Abstract:
One embodiment of the present invention is a unique turbine blade 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 turbine blades for gas turbine engines. 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. 61/581,541 filed Dec. 29, 2011, entitled GAS TURBINE ENGINE AND TURBINE BLADE, which is incorporated herein by reference. 
     
    
     FIELD OF THE INVENTION  
       [0002]    The present invention relates to gas turbine engines, and more particularly, gas turbine engines and turbine blades for gas turbine engines. 
       BACKGROUND  
       [0003]    Gas turbine engine turbine blades that effectively cool the blade tip and trailing edge 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 turbine blade 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 turbine blades for gas turbine engines. 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 a turbine blade in accordance with an embodiment of the present invention;  FIG. 2A  depicts some aspects of another non-limiting example of a turbine blade in accordance with an embodiment of the present invention. 
           [0008]      FIG. 3  illustrates some aspects of a non-limiting example of a turbine blade in accordance with an embodiment of the present invention. 
       
    
    
     DETAILED DESCRIPTION 
       [0009]    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. 
         [0010]    Referring to the drawings, and in particular  FIG. 1 , some aspects of a non-limiting example 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. 
         [0011]    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. 
         [0012]    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 of 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). 
         [0013]    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 . 
         [0014]    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 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  is 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. 
         [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, 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. 
         [0017]    Turbine rotor system  34  includes a plurality of blades (not shown in  FIG. 1 ) employed to extract energy from the high temperature high pressure gases in the engine  10  flowpath downstream of combustion system  20 . It is desirable to maintain the temperature of blades within certain temperature limits, e.g., based on the materials and coatings employed in or on the blades. In many cases, turbine blades are cooled by injecting cooling air into the blade. For many blades, the trailing edge, and in particular, the trailing edge portion at the blade tip is difficult to cool. A lack of adequate cooling may result in or increase the likelihood of oxidation and/or corrosion damage, and/or erosion of the blade tip trailing edge. In order to provide cooling to the trailing edge portion of the blade tip, embodiments of the present invention employ a novel tip cooling and squealer tip configuration. 
         [0018]    Referring to  FIG. 2 , some aspects of a non-limiting example of a turbine blade  50  in accordance with an embodiment of the present invention are illustrated. Turbine blade  50  includes a blade attachment feature  52 , a blade platform  54 , an airfoil body  56  culminating in a tip surface  58 , and a squealer tip  60  extending from tip surface  58 . Attachment feature  52  is configured to attached blade  50  to a turbine disk. Blade platform  54  extends from attachment feature  52 . Blade platform  54  is configured to form an inner gas flowpath boundary in conjunction with the blade platforms of the adjacent and other blades in the same turbine blade stage as the depicted blade  50 . In some embodiments, blade  50  may include one or more grooves  62  on each side of the blade for receiving interblade seals and/or dampers (not shown) for sealing between blades  50  and/or damping blades  50 . 
         [0019]    Airfoil body  56  extends radially outward of platform  54 , e.g., in a direction substantially perpendicular to engine centerline  48 . Airfoil body  56  includes a pressure side  64 , a suction side  66 , a leading edge  68  and a trailing edge  70 . Airfoil body  56  extends from a root portion  72  to a tip portion  74 , including a trailing edge tip portion  76 . Trailing edge tip portion  76  is formed, in part, by squealer tip  60 . Squealer tip  60  extends outwardly from tip surface  58 , e.g., radially outward in a direction substantially perpendicular to engine centerline  48 . Squealer tip  60  includes a pressure side rail portion  78  and a suction side rail portion  80 . Pressure side rail portion  78  and suction side rail portion  80  form therebetween a cavity  82  adjacent to and radially bounded on its bottom by tip surface  58 . Cavity  82  is disposed between pressure side rail portion  78  and suction side rail portion  80 . In one form, suction side rail portion  80  extends to trailing edge  70 . In one form, pressure side rail  78  does not extend to trailing edge  70 . In other embodiments, either or both of pressure side rail portion  78  and suction side rail portion  80  may or may not extend to trailing edge  70 . 
