Patent Publication Number: US-2016230566-A1

Title: Angled pedestals for cooling channels

Description:
STATEMENT OF GOVERNMENT RIGHTS 
     This invention was made with government support under contract no. N68335-13-C-0005 awarded by the Navy. The government has certain rights in the invention. 
    
    
     BACKGROUND 
     1. Field 
     The present disclosure relates to cooling systems, more specifically to pedestals disposed within cooling channels (e.g., within turbomachine components). 
     2. Description of Related Art 
     Certain cooling channels can include pedestals stretching into or across the channel perpendicular to the flow path. The pedestals provide additional surface area and increase heat transfer in the cooling channels. 
     Such conventional methods and systems have generally been considered satisfactory for their intended purpose. However, there is still a need in the art for improved cooling systems. The present disclosure provides a solution for this need. 
     SUMMARY 
     A turbomachine component includes a body defining a cooling inlet and a cooling outlet in fluid communication with each other through a cooling channel that defines a longitudinal axis, and at least one pedestal disposed within the cooling channel. The pedestal is angled within the cooling channel relative to the longitudinal axis. The turbomachine component can be a vane, a blade, a blade outer air seal, a combustor panel, or any other suitable component. The at least one pedestal can be angled between about 10 degrees to about 90 degrees relative to the longitudinal axis, or to any other suitable angle. 
     The pedestal can be angled tangentially circumferential within the cooling channel relative to an axial direction that is parallel to the longitudinal axis. The pedestal can be angled radially within the cooling channel relative to an axial direction that is perpendicular to the longitudinal axis. The pedestal can also be angled tangentially circumferentially and radially within the cooling channel simultaneously. The pedestal can include a circular, elliptical, or square cross-sectional shape, or any other suitable shape. 
     The pedestal can include a plurality of pedestals that are angled within the cooling channel relative to the longitudinal axis. At least two pedestals in the plurality of pedestals can have a different angle degree and/or angle direction relative to each other. 
     The plurality of pedestals can be disposed in a predetermined pattern. The predetermined pattern can include one or more rows or columns. 
     In at least one aspect of this disclosure, a method includes forming a turbomachine component having a body defining a cooling inlet and a cooling outlet in fluid communication with each other through a cooling channel that defines a longitudinal axis, wherein at least one pedestal is disposed within the cooling channel, wherein the pedestal is angled within the cooling channel relative to the longitudinal axis. Forming can include shaping the body into a turbomachine blade including an airfoil. Forming can include at least one of additively manufacturing the turbomachine component, using electric discharge machining (EDM) to manufacture the turbomachine component, or using a laser to manufacture the turbomachine component to produce, e.g., a negative of the airfoil. 
     In at least one aspect of this disclosure, a turbomachine includes a turbomachine component as described above. 
     These and other features of the systems and methods of the subject disclosure will become more readily apparent to those skilled in the art from the following detailed description taken in conjunction with the drawings. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       So that those skilled in the art to which the subject disclosure appertains will readily understand how to make and use the devices and methods of the subject disclosure without undue experimentation, embodiments thereof will be described in detail herein below with reference to certain figures, wherein: 
         FIG. 1  is a schematic, partial cross-sectional view of a turbomachine in accordance with this disclosure; 
         FIG. 2A  is a cross-sectional end view of an embodiment of a turbomachine component in accordance with this disclosure, shown with pedestals disposed in a cooling channel thereof that are angled relative to an axis A of the turbomachine such the pedestals diverge from axis A; 
         FIG. 2B  is a partial cross-sectional view of the turbomachine component of  FIG. 2A  taken along line  2 B- 2 B, showing the pedestals not angled in a radial direction of the turbomachine; 
         FIG. 2C  is a partial cross-sectional view of the turbomachine component of  FIG. 2A  taken along line  2 C- 2 C, showing the pedestals angled only circumferentially tangential relative to the turbomachine axis A; 
         FIG. 3A  is a cross-sectional elevation view of an embodiment of a turbomachine component in accordance with this disclosure, shown with pedestals disposed in a cooling channel thereof that are angled relative to axis A of the turbomachine opposite to that of the embodiment of  FIG. 2A  such the pedestals converge toward axis A; 
         FIG. 3B  is a partial cross-sectional view of the turbomachine component of  FIG. 3A  taken along line  3 B- 3 B, showing the pedestals not angled in a radial direction of the turbomachine; 
         FIG. 3C  is a partial cross-sectional view of the turbomachine component of  FIG. 3A  taken along line  3 C- 3 C, showing the pedestals angled only circumferentially tangential relative to the turbomachine axis A; 
         FIG. 4A  is a cross-sectional elevation view of an embodiment of a turbomachine component in accordance with this disclosure, shown with pedestals disposed in a cooling channel thereof that are not angled in the axial direction of axis A, i.e., the pedestals do not converge or diverge relative to axis A of turbomachine; 
         FIG. 4B  is a partial cross-sectional view of the turbomachine component of  FIG. 4A  taken along line  4 B- 4 B, showing the pedestals angled in the radial direction of the turbomachine; 
         FIG. 4C  is a partial cross-sectional view of the turbomachine component of  FIG. 4A  taken along line  4 C- 4 C, showing the pedestals angled in only the radial direction; 
         FIG. 5A  is a cross-sectional elevation view of an embodiment of a turbomachine component in accordance with this disclosure, shown with pedestals disposed in a cooling channel thereof that are not angled in the axial direction of axis A, i.e., the pedestals do not converge or diverge relative to axis A of turbomachine; 
         FIG. 5B  is a partial cross-sectional view of the turbomachine component of  FIG. 5A  taken along line  5 B- 5 B, showing the pedestals angled in the radial direction of the turbomachine; 
         FIG. 5C  is a partial cross-sectional view of the turbomachine component of  FIG. 5A  taken along line  5 C- 5 C, showing the pedestals angled in only the radial direction opposite to that of the embodiment of  FIG. 4A ; 
         FIG. 6A  is a cross-sectional elevation view of an embodiment of a turbomachine component in accordance with this disclosure, shown with pedestals disposed in a cooling channel thereof that are angled relative to both an axial and radial direction of the turbomachine, i.e., circumferentially tangential and radially angled; 
         FIG. 6B  is a partial cross-sectional view of the turbomachine component of  FIG. 6A  taken along line  6 B- 6 B, showing the pedestals angled in the radial direction of the turbomachine; 
         FIG. 6C  is a partial cross-sectional view of the turbomachine component of  FIG. 6A  taken along line  6 C- 6 C, showing the pedestals angled in both the axial and radial directions; 
         FIG. 7A  is a cross-sectional elevation view of an embodiment of a turbomachine component in accordance with this disclosure, shown with pedestals disposed in a cooling channel thereof that are angled relative to both an axial and radial direction of the turbomachine, i.e., circumferentially tangential and radially angled; 
         FIG. 7B  is a partial cross-sectional view of the turbomachine component of  FIG. 7A  taken along line  7 B- 7 B, showing the pedestals angled in the radial direction of the turbomachine; and 
         FIG. 7C  is a partial cross-sectional view of the turbomachine component of  FIG. 7A  taken along line  7 C- 7 C, showing the pedestals angled in both the axial and radial directions. 
     
