direct metal laser sintered coupons (simulating an engine internal surface) and tested under severe thermal cycling (up to 50 cycles) to demonstrate thermal resistivity and spalling resistance.
The first composite thermal barrier was prepared on a 50 mm square by 12 mm thick F75 cobalt-chrome (Co—Cr) block coupon. A metallic web of F75 Co—Cr was direct metal laser sintered onto the coupon. The metallic web was in the shape of a plurality of 1.15 mm diameter posts (each 1.5 mm tall on the coupon flat surface) each equally spaced apart in a triangular array. The coupon also included a raised edge along its perimeter 1.15 mm wide and 1.5 mm tall. The metallic web included 90% open frontal area (OFA) across its thickness (i.e., between the surface of the coupon and the termination ends of the plurality of posts). A 100 micron bond coat of nickel chromium aluminum yttrium (NiCrAlY) was applied to the surface of the coupon around the plurality of posts. YSZ was plasma sprayed over the NiCrAlY to fill the void space (defined by the OFA and the thickness of the plurality of the posts) around the metallic web. Excess YSZ extending above the termination ends of the plurality of posts was diamond ground so the YSZ was flush with the termination ends of the plurality of posts. The YSZ had porosity of about ≦1%.
The first composite thermal barrier coupon was tested by repeatedly heating and air cooling in increments of 10 cycles. The thermal barrier was then inspected for signs of damage or spalling of the insulation material from the void space around the metallic web. To record temperature and heat conduction across the thermal barrier during testing, a needle thermocouple was provided through a small hole in the bottom of the coupon to about 0.5 mm below the surface on which the thermal barrier was connected. During heating in each cycle, the thermal barrier coupon was direct flame heated for 30 seconds with a 2-stage Bethlehem Champion lampworking torch (using the inner burner ring only). The torch face was oriented normal to and 4 inches from the top of the thermal barrier. The torch was supplied with about 6 standard cubic feet per minute (SCFM) of natural gas and about 15 SCFM of pure oxygen gas. After 30 seconds of heating, the thermal barrier was cooled using a Vortec® Model 631 Cold Air Gun supplied with about 20 SCFM of air at about 12° C. The air gun was oriented normal to and 4 inches from the top of the thermal barrier. During each cycle, the thermal barrier coupon was cooled until the thermocouple read about 100° C. (i.e., about 3.5 minutes). The heating and cooling described above was considered a single cycle. After 10 cycles of each sequential heating and cooling operation, the first thermal barrier coupon was inspected under a microscope at about 10× magnification for visible signs of spalling, delamination of the thermal barrier from the coupon surface, and/or cracks in the metallic web or the insulation material (YSZ). The above described testing is considered severe thermal cycling and was conducted to simulate the most extreme conditions that a thermal barrier would experience in a combustion engine. Of course, in some engines a thermal barrier would not experience such extreme temperature swings and may exhibit improved performance over the results observed.
The above described testing of the first thermal barrier was repeated for 11 cycles. After the initial cycle, small visible cracks in the YSZ material were present between a majority of the plurality of posts. During the subsequent 10 cycles, the cracks in the YSZ material continued to propagate between all of the posts. Testing was stopped after 11 cycles despite no delamination or spalling of YSZ from the metallic web. The average peak temperature recorded by the thermocouple (at 0.5 mm below the surface on which the thermal barrier was connected) was about 467° C.
The second composite thermal barrier was prepared on a 50 mm square by 12 mm thick F75 Co—Cr block coupon. A metallic web of F75 Co—Cr was direct metal laser sintered onto the coupon similar to that pictured in
The testing described in Example 2 above was conducted on the second thermal barrier for 50 cycles. No visible signs of failure or delamination were observed. Testing was stopped after 50 cycles as it was determined the thermal barrier had provided superior resistance to spalling via thermal cycling. The average peak temperature recorded by the thermocouple (at 0.5 mm below the surface on which the thermal barrier was connected) was about 430° C. In was unexpected that the configuration of the metallic web of the second thermal barrier (titled hexagons) and the discrete portions of YSZ therebetween resulted in less thermal related cracking of the YSZ from thermal cycling.
