Patent Publication Number: US-2016230579-A1

Title: Rotor disk sealing and blade attachments system

Description:
FIELD OF INVENTION 
     The present disclosure relates to gas turbine engines, and, more specifically, to a rotor disk with integrated sealing and blade retention. 
     BACKGROUND 
     Gas turbine engines typically have alternating sets of rotors and stators in the compressor and turbine sections. The rotors may be disks that rotate adjacent to the stators. Sealing between the rotating rotors and the static stators may prevent gas-path air from moving between stages of a compressor or turbine outside of the gas path. A cover plate disposed on the rotating disks may provide sealing. The cover plate may be made separate from the rotor disk and disposed over the rotor disk. The cover plate may also lock a blade into the rotor disk. Adding a cover plate to each rotor in a turbine or compressor may increase the weight and cost of a turbine or compressor section, respectively. 
     SUMMARY 
     A rotor disk assembly comprises a circular body configured to rotate about an axis, a contoured slot formed partially through the circular body in an axial direction, and a protrusion extending radially from the circular body adjacent the contoured slot. 
     In various embodiments, the rotor disk assembly may further comprise a seal disposed on the protrusion. A rotating seal feature may extend from the circular body. The contoured slot may include squared edges. One of the squared edges may be parallel to a radial surface of the protrusion. A blade may be retained in the contoured slot. The blade may engage the protrusion to retain the blade axially within the contoured slot. 
     A turbine or compressor assembly is also provided. The turbine or compressor assembly may include a first disk configured to rotate about an axis, a first contoured slot formed partially through the first disk, a first protrusion adjacent to the first contoured slot and extending radially outward from the first disk, and a first blade disposed in the first contoured slot and configured to engage the first protrusion. 
     In various embodiments, the turbine or compressor assembly may further comprise a second disk aft of the first disk, a stator axially between the first disk and the second disk, and a brush seal extending radially inward from the stator. A landing may be coupled between the first disk and the second disk. The brush seal may extend toward the landing. A damper may be coupled between the stator and the brush seal. A second disk may be aft of the first disk, a stator may be axially between the first disk and the second disk, and a first knife seal may extend aft from the first disk towards an interface surface of the stator. A second knife seal may extend forward from the second disk towards the interface surface of the stator. The interface surface of the stator may include a honeycomb configured to deform in response to contact with the first knife seal and/or the second knife seal. The second disk may also include a second contoured slot formed partially through the second disk, a second protrusion adjacent to the first contoured slot and extending radially outward from the second disk, and a second blade disposed in the second contoured slot and configured to engage the second protrusion. The first protrusion may be aft of the first contoured slot and the second protrusion may be aft of the second contoured slot. 
     A disk sealing system is provided. The disk sealing system comprises a first disk including a first slot and a first protrusion configured to interface with a first blade, and a stator aft of the first disk. 
     In various embodiments, a first rotating seal feature extends aft from the first disk. The first rotating seal feature may have an annular shape. The stator may further comprise an interface surface and the rotating seal feature may contact the interface surface. The interface surface may comprise a honeycomb. A damper may extend radially inward from the stator and a seal may be disposed at an end of the damper. 
     The foregoing features and elements may be combined in various combinations without exclusivity, unless expressly indicated otherwise. These features and elements as well as the operation thereof will become more apparent in light of the following description and the accompanying drawings. It should be understood, however, the following description and drawings are intended to be exemplary in nature and non-limiting. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The subject matter of the present disclosure is particularly pointed out and distinctly claimed in the concluding portion of the specification. A more complete understanding of the present disclosure, however, may best be obtained by referring to the detailed description and claims when considered in connection with the figures, wherein like numerals denote like elements. 
         FIG. 1  illustrates an exemplary gas turbine engine, in accordance with various embodiments; 
         FIG. 2  illustrates a sealing system with rotating seal features formed integral with rotor disks, in accordance with various embodiments; 
         FIG. 3  illustrates a sealing system with a damper and brush seal between rotor disks, in accordance with various embodiments; 
         FIG. 4A  illustrates a partial cross section through a rotor disk having a retention slot to retain a blade on the rotor disk, in accordance with various embodiments; 
         FIG. 4B  illustrates a partial cross section through a rotor disk with a retention slot to retain a blade in an axial direction, in accordance with various embodiments; 
         FIG. 4C  illustrates a top view of a rotor disk comprising a retention slot with round corners, in accordance with various embodiments; 
         FIG. 5A  illustrates a partial cross section of a rotor disk assembly having a blade retained in the rotor disk, in accordance with various embodiments; and 
         FIG. 5B  illustrates a rotor disk assembly from forward looking aft and having a blade retained in the rotor disk, in accordance with various embodiments. 