    
    
     DETAILED DESCRIPTION 
     Reference will now be made to the drawings wherein like reference numerals identify similar structural features or aspects of the subject disclosure. For purposes of explanation and illustration, and not limitation, an illustrative view of an embodiment of a turbomachine blade in accordance with the disclosure is shown in  FIG. 2  and is designated generally by reference character  200 . Other embodiments and/or aspects of this disclosure are shown in  FIGS. 1 and 2B-7C . The systems and methods described herein can be used to enhance thermal transfer within components that have cooling channels. 
       FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as a two-spool turbofan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section  22  drives air along a bypass flow path B in a bypass duct defined within a nacelle  15 , while the compressor section  24  drives air along a core flow path C for compression and communication into the combustor section  26  then expansion through the turbine section  28 . Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. 
     The exemplary engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided and the location of bearing systems  38  may be varied as appropriate to the application. 
     The low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a first (or low) pressure compressor  44  and a first (or low) pressure turbine  46 . The inner shaft  40  is connected to the fan  42  through a speed change mechanism, which in exemplary gas turbine engine  20  is illustrated as a gear system  48  to drive the fan  42  at a lower speed than the low speed spool  30 . The high speed spool  32  includes an outer shaft  50  that interconnects a second (or high) pressure compressor  52  and a second (or high) pressure turbine  54 . A combustor  56  is arranged in exemplary gas turbine  20  between the high pressure compressor  52  and the high pressure turbine  54 . A mid-turbine frame  57  of the engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The mid-turbine frame  57  further supports bearing systems  38  in the turbine section  28 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via bearing systems  38  about the engine central longitudinal axis A which is collinear with their longitudinal axes. 
     The core airflow is compressed by the low pressure compressor  44  then the high pressure compressor  52 , mixed and burned with fuel in the combustor  56 , then expanded over the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  57  includes airfoils  59  which are in the core airflow path C. The turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. It will be appreciated that each of the positions of the fan section  22 , compressor section  24 , combustor section  26 , turbine section  28 , and fan gear system  48  may be varied. For example, gear system  48  may be located aft of combustor section  26  or even aft of turbine section  28 , and fan section  22  may be positioned forward or aft of the location of gear system  48 . 
     The engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the engine  20  bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine  46  has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine  20  bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor  44 , and the low pressure turbine  46  has a pressure ratio that is greater than about five (5:1). Low pressure turbine  46  pressure ratio is pressure measured prior to inlet of low pressure turbine  46  as related to the pressure at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. The geared architecture may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans. 
     A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section  22  of the engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft (10,668 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane 79 (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)]̂0.5. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.5 meters/second). 
     Referring to  FIGS. 2A-2C , a turbomachine component  200  includes a body  202  defining a cooling inlet  202   a  ( FIG. 2C ) and a cooling outlet  202   b  in fluid communication with each other through a cooling channel  201  that defines a longitudinal axis, and one or more pedestals  203  disposed within the cooling channel  201 . One or more of the pedestals  203  are angled within the cooling channel  201  relative to the longitudinal axis. Another way to describe the angle of the pedestals  203  is that the pedestals  203  or a portion thereof can meet one or both of the interior walls  201   a ,  201   b  that define the cooling channel  201  at a non-right angle. The pedestals  203  can be angled to any suitable angle degree (e.g., about 10 degrees to about 90 degrees). 
