Patent Application
20080219835
This invention generally relates to a gas turbine engine, and more particularly to an abradable component for a gas turbine engine.
Gas turbine engines typically include a compressor section, a combustor section and a turbine section. Air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to add energy to expand the air and accelerate the airflow into the turbine section. The hot combustion gases that exit the combustor section flow downstream through the turbine section, which extracts kinetic energy from the expanding gases and converts the energy into shaft horsepower to drive the compressor section.
The compressor section of the gas turbine engine typically includes multiple compression stages to obtain high pressure levels. Each compressor stage consists of a row of stationary airfoils called stator vanes followed by a row of moving airflows called rotor blades. The stator vanes direct incoming airflow for the next set of rotor blades.
Gas turbine engine operation and efficiency is affected by a number of factors which include component design, manufacturing tolerance, engine clearances and rub interactions. Cantilevered compressor stator vanes are known which are attached at their radial outward end (i.e., the stator vanes are mounted at an end adjacent to the engine casing). A radial inward end of each stator is unsupported and is positioned adjacent to a rotor seal land extending between adjacent rotor stages.
Attempts have been made to decrease the amount of clearance between the tips of the cantilevered stator vanes and the rotor seal lands. For example, cantilevered stator vanes are known in which stator tips rub against an abrasive section inlaid in the rotor seal land during initial running of the engine such that the build clearance between the stator vanes and the rotor seal lands are chosen accordingly. Typically, a build clearance of at least approximately 0.005″ is established between the two components. Thus, the build clearance is such that the rotor seal lands only contact the tips of the stator vanes during the maximum closure point in the flight cycle (i.e., the point of a flight cycle where the rotor blades and the stator vanes experience maximum growth as a result of thermal expansion). Therefore, during a majority of the flight cycle, airflow escapes between the stator vanes and the rotor seal lands and may recirculate resulting in inefficiency and instability of the gas turbine engine. Further, during the initial running of the engine, excessive rub interaction between the stator vanes and the abrasive section of the rotor seal land may result in vane tip damage, mushrooming, metal transfer to adjacent rotors, and rotor burn through.
Accordingly, it is desirable to provide improved rub interaction between adjacent components of a gas turbine engine having a reduced clearance defined therebetween to improve engine efficiency and stability.
A gas turbine engine component includes an airfoil having a radial outward end and a radial inward end. A seal member is positioned adjacent to the radial inward end of the airfoil. A tip of the radial inward end of the airfoil is coated with an abradable material. The seal member is coated with an abrasive material.
A gas turbine engine includes an engine casing and a compressor section, a combustor section and a turbine section within the engine casing. At least one of the compressor section and the turbine section includes an airfoil and a seal member adjacent to the airfoil. A tip of the airfoil is coated with an abradable material and the seal member is coated with an abrasive material.
The various features and advantages of this invention will become apparent to those skilled in the art from the following detailed description. The drawings that accompany the detailed description can be briefly described as follows.
The compressor sections 14, 16 also include multiple disks 28 which rotate about engine centerline axis A to rotate the rotor blades 26. Each disk 28 includes a disk rim 30. Each disk rim supports a plurality of rotor blades 26. A seal member, such as a rotor seal land 32, extends from each disk rim 30 between adjacent disk rims 30 of adjacent rows of rotor blades 26.
In one example, the stator vanes 24 are cantilevered stator vanes. That is, the stator vanes 24 are fixed to an engine casing 40 or other structure at their radial outward end 34 and are unsupported at a radial inward end 36. The radial inward end 36 is directly opposite of the radial outward end 34. An airfoil 25 extends between the opposite ends 34, 36. A tip 38 of the radial inward end 36 of each stator 24 extends adjacent to a rotor seal land 32 which extends between adjacent disk rims 30. The radial outward end 34 is mounted to the engine casing 40 which surrounds the compressor section 14, 16, the combustor section 18, and the turbine sections 20, 22. The tip 38 of each stator 24 may contact the rotor seal land 32 to limit re-circulation of airflow within the compressor.