         [0020]    Disposed at a leading edge tip portion  81  of airfoil body  56  is a leading edge rail portion  84  of squealer tip  60 . Pressure side rail portion  78  and suction side rail portion  80  are joined together by leading edge rail portion  84 . Leading edge rail portion  84  further forms cavity  82 , bounding cavity  82  at leading edge tip portion  81 . 
         [0021]    In one form, squealer tip  60  is offset from pressure side  64  of airfoil body  56 . In particular, in one form, pressure side rail portion  78  is offset along tip surface  58  of airfoil body  56  from pressure side  64 . In one form, squealer tip  60  extends outward from tip surface  58  at suction side  66 , i.e., suction side rail portion  80  is not offset from suction side  66 . In other embodiments, squealer tip  60  may be offset from suction side  66 , e.g., wherein suction side rail portion  80  is offset along tip surface  58  from suction side  66  of airfoil body  56 , e.g., as depicted in  FIG. 2A . In one form, a plurality of openings  86  are disposed in tip surface  58  in the tip surface  58  land extending between the pressure side  64  surface of airfoil body  56  and pressure side rail portion  78  and in the tip surface  58  land extending between leading edge  68  and leading edge rail portion  84 . In embodiments having an offset suction side rail portion, openings  88  may also be included in tip surface  58  in the tip surface  58  land extending between the suction side  66  surface of airfoil body  56  and suction side rail portion  80 . Openings  86  are configured to discharge air from tip surface  58 , e.g., cooling air and/or purge air. In one form, a plurality of openings  88  are disposed in tip surface  58  between pressure side rail portion  78  and suction side rail portion  80  of squealer tip  60 . Openings  88  are configured to discharge air from tip surface  58 , e.g., cooling air and/or purge air, into cavity  82 . Openings  86  and  88  are supplied with air via one or more internal passages  90  disposed within blade  50 , e.g., within airfoil body  56 . 
         [0022]    Squealer tip  60  includes a passage  92  extending between pressure side rail portion  78  and suction side rail portion  80 . Passage  92  is configured to expose trailing edge tip portion  76  to cavity  82 , and to distribute air from cavity  82  discharged by openings  88  to trailing edge tip portion  76 . In one form, passage  92  is in the form of a gap between pressure side rail portion  78  and suction side rail portion  80  adjacent to trailing edge tip portion  76 . 
         [0023]    Referring to  FIG. 3 , some aspects of a non-limiting example of a turbine blade  50  in accordance with an embodiment of the present invention are illustrated. In particular,  FIG. 3  illustrates examples of potential air flow from openings  86  and  88  that cool squealer tip  60  and trailing edge tip portion  76  during the operation of engine  10 . The air flow from openings  86  and  88  is illustrated using arrowed lines  94 . From the illustration of  FIG. 3 , it is seen that the air discharged from openings  86  flows along and over pressure side rail portion  78  and leading edge rail portion  84  of squealer tip  60 , providing film cooling to pressure side rail portion  78  and leading edge rail portion  84 , and removing heat from pressure side rail portion  78  and leading edge rail portion  84 . In some embodiments, air discharged from openings  86  may also flow along pressure side rail portion  78  toward suction side rail portion  80  at trailing edge tip portion  76 . Air discharged from openings  88  flows through cavity  82  and along and over suction side side rail portion  80  of squealer tip  60 , providing film cooling to suction side rail portion  80 , and removing heat from suction side rail portion  80 . In particular, it is seen from  FIG. 3  that some of the air discharged from openings  88  flows through passage  92  between pressure side rail portion  78  and suction side rail portion  80  of squealer tip  60 , providing film cooling to and removing heat from trailing edge tip portion  76 , including suction side rail portion  80  and trailing edge  70 . 
         [0024]    Embodiments of the present invention include a turbine blade for a gas turbine engine, comprising: an airfoil body having a pressure side, a suction side and a trailing edge tip portion, wherein the airfoil body culminates at a tip surface; and a squealer tip extending outwardly from the tip surface and having a pressure side rail portion and a suction side rail portion forming a cavity therebetween, wherein the squealer tip also includes a passage extending between the pressure side rail portion and the suction side rail portion configured to expose the trailing edge tip portion to the cavity; and wherein the pressure side rail portion is offset from the pressure side of the airfoil body. 