     As shown in  FIG. 2C , the pedestals  203  can include a circular cross-sectional shape, however, it is contemplated that any suitable cross-sectional shape (constant or variable) or combinations thereof can be used. For example, one or more of the pedestals  203  can include an elliptical, square, or other suitable cross-sectional shape and may have any suitable thickness or changing thickness along the length of the pedestals  203 . 
     The turbomachine component  200  can be a blade or vane including an airfoil or any other suitable component with cooling channels (e.g., a blade outer air seal, a combustor panel). As shown, the pedestals  203  can be angled within the cooling channel  201  relative to an axial direction (axis A) of a turbomachine (e.g., circumferentially tangential). Referring to  FIGS. 2B and 2C , the pedestals can be non-angled or perpendicular relative to a radial direction (e.g., longitudinal direction of the cooling channel  201 ) of the turbomachine, but can be angled with a forward-slash sweep such that the pedestals  203  diverge relative to axis A. 
     As shown, the turbomachine component  200  can include a plurality of pedestals  203  where at least two pedestals  203  in the plurality of pedestals  203  have a different angle degree and/or angle direction relative to each other. For example, as shown in  FIG. 2A , the pedestals  203  can have differing angles (e.g., angles θ, γ, α, and β) relative to the axis A. This can be due to a non-linear cooling channel  201  formed by the shape of the airfoil cross-section. Also as shown in  FIG. 2C , the plurality of pedestals  203  can be disposed in any suitable predetermined pattern. For example, the predetermined pattern can include one or more rows or columns of pedestals  203 . 
     Referring to  FIGS. 3A-3C , a turbomachine component  300  similar to component  200  can include a body  302  and pedestals  303  that are angled within the cooling channel  301  in a direction opposite to that of the embodiment of  FIGS. 2A-2C , e.g., such that the pedestals are circumferentially tangentially angled to converge relative to the axis A. 
     In certain embodiments, as shown in  FIG. 4A-4C , a turbomachine component  400  can include a body  402  having pedestals  403  disposed in a cooling channel  401  that are angled relative to a radial direction of a turbomachine. As shown in  FIG. 4B , the pedestals  403  can be angled radially outward to radially inward. Referring to  FIGS. 5A-5C , a turbomachine component  500  similar to component  400  can include a body  502  and pedestals  503  that are angled within the cooling channel  501  in a radial direction opposite to that of the embodiment of  FIGS. 4A-4C . 
     In certain embodiments, as shown in  FIG. 6A-6C , a turbomachine component  600  can include a body  602  having pedestals  603  disposed in a cooling channel  601  that are angled within the cooling channel  601  relative to both an axial direction and a radial direction of a turbomachine. Referring to  FIGS. 7A-7C , a turbomachine component  700  similar to component  600  can include a body  702  and pedestals  703  that are angled within the cooling channel  701  in an axial and/or radial direction that is different to that of the embodiment of  FIGS. 6A-6C . 
     In at least one aspect of this disclosure, a method includes forming a turbomachine component (e.g., component  200 ) as described above. Forming can include additively manufacturing the turbomachine component  200 . Forming can include shaping or otherwise forming the body  203  into a turbomachine blade including an airfoil. Other methods of manufacturing include conventional casting methods including electric discharge machining (EDM) or Laser of the core structure used to produce the negative of the airfoil. 
     The angled pedestals of any of the embodiments described herein allow for flow directional control and/or modification to enhance thermal control of components with cooling channels. For example, the angled pedestals  203  as shown in  FIGS. 2A-2C  can guide a forward-to-aft flow upward toward an upper surface defining the cooling channel. As will be appreciated by one having ordinary skill in the art, the direction and degree of angle of the pedestals can be selected to modify impingement on a desired portion of the cooling channel to regulate temperatures at certain portions of the turbomachine component as desired. 
     The methods and systems of the present disclosure, as described above and shown in the drawings, provide for enhanced components with cooling channels having superior properties including enhanced thermal efficiency. While the apparatus and methods of the subject disclosure have been shown and described with reference to embodiments, those skilled in the art will readily appreciate that changes and/or modifications may be made thereto without departing from the spirit and scope of the subject disclosure.