Patent Publication Number: US-9840920-B2

Title: Methods and apparatus for sealing a gas turbine engine rotor assembly

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application is a non-provisional application and claims priority to U.S. Provisional Patent Application Ser. No. 61/660,307 filed Jun. 15, 2012 for “TURBINE BLADE PLATFORM SEAL”, which is hereby incorporated by reference in its entirety. 
    
    
     BACKGROUND 
     The application described herein relates generally to gas turbine engines components, and more specifically to an apparatus for sealing the gap between adjacent turbine blade platforms. 
     A typical gas turbine engine has an annular axially extending flow path for conducting air sequentially through a compressor section, a combustion section, and a turbine section. The compressor section includes a plurality of rotating blades which add energy to the air. The air exits the compressor section and enters the combustion section. Fuel is mixed with the compressed air, and the resulting combustion gases mixture is ignited to add more energy to the system. The resulting products of the combustion then expand through the turbine section. The turbine section includes another plurality of rotating blades, which extract energy from the expanding air. A rotor shaft interconnecting the compressor section and turbine section transfers a portion of this extracted energy back to the compressor section. The remainder of the energy extracted may be used to power a load, for example, a fan, a generator, or a pump. 
     At least some known rotor assemblies include 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 to a tip, and also includes a dovetail that extends radially inward from a shank extending between the platform and the dovetail. The dovetail is coupled to the rotor blade within the rotor assembly to a rotor disk. 
     The sides of platform sections of adjacent blades in a row of blades abut each other to form a portion of the boundary defining the flow path for the air and combustion gases. Although it would be desirable to have adjacent platforms abut in a perfect sealing relationship, the necessity to accommodate thermal growth and machining tolerances results in a small gap being maintained between adjacent platforms. 
     In order to couple the dovetail to the rotor disk, the dovetail must be machined to be slightly smaller than the slot into which it is inserted. This causes small buffer cavities in front and behind the dovetail. During operation of the turbine, cooling air may leak from the front buffer cavity, across the top of the disk, to the buffer cavity behind the dovetail, through the gap between aft skirts of adjacent rotor blades and into the flow path of the combustion gases. Leakage of the air into the flow path of the hot combustion gases causes a loss in the engine cycle and therefore decreases the engine efficiency. It is desirable to reduce this leakage to decrease specific fuel consumption, therefore increasing engine efficiency. 
     Accordingly, there exists a need to provide an improved device for sealing the gap between turbine rotor blade platforms of adjacent rotating blades in a gas turbine engine. 
     BRIEF DESCRIPTION 
     In one aspect, a rotor assembly for use in a gas turbine engine having an axis of rotation is provided. The rotor assembly includes a plurality of rotor blades. Each rotor blade includes a platform extending between opposing side faces, a shank extending radially inward from the platform, and a slot at least partially defined in each of the opposing side faces. A sealing member is configured to be inserted into each slot of a first rotor blade of the plurality of rotor blades such that at least a portion of each sealing member extends beyond one of the opposing side faces. A second rotor blade of the plurality of rotor blades is coupled adjacent the first rotor blade such that at least a portion of one sealing member is inserted into a corresponding second slot on the second rotor blade. 
     In another aspect, a gas turbine engine having an axis of rotation is provided. The gas turbine engine comprises a rotating shaft and a rotor assembly coupled to the shaft. The rotor assembly includes a plurality of rotor blades, and each rotor blade includes a platform extending between opposing side faces, a shank extending radially inward from the platform, and a slot at least partially defined in each of the opposing side faces. A sealing member is configured to be inserted into each slot of a first rotor blade of the plurality of rotor blades such that at least a portion of each sealing member extends beyond one of the opposing side faces. A second rotor blade of the plurality of rotor blades is coupled adjacent the first rotor blade such that at least a portion of one sealing member is inserted into a corresponding second slot on the second rotor blade. 
