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

Full Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    The present application claims benefit of U.S. Provisional Patent Application No. 61/427,714, filed Dec. 28, 2010, entitled Gas Turbine Engine and Airfoil, which is incorporated herein by reference. 
     
    
     FIELD OF THE INVENTION 
       [0002]    The present invention relates to gas turbine engines, and more particularly, to airfoils for gas turbine engines. 
       BACKGROUND 
       [0003]    Gas turbine engine airfoils, particularly those that require cooling, 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 gas turbine engine. Another embodiment is a unique turbine engine airfoil. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and airfoils. 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 illustrates some aspects of a non-limiting example of a gas turbine engine in accordance with an embodiment of the present invention. 
           [0007]      FIG. 2  illustrates some aspects of a non-limiting example of an airfoil in accordance with an embodiment of the present invention. 
           [0008]      FIG. 3  illustrates some aspects of a non-limiting example of a cross section of the airfoil of  FIG. 2 . 
           [0009]      FIG. 4  illustrates some aspects of a non-limiting example of a cross section of the airfoil of  FIG. 2 . 
           [0010]      FIGS. 5A-5E  illustrate some aspects of a non-exhaustive group of non-limiting examples of different rib designs in accordance with embodiments 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 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. 
         [0013]    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  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, a wave rotor combustion system, a rotary valve combustion system, or a slinger combustion system, and may employ deflagration and/or detonation combustion processes. 
         [0014]    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 compressor blades (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) 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 . 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 . 
         [0015]    Discharge duct  26  extends between 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 a 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. 
         [0016]    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, and 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 combustor  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 as an engine exhaust flow, e.g., for providing thrust, such as for aircraft propulsion. 
         [0017]    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. The blades rotate with the corresponding rotor during the operation of engine  10 , which may increase the degree of difficulty in cooling the blade because the cooling air tends to migrate radially outward due to centrifugal force. Embodiments of the present invention includes an airfoil configured to mitigate and/or prevent the migration of cooling air due to centrifugal loading. 
         [0018]    Referring to  FIGS. 2-4 , a non-limiting example of some aspects of an airfoil  50  in accordance with an embodiment of the present invention is depicted. In one form, airfoil  50  is a turbine blade. In other embodiments, airfoil  50  may be a compressor blade. In still other embodiments, airfoil  50  may be a turbine or compressor vane. In one form, airfoil  50  is a dual wall airfoil. In other embodiments, airfoil  50  may be a single wall airfoil or an airfoil having more than two walls. Airfoil  50  includes a spar  52  and an outer skin  54 . In one form, airfoil  50  is formed of a conventional aerospace material, such as CMSX-4, available from Cannon Muskegon Corporation of Muskegon, Mich., USA. In other embodiments, other materials, conventional or otherwise, may be employed. 
         [0019]    In one form, spar  52  is hollow, having an internal volume that forms a cooling air supply passage  56 . In other embodiments, one or more other cooling air supply passages may be employed in addition to or in place of cooling air supply passage  56 . In other embodiments, cooling air supply passage  56  may be positioned adjacent to an inner wall other than spar  52 . Spar  52  includes a plurality of cooling air inlet openings  58  extending through the wall of spar  52 . Cooling air supply passage  56  is in fluid communication with cooling air inlet openings  58 . Cooling air supply passage  56  is operative to supply cooling air to cooling air inlet openings  58 . 
         [0020]    Outer skin  54  forms an outer wall for airfoil  50  on both the pressure side and the suction side of airfoil  50 . In one form, outer skin  54  extends around both the pressure side and the suction side. In other embodiments, outer skin  54  may be in the form of individual sheets, e.g., one outer wall for each of the pressure side and the suction side of airfoil  50 , e.g., illustrated in  FIG. 3  as an outer wall  54 A for the pressure side, and an outer wall  54 B for the suction side. Similarly, in one form, spar  52  extends around both the pressure side and the suction side. In other embodiments, spar  52  may be in the form of individual structures, e.g., one inner wall for each of the pressure side and the suction side of airfoil  50 , e.g., illustrated in  FIG. 3  as an inner wall  52 A for the pressure side, and an inner wall  52 B for the suction side. 
