Patent Publication Number: US-11041397-B1

Title: Non-metallic side plate seal assembly for a gas turbine engine

Description:
BACKGROUND 
     The present disclosure relates to a gas turbine engine and, more particularly, to a seal therefor. 
     Gas turbine engines typically include a compressor section to pressurize flow, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases. The combustion gases commonly exceed 2000 degrees F. (1093 degrees C.). 
     Cooling of engine components is performed via communication of cooling flow through airfoil cooling circuits. Gas path recirculation between static and rotating components may be caused by local, circumferential pressure variations. Bow waves from airfoil leading edges create higher static pressure locally in front of the airfoil and wakes that exit airfoils create local pressure and velocity gradients which interact with the down-stream airfoils. Due to limitations of blade platform overhangs, especially on high speed turbines, the circumferential pressure variation can extend past the flowpath edge, which may cause a cavity, defined as the space between a static and rotating body, to be exposed to cyclic pressure fluctuations. Such pressure fluctuations may cause hot gases to be pushed into the cavities, with potential detrimental effects such as excessive heating of the components. 
     To prevent or minimize the amount of hot gas ingestion, secondary cooling airflow system pressure may be increased to generate a net positive outflow. This increase in pressure may result in a significant loss in cycle efficiency. To minimize such cycle losses, the size of the cavity closest to the flowpath is minimized. Often this requires the rotating knife-edges of a turbine rotor to operate close to the flowpath, and a significant quantity of secondary cooling airflow to the knife edge seals and outer regions. 
     SUMMARY 
     A side plate seal assembly for a gas turbine engine according to one disclosed non-limiting embodiment of the present disclosure includes a multiple of non-metallic side plate seals that are arranged about an axis of the gas turbine engine to form a full hoop seal, each of the multiple of side plate seals comprise a retention surface and a knife edge seal surface that extends at an angle therefrom. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that the multiple of non-metallic side plate seals that are arranged about the axis each interface one to another via a shiplap interface. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that the multiple of non-metallic side plate seals are manufactured of a ceramic matrix composite (CMC). 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that the multiple of non-metallic side plate seals are manufactured of an organic matrix composite (OMC). 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that the knife edge seal surface extends from the retention surface at the angle between 130-160 degrees. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that the retention surface is generally planar. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that the retention surface tapers to an inner diameter surface. 
     A rotor assembly for a gas turbine engine according to one disclosed non-limiting embodiment of the present disclosure includes a rotor disk that defines an axis; a full hoop cover plate; and a non-metallic side plate seal assembly at least partially between the rotor disk and the full hoop cover plate, the non-metallic side plate seal assembly comprises a multiple of non-metallic side plate seals that are arranged about the axis. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that the rotor disk and full hoop cover plate are manufactured of a metallic alloy. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that the multiple of non-metallic side plate seals each interface one to another via a shiplap interface. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that the full hoop cover plate forms at least one knife edge seal and the non-metallic side plate seal assembly forms at least one knife edge seal, the non-metallic side plate knife edge seal outboard of the full hoop cover plate knife edge seal with respect to the axis. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that an outer diameter edge of a retention surface of the non-metallic side plate seal assembly abuts a platform of a rotor blade retained in the disk. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that the non-metallic side plate knife edge seal interfaces with a seal surface attached an inner vane platform, the inner vane platform downstream of the rotor disk. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that a lower surface that includes an inner diameter edge of the retention surface is sandwiched between the rotor disk and the full hoop cover plate. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that each of the multiple of non-metallic side plate seals are identical. 
     A gas turbine engine according to one disclosed non-limiting embodiment of the present disclosure includes a rotor disk along an engine axis; an inner vane platform adjacent to the rotor disk; a seal surface attached an inner vane platform; and a non-metallic side plate seal assembly, the non-metallic side plate seal assembly comprises a retention surface adjacent to the rotor disk and a knife edge seal surface that extends at an angle from the retention surface to interface with the seal surface. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that the non-metallic side plate seal assembly comprises a multiple of non-metallic side plate seals that are identical. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that the multiple of non-metallic side plate seals each interface one to another via a shiplap interface. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that an outer diameter edge of a retention surface of the non-metallic side plate seal assembly abuts a platform of a rotor blade retained in the disk. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes that a lower surface that includes an inner diameter edge of the retention surface is sandwiched between the rotor disk and a full hoop cover plate, wherein the full hoop cover plate forms at least one knife edge seal inboard of non-metallic side plate knife edge seal with respect to the engine axis. 
