Patent Publication Number: US-2022235664-A1

Title: Turbine airfoil incorporating modal frequency response tuning

Description:
BACKGROUND 
     1. Field 
     The present disclosure relates to airfoils used in gas turbine engines, and in particular, to a turbine airfoil having a tuned modal frequency response and a method for forming such a turbine airfoil. 
     2. Description of the Related Art 
     In a turbomachine, such as a gas turbine engine, air is pressurized in a compressor section and then mixed with fuel and burned in a combustor section to generate hot combustion gases. The hot combustion gases are expanded within a turbine section of the engine where energy is extracted to power the compressor section and to produce useful work, such as turning a generator to produce electricity. The hot combustion gases travel through a series of turbine stages within the turbine section. A turbine stage may include a row of stationary airfoils, i.e., vanes, followed by a row of rotating airfoils, i.e., turbine blades, where the turbine blades extract energy from the hot combustion gases for providing output power. 
     Modal frequency response tuning of turbine airfoils involves modifying the airfoils to move the natural vibrational frequency of the airfoils outside of an operating speed range of the gas turbine engine, to avoid resonance. Modal frequency response tuning of turbine airfoils may be challenging due to the limited number of design parameters available to steer the design away from the resonance. Furthermore, the currently available design parameters are simultaneously coupled to several modal frequencies, which may result in design tradeoffs. 
     SUMMARY 
     Briefly, aspects of the present disclosure are directed to modal frequency response tuning of turbine airfoils by modifying a geometry of an internal flow displacement element. 
     According to a first aspect, a turbine airfoil is provided. The turbine airfoil comprises an airfoil body formed by an outer wall, the outer wall comprising a pressure side wall and a suction side wall joined at a leading edge and at a trailing edge. The turbine airfoil further comprises a generally hollow flow displacement element positioned in an interior portion of the airfoil body and extending along a span-wise extent thereof. The flow displacement element defines an inactive cavity therewithin. The flow displacement element is spaced from the pressure side wall and the suction side wall to respectively define a first near-wall cooling flow channel and a second near-wall cooling flow channel. The flow displacement element comprises an outer surface facing said near-wall cooling flow channels and an inner surface facing the inactive cavity. The inner surface facing the inactive cavity comprises features configured to influence a mass and/or stiffness of the turbine airfoil, to thereby produce a predetermined modal frequency response of the turbine airfoil. 
     According to a second aspect, a method is provided for forming a turbine airfoil with a tuned modal frequency response for use in a turbine engine. The method comprises obtaining a first geometry of the turbine airfoil. The first geometry is a nominal geometry of the turbine airfoil defined by: an airfoil body formed by an outer wall comprising a pressure side wall and a suction side wall joined at a leading edge and at a trailing edge, and a generally hollow flow displacement element positioned in an interior portion of the airfoil body and extending along a span-wise extent thereof. The flow displacement element defines an inactive cavity therewithin. The flow displacement element is spaced from the pressure side wall and the suction side wall to respectively define a first near-wall cooling flow channel and a second near-wall cooling flow channel. The flow displacement element comprises an outer surface facing said near-wall cooling flow channels and an inner surface facing the inactive cavity. The method comprises determining a first natural frequency associated with the first geometry of the turbine airfoil and determining whether the first natural frequency occurs within a defined operating speed range of the turbine engine. The method further comprises determining a second geometry of the turbine airfoil, the second geometry differing from the first geometry in the provision of mass and/or stiffness influencing features on the inner surface of the flow displacement element facing the inactive cavity. The second geometry is associated with a second natural frequency that occurs outside the defined operating speed range of the turbine engine. The method then comprises manufacturing the turbine airfoil based on the determined second geometry. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The invention is shown in more detail by help of figures. The figures show preferred configurations and do not limit the scope of the invention. 
         FIG. 1  is a perspective view of a turbine airfoil; 
         FIG. 2  is a cross-sectional view along the section II-II of  FIG. 1 ; 
         FIG. 3  is a cross-sectional view along the section of  FIG. 2 , illustrating a nominal geometry of an internal flow displacement element; 
         FIG. 4  is a cross-sectional view, corresponding to the view of  FIG. 3 , illustrating a modified geometry of the internal flow displacement element according to one embodiment; 
         FIG. 5  is a cross-sectional view, corresponding to the view of  FIG. 2 , illustrating a modified geometry of the internal flow displacement element according to further embodiments; 
         FIG. 6  is a cross-sectional view along the section VI-VI of  FIG. 5 ; 
         FIG. 7  is a cross-sectional view, corresponding to the view of  FIG. 6 , illustrating a modified geometry of the internal flow displacement element according to a still further embodiment; and 
         FIG. 8  is a flowchart illustrating a method for manufacturing an airfoil according to one embodiment. 
