Abstract:
Systems and methods are disclosed herein for distributing cooling air in gas turbine engines. A tangential on board injector (“TOBI”) may supply cooling air to a turbine section of a gas turbine engine. The cooling air may be split into a first cooling air path and a second cooling air path. The first cooling air path may fluidly connect the TOBI and the interior of a first stage rotor blade. The second cooling air path may fluidly connect the TOBI and a cavity. The cavity may be located between a first disk and a second disk. The cooling air paths from a single cooling air source may thermally isolate portions of the turbine section.

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application is a nonprovisional of, and claims priority to, and the benefit of U.S. Provisional Application No. 62/014,335, entitled “SYSTEMS AND METHODS FOR DISTRIBUTING COOLING AIR IN GAS TURBINE ENGINES,” filed on Jun. 19, 2014, which is hereby incorporated by reference in its entirety. 
     
    
     FIELD 
       [0002]    The present disclosure relates generally to gas turbine engines. More particularly, the present disclosure relates to cooling air systems in gas turbine engines. 
       BACKGROUND 
       [0003]    Gas turbine engines with multiple turbine stages include interstage seal arrangements between adjacent stages for improved operating efficiency. The interstage seal arrangements confine the flow of hot combustion core gases within an annular path around and between stationary turbine stator blades, nozzles and also around and between adjacent rotor blades. 
         [0004]    The interstage seal arrangements may also serve to confine and direct cooling air to cool the turbine disks, the turbine blade roots, and also the interior of the rotor blades themselves as rotor blade cooling facilities higher turbine inlet temperatures, which results in higher thermal efficiency of the engine and higher thrust output. Multiple sources of cooling air may be directed to different portions of the turbine. The multiple sources may result in thermal gradients on the turbine disks and other turbine components. 
       SUMMARY 
       [0005]    A turbine section for a gas turbine engine may comprise a tangential on board injector (“TOBI”), a first stage rotor blade in fluid communication with the TOBI, and a first cavity located between a first disk and a second disk. The first cavity may be in fluid communication with the TOBI. 
         [0006]    In various embodiments, the first disk may comprise a disk arm having an orifice. The TOBI and the first cavity may be fluidly connected via the orifice. A segmented seal may be in contact with the first disk and the second disk. The first cavity may be bounded by the first disk, the second disk, and the segmented seal. The second disk may comprise a disk arm having an orifice. A second cavity may be aft of the second disk, and the second cavity may be in fluid communication with the TOBI. The second cavity may be fluidly connected to the TOBI via the orifice in the disk arm of the second disk. The TOBI may be configured to provide a single source of cooling air to the first stage rotor blade, the first disk, and the second disk. The turbine section may further comprise a seal between the first disk and the second disk. The seal may comprise a first radial span, a second radial span, a first axial span that extends between the first radial span and the second radial span, and a second axial span that extends between the first radial span and the second radial span. The first radial span, the second radial span, the first axial span, and the second axial span may form a torque box. 
         [0007]    A gas turbine engine may comprise a first cooling air path and a second cooling air path. The first cooling air path may be defined by a cooling air supply source and an interior of a first stage rotor blade. The second cooling air path may be defined by the cooling air supply source, an orifice in a disk arm of a first disk, and a first cavity bounded by the first disk, a second disk, and a seal located between the first disk and the second disk. 
         [0008]    In various embodiments, the orifice and the first cavity may be connected via a channel defined by a bore of the first disk and a disk arm of the second disk. The cooling air supply source may comprise a TOBI configured to supply cooling air to the first cooling air path and the second cooling air path. The TOBI may be a single source of the cooling air in the first cooling air path and the second cooling air path. The seal may comprise a segmented seal in contact with the first disk and the second disk. 
         [0009]    A method of cooling a turbine section of a gas turbine engine may comprise supplying cooling air to the turbine section, directing a first portion of the cooling air into a rotor blade of a first stage rotor assembly, and directing a second portion of the cooling air into a cavity between a disk of the first stage rotor assembly and a disk of a second stage rotor assembly. 
         [0010]    In various embodiments, the first portion of the cooling air and the second portion of the cooling air are supplied by a TOBI. The method may further comprise thermally isolating the disk of the first stage rotor assembly. The second portion of the cooling air may be directed through an orifice in a disk arm of the first stage rotor assembly. The cavity may be defined by the disk of the first stage rotor assembly, the disk of the second stage rotor assembly, and a segmented seal between the disk of the first stage rotor assembly and the disk of the second stage rotor assembly. 
