Patent Publication Number: US-10316673-B2

Title: CMC turbine blade platform damper

Description:
FIELD OF THE INVENTION 
     The present disclosure generally involves damping vibrations in a turbine. In particular embodiments, the damping system may be used to damp vibrations on the platforms of adjacent rotating blades made from ceramic matrix composite (CMC) materials using a CMC wedge damper. 
     BACKGROUND OF THE INVENTION 
     Turbines are widely used in a variety of aviation, industrial, and power generation applications to perform work. Each turbine generally includes alternating stages of peripherally mounted stator vanes and rotating blades. The stator vanes may be attached to a stationary component such as a casing that surrounds the turbine, and the rotating blades may be attached to a rotor located along an axial centerline of the turbine. A compressed working fluid, such as steam, combustion gases, or air, flows along a hot gas path through the turbine to produce work. The stator vanes accelerate and direct the compressed working fluid onto the subsequent stage of rotating blades to impart motion to the rotating blades, thus turning the rotor and performing work. 
     Each rotating blade generally includes an airfoil connected to a platform that defines at least a portion of the hot gas path. The platform in turn connects to a root that may slide into a slot in the rotor to hold the rotating blade in place. Alternately, the root may slide into an adaptor which in turn slides into the slot in the rotor. At operational speeds, the rotating blades may vibrate at natural or resonant frequencies that create stresses in the roots, adaptors, and/or slots that may lead to accelerated material fatigue. Therefore, various damper systems have been developed to damp vibrations between adjacent rotating blades. In some damper systems, a metal rod or damper is inserted between adjacent platforms, adjacent adaptors, and/or between the root and the adaptor or the rotor. At operational speeds, the weight of the damper seats the damper against the complementary surfaces to exert force against the surfaces and damp vibrations. 
     Higher operating temperatures generally result in improved thermodynamic efficiency and/or increased power output. Higher operating temperatures also lead to increased erosion, creep, and low cycle fatigue of various components along the hot gas path. As a result, ceramic material composite (CMC) materials are increasingly being incorporated into components exposed to the higher temperatures associated with the hot gas path. 
     However, as CMC materials become incorporated into the airfoils, platforms, and/or roots of rotating blades, the ceramic surfaces of the rotating blades more readily abrade with conventional metallic dampers. The increased abrasion of the CMC material by the metallic dampers may create additional foreign object debris along the hot gas path and/or reduce the mass of the dampers, reducing the damping force created by the dampers. Therefore, an improved system for damping vibrations in a turbine would be useful. 
     BRIEF DESCRIPTION OF THE INVENTION 
     Aspects and advantages of the invention will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the invention. 
     Damping systems are generally provided for a rotor blade platform. In one embodiment, the damping system includes a blade platform defining a damper pocket and a CMC wedge damper positioned within the damper pocket. The CMC wedge damper has at least one damper angled surface parallel to a longitudinal axis. The damper pocket comprises a pocket angled surface positioned about the at least one damper angled surface. 
     These and other features, aspects and advantages of the present invention will become better understood with reference to the following description and appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and, together with the description, serve to explain the principles of the invention. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       A full and enabling disclosure of the present invention, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended Figures, in which: 
         FIG. 1  is a schematic cross-sectional view of an exemplary gas turbine engine in accordance with an embodiment of the present disclosure; 
         FIG. 2  is an enlarged circumferential cross sectional side view of a high pressure turbine portion of a gas turbine engine in accordance with an embodiment of the present disclosure; 
         FIG. 3  is an axial view of two adjacent exemplary CMC rotor blade assemblies and an exemplary CMC; 
         FIG. 4  is a sectional view of the exemplary CMC damper of  FIG. 3  between the CMC rotor blade assemblies; 
         FIG. 5  is a plan view of an exemplary pocket formed between two adjacent platforms; 
         FIG. 6  is a perspective view of an exemplary embodiment of a CMC damper; 
         FIG. 7  is a perspective view of another exemplary embodiment of a CMC damper; 
         FIG. 8  is a perspective view an exemplary CMC turbine blade assembly showing the tabbed CMC damper; and 
         FIG. 9  is a perspective view of an exemplary triangular groove configured as a damper pocket on the CMC blade assembly platform. 
     
    
    
     Repeat use of reference characters in the present specification and drawings is intended to represent the same or analogous features or elements of the present invention. 
     DETAILED DESCRIPTION OF THE INVENTION 
     Reference now will be made in detail to embodiments of the invention, one or more examples of which are illustrated in the drawings. Each example is provided by way of explanation of the invention, not limitation of the invention. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made in the present invention without departing from the scope or spirit of the invention. For instance, features illustrated or described as part of one embodiment can be used with another embodiment to yield a still further embodiment. Thus, it is intended that the present invention covers such modifications and variations as come within the scope of the appended claims and their equivalents. 
