Patent Application
20070048140
This invention relates generally to gas turbine engines, and more specifically to methods and apparatus for assembling gas turbine engine components.
Accurate manufacturing of a component may be a significant factor in determining a fabricating time of the component. Specifically, when the component is a gas turbine engine blade, accurate manufacturing of the blade may be one of the most significant factors affecting an overall cost of fabrication of the gas turbine engine, as well as subsequent modifications, repairs, and inspections of the blade. For example, at least some known gas turbine engines include a compressor for compressing air which is mixed with a fuel and channeled to a combustor wherein the mixture is ignited within a combustion chamber for generating hot combustion gases. At least some known compressors include a rotor assembly that includes at least one row of circumferentially spaced rotor blades. Each rotor blade includes an airfoil that includes a pressure side, and a suction side connected together at leading and trailing edges. Each airfoil extends radially outward from a rotor blade platform. Each rotor blade also includes a dovetail that extends radially inward from a shank coupled to the platform. The dovetail is used to mount the rotor blade within the rotor assembly to a rotor disk or spool.
During operation, a pressure differential is created between the compressor blade pressure side and the compressor blade suction side which may result in an undesirable leakage flow between the upstream and downstream portions of the rotor. One such possible leakage path may form at an interconnection between each rotor blade and the rotor disk, where a gap may be defined between a blade base member, usually a dovetail design, and a rotor disk groove in which the rotor blades are carried.
Accordingly, at least one known gas turbine engine includes a silicone acetoxy sealant to facilitate sealing the blade base and the rotor disk. However, as engine performance requirements have increased, resulting in increased operating temperatures, however the known sealant may not withstand the increased operating temperatures for an extended period of time. As a consequence, the sealant degrades causing leakage to occur between the blade and the disk.
In one aspect, a method for assembling a gas turbine engine compressor having a plurality of stages and a plurality of blades coupled to each respective stage is provided. The method includes depositing a silicone oxime sealant onto at least a portion of a compressor blade, and coupling the compressor blade to a compressor disk such that the silicone oxime sealant is between the compressor blade and the compressor disk.
In another aspect, a gas turbine engine rotor assembly is provided. The gas turbine engine rotor assembly includes a rotor disk, a plurality of circumferentially-space rotor blades coupled to the rotor disk, and a silicone oxime sealant deposited onto at least a portion of the rotor blade such that the silicone oxime sealant is between the rotor blade and the disk.
In a further aspect, a gas turbine engine is provided. The gas turbine engine includes a rotor disk, a plurality of circumferentially-space rotor blades coupled to the rotor disk, and a silicone oxime sealant deposited onto at least a portion of the rotor blade such that the silicone oxime sealant is between the rotor blade and the disk.
As used herein, the terms “manufacture” and “manufacturing” may include any manufacturing process. For example, manufacturing processes may include grinding, finishing, polishing, cutting, machining, inspecting, and/or casting. The above examples are intended as exemplary only, and thus are not intended to limit in any way the definition and/or meaning of the terms “manufacture” and “manufacturing”. In addition, as used herein the term “component” may include any object to which a manufacturing process is applied. Furthermore, although the invention is described herein in association with a gas turbine engine, and more specifically for use with a compressor blade for a gas turbine engine, it should be understood that the present invention may be applicable to any component and/or any manufacturing process. Accordingly, practice of the present invention is not limited to the manufacture of compressor blades or other components of gas turbine engines.
A high pressure, multi-stage, axial-flow compressor 22 receives pressurized air from booster 20 and further increases the pressure of the air to a second, higher pressure level. The high pressure air flows to a combustor 24 and is mixed with fuel. The fuel-air mixture is ignited to raise the temperature and energy level of the pressurized air. The high energy combustion products flow to a first turbine 26 for driving compressor 22 through a first drive shaft 28, and then to a second turbine 30 for driving booster 20 through a second drive shaft 32 that is coaxial with first drive shaft 28. After driving each of turbines 26 and 30, the combustion products leave core engine 12 through an exhaust nozzle 34 to provide propulsive jet thrust.
Fan section 14 includes a rotatable, axial-flow fan rotor 36 that is driven by second turbine 30. An annular fan casing 38 surrounds fan rotor 36 and is supported from core engine 12 by a plurality of substantially radially-extending, circumferentially-spaced support struts 44. Fan rotor 36 carries a plurality of radially-extending, circumferentially spaced fan blades 42. Fan casing 38 extends rearwardly from fan rotor 36 over an outer portion of core engine 12 to define a secondary, or bypass airflow conduit. A casing element 39 that is downstream of and connected with fan casing 38 supports a plurality of fan stream outlet guide vanes 40. The air that passes through fan section 14 is propelled in a downstream direction by fan blades 42 to provide additional propulsive thrust to supplement the thrust provided by core engine 12.