Referring to
The tips 38 of the stator vanes 24 are coated with an abradable material 42. Therefore, the tips 38 are more abradable than the remaining portions of the stator vanes 24 (i.e., the base metal of the stator vanes 24 is less abradable than the abradable material 42). Correspondingly, the exterior surface 44 of each rotor seal land 32 is coated with an abrasive material 46. The abradable material 42 is designed to deteriorate when subjected to friction and the abrasive material 46 is designed to cause irritation to the abradable material 42. Therefore, the abrasive material 46 deteriorates at a slower rate than the abradable material 42. The actual thickness of the coatings of the abradable material 42 and the abrasive material 46 will vary based upon design specific parameters including but not limited to the size and type of the gas turbine engine 10.
In one example, the abrasive material 46 is Cubic Boron Nitride. In another example, the abrasive material is Zirconium Oxide. The Zirconium Oxide may be a Yttria stabilized Zirconium. In one example, the Yttria stabilized Zirconium includes Zirc Oxide stabilized with about 11-14% Yttria. In another example, the Yttria stabilized Zirconium includes Zirc Oxide stabilized with about 6-8% Yttria. In still another example, the stabilized Zirconium Oxide includes Zirc Oxide stabilized with about 18.5-21.5% Yttria. The term “about” as used in this description relative to the compositions refers to possible variations in the compositional percentages, such as normally accepted variations or tolerances in the art. In yet another example, the abrasive material is Aluminum Oxide.
The abradable material 42 includes Zirconium Oxide, in one example. In another example, the abradable material 42 includes the Yttria stabilized Zirconium. It should be understood that other materials may be utilized for the abradable material 42 and the abrasive material 46. A person of ordinary skill in the art having the benefit of this disclosure would be able to select appropriate materials for use as the abradable material 42 and the abrasive material 46. As can be appreciated by those of skill in the art, the Zirconium Oxide is capable of use both as the abrasive material 46 and the abradable material 42. The Zirconium Oxide (i.e., the abrasive material 46) applied to the rotor seal land 32 will abrade the Zirconium Oxide (i.e., the abradable material 42) applied to the tips 38 of the stator vanes 24 in this example.
In one example, the abradable material 42 and the abrasive material 46 are applied by thermal spray. In another example, where the abrasive material 46 includes Cubic Boron Nitride, the abrasive material 46 is applied by a electroplating. Other application methods are also contemplated as within the scope of the present invention.
Use of the abradable material 42 on the tip 38 of each stator 24 and the abrasive material 46 on the rotor seal lands 32 allows the clearance X defined between the stator vanes 24 and the rotor seal lands 32 to be reduced. During operation of the gas turbine engine 10, the components of the gas turbine engine 10 may experience thermal expansion, centrifugal loading, and high maneuver loads during high angle of attack, takeoff and landing flight conditions. The stator vanes 24 may rub against the rotor seal lands 32 while experiencing conditions of this type. During this rub interaction, the abradable material 42 of the stator vanes 24 rubs against the abrasive material 46 applied on the rotor seal lands 32 causing a portion of the abradable material to turn to harmless fine dust.
Minimal heat is generated during the rub interaction between the stator vanes 24 and the rotor seal lands 32. The tighter clearances between the stator vanes 24 and the rotor seal lands 32 reduce the recirculation of airflow within the gas turbine engine thereby improving efficiency and component stability. In one example, the stator vanes 24 are in perfect contact (i.e., line to line contact) with the rotor seal lands 32 during engine operation (See
Although the example components including the abradable and abrasive coatings as illustrated herein are disclosed in association with a compressor section of the gas turbine engine, it should be understood that any other adjacent components of a gas turbine engine, including but not limited to turbine stator vanes and components with slider seal type engagements, may include the abradable and abrasive materials to provide tighter clearances and improved rub interactions between the adjacent components at those tighter clearances. That is, the invention is no limited to compressor stator vanes and is applicable to any gas turbine engine component.
The foregoing description shall be interpreted as illustrative and not in any limiting sense. A worker of ordinary skill in the art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.