     As used herein, the terms “first”, “second”, and “third” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components. The terms “upstream” and “downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction to which the fluid flows. 
     As used herein, the terms “axial” or “axially” refer to a dimension along a longitudinal axis of an engine. The term “forward” used in conjunction with “axial” or “axially” refers to moving in a direction toward the engine inlet, or a component being relatively closer to the engine inlet as compared to another component. The term “aft” used in conjunction with “axial” or “axially” refers to moving in a direction toward the engine exhaust nozzle, or a component being relatively closer to the engine exhaust nozzle as compared to another component. 
     As used herein, the terms “radial” or “radially” refer to a dimension extending between a center longitudinal axis of the engine and an outer engine circumference. The use of the terms “proximal” or “proximally,” either by themselves or in conjunction with the terms “radial” or “radially,” refers to moving in a direction toward the center longitudinal axis, or a component being relatively closer to the center longitudinal axis as compared to another component. The use of the terms “distal” or “distally,” either by themselves or in conjunction with the terms “radial” or “radially,” refers to moving in a direction toward the outer engine circumference, or a component being relatively closer to the outer engine circumference as compared to another component. As used herein, the terms “lateral” or “laterally” refer to a dimension that is perpendicular to both the axial and radial dimensions. 
     All directional references (e.g., radial, axial, proximal, distal, upper, lower, upward, downward, left, right, lateral, front, back, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise) are only used for identification purposes to aid the reader&#39;s understanding of the present invention, and do not create limitations, particularly as to the position, orientation, or use of the invention. Connection references (e.g., attached, coupled, connected, and joined) are to be construed broadly and may include intermediate members between a collection of elements and relative movement between elements unless otherwise indicated. As such, connection references do not necessarily infer that two elements are directly connected and in fixed relation to each other. The exemplary drawings are for purposes of illustration only and the dimensions, positions, order and relative sizes reflected in the drawings attached hereto may vary. 
     Referring now to the drawings,  FIG. 1  is a schematic cross-sectional view of an exemplary high-bypass turbofan type engine  10 , herein referred to as “turbofan”, as may incorporate various embodiments of the present disclosure. As shown in  FIG. 1 , the turbofan  10  has a longitudinal or axial centerline axis  12  that extends therethrough for reference purposes. In general, the turbofan  10  may include a core turbine or gas turbine engine  14  disposed downstream from a fan section  16 . 
     The gas turbine engine  14  may generally include a substantially tubular outer casing  18  that defines an annular inlet  20 . The outer casing  18  may be formed from multiple casings. The outer casing  18  encases, in serial flow relationship, a compressor section having a booster or low pressure (LP) compressor  22 , a high pressure (HP) compressor  24 , a combustion section  26 , a turbine section including a high pressure (HP) turbine  28 , a low pressure (LP) turbine  30 , and a jet exhaust nozzle section  32 . A high pressure (HP) shaft or spool  34  drivingly connects the HP turbine  28  to the HP compressor  24 . A low pressure (LP) shaft or spool  36  drivingly connects the LP turbine  30  to the LP compressor  22 . The (LP) spool  36  may also be connected to a fan spool or shaft  38  of the fan section  16 . In particular embodiments, the (LP) spool  36  may be connected directly to the fan spool  38  such as in a direct-drive configuration. In alternative configurations, the (LP) spool  36  may be connected to the fan spool  38  via a speed reduction device  37  such as a reduction gear gearbox in an indirect-drive or geared-drive configuration. Such speed reduction devices may be included between any suitable shafts/spools within engine  10  as desired or required. 
     As shown in  FIG. 1 , the fan section  16  includes a plurality of fan blades  40  that are coupled to and that extend radially outwardly from the fan spool  38 . An annular fan casing or nacelle  42  circumferentially surrounds the fan section  16  and/or at least a portion of the gas turbine engine  14 . It should be appreciated by those of ordinary skill in the art that the nacelle  42  may be configured to be supported relative to the gas turbine engine  14  by a plurality of circumferentially-spaced outlet guide vanes  44 . Moreover, a downstream section  46  of the nacelle  42  (downstream of the guide vanes  44 ) may extend over an outer portion of the gas turbine engine  14  so as to define a bypass airflow passage  48  therebetween. 