         [0011]    The foregoing features and elements may be combined in various combinations without exclusivity, unless expressly indicated otherwise. These features and elements as well as the operation thereof will become more apparent in light of the following description and the accompanying drawings. It should be understood, however, the following description and drawings are intended to be exemplary in nature and non-limiting. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0012]    The subject matter of the present disclosure is particularly pointed out and distinctly claimed in the concluding portion of the specification. A more complete understanding of the present disclosure, however, may best be obtained by referring to the detailed description and claims when considered in connection with the drawing figures. 
           [0013]      FIG. 1  illustrates a schematic cross-section view of a gas turbine engine in accordance with various embodiments; 
           [0014]      FIG. 2  illustrates a cross-section view of a turbine section of a gas turbine engine in accordance with various embodiments; 
           [0015]      FIG. 3  illustrates a perspective view of a seal in accordance with various embodiments; and 
           [0016]      FIG. 4  illustrates a flowchart of a process for cooling a turbine section in accordance with various embodiments. 
       
    
    
     DETAILED DESCRIPTION 
       [0017]    The detailed description of various embodiments herein makes reference to the accompanying drawings, which show various embodiments by way of illustration. While these various embodiments are described in sufficient detail to enable those skilled in the art to practice the disclosure, it should be understood that other embodiments may be realized and that logical, chemical, and mechanical changes may be made without departing from the spirit and scope of the disclosure. Thus, the detailed description herein is presented for purposes of illustration only and not of limitation. For example, the steps recited in any of the method or process descriptions may be executed in any order and are not necessarily limited to the order presented. Furthermore, any reference to singular includes plural embodiments, and any reference to more than one component or step may include a singular embodiment or step. Also, any reference to attached, fixed, connected, or the like may include permanent, removable, temporary, partial, full, and/or any other possible attachment option. Additionally, any reference to without contact (or similar phrases) may also include reduced contact or minimal contact. 
         [0018]    Referring to  FIG. 1 , a gas turbine engine  100  (such as a turbofan gas turbine engine) is illustrated according to various embodiments. Gas turbine engine  100  is disposed about axial centerline axis  120 , which may also be referred to as axis of rotation  120 . Gas turbine engine  100  may comprise a fan  140 , compressor sections  150  and  160 , a combustion section  180 , and turbine sections  190 ,  191 . Air compressed in the compressor sections  150 ,  160  may be mixed with fuel and burned in combustion section  180  and expanded across turbine sections  190 ,  191 . Turbine sections  190 ,  191  may include high pressure rotors  192  and low pressure rotors  194 , which rotate in response to the expansion. Turbine sections  190 ,  191  may comprise alternating rows of rotary airfoils or blades  196  and static airfoils or vanes  198 . A plurality of bearings  115  may support spools in the gas turbine engine  100 .  FIG. 1  provides a general understanding of the sections in a gas turbine engine, and is not intended to limit the disclosure. The present disclosure may extend to all types of turbine engines, including turbofan gas turbine engines and turbojet engines, for all types of applications. 
         [0019]    The forward-aft positions of gas turbine engine  100  lie along axis of rotation  120 . For example, fan  140  may be referred to as forward of turbine section  190  and turbine section  190  may be referred to as aft of fan  140 . Typically, during operation of gas turbine engine  100 , air flows from forward to aft, for example, from fan  140  to turbine section  190 . As air flows from fan  140  to the more aft components of gas turbine engine  100 , axis of rotation  120  may also generally define the direction of the air stream flow. 
         [0020]    Referring to  FIG. 2 , a cross-section view of a high pressure turbine section (“HPT”)  254  of a gas turbine engine is illustrated according to various embodiments. The HPT  254  includes a two-stage turbine section with a first stage rotor assembly  260  and a second stage rotor assembly  262 , both of which are affixed to the outer shaft  250 . 
         [0021]    The rotor assemblies  260 ,  262  include a respective array of blades  264 ,  266  circumferentially disposed around a disk  268 ,  270 . Each blade  264 ,  266  includes a respective root  272 ,  274 , a platform  276 ,  278  and an airfoil  280 ,  282 . Each disk  268 ,  270  may comprise a bore  210 ,  212 , a neck  206 ,  208 , a rim  284 ,  286 , and a disk arm  230 ,  232 . Each blade root  272 ,  274  is received within a respective rim  284 ,  286  of the disk  268 ,  270 , and the airfoils  280 ,  282  extend radially outward toward a blade outer air seal (BOAS) assembly  281 ,  283 . 
