Patent Application
20240068371
The present disclosure generally relates to the field of turbomachines comprising high temperature components and to high resistance materials applied to such components, for example abrasive coatings and method of applying the same.
According to one embodiment, the present disclosure relates to axial, radial and mixed turbomachines, e.g. compressors and turbines, and more specifically to leakage control between the stationary and rotating components, and include abrasive materials applied to turbine rotor bucket or compressor rotor blade.
According to one embodiment, the present disclosure relates to abrasive coatings applied on rotor bucket tips to form a dynamic seal with the statoric part, called a shroud, to reduce the gas flow leakage and increase the efficiency of the gas turbine engine through the use of advanced materials and coatings with high temperature capability.
It is known that gas turbines generally include at least one stationary assembly extending over at least one rotor assembly. The rotor assembly includes at least one row of circumferentially spaced, rotatable, metallic turbine blades. The blades include metallic airfoils that extend radially outward from a rotatable hub to a metallic tip. Many of such metallic airfoils of rotor blades are fabricated from materials such as Nickel (Ni) based superalloys.
Stationary assemblies of turbomachines include surfaces that form metallic shrouds that may be routinely exposed to a hot gas flux. Some of such metallic surfaces include an applied metallic-based MCrAlY (where M=Co, Ni or Co/Ni, Cr=Chromium, Al=Aluminum and Y=Yttrium) coating and/or an applied ceramic thermal barrier coating that forms a shroud over the stationary assembly. Alternatively, some such metallic surfaces include applied ceramic matrix composites with, or without, a protective thermal barrier coating.
The metallic tips and the metallic shrouds define a tip clearance therebetween. However, such tip clearances are not suitable for high-temperature units that need high efficiencies. In order to reduce such tip clearances, gas turbines include abradable shrouds formed over the stationary assembly and the blade tips include an abrasive material formed thereon that has a greater hardness value than the blade material and the abradable coating. The abrasive material abrades the shroud coatings as the rotor assembly rotates within the stationary assembly. The abradable shroud coatings and the abrasive tips define a tip clearance therebetween. The tip clearance is small enough to facilitate reducing axial flow through the gas turbine that bypasses the blades, thereby facilitating increased efficiency and performance of the gas turbine. The tip clearance is also large enough to facilitate rub-free gas turbine operation through the range of available gas turbine operating conditions.
Various materials and processes have been suggested to provide a suitable abrasive tip cap on turbine stator and rotor blades. Typical abrasive materials used include silicon carbide, aluminum oxide, tantalum carbide and cubic boron nitride. The particles of abrasive material are usually incorporated with a metal matrix, including for example, nickel or cobalt-base alloys, to provide a sufficiently strong structure that can be bonded to the blade tip. However, the thickness of such a metal matrix is often limited because of the structural weakness of the abrasive composition.
In addition, some abrasive materials are damaged by high temperatures. As an example, for temperatures above approximately 927° C. (1700° F.), cubic boron nitride becomes unstable and is prone to oxidation. Also, while silicon carbide is better suited to survive temperatures in excess of approximately 927° C. (1700° F.), silicon carbide abrasives include free silicon that may attack the Ni/Co (Nickel/Cobalt) alloy substrates.
In some applications, it is conventional to apply the abrasive composition to the rotor blade tip using a thermal spray technique, such as plasma spraying or detonation gun spraying. Subsequent processes are typically necessary to provide the adhesion and structural integrity necessary for the abrasive composition to survive the hostile environment of a gas turbine. Such steps often include adhering the abrasive composition to the blade tip during a first heating and cooling cycle, and later depositing an additional quantity of the metal matrix over the abrasive composition through a second heating and cooling cycle, such as during hot isostatic pressing. As an alternative, it has also been suggested to melt the tip of the blade, such as with lasers, introduce the abrasive to the blade tip, and then re-solidify the blade tip.
While the above processes may be suitable for some turbine blade structures, turbine blade used in modern gas turbine engines are often fabricated from cast high temperature nickel-base superalloys having a single crystal microstructure. Single crystal blades are characterized by extremely high oxidation resistance and mechanical strength at elevated temperatures, which are necessary for the performance requirements of modern gas turbines. However, the single crystal microstructure must not be affected by the process by which the rotor blade abrasive tip caps are secured to the rotor blades. In particular, the process must not recrystallize the single crystal microstructure of the rotor blade, such that the high temperature properties of the rotor blade are lost or diminished. As a result, processes which entail melting the rotor blade tip to the single crystal rotor blade are entirely unacceptable. In addition, repeated thermal cycling of the rotor blade runs the risk of degrading the single crystal microstructure of the rotor blade.
Thus, it would be desirable to provide an abrasive composition which can be readily formed into an abrasive blade tip cap and which can be attached to a turbine rotor blade in a single heating and cooling cycle, under controlled temperature so as to minimize any degradation of the microstructure of a single crystal turbine rotor blade.