Compressor 50 includes an inlet 66 that defines a flow passageway 67 having a relatively large flow area, and an outlet 68 that defines a relatively smaller area flow passageway 69 through which the compressed air passes. An outer boundary of the flow passageway is defined by an outer annular casing 70 and an inner boundary of the flow passageway is defined by the blade platforms of respective blades 58, 64 carried by rotors 56, 60, and also by a stationary annular seal ring 72 that is carried at an inner periphery of each of the respective stator sections 52, 54. As shown, respective rotor disks 56, 60 are ganged together by a suitable disk-to-disk coupling arrangement (not shown), and the third stage disk is connected with a drive shaft 74 that is operatively connected with a turbine rotor (not shown).
Each stator section 52, 54 includes an annular abradable seal that is carried by a respective annular sealing ring 72 and that is adapted to be engaged by respective labyrinth seals carried by 56, 60 in order to minimize air leakage around the respective stators 52, 54. Sealing rings 72 also serve to confine the flow of air to the flow passageway defined by outer casing 70 and the radially innermost surfaces of the respective stator vanes 47.
Rotor blade 64 includes a base member 100 that has a shape that corresponds substantially with that of circumferential slot 84. Base member 100 as shown is in the form of a dovetail and includes an enlarged base portion 110 that is received in lateral recesses 112, 114 formed in rotor slot 84. Base member 100 also includes a recessed portion 116, 118 on each side to receive the inwardly-extending convex projections 92, 94 of rotor slot 84. A blade platform 120 is carried on base member 100 and extends in a generally transverse direction relative to the longitudinal axis of base member 100. Extending longitudinally from upper surface 119 of blade platform 120, and in a direction opposite to that of base member 100, is an airfoil portion 122, which is adapted to contact the gases that pass through engine 10.
In the exemplary embodiment, sealant 150 is deposited on blade lower surface 160. After a predetermined quantity of time sufficient to cure sealant 150 has elapsed, blade 58 is coupled to disk 60. According, sealant 150 substantially seals gap 162 such that airflow cannot be channeled through gap 162.
In the exemplary embodiment, sealant 150 is a room temperature vulcanizing silicone oxime sealant that is deposited onto at least a portion of the compressor blade 58. Oxime as used herein is defined as one in a class of chemical compounds with the general formula R1R2CNOH, where R1 is an organic side chain and R2 is either hydrogen, forming an aldoxime, or another organic group, forming a ketoxime, and can be formed by the action of hydroxylamine on aldehydes or ketones.
Moreover, during use, sealant 150 is deposited onto at least a portion of blade 58 as a thixotropic paste. Thixotropic as used herein, is defined a gel-like substance that becomes a fluid when subjected to either stirring or shaking, and then returns to a semi-solid state upon standing. Accordingly, sealant 150 is applied to at least a portion of blade 58 in a semi-fluidic state. Sealant 150 is then allowed to cure or harden onto blade 58. After sealant 150 has substantially cured, blade 58 is coupled to disk 60. In the exemplary embodiment, sealant 150 is a room temperature oxime-cure silicone sealant such as for example, Loctite™ 5920. Accordingly, sealant 150 is capable of sealing gap 162 and retaining its elastomeric properties up to temperatures of at least 600 degrees Fahrenheit.
Described herein is an exemplary sealant that facilitates reducing and/or eliminating the airflow between a high pressure compressor rotor disk and a compressor rotor blade. More specifically, the sealant is applied to a plurality of compressor blades that are coupled to the first three stages of a gas turbine engine compressor assembly. The sealant described herein is a room temperature vulcanizing silicone oxime sealant that is configured to withstand temperatures to at least 600 degrees Fahrenheit.
More specifically, the sealant described herein facilitates improving gas turbine engine performance by preventing airflow leakage between the compressor blades and the compressor rotor disk. For example, known materials used in such applications cannot withstand operating temperatures of greater than approximately 600 degrees Fahrenheit for an extended period of time. As a consequence, leakage occurs when the known sealants material degrades with time and temperature and effectively disappears, thus eliminating the airflow seal around the component. Whereas the sealant described herein is configured to withstand temperatures greater than 600 degrees Fahrenheit and thus increase engine performance over an extended period of time.
While the invention has been described in terms of various specific embodiments, those skilled in the art will recognize that the invention may be practiced with modification within the spirit and scope of the claims.