     
    
    
     DETAILED DESCRIPTION 
     The detailed description of exemplary embodiments herein makes reference to the accompanying drawings, which show exemplary embodiments by way of illustration. While these exemplary embodiments are described in sufficient detail to enable those skilled in the art to practice the exemplary embodiments of the disclosure, it should be understood that other embodiments may be realized and that logical changes and adaptations in design and construction may be made in accordance with this disclosure and the teachings herein. Thus, the detailed description herein is presented for purposes of illustration only and not limitation. The scope of the disclosure is defined by the appended claims. For example, the steps recited in any of the method or process descriptions may be executed in any order and are not necessarily limited to the order presented. 
     Furthermore, any reference to singular includes plural embodiments, and any reference to more than one component or step may include a singular embodiment or step. Also, any reference to attached, fixed, connected or the like may include permanent, removable, temporary, partial, full and/or any other possible attachment option. Additionally, any reference to without contact (or similar phrases) may also include reduced contact or minimal contact. Surface shading lines may be used throughout the figures to denote different parts but not necessarily to denote the same or different materials. 
     As used herein, “aft” refers to the direction associated with the tail (e.g., the back end) of an aircraft, or generally, to the direction of exhaust of the gas turbine. As used herein, “forward” refers to the direction associated with the nose (e.g., the front end) of an aircraft, or generally, to the direction of flight or motion. 
     As used herein, “distal” refers to the direction radially outward, or generally, away from the axis of rotation of a turbine engine. As used herein, “proximal” refers to a direction radially inward, or generally, towards the axis of rotation of a turbine engine. 
     In various embodiments, a seal and disk system with retention structure to retain a blade in a disk as well as sealing structure to seal the gas path may eliminate use of cover plates. Sealing structure formed integral with disks may be cheaper and lighter than cover plates. Similarly, a seal and damper extending from a stator to arms extending from the disk may be less expensive and lighter than cover plates. Additionally, a slot formed partially though the disk and aligned with a protrusion may retain a blade in the disk without a cover plate. Thus, the turbine or compressor section housing a disk as described in the present disclosure may be simplified and made lighter than a disk with a cover plate. 
     Referring to  FIG. 1 , a gas turbine engine  100  (such as a turbofan gas turbine engine) is illustrated according to various embodiments. Gas turbine engine  100  is disposed about axial centerline axis  120 , which may also be referred to as axis of rotation  120 . Gas turbine engine  100  may comprise a fan  140 , compressor sections  150  and  160 , a combustion section  180 , and a turbine section  190 . Air compressed in compressor sections  150 ,  160  may be mixed with fuel and burned in combustion section  180  and expanded across turbine section  190 . Turbine section  190  may include high-pressure rotors  192  and low-pressure rotors  194 , which rotate in response to the expansion. Turbine section  190  may comprise alternating rows of rotary airfoils or blades  196  and static airfoils or vanes  198 . A plurality of bearings  115  may support spools in the gas turbine engine  100 .  FIG. 1  provides a general understanding of the sections in a gas turbine engine, and is not intended to limit the disclosure. The present disclosure may extend to all types of turbine engines, including turbofan gas turbine engines and turbojet engines, for all types of applications. 
     The forward-aft positions of gas turbine engine  100  lie along axis of rotation  120 . For example, fan  140  may be referred to as forward of turbine section  190  and turbine section  190  may be referred to as aft of fan  140 . Typically, during operation of gas turbine engine  100 , air flows from forward to aft, for example, from fan  140  to turbine section  190 . As air flows from fan  140  to the more aft components of gas turbine engine  100 , axis of rotation  120  may also generally define the direction of the air stream flow. 
