Abstract:
One embodiment of the present invention is a vane assembly for a gas turbine engine. Another embodiment of the present invention is a damper seal that may be employed in conjunction with a vane assembly of a gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods and combinations for vane assemblies and the sealing and damping thereof. Further embodiments, forms, features, aspects, benefits and advantages of the present application shall become apparent from the description and figures provided herewith.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    The present application claims the benefit of U.S. Provisional Patent Application 61/290,601, filed Dec. 29, 2009, and is incorporated herein by reference. 
     
    
     GOVERNMENT RIGHTS 
       [0002]    The present application was made with United States government support under contract number N00019-04-C-0093 awarded by the United States Navy. The United States government may have certain rights in the present application. 
     
    
     FIELD OF THE INVENTION 
       [0003]    The present invention relates to a gas turbine engine, and more particularly, to a damper seal for a vane assembly of a gas turbine engine. 
       BACKGROUND 
       [0004]    Systems for compressing air and discharging the air to a combustor of a gas turbine engine remain an area of interest. Some existing systems have various shortcomings, drawbacks and disadvantages relative to certain applications. Accordingly, there remains a need for further contributions in this area of technology. 
       SUMMARY 
       [0005]    One embodiment of the present invention is a vane assembly for a gas turbine engine. Another embodiment of the present invention is a damper seal that may be employed in conjunction with a vane assembly of a gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods and combinations for vane assemblies and the sealing and damping thereof. Further embodiments, forms, features, aspects, benefits and advantages of the present application shall become apparent from the description and figures provided herewith. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0006]    The description herein makes reference to the accompanying drawings wherein like reference numerals refer to like parts throughout the several views, and wherein: 
           [0007]      FIG. 1  is a schematic depiction of a gas turbine engine in accordance with an embodiment of the present invention. 
           [0008]      FIG. 2  is a partial view of an outlet guide vane (OGV) employed in accordance with an embodiment of the present invention. 
           [0009]      FIG. 3  is a sectional view of the OGV of  FIG. 2  with a damper seal in accordance with an embodiment of the present invention. 
           [0010]      FIG. 4  depicts the OGV and damper seal of  FIG. 3  with the damper seal illustrated in an installed condition. 
       
    
    
     DETAILED DESCRIPTION 
       [0011]    For purposes of promoting an understanding of the principles of the invention, reference will now be made to the embodiments illustrated in the drawings, and specific language will be used to describe the same. It will nonetheless be understood that no limitation of the scope of the invention is intended by the illustration and description of certain embodiments of the invention. In addition, any alterations and/or modifications of the illustrated and/or described embodiment(s) are contemplated as being within the scope of the present invention. Further, any other applications of the principles of the invention, as illustrated and/or described herein, as would normally occur to one skilled in the art to which the invention pertains, are contemplated as being within the scope of the present invention. 
         [0012]    Referring now to the drawings, and in particular,  FIG. 1 , a non-limiting example of a gas turbine engine  10  in accordance with an embodiment of the present invention is schematically depicted. Gas turbine engine  10  is an axial flow turbofan engine, e.g., an aircraft propulsion power plant. In one form, gas turbine engine  10  is a turbofan engine. In other embodiments, gas turbine engine  10  may take other forms, including turbojet engines, turboprop engines, and turboshaft engines having axial, centrifugal and/or axi-centrifugal compressors and/or turbines. 
         [0013]    In the illustrated embodiment, gas turbine engine  10  includes a fan  12 , a compressor  14  with outlet guide vane (OGV)  16 , a diffuser  18 , a combustor  20 , a high pressure (HP) turbine  22 , a low pressure (LP) turbine  24 , an exhaust nozzle  26  and a bypass duct  28 . Diffuser  18  and combustor  20  are fluidly disposed between OGV  16  of compressor  14  and HP turbine  22 . LP turbine  24  is drivingly coupled to fan  12  via an LP shaft  30 . HP turbine  22  is drivingly coupled to compressor  14  via an HP shaft  32 . In one form, gas turbine engine  10  is a two-spool engine. In other embodiments, engine  10  may have any number of spools, e.g., may be a three-spool engine or a single spool engine. 
