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

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    The present application claims benefit of U.S. Provisional Patent Application No. 61/428,728, filed Dec. 30, 2010, entitled Gas Turbine Engine And Cooled Flowpath Component Therefor, which is incorporated herein by reference. 
     
    
     GOVERNMENT RIGHTS 
       [0002]    The present application was made with the United States government support under Contract No. FA 8650-07-C-2803, awarded by the U.S. Air Force. The United States government may have certain rights in the present application. 
     
    
     FIELD OF THE INVENTION 
       [0003]    The present invention relates to gas turbine engines, and more particularly, to gas turbine engines with cooled flowpath components. 
       BACKGROUND 
       [0004]    Cooled gas turbine engine flowpath components that effectively use a cooling fluid, such as cooling air, 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 unique gas turbine engine. Another embodiment of the present invention is a unique cooled gas turbine engine flowpath component. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and cooled gas turbine engine flowpath components. 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 
         [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  schematically illustrates some aspects of a non-limiting example of a gas turbine engine in accordance with an embodiment of the present invention. 
           [0008]      FIG. 2  illustrates some aspects of a non-limiting example of a flowpath component in accordance with an embodiment of the present invention. 
           [0009]      FIG. 3  illustrates some aspects of a non-limiting example of the flowpath component of  FIG. 2  in accordance with an embodiment of the present invention. 
           [0010]      FIG. 4  illustrates some aspects of a non-limiting example of a plurality of pins configured to transmit cooling air across the flowpath component of  FIG. 2 . 
           [0011]      FIGS. 5A and 5B  illustrate some aspects of a non-limiting example of a trailing edge portion of the flowpath component of  FIG. 2 . 
           [0012]      FIGS. 6A and 6B  illustrate cross sections depicting some aspects of non-limiting examples of cooling passages in a trailing edge portion of the flowpath component of  FIG. 2 ;  FIG. 6C  represents a composite cross section illustrating an overlay of the cooling passages of  FIGS. 6A and 6B . 
       
    
    
     DETAILED DESCRIPTION 
       [0013]    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. 
         [0014]    Referring to the drawings, and in particular  FIG. 1 , a non-limiting example of some aspects of a gas turbine engine  10  in accordance with an embodiment of the present invention is schematically 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 take other forms, and may be, for example, a turboshaft engine, a turbojet engine, a turboprop engine, or a combined cycle engine having a single spool or multiple spools. 
         [0015]    As a turbofan engine, gas turbine engine  10  includes a fan system  12 , a bypass duct  14 , a compressor system  16 , a diffuser  18 , a combustion system  20 , a turbine system  22 , a discharge duct  26  and a nozzle system  28 . Bypass duct  14  and compressor system  16  are in fluid communication with fan system  12 . Diffuser  18  is in fluid communication with compressor system  16 . Combustion system  20  is fluidly disposed between compressor system  16  and turbine system  22 . In one form, combustion system  20  includes a combustion liner (not shown) that contains a continuous combustion process. In other embodiments, combustion system  20  may take other forms, and may be, for example and without limitation, a wave rotor combustion system, a rotary valve combustion system or a slinger combustion system, and may employ deflagration and/or detonation combustion processes. 
         [0016]    Fan system  12  includes a fan rotor system  30 . In various embodiments, fan rotor system  30  includes one or more rotors (not shown) that are powered by turbine system  22 . Bypass duct  14  is operative to transmit a bypass flow generated by fan system  12  to nozzle  28 . Compressor system  16  includes a compressor rotor system  32 . In various embodiments, compressor rotor system  32  includes one or more rotors (not shown) that are powered by turbine system  22 . Each compressor rotor includes a plurality of rows compressor blades (not shown) that are alternatingly interspersed with rows of compressor vanes (not shown). Turbine system  22  includes a turbine rotor system  34 . In various embodiments, turbine rotor system  34  includes one or more rotors (not shown) operative to drive fan rotor system  30  and compressor rotor system  32 . Each turbine rotor includes a plurality of turbine blades (not shown) that are alternatingly interspersed with rows of turbine vanes (not shown). 
