Abstract:
One embodiment of the present invention is a unique variable camber vane system for a gas turbine engine. Another embodiment is a unique gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and variable camber vane systems. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.

Description:
GOVERNMENT RIGHTS 
       [0001]    The present application was made with the United States government support under Contract No. FA8650-07-C-2803, awarded by the United States Air Force. The United States government may have certain rights in the present application. 
     
    
     FIELD OF THE INVENTION 
       [0002]    The present invention relates to gas turbine engines, and more particularly, to gas turbine engines with variable camber vane systems. 
       BACKGROUND 
       [0003]    Gas turbine engines with variable camber vane systems 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 
       [0004]    One embodiment of the present invention is a unique variable camber vane system for a gas turbine engine. Another embodiment is a unique gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and variable camber vane systems. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0005]    The description herein makes reference to the accompanying drawings wherein like reference numerals refer to like parts throughout the several views, and wherein: 
           [0006]      FIG. 1  schematically depicts some aspects of a non-limiting example of a gas turbine engine in accordance with an embodiment of the present invention. 
           [0007]      FIG. 2  schematically depicts some aspects of a non-limiting example of a fan system for a gas turbine engine in accordance with an embodiment of the present invention. 
           [0008]      FIG. 3  depicts some aspects of a non-limiting example of a variable camber guide vane system in accordance with an embodiment of the present invention. 
           [0009]      FIG. 4  depicts some aspects of the variable camber guide vane system of  FIG. 3 . 
           [0010]      FIG. 5  depicts some aspects of a non-limiting example of a seal strip in accordance with an embodiment of the present invention. 
       
    
    
     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 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 depicted. In one form, gas turbine engine  10  is an aircraft propulsion power plant. In other embodiments, gas turbine engine  10  may be a land-based or marine engine. In one form, gas turbine engine  10  is a multi-spool turbofan engine. In other embodiments, gas turbine engine  10  may be a single or multi-spool turbofan, turboshaft, turbojet, turboprop gas turbine or combined cycle engine. 
         [0013]    Gas turbine engine  10  includes a fan system  12 , a compressor system  14 , a diffuser  16 , a combustion system  18  and a turbine system  20 . Compressor system  14  is in fluid communication with fan system  12 . Diffuser  16  is in fluid communication with compressor system  14 . Combustion system  18  is fluidly disposed between compressor system  14  and turbine system  20 . Fan system  12  includes a fan rotor system  22 . In various embodiments, fan rotor system  22  includes one or more rotors (not shown) that are powered by turbine system  20 . Compressor system  14  includes a compressor rotor system  24 . In various embodiments, compressor rotor system  24  includes one or more rotors (not shown) that are powered by turbine system  20 . Turbine system  20  includes a turbine rotor system  26 . In various embodiments, turbine rotor system  26  includes one or more rotors (not shown) operative to drive fan rotor system  22  and compressor rotor system  24 . Turbine rotor system  26  is driving coupled to compressor rotor system  24  and fan rotor system  22  via a shafting system  28 . In various embodiments, shafting system  28  includes a plurality of shafts that may rotate at the same or different speeds and directions. In some embodiments, only a single shaft may be employed. 
         [0014]    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 system  14 , and the balance is directed into a bypass duct (not shown). Compressor system  14  further pressurizes the air received from fan  12 , which is then discharged into diffuser  16 . Diffuser  16  reduces the velocity of the pressurized air, and directs the diffused airflow into combustion system  18 . Fuel is mixed with the pressurized air in combustion system  18 , which is then combusted. In one form, combustion system  18  includes a combustion liner (not shown) that contains a continuous combustion process. In other embodiments, combustion system  18  may take other forms, and may be, for example, a wave rotor combustion system, a rotary valve combustion system, or a slinger combustion system, and may employ deflagration and/or detonation combustion processes. The hot gases exiting combustor  18  are directed into turbine system  20 , which extracts energy in the form of mechanical shaft power to drive fan system  12  and compressor system  14  via shafting system  28 . The hot gases exiting turbine system  20  are directed into a nozzle (not shown), and provide a component of the thrust output by gas turbine engine  10 . 