         [0014]    Compressor  14  includes a plurality of blades and vanes  34  for compressing air. During the operation of gas turbine engine  10 , air is drawn into the inlet of fan  12  and pressurized by fan  12 . Some of the air pressurized by fan  12  is directed into compressor  14  and the balance is directed into bypass duct  28 . Bypass duct  28  directs the pressurized air to exhaust nozzle  26 , which provides a component of the thrust output by gas turbine engine  10 . Compressor  14  receives the pressurized air from fan  12 , which is compressed by blades and vanes  34 . 
         [0015]    The pressurized air discharged from compressor  14  is then directed downstream by OGV  16  to diffuser  18 , which diffuses the airflow, reducing its velocity and increasing its static pressure. The diffused airflow is then directed into combustor  20 . Fuel is mixed with the air in combustor  20 , which is then combusted in a combustion liner (not shown). The hot gases exiting combustor  20  are directed into HP turbine  22 , which extracts energy from the hot gases in the form of mechanical shaft power to drive compressor  14  via HP shaft  32 . The hot gases exiting HP turbine  22  are directed into LP turbine  24 , which extracts energy in the form of mechanical shaft power to drive fan  12  via LP shaft  30 . The hot gases exiting LP turbine  24  are directed into nozzle  26 , and provide a component of the thrust output by gas turbine engine  10 . 
         [0016]    Referring now to  FIG. 2 , OGV  16  is further described. In the depiction of  FIG. 2 , diffuser  18 , located just downstream from OGV  16 , is not shown for purposes of clarity of illustration. 
         [0017]    OGV  16  is a 360° compressor vane assembly having an outer band  36 , an inner band  38  and plurality of vanes  40 . Outer band  36  defines an outer flowpath wall OFW of OGV  16 . Inner band  38  defines an inner flowpath wall IFW of OGV  16 . Vanes  40  are airfoils, and are spaced apart from each other circumferentially. Vanes  40  extend in the radial direction between outer band  36  and inner band  38 . Each vane  40  has a tip end  42  and a root end  44 . 
         [0018]    OGV  16  is attached to a static structure (not shown) of gas turbine engine  10  at outer band  36 , e.g., via a bolted interface. In one form, OGV  16  is a unitary 360° casting. In other embodiments, OGV  16  may be formed from a plurality of circumferential vane segments that are assembled together, e.g., at installation into gas turbine engine  10 . 
         [0019]    Inner band  38  includes a plurality of bosses  46  and threaded bolt holes  48 . In one form, bosses  46  and threaded bolt holes  48  are circumferentially and alternatingly spaced apart around the inner periphery of inner band  38 . In other embodiments, other arrangements and/or spacing schemes may be employed. Inner band  38  is split between each vane  40  into segments. In one form, each segment extends from (includes) a single airfoil, i.e., vane  40 . In other embodiments, each segment may include more than one airfoil. In a particular form, inner band  38  is subdivided at partitions  50  into a plurality of circumferential inner band segments  52 , which may help reduce thermally induced stresses in OGV  16 . Partitions  50  are equally spaced around the circumference of inner band  38  in circumferential direction  54 . Each vane  40  is coupled to outer band  36  at tip end  42 , and is coupled to a respective inner band segment  52  at root end  44 . 
         [0020]    In one form, partitions  50  are located on both sides of each vane  40 , and hence each inner band segment  52  corresponds to a single vane  40 . In other embodiments, each inner band segment  52  may correspond with two or more vanes  40 , in which case a corresponding number of two or more vanes  40  are positioned between each pair of partitions  50 . In one form, each partition  50  is formed by electrical discharge machining (EDM) of inner band  38 , in particular using a wire EDM machine. In other embodiments, other methods of cutting or machining may be employed to form each partition  50 , for example, laser cutting, waterjet cutting and/or abrasivejet cutting. 
         [0021]    During the operation of gas turbine engine  10 , pressurized air passes through vanes  40  at a high rate of speed, which may induce a vibratory response into OGV  16 . For example, each inner band segment  52  and the corresponding vane  40  may behave as a cantilevered spring-mass system which may respond to excitation provided by the pressurized air being discharged through OGV  16  into diffuser  18 . In addition, air exiting OGV  16  may leak between the aft end of OGV  16  and diffuser  18 , thereby resulting in parasitic losses that may adversely affect the performance and efficiency of gas turbine engine  10 . 
         [0022]    Referring now to  FIG. 3 , a non-limiting example of a damper seal  56  in accordance with an embodiment of the present invention is depicted. In one form, damper seal  56  is configured for use in an inner band of a compressor vane assembly. In other embodiments, damper seal  56  may be configured for use in an outer band of a compressor vane assembly and/or inner and/or outer bands of turbine vane assemblies. 