     In yet another aspect, a method of assembling a rotor assembly for use with gas turbine engine having an axis of rotation is provided. The method comprises providing a plurality of rotor blades. Each rotor blade includes a platform extending between opposing side faces, a shank extending radially inward from the platform, a dovetail extending radially inward from the shank, and a slot at least partially defined in each of the opposing side faces. A sealing member is inserted into each slot of a first rotor blade of the plurality of rotor blades such that at least a portion of each sealing member extends beyond one of the opposing side faces. A second rotor blade of the plurality of rotor blades is coupled adjacent the first rotor blade such that at least a portion of one sealing member is inserted into a corresponding second slot on the second rotor blade. 
    
    
     
       BRIEF DESCRIPTION 
         FIGS. 1-8  show exemplary embodiments of the turbine blade platform seal as described herein. 
         FIG. 1  is a schematic view of the components of a known gas turbine engine. 
         FIG. 2A  is a side view of a rotor blade that may be used with the gas turbine engine shown in  FIG. 1 . 
         FIG. 2B  is an axial front view of a rotor blade that may be used with the gas turbine engine shown in  FIG. 1 . 
         FIG. 3  is a radial top view of a seal pin sealing a gap between two rotor blades. 
         FIG. 4A  is an axial forward looking view of a seal pin sealing the gap between two rotor blades. 
         FIG. 4B  is a close up portion of  FIG. 4A  illustrating a seal pin sealing the gap between two rotor blades. 
         FIG. 5  is a tapered seal pin with a radially outer radius greater than a radially inner radius. 
         FIG. 6  is a perspective view of a rotor blade with a spline seal coupled thereto. 
         FIG. 7  is an axial forward looking cross-sectional view of a spline seal housed within a slot formed by adjacent rotor blades to seal the gap between rotor blades. 
         FIG. 8  is a perspective view of a portion of a rotor blade having an open ended slot to receive a spline seal. 
     
    
    
     DETAILED DESCRIPTION 
     As combustion air flows through the gas turbine engine, the pressure of the air is relatively higher upstream of the rotor blades than it is downstream of the rotor blades. Because of the pressure differential, some of the air flowing through the turbine may leak through a gap that exists between adjacent rotor blades and cause the engine to perform less efficiently than if the gap were sealed to prevent leakage. Similar seals exist in other applications, but embodiments of the present invention apply the use of a seal in a rotational environment. 
     Referring now to the drawings, in which like numerals refer to like elements throughout the several views,  FIG. 1  shows a schematic view of the components of a known gas turbine engine  10 . Gas turbine engine  10  may include a compressor  15  coupled in flow communication with a combustor  25  further coupled in flow communication with a turbine  40 . Compressor  15  and turbine  40  are each coupled to a rotor shaft  50 . Turbine  40  is also coupled to an external load  45  via rotor shaft  50  or an additional rotor shaft. Shaft  50  provides an axis of rotation for engine  10 . 
     During operation, compressor  15  compresses an incoming flow of air  20 . Compressor  15  delivers the compressed flow of air  20  to a combustor  25 . Combustor  25  mixes the compressed flow of air  20  with a flow of fuel  30  and ignites the mixture to create a flow of combustion gases  35 . Although only a single combustor  25  is shown, gas turbine engine  10  may include any number of combustors  25 . The flow of combustion gases  35  is in turn delivered to a turbine  40 . The flow of combustion gases  35  drives the turbine  40  so as to produce mechanical work. The mechanical work produced in the turbine  40  drives a rotor shaft  50  to power compressor  15  and any additional external load  45  such as an electrical generator and the like. 
     Gas turbine engine  10  may use natural gas, various types of syngas, and other types of fuels. Gas turbine engine  10  may be one of any number of different gas turbines offered by General Electric Company of Schenectady, N.Y. or otherwise. Gas turbine engine  10  may have other configuration and may use other types of components. Other types of gas turbine engines also may be used herein. Multiple gas turbine engines  10 , other types of turbines, and other types of power generation equipment may be used herein together. 
       FIG. 2A  is a side view of a rotor blade  200  that may be used with gas turbine engine  10  (shown in  FIG. 1 ). When blades  200  are coupled within a rotor assembly, such as turbine  40  (shown in  FIG. 1 ), a predetermined platform gap (not shown in  FIG. 2 ) is defined between circumferentially adjacent rotor blades  200 . In the exemplary embodiment, blade  200  has been modified to include features that provide a seal between blades  200  to be described in further detail below. 