         [0015]    Referring to  FIG. 2 , a non-limiting example of some aspects of fan system  12  in accordance with an embodiment of the present invention is schematically depicted. Fan system  12  includes a variable guide vane system  40  having a variable inlet guide vane stage  42  and a variable outlet guide vane stage  44  disposed on either side of a rotating fan stage  46 . Variable inlet guide vane stage  42  is operative to guide air into rotating fan stage  46 , and to selectively vary the incidence angle of the air flow entering rotating fan stage  46 . Variable outlet guide vane stage  44  is operative to guide air exiting rotating fan stage  46 , and to selectively vary the incidence angle of the air flow exiting rotating fan stage  46 . Variable inlet guide vane stage  42  and variable outlet guide vane stage  44  are actuated by an actuation system (not shown). Although described herein as with respect to fan system  12 , it will be understood that variable guide vane system  40  may also or alternatively be employed as part of compressor system  14 . In addition, although variable guide vane system  40  includes both variable inlet and outlet guide vane stages, other embodiments may include only a variable inlet guide vane stage or a variable outlet guide vane stage. 
         [0016]    Referring to  FIGS. 3-5 , a non-limiting example of some aspects of variable inlet guide vane stage  42  in accordance with an embodiment of the present invention is illustrated. It will be understood that some embodiments of variable outlet guide vane stage  44  may be similar to variable inlet guide vane stage  42 , and hence, the following description of variable inlet guide vane stage  42  is also applicable to aspects of some embodiments of variable outlet guide vane stage  44 . Variable inlet guide vane stage  42  includes an outer band  50 , an inner band  52  and plurality of vanes  54 . Outer band  50  defines an outer flowpath wall of variable inlet guide vane stage  42 . Inner band  52  defines an inner flowpath wall of variable inlet guide vane stage  42 . Vanes  54  are airfoils that extend between outer band  50  and inner band  52 , and are spaced apart circumferentially. In one form, vanes  54  extend in the radial direction between outer band  50  and inner band  52 . In other embodiments, vanes  54  may extend between outer band  50  and inner band  52  at other angles. 
         [0017]    Each vane  54  includes an airfoil portion  56  and an airfoil portion  58 . Airfoil portion  56  extends between a tip portion  60  and a root portion  62 . In one form, airfoil portion  56  includes the trailing edge  64  of vane  54 . In other embodiments, airfoil portion  56  may be formed with a leading edge of vane  54  instead of trailing edge  64 , e.g., for use in variable outlet guide vane  44 . Airfoil portion  58  extends between a tip portion  66  and a root portion  68 . In one form, airfoil portion  58  includes the leading edge  70  of vane  54 . In other embodiments, airfoil portion  58  may be formed with a trailing edge instead of leading edge  70 , e.g., for use in variable outlet guide vane  44 . In one form, airfoil portion  56  is fixed, i.e., stationary. In other embodiments, airfoil portion  56  may be movable, e.g., pivotable about an axis so as to be able to vary the angle of the trailing edge of vane  54 . In one form, airfoil portion  58  is variable, being configured to pivot about a pivot axis  72  with respect to airfoil portion  56 , to provide a variable camber for vane  54 . In other embodiments, airfoil portion  58  may be fixed. In one form, airfoil portion  58  is coupled to an actuation system (not shown) that is operative to selectively position airfoil portion  58  at a desired incidence angle. In other embodiments, airfoil portion  56  may also or alternatively be coupled to an actuation system (not shown) that is operative to selectively position airfoil portion  56  at a desired incidence angle. 
         [0018]    Extending from airfoil portion  58  are pivot shafts  74  and  76 , which establish pivot axis  72 . Outer band  50  includes a plurality of spaced apart openings  78 . Inner band  52  includes a plurality of spaced apart openings  80 . Openings  78  and  80  are operative to receive pivot shafts  74  and  76 , respectively, and retain airfoil portions  58  in the engine axial, circumferential and radial direction. In one form, pivot shafts  74  and  76  retain airfoil portion  58  in outer band  50  and inner band  52  via anti-friction bushings  82  and  84 . Anti-friction bushings  82  and  84  are operative to provide bearing surfaces for pivot shafts  74  and  76 . Other embodiments may not include anti-friction bushings  82  and  84 . Airfoil portion  58  is operative to rotate in rotation directions  86  about pivot axis  72 . 