         [0025]    In a refinement, the passage is a gap between the pressure side rail portion and the suction side rail portion. 
         [0026]    In another refinement, the trailing edge tip portion includes a trailing edge; and wherein the suction side rail portion extends to the trailing edge. 
         [0027]    In yet another refinement, the trailing edge tip portion includes a trailing edge; and wherein the pressure side rail portion does not extend to the trailing edge. 
         [0028]    In still another refinement, the turbine blade further comprises at least one opening disposed in the tip surface between the pressure side of the airfoil body and the pressure side rail portion of the squealer tip, wherein the at least one opening is configured to discharge air from the tip surface. 
         [0029]    In yet still another refinement, the at least one opening is a plurality of openings. 
         [0030]    In a further refinement, the turbine blade further comprises at least one opening disposed in the tip surface between the pressure side rail portion of the squealer tip and the suction side rail portion of the squealer tip, wherein the at least one opening is configured to discharge air from the tip surface into the cavity. 
         [0031]    In a yet further refinement, the at least one opening is a plurality of openings. 
         [0032]    In a still further refinement, the passage is configured to distribute air from the cavity to the trailing edge tip portion. 
         [0033]    In a yet still further refinement, the airfoil body includes a leading edge tip portion; wherein the squealer tip includes a leading edge rail portion disposed at the leading edge tip portion, and wherein the pressure side rail portion of the squealer tip and the suction side rail portion of the squealer tip are joined together by the leading edge rail portion of the squealer tip. 
         [0034]    In another refinement, the leading edge tip portion includes a leading edge, further comprising an opening in the tip surface, wherein the opening is disposed between the leading edge of the airfoil body and the leading edge rail portion of the squealer tip; and wherein the opening is configured to discharge air from the tip surface. 
         [0035]    Embodiments of the present invention include a turbine blade for a gas turbine engine, comprising: an airfoil body having a pressure side, a suction side and a trailing edge tip portion, wherein the airfoil body culminates at a tip surface; and a squealer tip extending outwardly from the tip surface, wherein the squealer tip is offset from the pressure side of the airfoil body and extends outward from the tip surface at the suction side of the airfoil body; and wherein the squealer tip is configured to form a cavity therein and a passage exposing the cavity to the trailing edge tip portion. 
         [0036]    In a refinement, the squealer tip includes a pressure side rail portion and a suction side rail portion that form the cavity therebetween. 
         [0037]    In another refinement, the passage is a gap between the pressure side rail portion and the suction side rail portion. 
         [0038]    In yet another refinement, the trailing edge tip portion includes a trailing edge; and wherein the suction side rail portion extends to the trailing edge. 
         [0039]    In still another refinement, the airfoil body includes a leading edge tip portion; wherein the squealer tip includes a leading edge rail portion disposed at the leading edge tip portion, and wherein the pressure side rail portion of the squealer tip and the suction side rail portion of the squealer tip are joined together by the leading edge rail portion of the squealer tip. 
         [0040]    In yet still another refinement, the turbine blade further comprises a plurality of openings disposed in the tip surface between the pressure side of the airfoil body and the squealer tip, wherein the plurality of openings are configured to discharge air from the tip surface. 
         [0041]    In a further refinement, the turbine blade further comprises a plurality of openings disposed in the tip surface and positioned to discharge air from the tip surface into the cavity. 
         [0042]    In a yet further refinement, the passage is configured to distribute air from the cavity to the trailing edge tip portion. 
         [0043]    Embodiments of the present invention include a gas turbine engine, comprising: a compressor; a combustor in fluid communication with the compressor; and a turbine in fluid communication with the combustor, wherein the turbine includes a plurality of turbine blades, wherein at least one of the turbine blades includes: an airfoil body having a trailing edge tip portion; and means for cooling the trailing edge tip portion. 
         [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.