       FIG. 2  provides an enlarged cross sectioned view of the HP turbine  28  portion of the gas turbine engine  14  as shown in  FIG. 1 , as may incorporate various embodiments of the present invention. As shown in  FIG. 2 , the HP turbine  28  includes, in serial flow relationship, a first stage  50  which includes an annular array  52  of stator vane nozzles  54  (only one shown) axially spaced from an annular array  56  of turbine rotor blade assembly  58  (only one shown). The HP turbine  28  further includes a second stage  60  which includes an annular array  62  of stator vane nozzles  64  (only one shown) axially spaced from an annular array  66  of turbine rotor blades  68  (only one shown). The turbine rotor blade assemblies  58 ,  68  extend radially outwardly from and are coupled to the HP spool  34  ( FIG. 1 ). As shown in  FIG. 2 , the stator vane nozzles  54 ,  64  and the turbine rotor blade assemblies  58 ,  68  at least partially define a hot gas path  70  for routing combustion gases from the combustion section  26  ( FIG. 1 ) through the HP turbine  28 . 
     As further shown in  FIG. 2 , the HP turbine may include one or more shroud assemblies, each of which forms an annular ring about an annular array of rotor nozzles. For example, a shroud assembly  72  may form an annular ring around the annular array  56  of rotor blade assembly  58  of the first stage  50 , and a shroud assembly  74  may form an annular ring around the annular array  66  of turbine rotor blade assembly  68  of the second stage  60 . In general, shrouds of the shroud assemblies  72 ,  74  are radially spaced from blade tips  76 ,  78  of each of the rotor blade assembly  68 . A radial or clearance gap CL is defined between the blade tips  76 ,  78  and the shrouds. 
     Referring now to  FIG. 3 , a rotor blade assembly is depicted having a first CMC rotor blade assembly  102   a  and an adjacent second CMC rotor blade assembly  102   b . The first CMC rotor blade assembly  102   a  has an airfoil portion  102   a , a first platform portion  104   a , and a shank portion  106   a  with a dovetail attachment mechanism  108   a . Similarly, the second CMC rotor blade assembly  102   b  has an airfoil portion  102   b , a second platform portion  104   b , and a shank portion  106   b  with a dovetail attachment mechanism  108   b . Both the first and second CMC rotor blade assemblies  100   a ,  100   b  also include an axially upstream, or forward angel wing  107  and an axially downstream, or aft angel wing  109  (see  FIGS. 8 &amp; 9 ). 
     During engine operation, vibrations are induced in and between the first and second CMC rotor blade assemblies  100   a ,  100   b  including side-to-side, i.e., circumferential movement of the platform portions  104   a ,  104   b  that increase excitation stresses induced in the shank portions  106   a ,  106   b . A platform damping system  140  is positioned between adjacent portions of the platform portions  104   a ,  104   b . In the exemplary embodiment shown, CMC rotor blade assemblies  100   a ,  100   b  are unitarily formed as a single component via those CMC fabrication processes known in the art. However, in other embodiments, the CMC rotor blade assemblies  100   a ,  100   b  may be formed from separate components. 
       FIGS. 4 and 5  show an exemplary damping system  140  utilized between the adjacent first and second CMC rotor blade assemblies  100   a ,  100   b . As shown in  FIG. 4 , a CMC wedge damper  150  is generally shown defining a substantially triangular shape. More particularly, for the depicted embodiment of  FIG. 4 , the CMC wedge damper  150  has a rounded corner, equilateral-triangular cross section as viewed along the longitudinal axis. Further, the CMC wedge damper  150  is positioned within a damper pocket  142  defined recessed within a first side  120  of the first platform portion  104   a  of the first CMC rotor blade assembly  100   a . However, other shapes can be utilized with a corresponding pocket shape. For example,  FIG. 7  shows a CMC wedge damper  150  having a substantially triangular shape and including a pair of tabs  152 ,  154 , and the damper pocket  142  in  FIG. 5  illustrates an exemplary shape to house the CMC wedge damper  150  shown in  FIG. 7 . 