         [0021]    Outer skin  54  includes a plurality of cooling air exit openings  60 . Spar  52  forms an inner wall for airfoil  50  on both the pressure side and the suction side of airfoil  50 . Outer skin  54  and spar  52  are spaced apart from each other by a plurality of ribs  62 . In one form, ribs  62  extend between the outer wall formed by outer skin  54  and the inner wall formed by spar  52 . In other embodiments, ribs  62  may extend between other walls in addition to or in place of outer skin  54  and spar  52 . In one form, ribs  62  form a plurality of flow migration dams configured to reduce or prevent cooling air flow migration in a radially outward direction, e.g., due to centrifugal force during the rotation of airfoil  50  in the form of a turbine engine blade. In one form, ribs  62  are oriented horizontally in airfoil  50 . In other embodiments, ribs  62  may be oriented in other directions in addition to or in place of horizontal. In one form, airfoil  50  may have an attachment feature  64  configured to mechanically couple airfoil  50  to engine  10 . In one form, attachment feature  64  is operative to deliver cooling air to cooling air supply passage  56 . 
         [0022]    Each adjacent pair of ribs  62  form therebetween a cooling passage  66 . In one form, ribs  52  and cooling passages  66  are formed on both the pressure side and the suction side of airfoil  50 . In other embodiments, ribs  52  and cooling passages  66  may be formed only on either the pressure side or the suction side of airfoil  50 . In some embodiments, cooling passages  66  may also be formed between ribs  62  and end structures of airfoil  50 , e.g., at the root and tip of airfoil  50  (not shown). In one form, cooling passages  66  are bound by adjacent pairs of ribs  62  and by outer skin  54  and spar  52 . In other embodiments, cooling passages  66  may be bound by other walls in addition to ribs  62 . Cooling passages  66  are in fluid communication with cooling air inlet openings  58  and with cooling air exit openings  60 . In one form, each cooling passage  66  is in fluid communication with cooling air inlet openings  58  at one end, and with cooling air exit openings  60  at the opposite end. In other embodiments, cooling passages  66 , cooling air inlet openings  58  and cooling air exit openings  60  may be arranged otherwise. In one form, each cooling passage  66  adjacent to and in fluid communication with a single cooling air inlet opening  58  and with a single cooling air exit opening  60  and operative to receive cooling air from the single cooling air inlet opening  58  and the single cooling air exit opening  60 . In other embodiments, each cooling passage  66  may be adjacent to and in fluid communication with a plurality of cooling air inlet openings  58  and/or a plurality of cooling air exit openings  60 . 
         [0023]    During engine  10  operation, cooling air is delivered from cooling air supply passage  56  to cooling passages  66  via cooling air inlet openings  58 . The cooling air exits cooling passages  66  via cooling air exit openings  60 . In one form, cooling passages  66  are operative to flow cooling air to remove heat from outer skin  54  and spar  52 . In one form, cooling passages  66  extend continuously between the leading edge portion  68  of airfoil  50  and the trailing edge portion  70  of airfoil  50 . In other embodiments, cooling passages  66  may not extend continuously between leading edge portion  68  and trailing edge portion  70 . 
         [0024]    In one form, ribs  62  are configured to form vortexes  72  in cooling passages  66 . In a particular form, ribs  62  are configured to form a series of vortexes  72  in cooling passages  66 . In one form, ribs  62  are configured to form vortexes on each side of cooling passages  66 . e.g., the top and bottom of each cooling passage  66 . In other embodiments, ribs  62  may not be configured to form vortexes in cooling passages  66 . In one form, ribs  62  include a plurality of trip strips (turbulators)  74  extending from ribs  62  into cooling passages  66 . Trip strips  74  are configured to generate vortexes in the cooling air passing through cooling passages  66 . In other embodiments, trip strips  74  may not extend from ribs  62 , e.g., may be otherwise formed or extend within cooling passages  66 . 
         [0025]    In other embodiments, ribs  62  may be configured to form vortexes by virtue of having a particular shape, e.g., yielding a tortuous flowpath shape of cooling passages  66 , non-limiting examples of which are illustrated in  FIGS. 5A-5E . Other shapes may be employed in other embodiments. 
         [0026]    Embodiments of the present invention include an airfoil for a gas turbine engine, comprising: an outer wall having a plurality cooling air exit openings; an inner wall spaced apart from the outer wall, wherein the inner wall has a plurality of cooling air inlet openings; a plurality of flow migration dams, wherein the flow migration dams extend between the inner wall and the outer wall, the plurality of flow migration dams forming therebetween a plurality of cooling passages, wherein the cooling passages are in fluid communication with the cooling air inlet openings and with the cooling air exit openings; and a cooling air supply passage in fluid communication with the cooling air inlet openings, wherein the cooling air supply passage is operative to supply cooling air to the cooling air inlet openings. 
