Patent Application
20120189434
Gas turbine engines include compressor rotors having a plurality of rotating compressor blades. Minimizing the leakage of air, such as between tips of rotating blades and a casing of the gas turbine engine, increases the efficiency of the gas turbine engine because the leakage of air over the tips of the blades can cause aerodynamic efficiency losses. To minimize this, the gap at tips of the blades is set small and at certain conditions, the blade tips may rub against and engage an abradable seal at the casing of the gas turbine. The abradability of the seal material prevents damage to the blades while the seal material itself wears to generate an optimized mating surface and thus reduce the leakage of air.
Cantilevered vanes that seal against a rotor shaft are used for elimination of the air leakage and complex construction of vane inside diameter (ID) shroud, abradable seal and knife edges that are used in present gas turbine engines. Current cantilevered vane tip sealing experiences the difficulty that the tip gaps need to be set more open than desirable to prevent rub interactions that can cause rotor shaft coating spallation, vane damage or rotor shaft burn through due to thermal runaway events during rubs. Current materials have been found to lack the durability to prevent spallation and lack the abradability to prevent vane damage.
Blade outer seals do not have as many problems as inner seals, but do need to have the ability to resist fine particle erosion and have a suitable wear ratio between the seal and the airfoil.
It would be an advantage for an abradable coating for rotor that is capable of running against bare vane tips and have a desirable balance of wear between both the vane tips and the coating. The coating should also prevent catastrophic thermal runaway events, coating spallation and damage to the vanes.
The present invention comprises an abrasive coating forming a seal material on components of gas turbine engines. The present invention comprises an abrasive coating on the surface of the rotor to form a seal with the stator vanes and on the inside of the casing to form a seal with the rotor blades.
The abrasive coating contains ceramic particles in a composite matrix of hexagonal boron nitride (hBN) in nickel, cobalt, copper, iron or mixtures thereof. The ceramic particles are irregularly flattened shapes that are described as “splats” in the thermal spray field. The ceramic particles may be any ceramic that has a hardness of seven or more on the Mohs Scale for hardness, such as silica, quartz, alumina and zirconia.
The abrasive coating will often include a base bond coat layer. The bond coat may be MCr, MCrAl., MCrAlY or a refractory modified MCrAlY, where M is nickel, cobalt, iron or mixtures thereof.
When thermal protection is needed, there is also a layer between the abrasive coating and the bond coat comprising a ceramic layer that acts as a thermal barrier to protect the coated components. Ceramic layers include, for example, zirconia, hafnia, mullite, alumina.
Compressor 16 comprises stages of compressor vanes 26 and blades 28 arranged in low pressure compressor (LPC) section 30 and high pressure compressor (LPC) section 32. Turbine 20 comprises stages of turbine vanes 34 and turbine blades 36 arranged in high pressure turbine (HPT) section 38 and low pressure turbine (LPT) section 40. HPT section 38 is coupled to HPC section 32 via HPT shaft 42, forming the high pressure spool or high spool. LPT section 40 is coupled to LPC section 30 and fan 12 via LPT shaft 44, forming the low pressure spool or low spool. HPT shaft 42 and LPT shaft 44 are typically coaxially mounted, with the high and low spools independently rotating about turbine axis (centerline) CL.
Fan 12 comprises a number of fan airfoils circumferentially arranged around a fan disk or other rotating member, which is coupled (directly or indirectly) to LPC section 30 and driven by LPT shaft 44. In some embodiments, fan 12 is coupled to the fan spool via geared fan drive mechanism 46, providing independent fan speed control.
As shown in
In operation of turbine engine 10, incoming airflow F1 enters inlet 22 and divides into core flow FC and bypass flow FB, downstream of fan 12. Core flow FC propagates along the core flowpath through compressor section 16, combustor 18 and turbine section 20, and bypass flow FB propagates along the bypass flowpath through bypass duct 14.
LPC section 30 and HPC section 32 of compressor 16 are utilized to compress incoming air for combustor 18, where fuel is introduced, mixed with air and ignited to produce hot combustion gas. Depending on embodiment, fan 12 also provides some degree of compression (or pre-compression) to core flow FC, and LPC section 30 may be omitted. Alternatively, an additional intermediate spool is included, for example in a three-spool turboprop or turbofan configuration.