The third composite thermal barrier was prepared on a 50 mm square by 12 mm thick F75 Co—Cr block coupon. A metallic web of F75 Co—Cr was direct metal laser sintered onto the coupon. The metallic web was in the shape of a plurality of 1.15 mm diameter posts (each 1.5 mm tall on the coupon flat surface) each equally spaced apart in a triangular array. The coupon also included a raised edge along its perimeter 1.15 mm wide and 1.5 mm tall. The metallic web included 90% OFA across its thickness (i.e., between the surface of the coupon and the termination ends of the plurality of posts). A 100 micron bond coat of NiCrAlY was applied to the surface of the coupon around the plurality of posts. YSZ was plasma sprayed over the NiCrAlY to fill the void space (defined by the OFA and the thickness of the plurality of the posts) around the metallic web. Excess YSZ extending above the termination ends of the plurality of posts was diamond ground so the YSZ was flush with the termination ends of the plurality of posts. The YSZ had porosity of about ≦1%. Finally, a 0.5 mm thick metal alloy (95% nickel, 5% aluminum) continuous skin was plasma sprayed over the top of the YSZ in contact with the termination ends of the plurality of posts.
The testing described in Example 2 above was conducted on the third thermal barrier for 50 cycles. After less than about 10 cycles, the skin was visibly cracked at the center of the coupon. After about 35 cycles, the skin was severely cracked and began separating or delaminating from the coupon at the center thereof. Testing was stopped after 50 cycles and no cracking or delamination of the YSZ insulation material was observed. The average peak temperature recorded by the thermocouple (at 0.5 mm below the surface on which the thermal barrier was connected) was about 343° C. These results were unexpected compared to the first thermal barrier because the barrier provided increased thermal resistivity as shown by the lower peak temperature below the barrier. Also, the presence of the metallic skin in the thermal barrier resulted in less visible thermal related cracking of the YSZ from thermal cycling.
The fourth composite thermal barrier was prepared on a 50 mm diameter by 12 mm thick F75 Co—Cr cylindrical coupon. A metallic web of F75 Co—Cr was direct metal laser sintered onto the coupon. The metallic web was in the shape of a plurality of 1.15 mm diameter posts (each 1.5 mm tall on the coupon flat surface) each equally spaced apart in a triangular array. The coupon also included a raised edge along its perimeter 1.15 mm wide and 1.5 mm tall. The metallic web included 90% OFA across its thickness. No bond coat or YSZ was included. Instead the void space was filled with air as the insulation material. A 0.5 mm thick Co—Cr continuous skin was direct metal laser sintered over the top of the metallic web in contact with the termination ends of the plurality of posts.
The testing described in Example 2 above was conducted on the fourth thermal barrier for 50 cycles. After about 30 cycles, the skin began to visibly crack. Testing was stopped after 50 cycles with no severe cracking or delamination of the skin. The average peak temperature recorded by the thermocouple (at 0.5 mm below the surface on which the thermal barrier was connected) was about 528° C.
The fifth composite thermal barrier was prepared on a 50 mm diameter by 12 mm thick F75 Co—Cr cylindrical coupon. A metallic web of F75 Co—Cr was direct metal laser sintered onto the coupon. The metallic web was in the shape of a plurality of 1.15 mm diameter posts (each 1.5 mm tall on the coupon flat surface) each equally spaced apart in a triangular array. The coupon also included a raised edge along its perimeter 1.15 mm wide and 1.5 mm tall. The metallic web included 75% OFA across its thickness. No bond coat or YSZ was included. Instead the void space was filled with air as the insulation material. A 0.5 mm thick Co—Cr alloy continuous skin was direct metal laser sintered over the top of the metallic web in contact with the termination ends of the plurality of posts.
The testing described in Example 2 above was conducted on the fourth thermal barrier for 70 cycles. No visible signs of cracking were observed after 50 heating and cooling cycles. After about 70 cycles, the skin began to visibly crack at the center of the coupon. Testing was stopped after 70 cycles with no severe cracking or delamination of the skin. The average peak temperature recorded by the thermocouple (at 0.5 mm below the surface on which the thermal barrier was connected) was about 570° C.
In this prophetic example, the five composite thermal barriers (from Examples 2-6 described above) would be subjected to pressure, pulsation, and/or vibration testing to simulate exposure of the composite thermal barrier to the combustion reaction and movement of fluids inside the combustion chamber. Specifically, a hydro pulse system could deliver from about 1 psi to about 5000 psi pressurized hydraulic fluid (or similar fluid chosen to mimic combustion reactants) at about 30 Hertz (or another frequency chosen to mimic an internal combustion engine) normal to the face of the five composite thermal barriers (from Examples 2-6 described above). It is expected that these five thermal barriers would behave similar to surfaces within combustion engines that do not include thermal barriers. That is, these five composite thermal barriers would resist metal fatigue and would effectively limit spalling or delamination of the insulation material into the engine. The inventors expect that the second composite thermal barrier of Example 3 would spall a small portion of the YSZ material from the hexagonal cells into the engine during this prophetic test. However, the inventors expect that it would not be a catastrophic failure in which all of the YSZ material would spall into the engine simultaneously and damage crucial internal engine parts. The inventors also expect that the third, fourth, and fifth composite thermal barriers (of Examples 4-6) which include a skin would even more effectively resist this testing and limit spalling or delamination of material into the engine.