         [0019]    During the operation of engine  10 , air flows past vanes  54  in the general direction illustrated as direction  88 . Vane  54  has a pressure side  90  and a suction side  92 , wherein the pressure on pressure side  90  exceeds that of suction side  92 . The pressure differential between pressure side  90  and suction side  92  may vary, e.g., depending upon vane  54  camber and engine operating conditions. The pressure differential between pressure side  90  and suction side  92  provides an impetus to flow from pressure side  90  to suction side  92 , e.g., between airfoil portion  56  and airfoil portion  58 . It is desirable to reduce or prevent leakage between airfoil portion  56  and airfoil portion  58 , e.g., leakage flow from pressure side  90  to suction side  92 , e.g., in order to improve fan  12  and engine  10  efficiency. Accordingly, vanes  54  include a sealing arrangement  94  operative to seal between airfoil portion  56  and airfoil portion  58 . Sealing arrangement  94  includes a seal strip  96  arranged to seal against fluid flow between airfoil portion  56  and airfoil portion  58  during the operation of engine  10 , and to accommodate movement of one or both of airfoil portions  56  and  58 , e.g., rotation of airfoil portion  58  about pivot axis  72 , while sealing against fluid flow. 
         [0020]    In one form, seal strip  96  is a rigid structure that does not substantially deform in use or installation. In other embodiments, seal strip  96  may be a flexible structure. In one form, seal strip  96  is formed of a polymeric material, such as Vespel (commercially available from DuPont Engineering Polymers, located in Newark, Del., U.S.A.) and/or Torlon polyamide-imide (commercially available from Solvay Advanced Polymers, located in Alpharetta, Ga., U.S.A.). In other embodiments, seal strip  96  may be formed of other materials. In one form, seal strip  96  is disposed in a groove  98 . In one form, groove  98  is disposed in a face  100  of airfoil portion  56  that faces airfoil portion  58 . In one form, seal strip  96 , groove  98  and face  100  extend between tip portion  60  and root portion  62  of airfoil portion  56 . In other embodiments, seal strip  96 , groove  98  and/or face  100  may extend only partially between tip portion  60  and root portion  62 . Face  100  is formed with a radius  102  centered on pivot axis  72 . In one form, face  100  is formed integrally with airfoil portion  56 . In other embodiments, face  100  may be formed separately and affixed to airfoil portion  56 . In one form, seal strip  96  is partially installed in groove  98 , that is, leaving a portion  108  of seal strip  96  extending beyond face  100  of airfoil portion  56 . Seal strip  96  has a width  104  greater than a width  106  of groove  98 , and is installed into groove  98  with an interference fit, e.g., 0.001-0.002 inch. The amount of interference may vary with the needs of the application. 
         [0021]    Airfoil portion  58  includes a crown  110  facing face  100  of airfoil portion  56 . In one form, crown  110  is formed integrally with airfoil portion  58 . In other embodiments, crown  110  may be formed separately and affixed to airfoil portion  58 . Crown  110  is formed with a radius  112  centered on pivot axis  72 . In one form, crown  110  extends between tip portion  66  and root portion  68  of airfoil portion  58 , and is positioned opposite groove  98 . In other embodiments, crown  110  may extend only partially between tip portion  66  and root portion  68 . In one form, face  100  of airfoil portion  56  is concave, and is operative to receive therein crown  110  opposite groove  98  in a nested arrangement. In other embodiments, face  100  may be convex. In one form, crown  110  of airfoil portion  58  is convex, and is operative to be received into face  100  in a nested arrangement. In other embodiments, crown  110  may be convex, e.g., an inverted crown. Although the depicted embodiment includes groove  98  and seal strip  96  being located in face  100 , in other embodiments, groove  98  and seal strip  96  may be located in crown  110 . 
         [0022]    Seal strip  96  includes a rubbing surface  114 . In one form, rubbing surface  114  is disposed opposite radius  112  of crown  110 , and is operative to contact and seal against radius  112  of crown  110  of airfoil portion  58 . During movement of airfoil portion  58 , e.g., when changing the camber of vane  54  by rotating airfoil portion  58  about pivot axis  72 , rubbing surface  114  may rub against crown  110 , e.g., until wear of seal strip  96  resulting from rotation of airfoil portion  58  reduces or eliminates contact between seal strip  96  and crown  110 . In other embodiments, rubbing surface  114  may be configured to be in close proximity to crown  110 , but without any rubbing contact. In still other embodiments, seal strip  96  may be installed in crown  110 , and rubbing surface  114  may be configured to seal against face  100 . 