     
    
    
     DETAILED DESCRIPTION 
     In the following detailed description of the various embodiments, reference is made to the accompanying drawings that form a part hereof, and in which is shown by way of illustration, and not by way of limitation, a specific embodiment in which the invention may be practiced. It is to be understood that other embodiments may be utilized and that changes may be made without departing from the spirit and scope of the present invention. 
     Modal frequency response tuning may be achieved via airfoil design modifications, such as by modifying the airfoil chord, camber, twist, etc., blade shank redesign, internal rib positioning, among others. All of these design aspects are directly coupled to thermal and/or aerodynamic efficiency of the engine. The present inventors recognize that a modification of one or more of the above design aspects often leads to a compromise in performance and/or mechanical integrity of the engine. 
     The embodiments described herein eliminate at least some of the tradeoffs between aerodynamic efficiency, coolant reduction and modal frequency response tuning that are faced in the state of the art, offering a design strategy that allows independent efficiency optimization in each of the above-mentioned categories. The above may be achieved by providing mass and/or stiffness influencing features in an internal flow displacement element of a turbine airfoil, to tune a modal frequency response, without altering the geometry of the internal cooling flow channels and the external shape of the turbine airfoil. The concepts described herein may be extended to multiple stages of blades and vanes to provide significant aero gains and increase in turbine overall efficiency. 
     Referring now to  FIG. 1 , an example turbine airfoil  10  is illustrated. In this example, the airfoil  10  is a turbine blade of a gas turbine engine. It should however be noted that aspects of the invention could also be incorporated into a stationary vane of a gas turbine engine. The illustrated airfoil  10  includes an airfoil body  12  formed from an outer wall  18  adapted for use, for example, in a turbine section of an axial flow gas turbine engine. The outer wall  18  extends span-wise along a radial direction of the turbine engine and includes a generally concave shaped pressure side wall  20  and a generally convex shaped suction side wall  22 . The pressure side wall  20  and the suction side wall  22  are joined at a leading edge  24  and at a trailing edge  26 . As illustrated, the airfoil body  12  extends radially outward from a platform  14  toward a tip  15  of the turbine airfoil  10 . A root portion  16  extends radially inward from the platform  14 , for coupling the turbine airfoil  10  to a rotor disc (not shown). In the present description, a radial or span-wise direction is understood to be a direction extending from the root portion  16  toward the tip  15  of the turbine airfoil. 
       FIGS. 2 and 3  are cross-sectional views depicting a nominal or first geometry of the turbine airfoil  10 . Referring to  FIG. 2 , the turbine airfoil  10  has a chordal axis  17  extending centrally between the pressure side wall  20  and the suction side wall  22  from the leading edge  24  to the trailing edge  26 . A chord-wise direction may be defined as a direction parallel to the chordal axis  17 . As illustrated, the airfoil body  12  comprises an interior portion which may receive a cooling fluid, such as air from a compressor section (not shown), via one or more cooling fluid supply passages formed through the root portion  16 . In the present embodiment, the turbine airfoil  10  comprises a plurality of span-wise extending partition walls  28  positioned in the interior portion of the airfoil body  12 . The partition walls  28  connect the pressure side wall  20  and the suction side wall  22  and are spaced along a chord-wise direction to define radial cooling flow channels  19  therebetween. The turbine airfoil  10  comprises at least one flow displacement element  30  positioned in the interior portion of the airfoil body  12  and extending along a span-wise extent thereof. In the shown example, the flow displacement element  30  is positioned between a pair of adjacent partition walls  28 . The flow displacement element  30  is generally hollow, defining an internal cavity  40  within. The flow displacement element  30  is spaced from the pressure side wall  20  and the suction side wall  22  to respectively define a first near-wall cooling flow channel  92  and a second near-wall cooling flow channel  94 . In the shown embodiment, a connecting channel  96  is formed between the flow displacement element  30  and an adjacent partition wall  28 , which connects the near-wall cooling flow channels  92  and  94  to form a radial flow pass with a C-shaped cross-section. 
     The flow displacement element  30  may be manufactured integrally with the airfoil body. In the shown embodiment, the turbine airfoil  10  is provided with a pair of connector ribs  32 ,  34  that respectively connect the flow displacement  30  element to the pressure side wall  20  and to the suction side wall  22  along a span-wise extent. As a result, a pair of C-shaped radial flow passes of symmetrically opposed cross-sections are formed on either side of the connector ribs  32 ,  34 . In other embodiments, additionally or alternately, connector ribs may be provided to connect the flow displacement element  30  one or more of the partition walls  28 . 