Combustion gas exits combustor 18 and enters HPT section 38 of turbine 20, encountering turbine vanes 34 and turbine blades 36. Turbine vanes 34 turn and accelerate the flow, and turbine blades 36 generate lift for conversion to rotational energy via HPT shaft 50, driving HPC section 32 of compressor 16 via HPT shaft 50. Partially expanded combustion gas transitions from HPT section 38 to LPT section 40, driving LPC section 30 and fan 12 via LPT shaft 44. Exhaust flow exits LPT section 40 and turbine engine 10 via exhaust nozzle 24.
The thermodynamic efficiency of turbine engine 10 is tied to the overall pressure ratio, as defined between the delivery pressure at inlet 22 and the compressed air pressure entering combustor 18 from compressor section 16. In general, a higher pressure ratio offers increased efficiency and improved performance, including greater specific thrust. High pressure ratios also result in increased peak gas path temperatures, higher core pressure and greater flow rates, increasing thermal and mechanical stress on engine components.
Bond coat 62 is thin, up to 10 mils (254 microns), more specifically ranging from about 3 mils to about 7 mils (about 76 to about 178 microns). Abrasive coating 66 may be about the same thickness as bond coat 64, again ranging from about 3 mils to about 7 mils (about 76 to about 178 microns), while some applications that have larger variation in tip clearance may require a thicker abrasive layer. Abrasive layer 66 may be as thick as 300 mils (7620 microns) in some applications.
The bond coat may be MCr, MCrAl., MCrAlY or a refractory modified MCrAlY, where M is nickel, cobalt, iron or mixtures thereof. For example, bond coat 62 may be 15-40% Cr 6-15% Al, 0.61 to 1.0%. Y and the balance is cobalt, nickel or iron and combinations thereof.
Top abrasive layer 66 is a low strength abradable composite matrix of a metal alloy such as Ni, Co, Cu, Al MCrAlY loaded with hexagonal boron nitride (hBN) into which flat ceramic particles have been added by thermal spraying. The amount of Ni to hBN in the abradable matrix ranges from about 30% to about 60% by volume, and more specifically about 40% to about 50% Ni by volume, with the balance being hBN. The Ni alloy, hBN (ahBN) and ceramic may be deposited as a coating by individually feeding the powders to one or more spray torches or by blending the two powders and air plasma spraying (APS). Other spray processes would also be effective, such as combustion flame spray, HVOF, HVAF, LPPS, VPS, HVPS and the like. As part of the coating is a quantity of ceramic that at least partially melts during the spray process to form disc like flat particles, or splat particles.
The ceramic particles may be any ceramic that has a hardness of seven or more on the Mohs Scale for hardness, such as silica, quartz, alumina and zirconia and that at least partially melts at the spray temperatures. The amount of ceramic in coating 66 ranges from about 1% to about 10% by volume. The amount of metal alloy will range from about 30% to about 60% and more specifically about 40% to about 50% Ni by volume, and the balance of 30% to about 70% by volume of hBN. During the spray application of coating 66, the porosity of coating 66 is controlled to be less than about 10% and even below 5% to decrease the aerodynamic effect.
Abrasive layer 66 may also be deposited on an intermediate thermally insulating layer to further protect the rotor shaft from burn through during excessive vane contact.
Optional ceramic layer 64, shown in
As can be seen from
Coating 66 and 76 has a high abradability during fast and/or deep rubs to prevent catastrophic runaway events and damage to turbine components. During low speed rub interactions when frictional heating is low, the ceramic particles result in the desired wear of airfoil tips. When the interaction rate and rub forces increase for any reason, including local vane material transfer, thermal growth and high interaction rates, rub forces may climb only to a limit. Coating 66 and 76 is designed to have a low enough strength to limit rub forces on the airfoils by abrading at contact pressures of less than about 1,000 psi. In one case, 1,000 psi coating strength relates to about 20 pounds per vane loading of compressor stators. Because the bulk coating must meet the durability requirements of the environment, such as the high G environment of the shaft outside diameter in a cantilevered vane sealing application, the abradable coating 66 and 76 has a strength of greater than about 300 psi. The dual nature of coating 66 and 76 provides high abradability when interaction rates and rub forces increase while also cutting the airfoil when interaction rates are low and the ceramic particles dominate the rub interaction.
While the invention has been described with reference to an exemplary embodiment(s), it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment(s) disclosed, but that the invention will include all embodiments falling within the scope of the appended claims.