     When coupled within rotor assembly  40 , each rotor blade  200  is coupled to a rotor disk (not shown) that is rotatably coupled to a rotor shaft, such as shaft  50  (shown in  FIG. 1 ). In an alternative embodiment, blades  200  are mounted within a rotor spool (not shown). In the exemplary embodiment, circumferentially adjacent blades  200  are identical and each extends radially outward from the rotor disk and includes an airfoil  202 , a platform  204 , a shank  206 , and a dovetail  208 . In the exemplary embodiment, airfoil  202 , platform  204 , shank  206 , and dovetail  208  are collectively known as a blade. 
       FIGS. 2A and 2B  illustrate a leading edge  210  and a trailing edge  212  of airfoil  202 . Leading edge  210  is on the forward side of airfoil  202 , and trailing edge  212  is on the aft side. As used herein, “forward” and “upstream” are used to refer to the inlet end of a turbine in a gas turbine engine, and “aft” and “downstream” are used to refer the to the opposite, outlet, end of a turbine in a gas turbine engine. 
     Platform  204  extends between airfoil  202  and shank  206  such that each airfoil  202  extends radially outward from each respective platform  204 . Shank  206  extends radially inwardly from platform  204  to dovetail  208 , and dovetail  208  extends radially inwardly from shank  206  to facilitate securing rotor blades  200  to the rotor disk. Platform  204  also includes a forward skirt  214  and an aft skirt  216  that are connected together with first slash face side  218  and an opposite second slash face side  220 . First slash face side  218  of shank  206  may include a cavity  222  for receiving a moveable element, for example, a moveable seal. It is contemplated that the moveable seal may be a seal pin  224 . 
       FIGS. 3-4B  show seal pin  224  within cavity  222  and operating to provide a seal configured to prevent cooling air from leaking between aft skirts  216  of adjacent rotor blades  200 . When rotor blades  200  are coupled within rotor assembly  40 , a platform gap  300  is defined between adjacent rotor blade platforms  204 . Centrifugal forces of rotating rotor assembly  40  cause seal pin  224  to seal platform gap  300  as described in further detail below. Cavity  222  is defined by a back surface  302 , a forward side surface  306 , an aft side surface  304 , a radially inner surface  402 , and a radially outer surface  404 . Back surface  302  and radially inner surface  402  are rounded in order to limit binding the movement of the ends of seal pin  224  within cavity  222 . Side surfaces  304  and  306  are angled such that they are wider at the opening of cavity  222  than where they connect to back surface  302 . Seal pin  224  contacts top surface  302  due to centrifugal force acting upon seal pin  224 . Top surface  404  is angled such that it directs seal pin  224  to fall toward the second slash face side  220  of adjacent rotor blade  200 . 
     Seal pin  224  is substantially circular in cross-section and extends radially within cavity  222 . In the exemplary embodiment, seal pin  224  has a diameter of approximately 0.04 inches. However, because the dimensions of rotor blade  200  may vary, depending on the engine size in which it is used, seal pin  224  may have any diameter sufficient to facilitate operation of rotor assembly  40  as described herein. Seal pin  224  is rounded at each of the two ends (best shown in  FIG. 4A ) to reduce binding with top surface  404  and bottom surface  402  during movement from a first position to a second position (shown in  FIG. 4A ). 
     Cavity  222  extends far enough into shank  206  to allow seal pin  224  to be housed substantially entirely within cavity  222 . In other words, seal pin  224  may include a maximum outside diameter that is less than the distance between the deepest portion of cavity  222  and a plane extending along first slash face side  218  of rotor blade  100 . Thus, seal pin  224  may be sufficiently recessed within cavity  222  to provide clearance for sliding an adjacent rotor blade into rotor disk. 
     Although only a single seal pin  224  is illustrated for rotor blade  200 , seal pin  224  may be positioned between each of opposing rotor blades  200  of a turbine stage. For example, a first turbine stage including seventy-two rotor blades  200  may include seventy-two seal pins  224 . 