     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 appreciated; however, the following description and drawings are intended to be exemplary in nature and non-limiting. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Various features will become apparent to those skilled in the art from the following detailed description of the disclosed non-limiting embodiments. The drawings that accompany the detailed description can be briefly described as follows: 
         FIG. 1  is a schematic cross-section of an example gas turbine engine architecture. 
         FIG. 2  is an schematic cross-section of an engine turbine section including a side plate seal assembly. 
         FIG. 3  is a partial perspective view of the side plate seal assembly. 
         FIG. 4  is a partial perspective view of the side plate seal assembly illustrating the segments thereof. 
         FIG. 5  is a perspective view of one segment of the side plate seal assembly. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as a two-spool turbo fan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . The fan section  22  drives air along a bypass flowpath while the compressor section  24  drives air along a core flowpath for compression and communication into the combustor section  26  then expansion through the turbine section  28 . Although depicted as a turbofan in the disclosed non-limiting embodiment, the concepts described herein may be applied to other turbine engine architectures such as turbojets, turboshafts, and three-spool (plus fan) turbofans. 
     The engine  20  generally includes a low spool  30  and a high spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine case structure  36  via several bearing structures  38 . The low spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a low pressure compressor (“LPC”)  44  and a low pressure turbine (“LPT”)  46 . The inner shaft  40  drives the fan  42  directly or through a geared architecture  48  to drive the fan  42  at a lower speed than the low spool  30 . An exemplary reduction transmission is an epicyclic transmission, namely a planetary or star gear system. 
     The high spool  32  includes an outer shaft  50  that interconnects a high pressure compressor (“HPC”)  52  and high pressure turbine (“HPT”)  54 . A combustor  56  is arranged between the high pressure compressor  52  and the high pressure turbine  54 . The inner shaft  40  and the outer shaft  50  are concentric and rotate about the engine central longitudinal axis A which is collinear with their longitudinal axes. 
     Core flow is compressed by the LPC  44  then the HPC  52 , mixed with the fuel and burned in the combustor  56 , then the combustion gasses are expanded over the HPT  54  and the LPT  46 . The turbines  46 ,  54  rotationally drive the respective low spool  30  and high spool  32  in response to the expansion. The main engine shafts  40 ,  50  are supported at a plurality of points by bearing assemblies  38  within the engine case structure  36 . 
     With reference to  FIG. 2 , an enlarged schematic view of a portion of the turbine section  28  is shown by way of example. A full ring shroud assembly  60  within the engine case structure  36  supports a blade outer air seal (BOAS) assembly  62 . The blade outer air seal (BOAS) assembly  62  contains a multiple of circumferentially distributed BOAS  64  proximate to a rotor assembly  66 . The full ring shroud assembly  60  and the blade outer air seal (BOAS) assembly  62  are axially disposed adjacent to a stationary vane ring  68 . The vane ring  68  includes an array of vanes  70  between a respective inner vane platform  72  and an outer vane platform  74 . The stationary vane ring  68  may be mounted to the engine case structure  36  by a multiple of segmented hooked rails  76  that extend from the outer vane platform  74 . The vane rings  68  align the flow while the rotor assembly  66  collects the energy of the working medium combustion gas flow to drive the turbine section  28  which in turn drives the compressor section  24 . One rotor assembly  66  and one downstream stationary vane ring  68  are described in detail as representative of any number of multiple engine stages. 