         [0023]    Damper seal  56  includes a friction damper portion  58  and an air seal portion  60 . Friction damper portion  58  extends circumferentially along inner band  38  in circumferential direction  54  (see  FIG. 2 ). In one form, friction damper portion  58  is a continuous strip, e.g., a continuous strip formed into a ring. In one form, friction damper portion  58  is a continuous strip formed into a ring, and welded together at its ends. In other embodiments, the ends of the strip may not be welded together. In other embodiments, friction damper portion  58  may be formed by joining together a plurality of individual segments, or may be otherwise formed as a continuous ring. In still other forms, friction damper portion  58  may be discontinuous, e.g., and may include one or more continuous ring portions having damper segments extending therefrom that are distributed circumferentially in circumferential direction  54  along inner band  38 . 
         [0024]    Friction damper portion  58  is structured to contact each inner band segment  52 . Friction damper portion  58  provides friction damping of inner band segments  52  based on the contact, e.g., in the form of friction losses due to sliding contact between inner band segments  52  and friction damper portion  58 . In other embodiments, it is alternatively contemplated that friction damper portion  58  contacts only certain inner band segments. Contact between friction damper portion  58  and inner band segments  52  may be maintained, for example, by providing friction damper portion  58  with an outer circumference that is greater than the inner circumference of inner band  38 . 
         [0025]    In one form, air seal portion  60  extends from friction damper portion  58  in an axial direction  62  that is substantially perpendicular to circumferential direction  54 . Axial direction  62  is parallel to the axis of rotation of engine  10  main rotor components, e.g., fan  12 , compressor  14 , HP turbine  22  and LP turbine  24 . In other embodiments, air seal portion extends from friction damper portion in radial and/or axial directions. Air seal portion  60  is structured to seal against diffuser  18 , which is spaced apart from OGV  16  downstream in axial direction  62 . In one form, air seal portion  60  is structured in the form of a bellows  64  having two convolutions  66  and  68  that extend in axial direction  62 , and is compressible in axial direction  62 . In other embodiments, air seal portion  60  may take other forms, including bellows having a greater or lesser number of convolutions, and including forms other than bellows. 
         [0026]    In one form, air seal portion  60  is integral with friction damper portion  58 . Friction damper portion  58  includes a cylindrical surface  70  that extends substantially in axial direction  62 , although other surface forms may alternatively be employed. In the present embodiment, air seal portion  60  and friction damper portion  58  are formed from sheet metal, e.g., a common strip of material. It is alternatively contemplated that air seal portion  60  and friction damper portion  58  may be formed separately and subsequently joined together, e.g., via welding, brazing, bolting, or other suitable joining methodology. 
         [0027]    In one form, damper seal  56  is attached to inner band  38  using bosses  46  and bolt holes  48 . In particular, damper seal  56  includes a plurality of holes  72  corresponding in location to bosses  46  and bolt holes  48 . Holes  72  adjacent bosses  46  are slightly smaller in diameter than bosses  46  so as to create an interference fit, e.g., of approximately 0.002 inch, although any suitable interference fit may be employed in other embodiments. Holes  72  adjacent to bolt holes  48  are sized to allow passage therethrough of bolts (not shown) to further secure damper seal  56  to inner band  38 . In other embodiments, damper seal  56  may be attached to inner band  38  using other suitable attachment methods, e.g., including other types of mechanical fasteners, clips, etc., and/or brazing and/or welding. 
         [0028]    Referring now to  FIG. 4 , OGV  16  and damper seal  56  are depicted in the installed condition, wherein air seal portion is compressed between OGV  16  and diffuser  18 , thus sealing the gap  74  disposed between OGV  16  and diffuser  18 . 
         [0029]    During the operation of gas turbine engine  10 , the excitation of OGV  16 , in particular, vanes  40  and inner band segments  52 , may result in a reduced vibratory response in OGV  16  due to the friction damping generated by the contact of friction damper portion  58  with inner band segments  52  of inner band  38 . In addition, leakage of compressed air between OGV  16  and diffuser  18  may be reduced or eliminated by air seal portion  60 , which extends from OGV  16  to diffuser  18 . Sealing contact between damper seal  56  and diffuser  18  is maintained by virtue of the compressive stresses in air seal portion  60 , in particular, convolutions  66  and  68  of bellows  64 . 