         [0017]    Turbine rotor system  34  is drivingly coupled to compressor rotor system  32  and fan rotor system  30  via a shafting system  36 . In various embodiments, shafting system  36  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. Turbine system  22  is operative to discharge an engine  10  core flow to nozzle  28 . 
         [0018]    In one form, fan rotor system  30 , compressor rotor system  32 , turbine rotor system  34  and shafting system  36  rotate about an engine centerline  48 . In other embodiments, all or parts of fan rotor system  30 , compressor rotor system  32 , turbine rotor system  34  and shafting system  36  may rotate about one or more other axes of rotation in addition to or in place of engine centerline  48 . 
         [0019]    Discharge duct  26  extends between a bypass duct discharge portion  38 , a discharge portion  40  of turbine system  22  and engine nozzle  28 . Discharge duct  26  is operative to direct bypass flow and core flow from bypass duct discharge portion  38  and turbine discharge portion  40 , respectively, into nozzle system  28 . In some embodiments, discharge duct  26  may be considered a part of nozzle  28 . Nozzle  28  in fluid communication with fan system  12  and turbine system  22 . Nozzle  28  is operative to receive the bypass flow from fan system  12  via bypass duct  14 , and to receive the core flow from turbine system  22 , and to discharge both as an engine exhaust flow, e.g., a thrust-producing flow. In other embodiments, other nozzle arrangements may be employed, including separate nozzles for each of the core flow and the bypass flow. 
         [0020]    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  16  as core flow, and some of the pressurized air is directed into bypass duct  14  as bypass flow, which is discharged into nozzle  28  via discharge duct  26 . Compressor system  16  further pressurizes the portion of the air received therein from fan  12 , which is then discharged into diffuser  18 . Diffuser  18  reduces the velocity of the pressurized air, and directs the diffused core airflow into combustion system  20 . Fuel is mixed with the pressurized air in combustion system  20 , which is then combusted. The hot gases exiting combustion system  20  are directed into turbine system  22 , which extracts energy in the form of mechanical shaft power sufficient to drive fan system  12  and compressor system  16  via shafting system  36 . The core flow exiting turbine system  22  is directed along an engine tail cone  42  and into discharge duct  26 , along with the bypass flow from bypass duct  14 . Discharge duct  26  is configured to receive the bypass flow and the core flow, and to discharge both into nozzle  28  as an engine exhaust flow, e.g., for providing thrust, such as for aircraft propulsion. 
         [0021]    Compressor rotor system  32  includes a plurality of blades and vanes (not shown) employed to add energy to the gases prior to combustion. Turbine rotor system  34  includes a plurality of blades and vanes (not shown) employed to extract energy from the high temperature high pressure gases in the flowpath. It is desirable to maintain the temperature of blades and vanes within certain temperature limits, e.g., based on the materials and coatings employed in the blades and vanes. In many cases, blades and vanes are cooled by injecting cooling air into the blade or vane. 
         [0022]    Referring to  FIG. 2 , some aspects of a non-limiting example of a cooled flowpath component  50  in accordance with an embodiment of the present invention is illustrated. As used herein, a “flowpath component” is a component of engine  10  that is at least partially disposed within or exposed to core gas flow in engine  10  and/or forms at least in part the flowpath boundary in engine  10  that contains the core gas flow. In one form, flowpath component  50  is a turbine flowpath component. In a particular form, flowpath component  50  is a turbine vane airfoil, referred to herein as airfoil  50 . In other embodiments, flowpath component  50  may take other forms, and may be, for example and without limitation, a turbine blade, a strut, a blade platform or shroud, or may be a compressor or other flowpath component. In one form, airfoil  50  includes a leading edge  52 , a trailing edge  54 , a tip  56  and a hub  58 . 
