The present invention relates generally to couplings between components in a turbine engine, and, more particularly, to bicast couplings between turbine engine components, such as turbine blades and snubber assemblies.
A turbomachine, such as a steam or gas turbine is driven by a hot working gas flowing between rotor blades arranged along the circumference of a rotor so as to form an annular blade arrangement, and energy is transmitted from the hot working gas to a rotor shaft through the rotor blades. As the capacity of electric power plants increases, the volume of flow through industrial turbine engines has increased more and more and the operating conditions (e.g., operating temperature and pressure) have become increasingly severe. Further, the rotor blades have increased in size to harness more of the energy in the working gas to improve efficiency. A result of all the above is an increased level of stresses (such as thermal, vibratory, bending, centrifugal, contact and torsional) to which the rotor blades are subjected.
In order to limit vibrational stresses in the blades, various structures may be provided to the blades to form a cooperating structure between blades that serves to dampen the vibrations generated during rotation of the rotor. For example, mid-span snubber structures, such as cylindrical standoffs, may be provided extending from mid-span locations on the blades for engagement with each other. Two mid-span snubber structures are typically located at the same height on either side of a blade with their respective contact surfaces pointing in opposite directions. The snubber contact surfaces on adjacent blades are separated by a small space when the blades are stationary. However, when the blades rotate at full load and untwist under the effect of the centrifugal forces, snubber surfaces on adjacent blades come in contact with each other to dampen vibrations by friction at the contacting snubber surfaces.
In accordance with one aspect of the invention, a turbine blade assembly is provided in a turbine engine. The turbine blade assembly comprises a turbine blade having a pressure sidewall and an opposed suction sidewall and a first snubber assembly associated with one of the pressure sidewall and the suction sidewall. The first snubber assembly comprises a first base portion extending outwardly from the one of the pressure sidewall and the suction sidewall, and a first snubber portion. The first base portion is integrally cast with the turbine blade and includes first connection structure. The first snubber portion is bicast onto the first base portion and includes second connection structure that interacts with the first connection structure to substantially prevent separational movement between the first base portion and the first snubber portion.
In accordance with a second aspect of the invention, a coupling is provided between two components in a turbine engine. The coupling comprises a first component formed by a first casting procedure and including first connection structure comprising one of a continuous annular ridge and a continuous annular groove, and a second component bicast onto the first component during a second casting procedure and having second connection structure comprising the other of a continuous annular ridge and a continuous annular groove. The second connection structure interacts with the first connection structure to substantially prevent separational movement between the first component and the second component.
In accordance a third aspect of the invention, a method is provided for forming a turbine blade assembly. A first casting procedure is performed to form: a turbine blade having a pressure sidewall and an opposed suction sidewall; and a base portion of a snubber assembly, the base portion extending from one of the pressure sidewall and the suction sidewall. A second casting procedure is performed by casting a snubber portion onto the base portion such that first connection structure of the base portion interacts with second connection structure of the snubber portion to substantially prevent separational movement between the base portion and the snubber portion.
While the specification concludes with claims particularly pointing out and distinctly claiming the present invention, it is believed that the present invention will be better understood from the following description in conjunction with the accompanying Drawing Figures, in which like reference numerals identify like elements, and wherein:
In the following detailed description of the preferred embodiment, reference is made to the accompanying drawings that form a part hereof, and in which is shown by way of illustration, and not by way of limitation, a specific preferred embodiment in which the invention may be practiced. It is to be understood that other embodiments may be utilized and that changes may be made without departing from the spirit and scope of the present invention.
A tab 48 may extend from the pressure and/or suction sides of the end portion 42 to function in cooperation with an associated vane platform to define an origin for differential expansion and contraction of the platform in the chordwise dimension. Tab 48 may be located for example at a mid-chord position or at a maximum airfoil thickness position as shown in
The taper angle 44 may vary around the airfoil to accommodate varying amounts of differential contraction of the platform 50 and collar 52 at different points around the curvature of the airfoil. The taper angle on the pressure side 36 may be less than on the suction side in order to equalize pressure on the various contact surfaces. In an exemplary engineering model, a taper angle of 3-5 degrees on the pressure side and 50% greater than the pressure side taper angle on the suction side was found to be advantageous—for example, 4 degrees on the pressure side and 6 degrees on the suction side. The optimum angles depend on the airfoil shape.
The combination of stress relief slots 70, 72, spanwise clearance gap 55, and varying taper angles 44 provides substantially uniformly distributed contact pressures in the connection over a range of operating temperatures and differential thermal expansion conditions. The connection allows a limited range of relative movement, maintains a gas seal along the contact surfaces, minimizes vibration, minimizes stress concentrations, and provides sufficient contact area and pressure for rigidity and stability of the vane ring assembly.
The use of bi-casting enables less costly repair should the platform become damaged in service. The platform can be cut off, saving the high-value airfoil, and then a new replacement platform can be bi-cast onto the airfoil.
Referring now to
Each of the blades 114a, 114b further includes first and second snubber assemblies 124, 126 located generally mid-span between the blade root 116 and the blade tip 118 of each of the blades 114a, 114b.