In one aspect, the subject matter disclosed herein is directed to an abrasive material preform configured to be fixedly coupled to a gas turbine rotor blade through a single heating and cooling cycle under controlled temperature.
In another aspect, the subject matter disclosed herein is directed to a method for producing such an abrasive material preform.
In yet another aspect, the subject matter disclosed herein is directed to a method for attaching such an abrasive material preform to a gas turbine blade in a single heating and cooling cycle to preserve the microstructure of a single crystal rotor blade and the stability of the abrasive material.
A more complete appreciation of the disclosed embodiments of the invention and many of the attended advantages thereof will be readily obtained as the same becomes better understood by reference to the following detailed description when considered in connection with the accompanying drawings, wherein:
In one aspect, the subject matter disclosed herein is directed to an abrasive material preform 11 configured to be fixedly coupled to a gas turbine rotor blade 10 through a single heating and cooling cycle under controlled temperature to realize a gas turbine blade 10 coated with an abrasive material preform 11 as shown in
According to one aspect, the subject matter disclosed herein is more specifically directed to a pre-sintered abrasive material preform 11 composed of a homogeneous mixture of a superalloy base material and braze alloy powders configured to be tack welded on a blade tip and then vacuum brazed, to realize a gas turbine blade 10 coated with an abrasive material preform 11 as shown in
In the present disclosure, the term powder is used according to its generally known meaning, to identify fine, dry, solid particles with mesh size between few to thousands of microns.
Additionally, in the present disclosure, the term sintering is also used according to its generally known meaning, to identify a process of compacting and forming a solid mass of material by heat or pressure without melting it to the point of liquefaction.
The term “preform” is used in the present disclosure to identify a preliminarily shaped component.
According to an exemplary embodiment, a pre-sintered preform can be a sintered powder metallurgy product composed of a bonding layer 12 composed of a homogeneous mixture of superalloy base material and braze alloy powders and of a top layer 13 or abrasive layer 13 composed of abrasive powders, also called abrasive grits, with a composition within the ranges of Table 1.
The metallic and abrasive powders are chosen to withstand high temperatures in gas turbine section. In particular, the abrasive grits ensure both short term cutting capability and thermal stability, assuring the clearance maintenance over time.
Powder particle size shall meet the following requirements:
In an exemplary embodiment of the system, the composition of the nickel braze alloy powder is referred to in Table 2.
In an exemplary embodiment of the system, the composition of the nickel based superalloy powder is referred to in Table 3.
According to an exemplary embodiment, a pre-sintered preform 11 is realized through the process shown in
According to an exemplary embodiment, a pre-sintered preform 11 is coupled to a gas turbine blade tip through the process shown in
In particular, as shown in
Alternatively, according to an exemplary embodiment, as shown in
According to an exemplary embodiment, the brazing step 60 of blade 10 with previously tack welded 50 preform 11 is carried out at 1200-1220° C. at a pressure lower than 5×10E-4 torr. According to an exemplary embodiment, also shown in
According to an exemplary embodiment, the brazing step 60 of blade 10 has to follow the following thermal cycle:
An important advantage of the exemplary embodiment of the presintered preforms is the possibility of using such preforms at high temperature, tested up to 980° C. metal temperature. The pre-sintered preforms can also be produced as net shape preforms, in order to reduce waste and be flexible for the application on axial, radial and mixed turbomachines.
An additional application of the pre-sintered preforms according to the exemplary embodiments herein disclosed might be an assembly of combustion liner and transition piece which slide past each other, the transition piece channelling the high-temperature gas from the combustion liner to a first statoric nozzle of a gas turbine.
Another application of the pre-sintered preforms according to the exemplary embodiments herein disclosed on gas turbine blades might be angel wing seals between a rotor blade and nozzle in a turbine, which inhibits ingestion of hot gas from a hot gas flow through the turbine into turbine wheel spaces.
Still another application of the pre-sintered preforms according to the exemplary embodiments herein disclosed is to realize sealing among rotating turbine components, stationary nozzles, and casing of a gas turbine, such as on J-seals. It is known that J-seals are an integral part of efficient steam turbine operation. The failure of a J-seal can cause significant damage to a turbine rotor as material migrates downstream. For that reason, plant staff must conduct inspections of steam path systems to identify potential problems during regularly scheduled outages in order to check the integrity of the sealing. Steam turbine efficiency relies heavily on integrity and performance of steam path stage-to-stage seals. Using abrasive pre-sintered preforms according to the exemplary embodiments herein disclosed can result in a significant advantage in sealing among rotating turbine components, stationary nozzles, and casing by allowing for a long-lasting integrity of seals.
| Number | Date | Country | Kind |
|---|---|---|---|
| 102021000000626 | Jan 2021 | IT | national |
| Filing Document | Filing Date | Country | Kind |
|---|---|---|---|
| PCT/EP2022/025007 | 1/10/2022 | WO |