         [0023]    Referring to  FIG. 3 , airfoil  50  is defined by a pressure side PS and a suction side SS. Airfoil  50  includes a spar  60  and a coversheet  62 . Coversheet  62  is configured for mating engagement with spar  60 . In one form, coversheet  62  is diffusion bonded to spar  60  on both pressure side PS and suction side SS. In other embodiments, coversheet  62  may be affixed to spar  60  using other bonding techniques and/or other joining methods, e.g., including welding, brazing or other material joining methods. In one form, coversheet  62  is configured to almost fully enclose spar  60 , leaving a portion of spar  60  uncovered on pressure side PS adjacent trailing edge  54 . In other embodiments, coversheet  62  may be configured to completely cover spar  60 , or may be configured to partially cover spar  60  to a greater or lesser extent than that illustrated in  FIG. 3 . 
         [0024]    Spar  60  and coversheet  62  are configured to form a gap  64  between spar  60  and coversheet  62  via the use of ribs and other standoff devices  66 , e.g., formed in spar  60  and coversheet  62 . Spar  60  includes a perimetrical wall  68  that extends around the perimeter of spar  60  and defines a leading edge cooling air supply cavity  70  and a trailing edge cooling air supply cavity  72  separated by a continuous rib  74  that extends between tip  56  and hub  58 . A plurality of apertures  76  in spar  60  extend through wall  68 . Apertures  76  are configured to deliver cooling air from cooling air supply cavities  70  and  72  into gap  64 . Gap  64  is operative as a distribution annulus to deliver cooling air about the outer periphery of spar  60  and about the inner periphery of coversheet  62 . Coversheet  62  includes a plurality of openings  78  spaced apart about the periphery of coversheet  62  for providing film cooling of coversheet  62 . Although a portion of the pressure side PS and suction side SS are not illustrated in  FIG. 3 , it will be understood that spar  60 , coversheet  62 , gap  64  and perimetrical wall  68  extend continuously, e.g., at a desired aerodynamic profile, between leading edge  52  and the illustrated portion of pressure side PS, bridging the gap  61  illustrated in  FIG. 3 ; and that spar  60 , coversheet  62 , gap  64  and perimetrical wall  68  extend continuously, e.g., at a desired aerodynamic profile, between leading edge  52  and the illustrated portion of suction side SS, bridging the gap  63  illustrated in  FIG. 3 , thereby closing off the periphery of leading edge cooling air supply cavity  70 . 
         [0025]    Referring to  FIG. 4  in conjunction with  FIG. 3 , disposed in the trailing edge portion of airfoil  50  are a plurality of hollow pins  80 . Hollow pins  80  are illustrated in cross-section in  FIG. 4 , with spar  60  and coversheet  62  removed for clarity. Pins  80  are oval in cross-sectional shape. As used herein, “oval” includes both elliptical shapes, and “racetrack” shapes, e.g., in the form of a rectangle with rounded corners. Pins  80  bridge cooling air supply cavity  72 , extending between pressure side PS of spar  60  and suction side SS of spar  60 . Pins  80  are configured to transmit cooling air directly, i.e., in a straight line, from gap  64  on suction side SS to gap  64  on pressure side PS via oval openings  82  extending through pins  80 , which are in fluid communication with gap  64  on both suction side SS and pressure side PS. Pins  80  are spaced apart in the direction from tip  56  to hub  58  by a sufficient amount to allow the flow of cooling air in cooling air supply cavity  72  to flow in sufficient quantity between pins  80  in the direction of trailing edge  54  to provide a desired amount of cooling air for cooling airfoil  50  in the vicinity of trailing edge  54 . Pins  80  are oval in shape in order to maximize the area of openings  82  for transmitting cooling air, while also maximizing the flow area between pins  80  to accommodate the flow therethrough for cooling airfoil  50  in the vicinity of trailing edge  54 . In other embodiments, pins  80  may have other shapes. 