     The cavity  40  defined within the flow displacement element  30  is an inactive cavity. That is to say that the cavity  40  is a dead space that does not allow any active flow of fluids. The flow displacement element  30  thereby serves to reduce the cross-sectional area of the radial flow of the cooling fluid and to displace the cooling fluid toward the pressure side  20  and the suction side  22 , i.e., move a major part of the cooling fluid into the near-wall cooling flow channels  92  and  94 . As shown in  FIG. 3 , the inactive cavity  40  extends span-wise from a first end  42  to a second end  44 . In the shown embodiment, the first end  42  is located proximate to the tip  15  and the second end  44  is located at the root portion  16 . The inactive cavity  40  is open at the first end  42 , and closed at the second end  44 . The near-wall cooling flow channels  92  and  94  are connected to a cooling fluid inlet  21  located at the root portion  16 . The flow displacement element  30  comprises an outer surface  36  facing the near-wall cooling flow channels  92 ,  94  and an inner surface  38  facing the inactive cavity  40 . In the nominal geometry shown in  FIGS. 2 and 3  the outer surface  36  and the inner surface  38  of the flow displacement element  30  are generally parallel to each other, such that a wall thickness t of the flow displacement element  30 , defined between the outer surface  36  and the inner surface  38 , is substantially constant along the span-wise and chord-wise directions. 
     In accordance with the present disclosure, a modal frequency response of the turbine airfoil may be tuned by modifying the first (nominal) geometry of the turbine airfoil to form a second (adapted) geometry having a predetermined modal frequency response. In particular, the second geometry may be determined such that the turbine airfoil has a natural (modal) frequency that occurs outside a defined normal operating speed range of the turbine engine. The second geometry essentially differs from the first geometry in the provision of mass and/or stiffness influencing features  50  on the inner surface  38  of the flow displacement element  30  facing the inactive cavity  40 . Examples of such mass and/or stiffness influencing features  50  are illustrated referring to  FIG. 4-7 . 
     In a first embodiment of an adapted geometry shown in  FIG. 4 , the mass and/or stiffness influencing features  50  are realized by providing a contour to the inner surface  38  of the flow displacement element  30 , such that the flow displacement element  30  has an altered wall thickness t at one or more locations in relation to the first geometry of the turbine airfoil  10  (see  FIG. 3 ). The view shown in  FIG. 4  illustrates a two-dimensional contouring of the surface  38 , whereby the wall thickness t of the flow displacement element  30  is varied in the span-wise direction. In an alternate embodiment, the surface  38  may be provided with a two-dimensional contouring, whereby the wall thickness t of the flow displacement element  30  is varied in the chord-wise direction. In a still alternate embodiment, the surface  38  may be provided with a three-dimensional contouring, whereby the wall thickness t of the flow displacement element  30  is varied in both the span-wise and chord-wise directions. The wall thickness distribution of the flow displacement element  30  may be designed to customize both the stiffness and mass of the turbine airfoil  10  and drive the modal frequencies up or down to avoid a resonance within the defined normal operating speed range of the turbine engine. In one embodiment, the inner surface  38  of the flow displacement element may be contoured by determining thicknesses t 1 , t 2 , t 3  at various span-wise (and/or chord-wise) sections, to achieve a defined modal frequency response, and joining the sections by a smooth function curve. In other embodiments, the inner surface  38  of the flow displacement element  30  may be three-dimensionally contoured to define one or more crests (hills), one or more troughs (valleys), or a combination of crests and troughs. 
     Turning now  FIGS. 5 and 6 , a second embodiment of an adapted geometry of a turbine airfoil is illustrated that employs multiple different types of mass and/or stiffness influencing features  50 . A first type of mass and/or stiffness enhancing feature  50  depicted herein includes one or more chord-wise extending protrusions  60  on the inner surface  38  of the flow displacement element  30 . The chord-wise extending protrusions  60  have a significantly higher influence on the stiffness than the mass of the turbine airfoil  10 , and may hence be referred to as stiffening bands. In another variant (not shown), the mass and/or stiffness influencing features may comprise one or more chord-wise extending depressions or grooves that may serve to reduce the overall stiffness of the turbine airfoil. A second type of mass and/or stiffness influencing features  50  depicted herein includes one or more rods  70  located at an interior of the flow displacement element  30 . Each rod  70  extends between a first end  72  and a second end  74 . The first and second ends  72 ,  74  of each rod  70  are connected to the inner surface  38  of the flow displacement element  30  at a first and a second location respectively. As a result, the rods  70  also have a significantly higher influence on the stiffness than the mass of the turbine airfoil  10 , and may hence be referred to stiffening rods. As shown in  FIG. 5  the stiffening rods  70  may be arranged in a chord-wise extending array in the interior of the flow displacement element  30 . Alternately or additionally, the stiffening rods  70  may be arranged in a span-wise extending array in the interior of the flow displacement element  30 . In the illustrated embodiment, the first and second ends  72 ,  74  of the stiffening rods  70  are positioned on the stiffening bands  60 . 