     With reference to  FIG. 2 , sealing system  200  is shown with forward rotor disk  202  and aft rotor disk  204 . Forward rotor disk  202  may comprise blade platform  206  to support a blade. Aft rotor disk  204  may comprise a blade platform  208  to retain a blade. Stator  210  includes vane  212  and interface surface  218 . Rotating seal feature  216  may extend axially from forward rotor disk  202  towards interface surface  218  of stator  210 . In various embodiments, rotating seal feature  216  may be a knife edge seal. Rotating seal feature  216  may make contact with interface surface  218 . On a “green” run (i.e., first engine start up), rotating seal feature  216  may contact interface surface  218  as rotating seal feature  216  rotates with forward rotor disk  202 . Interface surface  218  may be a honeycomb surface and may deform as rotating seal feature  216  contacts interface surface  218 . 
     In various embodiments, a rotating seal feature  214  may also extend forward from aft rotor disk to interface surface  218  of stator  210 . Rotating seal feature  214  may make contact with interface surface  218  and contact interface surface  218  on the green run. Interface surface  218  may deform in response to rotating seal feature  214  contacting interface surface  218 . 
     In various embodiments, rotating seal feature  216  and rotating seal feature  214  may be formed integrally with forward rotor disk  202  and aft rotor disk  204 , respectively. Thus, rotating seal feature  216  and rotating seal feature  214  along with forward rotor disk  202  and aft rotor disk  204  may be made from a titanium alloy or a high-performance nickel based alloy (e.g., one of the nickel alloys available under the trade name INCONEL). The contour of rotating seal feature  216  and rotating seal feature  214  may be machined by turning. Rotating seal feature  216  and rotating seal feature  214  may be annular in shape with a portion of the rotating seal feature connecting to forward rotor disk  202  or aft rotor disk  204 . Rotating seal feature  216  and rotating seal feature  214  may seal turbine cavities from the gas path. 
     With reference to  FIG. 3 , a sealing system  240  comprising a seal  254  is shown, in accordance with various embodiments. Stator  250  may have damper  252  and seal  254  extending into the space between forward rotor disk  242  and aft rotor disk  244  and between forward platform  246  and aft platform  248 . Damper  252  may function as a seal having an annular wall and interface with seal  254 . Seal  254  may extend to landing  256  built onto arms  258  that attach forward rotor disk  242  to aft rotor disk  244 . Seal  254  may seal stages of the turbine or compressor from one another. Damper  252  may dampen vibration modes and provide support for seal  254  at an end of damper  252 . In various embodiments, seal  254  may be a brush seal, labyrinth seal, or non-contacting compliant seals. If seal  254  is a brush seal, for example, bristles from the brush seal may extend to and contact landing  256 . Seal  254  and damper  252  may form an annular seal structure with one a distal portion of damper  252  anchored to stator  250 . 
     With reference to  FIG. 4A , a partial cross section of a rotor disk  280  is shown with protrusion  284  to retain a blade. Rotor disk  280  may be integrated into the sealing systems depicted in  FIGS. 2 and 3 . Rotor disk  280  has a circular body portion  282  with protrusion  284  at a distal end of circular body portion  282 . Protrusion  284  may extend radially outward from rotor disk  280 . The distal end of rotor disk  280  has an axial length D 1 . Protrusion  284  of rotor disk  280  has an axial length D 2 . The ratio of D 1 /D 2  may be determined by structural requirements of different applications. In various embodiments, the ratio of D 1  to D 2  may be in the range from two to eight. Circumferential surface  286  of rotor disk  280  may serve as an interface surface for a blade to be attached to a distal end of rotor disk  280 . Radial surface  288  defined by a boundary of protrusion  284  may include a seal  290 . Seal  290  may be disposed between a later installed blade (i.e., installed on rotor disk  280 ) and a surface of rotor disk  280  to seal cooling air. The blade may be installed in contoured slot  300 , shown by ghosted lines. 
     With reference to  FIG. 4B , rotor disk  280  viewed in the direction from a high pressure side to a low pressure side (forward to aft in a turbine or aft to forward in a compressor) is shown, in accordance with various embodiments. Rotor disk  280  comprises a contoured slot  300  to interface with a turbine blade. Protrusion  284  extends above circumferential surface  286 . Seal  290  in radial surface  288  of protrusion  284  is configured to interface with a blade in rotor disk  280 . 