Two conventional thermal barriers were prepared on coupons (simulating an engine internal surface) and tested under severe thermal cycling (up to 50 cycles) to demonstrate the superior thermal resistivity and spalling resistance of the composite thermal barriers of the present disclosure (and demonstrated in Example 2 above).
The first comparative thermal barrier was prepared on a 50 mm diameter by 12 mm thick F75 Co—Cr cylindrical coupon. No metallic web was applied thereto. A 100 micron bond coat of NiCrAlY was applied to the entire surface of the coupon. A YSZ layer was plasma sprayed over the NiCrAlY and diamond ground down to uniform 2.0 mm layer thereon. No metallic skin was applied.
The testing described in Example 2 above was conducted on the first comparative thermal barrier for 11 cycles. After the initial heating-cooling cycle, small visible cracks in the YSZ material were visible. After the subsequent 2 cycles, the cracks in the YSZ material continued to propagate and the YSZ began to delaminate from the surface of the coupon. Testing was stopped after 11 cycles as the YSZ almost completely delaminated from the coupon and it was determined that the thermal barrier had failed. The average peak temperature recorded by the thermocouple (at 0.5 mm below the surface on which the YSZ was connected) was about 369° C.
The second comparative thermal barrier was prepared on a 50 mm diameter by 12 mm thick F75 Co—Cr cylindrical coupon. No metallic web was applied thereto. A 100 micron bond coat of NiCrAlY was applied to the entire surface of the coupon. A YSZ layer was plasma sprayed over the NiCrAlY and diamond ground down to uniform 1.0 mm layer thereon. No metallic skin was applied.
The testing described in Example 2 above was conducted on the second comparative thermal barrier for 50 cycles. After the first 2 cycles, extensive “mud puddle” cracks across the surface of the YSZ material were visible. During the subsequent about 40 cycles, the “mud puddle” cracks in the YSZ material continued to propagate. Testing was stopped after 50 cycles. The average peak temperature recorded by the thermocouple (at 0.5 mm below the surface on which the YSZ was connected) was about 478° C.
In a third comparative example, a 50 mm diameter by 12 mm thick F75 Co—Cr cylindrical coupon was prepared to simulate an engine surface without any thermal barrier. The testing described in Example 2 above was conducted on the third comparative thermal barrier for 1 cycle. The average peak temperature recorded by the thermocouple (at 0.5 mm below the surface on which the flame was directed) was about 705° C.
In this prophetic comparative example, the 2 thermal barriers (from Comparative Examples 1-2 described above) would be subjected to pressure, pulsation, and/or vibration testing described in Prophetic Example 7 above. It is expected that these two composite thermal barriers would behave similar to conventional thermal barriers inside of combustion engines. That is, with the brittle ceramic material directly exposed to the pressurized fluid, the inventors expect rapid failure of the conventional thermal barrier and immediately delamination of the ceramic from the coupon surface.
As used herein, the singular forms “a,” “an,” and “the” include plural referents unless the context clearly dictates otherwise. Ranges can be expressed herein as from “about” one particular value, and/or to “about” another particular value. When such a range is expressed, examples include from the one particular value and/or to the other particular value. Similarly, when values are expressed as approximations, by use of the antecedent “about,” it will be understood that the particular value forms another aspect. It will be further understood that the endpoints of each of the ranges are significant both in relation to the other endpoint, and independently of the other endpoint.
It is also noted that recitations herein refer to a component of the present disclosure being “configured” or “adapted to” function in a particular way. In this respect, such a component is “configured” or “adapted to” embody a particular property, or function in a particular manner, where such recitations are structural recitations as opposed to recitations of intended use. More specifically, the references herein to the manner in which a component is “configured” or “adapted to” denotes an existing physical condition of the component and, as such, is to be taken as a definite recitation of the structural characteristics of the component.
It will be apparent to those skilled in the art that various modifications and variations can be made to the present disclosure without departing from the spirit and scope hereof. Since modifications combinations, sub-combinations and variations of the disclosed embodiments incorporating the spirit and substance of the disclosure may occur to persons skilled in the art, the present disclosure should be construed to include everything within the scope of the appended claims and their equivalents.
This application claims the benefit of priority under 35 U.S.C. §119 of U.S. Provisional Application Ser. No. 62/235,008 filed on Sep. 30, 2015, the content of which is relied upon and incorporated herein by reference in its entirety.
| Number | Date | Country | |
|---|---|---|---|
| 62235008 | Sep 2015 | US |