         [0027]    In a refinement, the airfoil further comprises a one or more trip strips in one or more cooling passages configured to generate one or more vortexes. 
         [0028]    In another refinement, the one or more trip strips extend from the flow migration dams. 
         [0029]    In yet another refinement, the one or more trip strips include a series of trip strips in each cooling passage, wherein the series of trip strips is configured to generate a series of vortexes in each cooling passage. 
         [0030]    In still another refinement, the one or more trip strips extend from the flow migration dams. 
         [0031]    In yet still another refinement, the migration dams are oriented horizontally. 
         [0032]    In a further refinement, the cooling passages each have a first end and a second end opposite the first end, and wherein the cooling passages are in fluid communication with the cooling air inlet openings at the first ends, and in fluid communication with the cooling air exit openings at the second ends. 
         [0033]    In a yet further refinement, the flow migration dams are configured to reduce cooling air flow migration in a radially outward direction due to centrifugal force. 
         [0034]    In a still further refinement, the cooling air supply passage is disposed adjacent to the inner wall. 
         [0035]    In a yet still further refinement, the inner wall forms a spar for the airfoil. 
         [0036]    In an additional refinement, the airfoil is configured as a dual wall airfoil. 
         [0037]    In another additional refinement, the outer wall and the inner wall are disposed on a pressure side of the airfoil, further comprising: a second outer wall disposed on a suction side of the airfoil, the second outer wall having a second plurality cooling air exit openings; a second inner wall disposed on a suction side of the airfoil and spaced apart from the second outer wall, wherein the second inner wall has a second plurality of cooling air inlet openings; a second plurality of flow migration dams, wherein the flow migration dams extend between the second inner wall and the second outer wall, the second plurality of flow migration dams forming therebetween a second plurality of cooling passages, wherein the second cooling passages are in fluid communication with the second cooling air inlet openings and with the second cooling air exit openings, wherein the cooling air supply passage is disposed between the inner wall and the second inner wall, and is in fluid communication with the second cooling air inlet openings, wherein the cooling air supply passage is operative to supply cooling air to the second cooling air inlet openings. 
         [0038]    In yet another additional refinement, the airfoil has a leading edge portion and a trailing edge portion; and wherein the flow migration dams extend continuously between the leading edge portion and the trailing edge portion. 
         [0039]    Embodiments include a turbine engine, comprising: an airfoil, the airfoil including: a hollow spar structure having a plurality of cooling air inlet openings, wherein an internal volume in the hollow spar structure forms a cooling air supply passage operative to deliver cooling air to the cooling air inlet openings; an outer skin spaced apart from the hollow spar structure by a plurality of ribs, wherein the plurality of ribs form a plurality of cooling passages, each cooling passage being defined between an adjacent pair of ribs, wherein the cooling air inlet openings are in fluid communication with the cooling passages; and wherein the outer skin includes a plurality of cooling air exit openings in fluid communication with the plurality of cooling passages. 
         [0040]    In a refinement, the ribs are configured to form vortexes in the cooling passages. 
         [0041]    In another refinement, the ribs are configured to form a series of vortexes in each cooling passage. 
         [0042]    In yet another refinement, the ribs are configured to form vortexes on each side of the cooling passages. 
         [0043]    In still another refinement, one or more ribs include one or more trip strips extending from the one or more ribs, and wherein the one or more trip strips are configured to generate one or more vortexes. 
         [0044]    In a further refinement, the ribs are configured to prevent a migration of flow of cooling air between the hollow spar structure and the outer skin in a radial direction. 
         [0045]    Embodiments of the present invention include an airfoil for a turbine engine, comprising: an outer wall having a plurality cooling air exit openings; an inner wall spaced apart from the outer wall, wherein the inner wall has a plurality of cooling air inlet openings; a cooling air supply passage in fluid communication with the cooling air inlet openings, wherein the cooling air supply passage is operative to supply cooling air to the cooling air inlet openings, and means for cooling the outer wall without allowing flow migration in a radially outward direction, wherein the means for cooling is in fluid communication with both the cooling air inlet openings and the cooling air exit openings. 
         [0046]    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.

Technology Classification (CPC): 8