     The rotor assembly  66  includes an array of blades  84  circumferentially disposed around a disk  86 . While the description below refers to “blades” in the turbine section, the seal configurations are applicable to both buckets and blades in the respective turbine and compressor sections of turbomachines. It will be appreciated that the term “bucket” usually refers to the airfoil-shaped components employed in the turbine section(s) of turbomachines, while the term “blade” usually refers to the airfoil-shaped components typically employed in the compressor section of the machines. 
     Each blade  84  includes a root  88 , a platform  90  and an airfoil  92 . The blade roots  88  are received within a respective slot  94  in the disk  86  and the airfoils  92  extend radially outward such that a tip  96  of each airfoil  92  is closest to the blade outer air seal (BOAS) assembly  62 . The airfoil  92  defines a blade chord between a leading edge  98 , which may include various forward and/or aft sweep configurations, and a trailing edge  100 . A first sidewall that may be convex to define a suction side, and a second sidewall that may be concave to define a pressure side are joined at the leading edge  98  and at the axially spaced trailing edge  100 . The tip  96  extends between the sidewalls opposite the platform  90 . 
     The blade outer air seal (BOAS) assembly  62 , the platform  90 , the inner vane platform  72  and the outer vane platform  74  define the working medium combustion gas flow in a primary flow path P. The blade outer air seal (BOAS) assembly  62  and the outer vane platform  74  define an outer boundary of the flow path P. The platform  90  and the inner vane platform  72  bound the inner portion of the flow path P. 
     A full hoop inner air seal  78  that extends from the inner vane platform  72  provides one or more seal surfaces  80  that seal with the rotor assembly  66  to further contain the inner portion of the flow path P. The rotor assembly  66  includes a full hoop cover plate  82  with respective knife edges  81  that interface with the seal surfaces  80 . The full hoop cover plate  82  may be manufactured of alloys such as Inconel 625, Inconel 718 and Haynes  230  which have specific benefit for high temperature environments, such as, for example, environments typically encountered by aerospace and gas turbine engine. 
     A side plate seal assembly  110  also interfaces with a seal surface  112  that attaches to, or extends from, the inner vane platform  72 . The seal surfaces  80 ,  112  may be manufactured of a honeycomb material in which the honeycombs of these honeycomb structures may be open in the direction toward the knife edge seal projections. 
     The side plate seal assembly  110  is formed from a multiple of side plate seal segments  120  ( FIG. 3 ) that are manufactured of a non-metallic material such as ceramic matrix composite (CMC) or organic matrix composite (OMC). The ceramic matrix composite (CMC) or organic matrix composite (OMC) material typically includes prepreg ceramic plys that include prepreg ceramic fiber tows, the tows in each ply lying adjacent to one another in a planar arrangement such that each ply has a unidirectional orientation. Examples of CMC materials include, but are not limited to, carbon-fiber-reinforced carbon (C/C), carbon-fiber-reinforced silicon carbide (C/SiC), silicon-carbide-fiber-reinforced silicon carbide (SiC/SiC), alumina-fiber-reinforced alumina (Al 2 O 3 /Al 2 O 3 ), organic matrix composite (e.g. carbon fiber epoxy) or combinations thereof. The CMC may have increased elongation, fracture toughness, thermal shock, dynamic load capability, and anisotropic properties as compared to a monolithic ceramic structure. Other Ceramic matrix composite (CMC) materials may utilize tackified ceramic fabric/fibers whereby the fibers have not been infiltrated with matrix material, 3D weave architectures of dry fabrics, and others. Although CMCs are primarily discussed in the disclosed embodiment, other such non-metallic materials may also be utilized to form the segments. 
     Manufacture of the CMC typically includes laying up pre-impregnated composite fibers having a matrix material already present (prepreg) to form the geometry of the part (pre-form), autoclaving and burning out the pre-form, infiltrating the burned-out pre-form with the melting matrix material, then final machining and treatments of the pre-form. Infiltrating the pre-form may include depositing the ceramic matrix out of a gas mixture, pyrolyzing a pre-ceramic polymer, chemically reacting elements, sintering, generally in the temperature range of 1700-3000 F (925-1650 C), or electrophoretically depositing a ceramic powder. 