         [0030]    Embodiments of the present invention include a vane assembly for a gas turbine engine. The vane assembly may include an outer band, an inner band, a plurality of airfoils, and a damper seal. The inner band may be subdivided into a plurality of circumferential segments. The plurality of airfoils may be spaced apart circumferentially and extend between the outer band and the inner band. Each airfoil may have a tip end and a root end, and may be is coupled to the outer band at the tip end, and coupled to a respective segment of the inner band at the root end. The damper seal which may include a friction damper portion extending along the inner band in the circumferential direction. The friction damper may be in contact with at least two of the circumferential segments and may be structured to provide friction damping of at least two circumferential segments based on the contact. The damper seal may also include an air seal portion extending from the friction damper portion in an axial direction substantially perpendicular to the circumferential direction. The air seal may be structured to seal against an engine component that is spaced apart from the vane assembly in the axial direction. 
         [0031]    In one refinement of the embodiment the air seal portion is integral with the friction damper portion. 
         [0032]    In another refinement of the embodiment the friction damper portion is a continuous strip extending circumferentially along the inner band. 
         [0033]    In another refinement of the embodiment the friction damper portion is structured to contact each the circumferential segment. 
         [0034]    In another refinement of the embodiment the inner band is split between each airfoil, and each segment extends from a single airfoil. 
         [0035]    In another refinement of the embodiment the air seal portion is structured as a bellows. 
         [0036]    In another refinement of the embodiment the air seal portion includes at least two convolutions extending in the axial direction. 
         [0037]    In another refinement of the embodiment the vane assembly is a compressor vane assembly. 
         [0038]    In another refinement of the embodiment the engine component is a diffuser located downstream of a compressor of the gas turbine engine. 
         [0039]    In another refinement of the embodiment the outer band defines an outer flowpath wall and the inner band defines an inner flowpath wall. 
         [0040]    In another refinement of the embodiment the friction damper portion and the air seal portion are formed from sheet metal. 
         [0041]    In another refinement of the embodiment the damper seal is at least one of bolted and pinned to the inner band. 
         [0042]    Another embodiment of the present invention may include a damper seal for the vane assembly of a gas turbine engine. The damper seal may include a friction damper portion having a surface structured to contact a segment of a vane assembly to provide friction damping of the segment. The damper seal may also include an air seal portion structured to seal against a gas turbine engine component that is spaced apart from the segment in an axial direction, and the air seal portion may be integral with the friction damper portion. 
         [0043]    In one refinement of the embodiment the friction damper and the air seal are formed as a continuous ring. 
         [0044]    In another refinement of the embodiment the damper seal is formed from sheet metal. 
         [0045]    In another refinement of the embodiment the air seal portion is compressible in the axial direction. 
         [0046]    In another refinement of the embodiment the air seal portion is structured as a bellows. 
         [0047]    In another refinement of the embodiment the air seal portion includes at least two convolutions extending in the axial direction. 
         [0048]    In another refinement of the embodiment the surface extends in the axial direction. 
         [0049]    Another embodiment may include a damper seal for a vane assembly of a gas turbine engine. The damper seal may include means for providing friction damping of a plurality of segments of the vane assembly; and means for sealing against a gas turbine engine component that may be spaced apart from the segments in an axial direction, wherein and the means for sealing is integral with the means for providing friction damping. 
         [0050]    While the invention has been described in connection with what is presently considered to be the most practical and preferred embodiment, it is to be understood that the invention is not to be limited to the disclosed embodiment(s), but on the contrary, is intended to cover various modifications and equivalent arrangements included within the spirit and scope of the appended claims, which scope is to be accorded the broadest interpretation so as to encompass all such modifications and equivalent structures as permitted under the law. Furthermore it should be understood that while the use of the word preferable, preferably, or preferred in the description above indicates that feature so described may be more desirable, it nonetheless may not be necessary and any embodiment lacking the same may be contemplated as within the scope of the invention, that scope being defined by the claims that follow. In reading the claims it is intended that when words such as “a,” “an,” “at least one” and “at least a portion” are used, there is no intention to limit the claim to only one item unless specifically stated to the contrary in the claim. Further, when the language “at least a portion” and/or “a portion” is used the item may include a portion and/or the entire item unless specifically stated to the contrary.