     In operation, seal pin  224  initially sits at the bottom of cavity  222  such that the radially inner end of seal pin  224  is adjacent to bottom surface  402 . As rotor assembly  40  begins to rotate, centrifugal force slides seal pin  224  in a radially outward direction within cavity  222 . As seal pin  224  comes into contact with top surface  404 , the angle of top surface  404  forces seal pin  224  to fall against the flat second slash face surface  220  of the adjacent rotor blade  200 , forming a seal. To facilitate this seal top surface  404  has an angle of approximately 19 degrees. However, because the dimensions of rotor blade  200  may vary, depending on the engine size in which it is used, top surface  404  may have any angle sufficient to force seal pin  224  to fall against the flat second slash face surface  220  of the adjacent rotor blade  200 . In order to accommodate the angles defining the walls of cavity  222 , platform  204 , shank  206 , and slash face sides  220  and  218  are manufactured with a tilt of approximately 4 degrees from radially vertical. However, because the dimensions of rotor blade  200  may vary, depending on the engine size in which it is used, slash face sides  220  and  218  may have any angle sufficient to facilitate seal pin  224  in forming a seal. This slash face angle causes seal pin  224  to fall against the flat second slash face side  220  of the adjacent rotor blade  200 , such that the entire length of seal pin  224  is in contact with second slash face  220  to provide a continuous seal. Without the slash face angle, the moment caused by the rotating disc would cause only the radially outer tip of seal pin  224  to contact second slash face surface  220  of the adjacent rotor blade  200  while the radially inner end of pin  224  would remain within cavity  222 , and a seal would not be formed. 
     In another embodiment,  FIG. 5  shows a tapered seal pin  500  with a radially outer radius greater than a radially inner radius that functions in a similar manner as seal pin  224 . Tapered seal pin  500  may be used within the same cavity as shown in  FIGS. 3-4B . 
     Tapered seal pin  500  is substantially circular in cross-section and extends radially within cavity  222 . In the exemplary embodiment, tapered seal pin  500  has a radially outer diameter of approximately 0.08 inches and a radially inner diameter of approximately 0.04 inches. However, because the dimensions of rotor blade  200  may vary, depending on the engine size in which it is used, tapered seal pin  500  may have any diameter sufficient to permit passage of an adjacent rotor blade  200  during assembly. Tapered seal pin  500  is rounded at each of the two ends, for example, to reduce binding with top surface  404  and bottom surface  402  during movement from a first position to a second position (shown in  FIG. 4A ). 
     Centerline axis reference line  502  travels through a center of gravity  506  of tapered seal pin  500  to the centerline of engine  10  such that reference line  502  enters tapered seal pin  500  at the center of the radially outer tip and exits at the center of the radially inner tip. A second reference line  504  also travels through center of gravity  506  of tapered seal pin  500 , but reference line  504  is perpendicular to centerline of engine  10 . Phi is the angle measured between reference lines  502  and  504  at center of gravity  506  of tapered seal pin  500 . An angle where phi is greater than zero is required to cause tapered seal pin  500  to slide up cavity  222  and fall against the adjacent rotor blade  200 , described in further detail below. If phi is less than zero, then the moment created by the rotating disc causes the radially inner portion of tapered seal pin  500  to rotate away from the adjacent blade, and a seal is not formed. 
     Although only a single tapered seal pin  500  is illustrated for rotor blade  200 , it is contemplated that a tapered seal pin  500  may be positioned between each of opposing rotor blades  200  of a turbine stage. For example, a first turbine stage including seventy-two rotor blades  200  may include seventy-two tapered seal pins  500 . 