     Referring to  FIG. 7 , a third embodiment of an adapted geometry of a turbine airfoil is illustrated. Herein the mass and/or stiffness influencing features  50  comprise one or more cantilevered rods  80  located at an interior of the flow displacement element  30 . Each cantilevered rod  80  extends from a first end  82  connected to the inner surface  38  of the flow displacement element  30  to a second end  84 , the second end  84  being a free end. In this case, the cantilevered rods  80  have a significantly higher influence on the mass than the stiffness of the turbine airfoil  10 . 
     In various embodiments, different types of mass and/or stiffness influencing features  50 , such as a contoured inner surface  38 , one or more stiffening bands  60 , one or more stiffening rods  70  and one or more cantilevered rods  80  may be employed individually or as a combination of two or more different types of features to allow greater flexibility in designing a turbine airfoil with a predetermined modal frequency response. 
       FIG. 8  is a flowchart illustrating a method  100  for manufacturing a turbine airfoil according to one embodiment. Block  102  of the method  100  comprises obtaining a first geometry of the turbine airfoil, the first geometry being a nominal geometry of the turbine airfoil. The nominal geometry of the turbine airfoil may be obtained, for example, based on aerodynamic and heat transfer considerations, among other factors. To this end, the nominal geometry may specify an external geometry of the airfoil body and an internal geometry of the cooling channels of the turbine airfoil, among others. An example of a nominal geometry for a turbine airfoil is illustrated referring to  FIG. 2-3 . Block  104  of the method  100  includes determining at least one first natural frequency, corresponding to one or more vibration modes, associated with the first geometry of the turbine airfoil. Block  106  of the method  100  involves determining whether the above-determined natural frequency, for any of the vibration modes, occurs within a defined operating speed range of the turbine engine, i.e., determining whether a condition for resonance is met. In general, it is desirable to avoid one or more forced response drivers, such as, combustor can count, upstream vane count, etc., at all speeds within the defined speed range of the turbine engine. If a condition for resonance is determined, block  108  of the method  100  involves determining a second (adapted) geometry of the turbine airfoil  10 . The second geometry differs from the first geometry in the provision of mass and/or stiffness influencing features  50  on the inner surface  38  of the flow displacement element  30  facing the inactive cavity  40 . Examples of adapted geometries are illustrated referring to  FIG. 4-7 . The second (adapted) geometry is determined such that the associated second natural frequencies, for any of the vibration modes, occurs outside the defined operating speed range of the turbine engine. Block  110  of the method  100  involves manufacturing the turbine airfoil  10  based on the determined second geometry. 
     The above-described method provides significantly reduced complexity in relation to a conventional method of frequency response tuning of a turbine airfoil that would typically require a reevaluation of aerodynamic and cooling performance before the tuning change is accepted. By providing mass and/or stiffness influencing features  50  on a surface  38  facing an inactive cavity  40 , it may be ensured that the near-wall cooling flow channels  92 ,  94  are essentially unaltered in the second geometry of the turbine airfoil  10  in relation to the first geometry of the turbine airfoil  10 . The described embodiments do not depend on any external modification to the airfoil geometry such as chord, camber, twist, root redesign, etc., whereby the external geometry of the airfoil body  12  may also be unaltered in the second geometry of the turbine airfoil  10  in relation to the first geometry of the turbine airfoil  10 . The embodiments described thus eliminate at least some of the tradeoffs between aerodynamic efficiency, coolant reduction and modal frequency response tuning, as faced in the state of the art. 
     In an example embodiment, the flow displacement element  30  may be formed integrally with the airfoil body  12 . In this case, the manufacturing process may involve any technique that does not require post manufacturing assembly as in the case of inserts. In one example, the flow displacement element  30  may be cast integrally with the airfoil body  12 , for example from a ceramic casting core. Other manufacturing techniques may include, for example, additive manufacturing processes such as 3-D printing. This allows the inventive design to be used for highly contoured airfoils, including 3-D contoured blades and vanes. 
     While specific embodiments have been described in detail, those with ordinary skill in the art will appreciate that various modifications and alternative to those details could be developed in light of the overall teachings of the disclosure. Accordingly, the particular arrangements disclosed are meant to be illustrative only and not limiting as to the scope of the invention, which is to be given the full breadth of the appended claims, and any and all equivalents thereof