     In the embodiments shown in  FIGS. 4 and 5 , the damper pocket  142  is defined recessed within the first side  144  of the first blade platform portion  104   a . Through this positioning, the damper pocket  142  can allow for sufficient clearance between each pocket angled surface  148  and damper angled surface  146   a ,  146   b  of the CMC wedge damper  150  to allow for movement of the CMC wedge damper  150  within the damper pocket  142 . When the CMC wedge damper  150  is propelled radially outward in the damper pocket  142  by centrifugal force, e.g., during operation of turbofan  10  ( FIG. 1 ), the CMC wedge damper  150  assumes a consistent equilibrium position with one damper angled surface  146   a  slidingly engaging a pocket angled surface  148  (i.e., the outer pocket angled surface) and another or second damper angled surface  146   b  of the CMC wedge damper  150  slidingly engaging a radial surface  158  of a second side  145  of the second platform portion  104   b  of the adjacent second CMC rotor blade assembly  100   b . In this way, vibrational energy in the CMC rotor blade assemblies  100   a ,  100   b  may be dissipated or absorbed by the CMC wedge damper  150 . For this embodiment, the radial surface  158  of the second blade platform  104   b  extends substantially parallel to a radial direction R as shown best in  FIG. 4 . Further, as shown in  FIG. 4 , when the CMC wedge damper  150  is propelled radially outward within the damper pocket  142 , the second angled surface  146   b  of the CMC wedge damper  150  is oriented substantially parallel to the radial direction R and slidingly engages the radial surface  158  of the second blade platform  104   b . The damper pocket  142  in  FIG. 5  illustrates an exemplary shape to house the wedge damper shown in  FIG. 7 , as noted above. 
       FIGS. 6 &amp; 7  illustrate exemplary CMC wedge dampers  150 . Each of the CMC wedge dampers  150  generally have a triangular shaped body extending along a longitudinal axis L with at least one damper angled surface  146  parallel to the longitudinal axis L. As seen in  FIG. 7 , the CMC wedge damper  150  can also have a leading tab  152 , a trailing tab  154 , and/or a notched corner  156 . For the depicted embodiment of  FIG. 7 , the CMC wedge damper  150  defines the notched corner  156  extending between the leading tab and the trailing tab along a longitudinal direction extending parallel to the longitudinal axis L. The leading tab  152 , trailing tab  154  and notched corner  156  can be configured to offset the center of gravity of the wedge damper  150  to help control the positioning of the CMC wedge damper  150  during use. Additionally, the leading tab  152  and trailing tab  154  can prevent the CMC wedge damper  150  from sliding out of the pocket in the longitudinal direction during turbine operation. As shown, the trailing tab  154  and leading tab  152  have at least one tab angled surface  155  transverse to the longitudinal axis L. The leading tab  152  can be formed with at least one of a contact prong and a rounded crown. The trailing tab  154  can also be formed to have a protrusion. However, other shapes can be utilized for the tabs  152 ,  154 . Additionally, other features can be utilized on or within the body of the wedge damper  150 . 
     Referring to  FIGS. 8 &amp; 9 , two embodiments of the damping system  140  are shown that each include a triangular groove  162  formed into a flat vertical face  164  of first platform portion  104   a . Triangular groove  162  can be formed by any method that enables operation of the damping system  140  as described herein. In the exemplary embodiment, CMC wedge damper  150  nests in the triangular groove  162  when installed. CMC wedge damper  150  is sized, configured, and oriented on a flat vertical face  164  of a first platform portion  104   a  to be at least partially received and retained within triangular groove  162  of an adjacent flat vertical face  164  of second platform portion  104   b . Both the damper pocket  142  and triangular groove  162  are sized to receive and retain CMC wedge damper  150  without coupling methods such as welding, brazing, and fastener hardware. 
     In particular embodiments, the CMC wedge damper  150  is constructed from a CMC material that is similar to and/or compatible with the CMC material of the CMC rotor blade assemblies  100   a ,  100   b . For example, the CMC material may be a silicon based, non-oxide ceramic matrix composite. As used herein, “CMCs” refers to silicon-containing, or oxide-oxide, matrix and reinforcing materials. Some examples of CMCs acceptable for use herein can include, but are not limited to, materials having a matrix and reinforcing fibers comprising non-oxide silicon-based materials such as silicon carbide, silicon nitride, silicon oxycarbides, silicon oxynitrides, and mixtures thereof. Examples include, but are not limited to, CMCs with silicon carbide matrix and silicon carbide fiber; silicon nitride matrix and silicon carbide fiber; and silicon carbide/silicon nitride matrix mixture and silicon carbide fiber. Furthermore, CMCs can have a matrix and reinforcing fibers comprised of oxide ceramics. 
     An example of the damping performance of the CMC wedge damper  150  illustrates a new class of turbine blade vibratory damping response as compared to current metal dampers. Modeling results for the new CMC wedge damper determined that scaling up the damper stiffness to simulate the CMC material with a modulus ratio of 40.3/13=3.1, and scaling down the mass of the damper to simulate the CMC material with a density ratio of 0.102/0.317=0.32, the CMC wedge damper provided at least four times the undamped critical location vibratory response stress reduction of an otherwise identical damper but for being made from metals comprising superalloys of aluminum, iron, nickel, titanium, cobalt, chromium or mixtures thereof. These results apply for an undamped critical location stress of at least 4000 psi. 
     This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.