     In operation, tapered seal pin  500  initially sits at the bottom of cavity  222  such that the radially inner end of seal pin  224  is adjacent to bottom surface  402 . As rotor assembly  40  begins to rotate, centrifugal force slides tapered seal pin  500  in a radially outward direction within cavity  222 . As tapered seal pin  500  comes into contact with top surface  404 , the angle of top surface  404  forces tapered seal pin  500  to fall against the flat second slash face surface  220  of the adjacent rotor blade  200 , forming a seal. To facilitate tapered seal pin  500  forming a seal, top surface  404  has an angle of approximately 19 degrees. However, because the dimensions of rotor blade  200  may vary, depending on the engine size in which it is used, top surface  404  may have any angle sufficient to force tapered seal pin  500  to fall against the flat second slash face surface  220  of the adjacent rotor blade  200 . In the present embodiment, the taper of tapered seal pin  500  allows a seal to be formed against second slash face surface  220  of the adjacent rotor blade  200  without requiring platform  204 , shank  206 , and slash face sides  220  and  218  to be manufactured with a slash face angle. 
     Tapered seal pin  500  allows a seal to be created in platform gap  300  without modifying the angle of platform  204 , shank  206 , and slash face sides  220  and  218 . A seal is still created in platform gap  300  with platform  204 , shank  206 , and slash face sides  220  and  218  in a substantially vertical formation. 
       FIG. 6  shows a perspective view of yet another embodiment of the present invention where a spline seal  600  bridges gap  300  between adjacent circumferential rotor blades  200  of rotor assembly  40 . In the exemplary embodiment, blade  200  has been modified to include features that provide a seal between blades  200  to be described in further detail below. Spline seals are known to be used in turbines for sealing the gaps between the shrouds of adjacent stationary vanes. However, stationary vanes are not subject to centrifugal forces during operation of the turbine as such are rotor blades. Embodiments of the present invention apply the use of spline seal  600  in a rotational environment, such as rotor assembly  40 . In the exemplary embodiment, spline seal  600  may be a thin rectangular member having a height of approximately 0.3715 inches, a width of approximately 0.15 inches, and a thickness of approximately 0.01 inches in the axial direction. However, because the dimensions of rotor blade  200  may vary, depending on the engine size in which it is used, spline seal  600  may have any dimensions sufficient to prevent leakage of air through gap  300  between adjacent rotor blades  200 . Spline seal  600  may be formed of a high temperature alloy material having a forward surface  602  and an aft surface  604 . 
     In the exemplary embodiment, circumferentially adjacent blades  200  are identical and each extends radially outward from the rotor disk and includes an airfoil  202 , a platform  204 , a shank  206 , and a dovetail  208 . In the exemplary embodiment, airfoil  202 , platform  204 , shank  206 , and dovetail  208  are collectively known as a blade. Platform  204  extends between airfoil  202  and shank  206  such that each airfoil  202  extends radially outward from each respective platform  204 . Shank  206  extends radially inwardly from platform  204  to dovetail  208 , and dovetail  208  extends radially inwardly from shank  206  to facilitate securing rotor blades  200  to the rotor disk. 
     An aft portion of platform  204 , such as aft skirt  216 , includes a radially outward portion of a slot  608  that is machined into platform  204  to accept the radially outward portion of spline seal  600  near aft skirt  216 . A seal support structure  606  extends outward from shank  206  and includes a radially inward portion of slot  608  configured to accept the radially inward portion of spline seal  600 . Seal support structure  606  is positioned radially inward of platform  204  such that spline seal  600  may be inserted into slot  608  defined by seal support structure  606  and platform  204 . 
       FIG. 7  is a forward looking axial view of spline seal  600  housed within slot  608  formed by adjacent rotor blades  200  to seal gap  300  between rotor blades  200 . Rotor blade  200  includes identical structure on opposing sides such that opposing sides both include seal support structure  606  and platform  204 , which define slot  608 . Adjacent rotor blades  200  are identical such that adjacent rotor blades  200  each include opposing sides both having seal support structure  606  and platform  204 , which define slot  608 . Spline seal  600  is inserted into slot  608  in rotor blade  200  such that a portion of spline seal extends beyond the vertical plane defined by the side of platform  204 . Adjacent rotor blade  200  is then coupled to rotor blade  200  having spline seal  600  such that gap  300  is formed between adjacent rotor blades  200 . The portion of spline seal  600  extending beyond rotor blade is inserted into an identical slot  608  on adjacent rotor blade  200 , such that spline seal  600  bridges gap  300  and is fully contained within slot  608 , thus interlocking adjacent rotor blades  200 . 