         [0026]    Referring to  FIGS. 5A and 5B , some aspects of a non-limiting example of a trailing edge portion  84  of airfoil  50  is illustrated. Formed in trailing edge portion  84  in coversheet  62  are a plurality of pedestals  86 . In other embodiments, pedestals  86  may be formed in other regions of coversheet  62 . Pedestals  86  may be formed by any convenient means, for example but not limited to, material removal means such as chemical or electrochemical machining, the use of freeform manufacturing techniques to form coversheet  62 , micromachining and/or electrical discharge machining; or material addition means, such as vapor deposition, selective laser sintering, and/or one or more other freeform fabrication techniques. The shape of pedestals  86  may vary with the needs of the application. Pedestals  86  are configured for attachment to spar  60 . 
         [0027]    Pedestals  86  extend toward spar  60  from a base  88  to a plateau  90  in contact with spar  60 . In one form, pedestals  86  are formed in coversheet  62  in a trailing edge portion of airfoil  50 . In other embodiments, pedestals  86  may be formed in other locations of airfoil  50  in addition to or in place of the trailing edge portion of airfoil  50 . In addition, in other embodiments, pedestals  86  may be formed completely or partially in spar  60 . Pedestals  86  are configured to form therebetween a cooling circuit  92  for the flow of cooling air to cool trailing edge  54 . In one form, cooling circuit  92  is operative to discharge cooling air from trailing edge  54  via exit slots  94  formed between adjacent pedestals  86  at trailing edge  54 . Cooling circuit  92  is defined between coversheet  62 , spar  60  and pedestals  86 . Pedestals  86  are bonded at plateau  90  to spar  60 , e.g., in the manner set forth above with respect to coversheet  62 . By forming pedestals  86  in coversheet  62 , the bond interface between coversheet and spar is moved inward into airfoil  50 , which places the bond interface at a cooler location in airfoil  50  than had the pedestals been formed on spar  60  and a thin coversheet, e.g., having a thickness  96 , been bonded to such pedestals. This placement of the bond interface at a cooler location may increase the life of the bond joint between coversheet  62  and spar  60 . 
         [0028]    Referring to  FIG. 6A-6C , a plurality of internal cooling air passages  100  is interleaved with a plurality of internal cooling air passages  102  along the direction between tip  56  and hub  58  of airfoil  50 . That is, as viewed along the direction between tip  56  and hub  58 , cooling air passages  100  and cooling air passages  102  are alternatively arranged, e.g., one on top of the other. In one form, cooling air passages  100  and  102  are disposed within spar  60 . In other embodiments, cooling air passages may be arranged differently. Cooling air passages  100  and cooling air passages  102  are configured to deliver cooling air for different locations of trailing edge portion  84  of airfoil  50 . Cooling air passages  100  and cooling air passages  102  may be formed by any conventional or other means. As illustrated in  FIG. 6A , cooling passage  100  penetrates the wall of spar  60  and receives cooling air  104  from supply cavity  72  on suction side SS. Cooling air passage  100  directs cooling air  104  along the inside of coversheet  62  on suction side SS, providing cooling for that portion of coversheet  62 . Cooling air passage  100  then transfers cooling air  104  from suction side SS through spar  60  toward pressure side PS to deliver the cooling air via a discharge opening  106  on pressure side PS for film cooling of trailing edge portion  84  of airfoil  50 . 
         [0029]    As illustrated in  FIG. 6B , cooling air passage  102  penetrates the wall of spar  60  and receives cooling air  108  from cavity  72  on pressure side PS, and delivers cooling air  108  to cooling circuit  92  on suction side SS of airfoil  50  for cooling of trailing edge portion  84  of airfoil  50 . Cooling air  108  is discharged through exit slots  94  on trailing edge  54  of airfoil  50 . As illustrated in  FIG. 6C , cooling passages  100  and  102  cross over each other to deliver cooling air to opposite sides of airfoil  50 . 
         [0030]    Embodiments of the present invention include a turbine flowpath component for a gas turbine engine, comprising: a spar; and a coversheet configured to at least partially enclose the spar, and configured for mating engagement with the spar; wherein the coversheet includes a plurality of pedestals formed therein and extending toward the spar; and wherein the plurality of pedestals are configured to form a cooling circuit for cooling air. 
         [0031]    In a refinement, the turbine flowpath component is defined by a pressure side and a suction side; wherein the coversheet is configured for engagement with the spar on the suction side. 
