Patent Application
20080159869
This application relates generally to gas turbine engine rotor blades and, more particularly, to methods and apparatus for fabricating a rotor assemblies.
Known gas turbine engine compressor rotor blades include airfoils having a leading edge, a trailing edge, a pressure side, a suction side, a root portion, and a tip portion. The pressure and suction sides connect at the airfoil leading and trailing edges, and span radially between the root and tip portions. An inner flow-path is defined at least partially by the root portion, and an outer flow-path is defined at least partially by a stationary casing coupled radially outward from the rotor blades. At least some known stationary casings include an abradable material that is spaced circumferentially within the casing and radially outward from the blade tip portion. At least some known compressors, for example, include a plurality of rows of rotor blades that extend radially and orthogonally outward from a rotor disk.
At least some known compressor rotor blades are coupled in a converging flow-path that may be susceptible to high airfoil radial loading and vibratory stresses generated by blade dynamic responses if the airfoil tips rub against the abradable casing. More specifically, such loading and stresses may be generated as a result of the rotor blade deflecting and rubbing the abradable casing. The blade dynamic response generally causes the airfoils to assume a first flex mode shape which results in high airfoil stresses at a peak location near the root portion of the airfoil. Moreover, generally the effect of tip rubs may be more severe to the airfoil when the suction side contacts the abradable casing rather than the pressure side.
In one aspect, a method for assembling a rotor assembly is provided. The method comprises providing a rotor blade including a first sidewall, a second sidewall, where the first and second sidewalls are connected at a leading edge and a trailing edge and extend in span from a root portion to a tip portion, removing blade material from the tip portion to form a tip portion rake angle that enables the tip portion to extend obliquely between the first and second sidewalls, and coupling the rotor blade to a shaft such that during tip rubs the tip portion rake angle facilitates reducing radial loading induced to the blade during tip rubs.
In another aspect, an airfoil for use in a rotor assembly is provided. The airfoil comprises a first sidewall, a second sidewall coupled to the first sidewall at a leading edge and at a trailing edge, a root portion, and a tip portion extending obliquely between the first and second sidewalls at an angle that facilitates reducing radial loading induced to the airfoil during tip rubs.
In a further aspect, a rotor assembly for use in a gas turbine engine is provided. The rotor assembly comprises a rotor shaft, and a plurality of rotor blades coupled to the rotor shaft such that each rotor blade comprises an airfoil portion comprising a first sidewall, a second sidewall coupled to the first sidewall at a leading edge and at a trailing edge, a root portion, and a tip portion extending obliquely between the first and second sidewalls at an angle that facilitates reducing radial loading induced to the airfoil during tip rubs.
The present invention provides an exemplary apparatus and method for fabricating a compressor rotor blade for a gas turbine engine. Specifically, in the exemplary embodiment, a booster compressor rotor blade is provided that includes a first sidewall, a second sidewall, a root portion and a tip portion. In the exemplary embodiment, the tip portion is oriented to facilitate reducing radial and axial loads induced to the rotor blade during pre-defined engine operations.
Although the present invention described herein is described in connection with the turbine engine shown in
During operation, air enters engine 10 through intake side 28 and flows through fan assembly 13 and compressed air is supplied from fan assembly 13 to booster compressor 14 and high pressure compressor 22. The plurality of rotor blades 40 compress the air and deliver the compressed air to core gas turbine engine 16. Airflow is further compressed by the high-pressure compressor 22 and is delivered combustor 24. Airflow from combustor 24 drives rotating turbines 18 and 26 and exits gas turbine engine 10 through exhaust side 30.
In the exemplary embodiment, an abradable material 32 is coupled to a casing circumferentially about rotor blades 40. Platform 55 defines an inner boundary 34 of a flow-path 35 extending through booster compressor 14, and abradable material 32 defines a radially outer boundary 36 of flow-path 35. In an alternative embodiment, inner boundary 34 may be defined by a rotor disk 20 (shown in
During normal engine operations, rotor disk 20 rotates within an orbiting diameter that is substantially centered about longitudinal axis 12. Accordingly, rotor blades 40 rotate about longitudinal axis 12 such that clearance gap 33 is substantially maintained and more specifically such that tip portion 60 remains a distance D1 from abradable material 32, with the exception of minor variations due to small engine 10 imbalances. Clearance gap 33 is also sized to facilitate reducing an amount of air i.e., tip spillage, that may be channeled past tip portion 60 during engine operation.