     Each of the multiple of side plate seal segments  120  include a shiplap interface  102 ,  104  ( FIG. 4 ) to form the full ring side plate seal assembly  110  that is retained between the full hoop cover plate  82  and the disk  86 . The multiple of side plate seal segments  120  may be identical segments to facilitate manufacture as well as accommodate thermal growth of the adjacent alloy full hoop cover plate  82  and disk  86 . 
     Each of the multiple of side plate seal segments  120  includes a retention surface  122  and a knife edge seal surface  124  that extends at an angle T thereto. The retention surface  122  is generally planar. The knife edge seal surface  124  extends from the retention surface  122  with a significant radius  125  to facilitate manufacture. An upper surface  126  that includes an outer diameter edge  128  of the retention surface  122  forms an angle W with the knife edge seal surface  124 . In one example, the angle T may be from 130-160 degrees and more specifically 145 degrees, and the angle W may be from 30-60 degrees and more specifically 43 degrees ( FIG. 5 ). A lower surface  130  that includes an inner diameter edge  132  of the retention surface  122  provides an inner axial constraint surface. In one embodiment, the lower surface  130  tapers to the inner diameter edge  132  of the retention surface  122  to further reduce centrifugal load on the full ring side plate seal assembly  110  during engine operation. 
     The retention surface  122  formed by the multiple of side plate seal segments  120  forms a full hoop plate that when assembled around the axis A, is retained between the full hoop cover plate  82  and the disk  86 . The outer diameter edge  128  of the retention surface  122  may also abut the platform  90  to further retain the side plate seal assembly  110  against the centrifugal loads during engine operation. That is, the side plate seal assembly  110  is retained under the platforms  90  formed by the adjacent blades  84  during engine operation. 
     The knife edge seal surface  124  formed by the multiple of side plate seal segments  120  forms an annular array of knife edge seal edge  134  that rides along the seal surface  112 . The knife edge seal surface  124  extends from the retention surface  122  and may thereby replace the outer most seal region of the rotating full hoop cover plate  82 . The non-metallic side plate seal assembly  110  is capable of withstanding the hot gas recirculation and pumping with minimal secondary flow and thereby further protects the metallic full hoop cover plate  82 . Replacing the outermost region of the full hoop cover plate  82  greatly reduces the thermal load and temperature of the full hoop cover plate  82 , allowing a lighter and more durable full hoop cover plate  82 . 
     The segmented side plate seal assembly  110  permits a relatively smaller outer cavity  150  ( FIG. 2 ) that is operable at much higher temperatures as compared to inner cavities  152 ,  154 ,  156  without increased cooling airflow. The relatively smaller outer cavity  150  is the first impediment to hot gas ingestion and essentially shields the inboard static and rotating structures from high temperature core airflow. The low density of the CMC side plate seal assembly  110  greatly reduces the centrifugal load on the rotor assembly  66  compared to a cast metal alloy design. The ability of CMC structures to be woven with 2D and 3D enables the compressive load, applied at the outer edge, to be carried with low risk of delamination. The density and fiber architecture enables a relatively long projecting knife edge seal surface  124  from the side-plate, which maximizes the ability to seal over large axial translation of the rotor relative to the static structure, insuring a stable seal interface The knife edge seal surface  124  can resist  2200 - 2500 F exposure mainly due to the inherent capability of SiC—SiC combined with the very low stress state in the knife edge seal surface  124 . The ship-lap interfaces are readily manufactured by conventional grinding techniques. When combined with the relatively low coefficient of thermal expansion, the intersegment gaps between each segment can be minimized, because the risk of binding due to rapid heating relative to the rotor disk is avoided. 
     Although particular step sequences are shown, described, and claimed, it should be appreciated that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present disclosure. 
     The foregoing description is exemplary rather than defined by the limitations within. Various non-limiting embodiments are disclosed herein, however, one of ordinary skill in the art would recognize that various modifications and variations in light of the above teachings will fall within the scope of the appended claims. It is therefore to be appreciated that within the scope of the appended claims, the disclosure may be practiced other than as specifically described. For that reason, the appended claims should be studied to determine true scope and content.