         [0032]    In another refinement, the turbine flowpath component further comprises a first plurality of cooling air passages interleaved with a second plurality of cooling air passages, wherein the first plurality of cooling air passages delivers cooling air to a first location; and wherein the second plurality of cooling air passages delivers cooling air to a second location different from the first location. 
         [0033]    In yet another refinement, the first plurality of cooling air passages is configured to deliver cooling air to the cooling circuit. 
         [0034]    In still another refinement, the turbine flowpath component is defined by a pressure side and a suction side; and wherein the second plurality of cooling air passages is configured to transfer cooling air from the suction side to the pressure side. 
         [0035]    In yet still another refinement, the second plurality of cooling air passages are configured to discharge cooling air for film cooling on the pressure side. 
         [0036]    In a further refinement, the turbine flowpath component is defined by a pressure side and a suction side, wherein the spar forms a cavity between the pressure side and the suction side, further comprising a plurality of hollow pins extending between the pressure side and the suction side and bridging the cavity, wherein the hollow pins are oval in cross-sectional shape. 
         [0037]    In a yet further refinement, the hollow pins are configured to transmit cooling air directly from the suction side to the pressure side. 
         [0038]    In a still further refinement, the pedestals are bonded to the spar. 
         [0039]    Embodiments of the present invention include a gas turbine engine, comprising: a turbine having a turbine flowpath component, wherein the turbine flowpath component is defined by a first side and a second side opposite the first side; and wherein the turbine flowpath component includes a first plurality of internal cooling air passages and a second plurality of internal cooling air passages, wherein the first plurality of internal cooling air passages is interleaved with the second plurality of internal cooling air passages; wherein the first plurality of internal cooling air passages crosses over the second plurality of internal cooling air passages; wherein the first plurality of internal cooling air passages is configured to deliver cooling air toward the first side; and wherein the second plurality of internal cooling air passages is configured to deliver cooling air toward the second side. 
         [0040]    In a refinement, the gas turbine engine further comprises a spar, wherein the first plurality of internal cooling air passages and the second plurality of internal cooling air passages are disposed within the spar. 
         [0041]    In another refinement, the spar includes a wall defining a cooling air supply cavity, wherein at least one of the first plurality of internal cooling air passages and the second plurality of internal cooling air passages penetrates the wall and is configured to deliver cooling air from the cooling air supply cavity through the wall to a respective at least one of the first side and the second side. 
         [0042]    In yet another refinement, the gas turbine engine further comprises a coversheet configured to at least partially enclose the spar; and a cooling circuit defined between the spar and the coversheet at the first side, wherein the first plurality of internal cooling air passages is configured to supply cooling air to the cooling circuit. 
         [0043]    In still another refinement, the coversheet includes a plurality of pedestals formed therein on the first side; wherein the pedestals extend toward the spar and define the cooling circuit between the spar and the coversheet. 
         [0044]    In yet still another refinement, the turbine flowpath component has a trailing edge; and wherein the pedestals are disposed adjacent to the trailing edge. 
         [0045]    In a further refinement, the coversheet includes an opening on the second side; and wherein the second plurality of internal cooling air passages is configured to supply cooling air to the opening. 
         [0046]    In a yet further refinement, the gas turbine engine further comprises a plurality of hollow pins extending between the first side and the second side and bridging the cooling air supply cavity, wherein the hollow pins are oval in cross-sectional shape; and wherein the hollow pins are configured to transmit cooling air directly from the first side to the second side. 
         [0047]    In a still further refinement, the turbine flowpath component is configured as a turbine airfoil. 
         [0048]    In a yet still further refinement, the turbine flowpath component is configured as a turbine vane. 
         [0049]    Embodiments of the present invention include a flowpath component for a gas turbine engine, comprising: a spar defined by a first side, a second side, and a cooling air supply cavity formed between the first side and the second side; and means for cooling the turbine flowpath component using cooling air supplied by the cooling air supply cavity. 
         [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.