     In operation, spline seal  600  initially sits at a radially inner portion of slot  608  such that a radially inner end  610  of spline seal  600  is in contact with a radially inner surface  609  of slot  608  on support structure  606  of adjacent rotor blades  200 . Slot  608  is angled such that, as rotor assembly  40  begins to rotate, centrifugal force causes spline seal  600  to move in a radially outward direction within slot  608 . A radially outer end  612  of spline seal  600  contacts a radially outer surface  611  of slot  608 , which acts to restrict further movement of spline seal  600  and keep spline seal  600  positioned within slot  608  to prevent the leakage of air between adjacent rotor blades  200 . Sealing is achieved when air pressure from the forward side of rotor blade  200  presses spline seal  600  into contact with the aft surfaces of slot  608 . This final position of spline seal  600  positions spline seal  600  to prevent leakage and also provides support to spline seal  600  to prevent buckling from the sustained high loads acting on forward seal surface  602  during operation. 
       FIG. 8  is a perspective view of a portion of rotor blade  200  having an open ended slot  802  to receive a spline seal  800 . Spline seals are known to be used in turbines for sealing the gaps between the shrouds of adjacent stationary vanes. However, stationary vanes are not subject to centrifugal forces during operation of the turbine as such are rotor blades. Embodiments of the present invention apply the use of a spline seal  800  in a rotational environment. Spline seal  800  may be a thin rectangular member having a height of approximately 0.3715 inches, a width of approximately 0.15 inches, and a thickness greater at the radially outer end than at the radially inner end. However, because the dimensions of rotor blade  200  may vary, spline seal  800  may have any dimensions sufficient to prevent leakage of air through gap  300  between adjacent rotor blades  200 . Spline seal  800  may be formed of a high temperature alloy material having a forward surface  806  and an aft surface  808 . 
     In the exemplary embodiment, circumferentially adjacent blades  200  are identical and each extends radially outward from the rotor disk and includes an airfoil  202 , a platform  204 , a shank  206 , and a dovetail  208 . In the exemplary embodiment, airfoil  202 , platform  204 , shank  206 , and dovetail  208  are collectively known as a bucket. Platform  204  extends between airfoil  202  and shank  206  such that each airfoil  202  extends radially outward from each respective platform  204 . Shank  206  extends radially inwardly from platform  204  to dovetail  208 , and dovetail  208  extends radially inwardly from shank  206  to facilitate securing rotor blades  200  to the rotor disk. 
     Slot  802 , having a retention feature  804  at the radially outer portion, is machined into an aft portion of platform  204  to accept the radially outward portion of spline seal  800 . The greater thickness of the radially outer portion of spline seal  800  fits into retention feature  804  of slot  802  such that spline seal  800  is locked in place. Slot  802  is open-ended at its radially inner portion such that retention feature  804  is the sole method of securing spline seal  800  in place. Spline seal  800  is supported by aft seal surface  808  being in contact with the aft surface of slot  802 , such that during operation, combustion gases press against forward seal surface  806  of spline seal  800  to secure aft surface  808  against the aft surface of slot  802 . This final position of spline seal  800  places spline seal  800  in a location to prevent leakage and also provides support to spline seal  800  to prevent buckling from the sustained high loads acting on forward seal surface  806  during operation. 
     The seal pin  224 , tapered seal pin  500 , and spline seals  600  and  800  each provide an effective seal across gap  300  between adjacent rotor blades  200  thereby preventing the leakage of air under blade platforms  204  and increasing the efficiency of the engine. 
     Exemplary embodiments of turbine blade platform seals are described above in detail. The seals are not limited to the specific embodiments described herein, but rather, components of systems may be utilized independently and separately from other components described herein. For example, the seals may also be used in combination with other turbine systems, and are not limited to practice with only the turbine engine systems as described herein. Rather, the exemplary embodiment can be implemented and utilized in connection with many other turbine engine applications. 
     Although specific features of various embodiments of the invention may be shown in some drawings and not in others, this is for convenience only. In accordance with the principles of the invention, any feature of a drawing may be referenced and/or claimed in combination with any feature of any other drawing. 
     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 have 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 language of the claims.