In the event of a deflection of blade 40, as shown hidden in
In the exemplary embodiment, blade 140 has a stacking axis 80. Moreover, in the exemplary embodiment, stacking axis 80 extends through blade 140 in a span-wise direction from root portion 54 to tip portion 160. Generally, and in some embodiments, axis 80 is substantially parallel with a line (not shown) extending through blade 140 in a span-wise direction which is substantially centered along a chord-wise cross-section (not shown) of airfoil 42. Tip surface 162 extends obliquely between airfoil sides 44 and 46. More specifically, tip surface 162 is oriented at a rake angle Θ. Rake angle Θ of tip surface 162 is measured with respect to a plane 82 extending through rotor blade 140 substantially perpendicular to stacking axis 80. Plane 82, as described in more detail below, facilitates the fabrication and orientation of tip surface 162. In one embodiment, during a fabrication process, plane 82 is established using a plurality of datum points defined on an external surface of blade 140. Alternatively, blade tip surface 162 may be oriented at any rake angle Θ that enables blade 140 to function as described herein.
In the exemplary embodiment, the orientation of tip surface 162, as defined by rake angle Θ, causes the clearance gap 33 to be non-uniform across blade tip portion 160. Specifically, in the exemplary embodiment, because tip surface 162 is oriented at rake angle Θ, a height D1 of clearance gap 33 at convex edge 166 is greater than a height D2 of clearance gap 33 at concave edge 164. In the exemplary embodiment, surface 162 is formed via a raking process. Alternatively, surface 162 may be formed at rake angle Θ using any other known fabricating process, including but not limited to, a machining process.
In the exemplary embodiment, an existing blade 40 may be modified to include tip portion 160. Specifically, excess blade material from an existing blade tip portion 60 is removed via a raking process to form tip portion 160 with a corresponding rake angle Θ that facilitates prevention of convex edge 166 contact with abradable material 32 during a maximum blade dynamic response. More specifically, in the exemplary embodiment, rake angle Θ is between about 5° to about 15°. In an alternative embodiment, blade 140 is formed with tip portion 160 having rake angle Θ via a known casting process, such that tip portion 160 is formed with a desired rake angle Θ.
During normal engine operations, the rotor disk 20 rotates within an orbiting diameter that is substantially centered about longitudinal axis 12. Accordingly, rotor blades 140 rotate about longitudinal axis 12, and a sufficient clearance gap 33 is maintained between rotor blade tip portion 160 and abradable material 32. In the event blade 140 is deflected, tip portion 160 may inadvertently rub abradable material 32. As shown as hidden in
In the exemplary embodiment, rake angle Θ is selected to facilitate preventing blade tip surface 162 from contacting the abradable material 32. Rather, because of rake angle Θ, during tip rubs, generally only concave edge 164 will contact the abradable material 32, and moreover, the contact will be at an angle which facilitates edge 164 cutting and removing material 32 rather than jamming into the material 32. As a result, radial blade loading and the blade dynamic response are facilitated to be reduced.
The above-described rotor blade facilitates reducing radial and axial loading induced to the blade during inadvertent tip rubs between the rotor blades and the abradable material. Specifically, the tip portion is oriented at a rake angle that enables the concave edge to contact the abradable material rather than the convex edge of the airfoil. Contact with the concave edge facilitates reducing radial and axial forces induced to the blade, as well as the flex and vibration of the blade. Reduction of blade flex and vibrations induced to the blade reduces the dynamic response of the blade and the likelihood of material fatigue at the first flex stress location. As such, a useful life of the blade is facilitated to be increased in a cost-effective and reliable manner.
Exemplary embodiments of rotor blades are described above in detail. The rotor blades are not limited to the specific embodiments described herein, but rather, components of each assembly may be utilized independently and separately from other components described herein. For example, each rotor blade component can also be used in combination with other blade system components, with other gas and non-gas turbine engines.
While the invention has been described in terms of various specific embodiments, those skilled in the art will recognize that the invention can be practiced with modification within the spirit and scope of the claims.