The first snubber assembly 124 associated with the first blade 114a will now be described, it being understood that the first snubber assemblies 124 of the other blades 114 are substantially identical to the first snubber assembly 124 described herein. As shown in
As will be described herein, the first base portion 128 is cast integrally with the first blade 114a during a first casting procedure. The first base portion 128 and the first blade 114a may be formed, for example, from an equiaxed material or a directionally solidified material. More specifically, the base portion 128 and the first blade 114a may be formed, for example from a nickel based superalloy, such as CM247CC or CM247DS.
The first snubber portion 130 is cast onto the first base portion 128 of the first snubber assembly 124 during a second casting procedure, also referred to herein as a bicasting procedure or a bicast. The first snubber portion 130 is formed from a material that preferably has good oxidation, corrosion, and/or creep resistance characteristics. It is noted that the first snubber portion 130 may preferably be formed from the same/similar material as the first base portion 128 and the first blade 114a such that each of these components has the same or similar coefficient of thermal expansion, although the first snubber portion 130 may be formed from a different material than the first base portion 128 and the first blade 114a.
As shown in
As shown most clearly in
A second end portion 130B of the first snubber portion 130 in the embodiment shown defines a first angled surface 154a. The first angled surface 154a is spaced from a corresponding angled surface 154b of a snubber portion 130 of a snubber assembly 124 of the adjacent second blade 114b, such that a first space S1 or gap is formed therebetween, see
Referring now to
In accordance with an aspect of the present invention, the first base portion 128 of the first snubber assembly 124 includes an opening 166 formed therein, which opening 166 is in communication with the cooling passage 162 within the first blade 114a. The opening 166 may comprise a central bore formed in the first base portion 128, which may be formed during the first casting procedure or subsequent to the first casting procedure during a machining operation. The opening 166 receives cooling fluid from the cooling passage 162 for cooling the first snubber assembly 124.
The first snubber portion 130 in turn includes a cooling passageway 168 defined by the inner wall 144, which cooling passageway 168 is in communication with the opening 166 in the first base portion 128, see
The first snubber portion 130 also includes at least one cooling fluid outlet 170 that discharges cooling fluid in the cooling passageway 168 from the first snubber assembly 124, see
The second snubber assembly 126 associated with the first blade 114a extends from the suction sidewall 124 of the first blade 114a but is otherwise generally a mirror image of the first snubber assembly 124 associated with the first blade 114a. The second snubber assembly 126 (and the second snubber assemblies 126 associated with the remaining blades 114) will thus not be described in detail herein.
It is noted that while the illustrated first blade 114a includes the cooling circuit 160, the present invention is not intended to be limited to blades including cooling circuits. For example, the snubber assemblies 124, 126 described herein, which are formed by first and second casting procedures, could also be used with blades that are not cooled by internal cooling circuits. In such a case, the snubber assemblies 124, 126 would not need to include the openings 166 or cooling passageways 168.
During operation of the engine, centrifugal forces are exerted on the blades 114 and first and second snubber structures 124, 126 as a result of the rotation of the blades 114 with the rotor 110. These centrifugal forces cause the blades 114 to “untwist”, which causes the first and second angled surfaces 154a, 154b of the respective snubber structures 124, 126 to move toward each other to engage each other with a damping force. It should be noted that it is desirable to configure the snubber structures 124, 126 to produce a damping force that is sufficient to produce damping at the interface between the snubber structures 124, 126 to control blade vibration.
Referring to
The first blade 114a and the base portions 128 of the first and second snubber assemblies 124, 126 are formed during a first casting procedure. The resulting structure is illustrated in
Thereafter, the snubber portions 130 of the first and second snubber assemblies 124, 126 are formed during a second casting procedure, wherein the second connection structures 140 of the snubber portions 130 of the first and second snubber assemblies 124, 126 interact with the first connection structures 136 of the base portions 128 of the first and second snubber assemblies 124, 126 to substantially prevent separational movement between the snubber portions 130 and the base portions 128 of the first and second snubber assemblies 124, 126.
By forming the snubber assemblies 124, 126 via first and second casting procedures, difficulties associated with the formation of prior art snubber assemblies are avoided. For example, prior art attempts to cast the entire snubber assembly with the blade during a single casting procedure have resulted in snubber assemblies having significant defects or deformities, which must be repaired to be suitable for use. In addition prior art attempts at casting a base portion of a snubber assembly with the blade during a first casting procedure and then affixing a snubber portion to the base portion by welding, brazing, or other technique using a bonding material are expensive and often result in damage to the blade, base portion, and/or snubber portion. These difficulties of the prior art are avoided by the formation of snubber assemblies using two casting procedures as disclosed herein.
While particular embodiments of the present invention have been illustrated and described, it would be obvious to those skilled in the art that various other changes and modifications can be made without departing from the spirit and scope of the invention. It is therefore intended to cover in the appended claims all such changes and modifications that are within the scope of this invention.
This application is a Continuation-In-Part of U.S. patent application Ser. No. 12/752,460, filed Apr. 1, 2010 and entitled “TURBINE AIRFOIL TO SHROUD ATTACHMENT” by Christian X. Campbell et al., the entire disclosure of which is incorporated by reference herein.
This invention was made with U.S. Government support under Contract Number DE-FC26-05NT42644 awarded by the U.S. Department of Energy. The U.S. Government has certain rights to this invention.
Number | Date | Country | |
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Parent | 12752460 | Apr 2010 | US |
Child | 14071687 | US |