     With reference to  FIG. 4C , a top view of contoured slot  300  in rotor disk  280  is shown, in accordance with various embodiments. Contoured slot  300  extends partially through rotor disk  280 . Protrusion  284  at a low pressure side of rotor disk  280  may retain a blade in contoured slot  300 . Contoured slot  300  may be formed with a contoured disk that leaves rounded edges  310  in contoured slot  300 . Contoured slot  300  may be adjacent to protrusion  284  so that a line extending from contoured slot  300  at the aft most point of contoured slot  300  may be coplanar with radial surface  288  of protrusion  284 . Rounded edges may be removed or left in place depending on the desired shape of the blade to be retained in contoured slot  300 . Upon removing rounded edges, contoured slot  300  may have squared edges  312  and  314  with squared edge  314  parallel to radial surface  288  of protrusion  284 . 
     In various embodiments, contoured slot  300  may be formed using electrochemical machining (ECM), electrical discharge machining (EDM), and/or super abrasive machining (SAM). Contoured slot  300  may also be formed using conventional milling techniques. In various embodiments, SAM is carried out using a grind wheel having a contour similar to the contour of contoured slot  300  (as shown in  FIG. 4B ). EDM or ECM may be used to remove rounded edges  310  as desired. 
     In various embodiments, and with reference to  FIGS. 5A and 5B , a blade  320  is shown installed in rotor disk  280 . Blade  320  may have surface  322  to interface with radial surface  288  and seal  290 . Blade platform  324  may extend axially from protrusion  284 . Blade  320  may also include surface  326  to rest on and interface with circumferential surface  286  of rotor disk  280 . Blade  320  may extend into contoured slot  300  ( FIG. 4B ) with the surface of blade  320  having a contour matched to contoured slot  300 . Contoured slot  300  and protrusion  284  may retain blade  320  axially in rotor disk  280  during use without requiring a contour plate or other extra component to retain blade  320 . Protrusion  284  may be on a low pressure side of blade  320  so that the pressure differential between a high pressure side and low pressure side of blade  320  tends to force blade  320  into protrusion  284 . 
     Benefits and other advantages have been described herein with regard to specific embodiments. Furthermore, the connecting lines shown in the various figures contained herein are intended to represent exemplary functional relationships and/or physical couplings between the various elements. It should be noted that many alternative or additional functional relationships or physical connections may be present in a practical system. However, the benefits, advantages, and any elements that may cause any benefit or advantage to occur or become more pronounced are not to be construed as critical, required, or essential features or elements of the disclosure. The scope of the disclosure is accordingly to be limited by nothing other than the appended claims, in which reference to an element in the singular is not intended to mean “one and only one” unless explicitly so stated, but rather “one or more.” Moreover, where a phrase similar to “at least one of A, B, or C” is used in the claims, it is intended that the phrase be interpreted to mean that A alone may be present in an embodiment, B alone may be present in an embodiment, C alone may be present in an embodiment, or that any combination of the elements A, B and C may be present in a single embodiment; for example, A and B, A and C, B and C, or A and B and C. 
     Systems, methods and apparatus are provided herein. In the detailed description herein, references to “various embodiments”, “one embodiment”, “an embodiment”, “an example embodiment”, etc., indicate that the embodiment described may include a particular feature, structure, or characteristic, but every embodiment may not necessarily include the particular feature, structure, or characteristic. Moreover, such phrases are not necessarily referring to the same embodiment. Further, when a particular feature, structure, or characteristic is described in connection with an embodiment, it is submitted that it is within the knowledge of one skilled in the art to affect such feature, structure, or characteristic in connection with other embodiments whether or not explicitly described. After reading the description, it will be apparent to one skilled in the relevant art(s) how to implement the disclosure in alternative embodiments. 
     Furthermore, no element, component, or method step in the present disclosure is intended to be dedicated to the public regardless of whether the element, component, or method step is explicitly recited in the claims. No claim element herein is to be construed under the provisions of 35 U.S.C. 112(f), unless the element is expressly recited using the phrase “means for.” As used herein, the terms “comprises”, “comprising”, or any other variation thereof, are intended to cover a non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements does not include only those elements but may include other elements not expressly listed or inherent to such process, method, article, or apparatus.