         [0023]    Rubbing surface  114  is preformed prior to installation into airfoil portion  56 , e.g., machined. In one form, rubbing surface  114  is configured as a radius  116  centered about pivot axis  72 , e.g., the same radius as radius  112  of crown  110 . In other embodiments, radius  116  may be the same radius as radius  102  of face  100  or any other radius suitable for the application. In still other embodiments, other shapes for rubbing surface  114  may be employed. In one form, rubbing surface  114  is concave. In other embodiments, rubbing surface  114  may take other forms, and may be, for example, convex. 
         [0024]    Embodiments of the present invention include a variable camber vane system for a gas turbine engine, comprising: a first airfoil portion having a first tip portion, a first root portion, a face extending at least partially between the first tip portion and the first root portion, and a groove in the face extending at least partially between the first tip portion and the first root portion, wherein the groove has a groove width; a second airfoil portion arranged to rotate with respect to the first airfoil portion about a pivot axis, wherein the second airfoil portion includes a second tip portion; a second root portion; and a crown extending at least partially between the second tip portion and the second root portion, wherein the crown includes a crown radius centered about the pivot axis and positioned opposite the groove; and a seal strip having a seal width greater than the groove width and a rubbing surface opposite the crown radius, wherein the seal strip is at least partially disposed in the groove with an interference fit; and wherein the seal strip is arranged to seal against fluid flow between the first airfoil portion and the second airfoil portion. 
         [0025]    In a refinement, the seal strip is a rigid structure. 
         [0026]    In another refinement, the rubbing surface has a rubbing surface radius the same as the crown radius. 
         [0027]    In yet another refinement, the crown is formed integrally with the second airfoil portion. 
         [0028]    In still another refinement, the face is formed integrally with the first airfoil portion. 
         [0029]    In yet still another refinement, the face is concave and operative to receive the crown therein. 
         [0030]    In a further refinement, the first airfoil portion is stationary. 
         [0031]    In a yet further refinement, the first airfoil portion and the second airfoil portion form at least part of an inlet guide vane having a fixed leading edge and a variable trailing edge; wherein the first airfoil portion includes the leading edge; and wherein the second airfoil portion includes the trailing edge. 
         [0032]    In a still further refinement, the first airfoil portion and the second airfoil portion form at least part of an outlet guide vane having a variable leading edge and a fixed trailing edge; wherein the first airfoil portion includes the leading edge; and wherein the second airfoil portion includes the trailing edge. 
         [0033]    Embodiments of the present invention include a gas turbine engine, comprising: at least one of a fan and a compressor having a variable camber vane system, the variable camber vane system including: at least two airfoil portions adapted to vary a camber of the variable camber vane system, wherein a first of the airfoil portions includes a groove and a second of the airfoil portions includes a crown having a crown radius; and a seal strip at least partially disposed in the groove with an interference fit, wherein the seal strip includes a rubbing surface opposite the crown radius and operative to seal against fluid flow between the first of the airfoil portions and the second of the airfoil portions. 
         [0034]    In a refinement, the rubbing surface contacts the crown at the crown radius. 
         [0035]    In another refinement, the seal strip is formed of a polymer material. 
         [0036]    In yet another refinement, the seal strip is formed of at least one of Vespel and Torlon. 
         [0037]    In still another refinement, the at least two airfoil portions form an inlet guide vane. 
         [0038]    In a further refinement, the at least two airfoil portions form an outlet guide vane. 
         [0039]    Embodiments include a gas turbine engine, comprising: at least one of a fan and a compressor having a variable camber vane system, the variable camber vane system including: at least two airfoil portions adapted to vary a camber of the variable camber vane system, wherein a first of the airfoil portions includes a groove; and wherein a second of the airfoil portions includes a crown having a crown radius; and a seal strip disposed in the groove; wherein the seal strip has a rubbing surface radius preformed thereon and configured for sealing engagement with the crown. 
         [0040]    In a refinement, the seal strip is a rigid structure formed of a polymer. 
         [0041]    In another refinement, the crown radius is convex, and the rubbing surface radius is concave. 
         [0042]    In yet another refinement, the seal strip is fitted in the groove with an interference fit. 
         [0043]    In still another refinement, the crown is nested within the first of the airfoil portions opposite the groove. 
         [0044]    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.