         [0022]    The blades  264 ,  266  are disposed in the core flow path that is pressurized in the compressor sections, then heated to a working temperature in the combustion section  180 . The platforms  276 ,  278  separate a gas path side inclusive of the airfoil  280 ,  282  and a non-gas path side inclusive of the root  272 ,  274 . 
         [0023]    A shroud assembly  288  within the engine case structure between the blade stages directs the hot gas core airflow in the core flow path from the first stage blades  264  to the second stage blades  266 . The shroud assembly  288  may at least partially support the BOAS assemblies  281 ,  283  and includes an array of vanes  290  that extend between a respective inner vane platform  292  and an outer vane platform  294 . The outer vane platforms  294  may be supported by the engine case structure and the inner vane platforms  292  support an abradable annular seal  296  to seal the hot gas core airflow in the axial direction with respect to a segmented interstage seal assembly  300 . 
         [0024]    The segmented interstage seal assembly  300  includes a plurality of individual seal segments  302  (with reference also to  FIG. 3 ) disposed between the first and second rotor assemblies  260 ,  262  for sealing between the axially flowing hot gas core airflow and a radially inner first cavity C 1  between a respective neck  206 ,  208  and bore  210 ,  212  of the disks  268 ,  270 . The multiple seal segments  302  thereby eliminate hoop stress in the segmented interstage seal assembly  300 . Each seal segment  302  may be cast of a material such as Inconel®  625  (an austenitic nickel-chromium-based superalloy) or any other suitable material to provide increased knife edge temperature capability. Such materials facilitate reduced transient load variation into the rim  284 ,  286  by minimization, if not elimination, of the thermally induced growth relative to the disks  268 ,  270 . 
         [0025]    Cooling air may be directed to the turbine section via a tangential onboard injector (“TOBI”)  220 . The TOBI  220  may direct cooling air from the high pressure compressor to the HPT  254 . In various embodiments, the cooling air from the TOBI  220  may be injected tangentially, such that the cooling air contains a circumferential velocity. Thus, the direction of the cooling air exiting the TOBI  220  may correspond to the direction of rotation of the first stage HPT blades  264 . 
         [0026]    In various embodiments, the cooling air from the TOBI  220  may be divided into a blade path A and a cavity path B. An annular seal  248  may divide the cooling air into the blade path A and the cavity path B. The cooling air following the blade path A may enter the first blade  264  and flow into an interior of the airfoil  280  in order to cool the airfoil  280 . The cooling air may then exit the blade  264  through film holes in the airfoil  280 . 
         [0027]    The cooling air following the cavity path B may be directed inward along neck  206  toward the outer shaft  250 . A disk arm  230  of the first disk may comprise an orifice  231 . The cooling air following the cavity path B may flow through the orifice  231 , through a channel  234  between bore  210  and a disk arm  232  of the second disk  270 , and into a first cavity C 1 , defined as the region bounded by the first disk  268 , the second disk  270 , and the segmented seal assembly  300 . In some cases, first cavity C 1  may be referred to as the “1-2 cavity.” 
         [0028]    A disk arm  232  of the second disk  270  may comprise an orifice  233 . A portion of the cooling air flowing through the orifice  231  in the first disk may flow through the orifice  233  in the disk arm  232  of the second disk  270 . The air flowing though the orifice  233  in the disk arm  232  of the second disk  270  may flow through a channel  236  between bore  212  and outer shaft  250 , and into a second cavity C 2  located aft of the second disk  270 . The second cavity C 2  may generally be defined by second disk  270  and an aft wall  256  of HPT  254 . In various embodiments, any suitable component may be in the location of aft wall  256 , such as a bearing compartment. 
         [0029]    The configuration of the HPT  254  thus fluidly connects the blade  264  of the first stage rotor assembly  260 , the first cavity C 1 , and the second cavity C 2  to a single source of cooling air. Thus, the blade  264 , the first cavity C 1 , and the second cavity C 2  bay be in fluid communication with one another. In various embodiments, the single source of cooling air may be the TOBI  220 . Cooling the blade  264  of the first stage rotor assembly  260 , the first cavity C 1 , and the second cavity C 2  from a single source of cooling air may decrease thermal gradients on the disks  268 ,  270 , which may decrease low cycle fatigue and increase the product lifetime of the disks  268 ,  270 . Furthermore, the segmented seal assembly  300  may limit the amount of hot gas entering the first cavity C 1 , which further thermally isolates the disks  268 ,  270 . 
         [0030]    Each seal segment  302  is radially supported on a respective pilot diameter  214 ,  216  formed by the respective rim  284 ,  286  of the disk  268 ,  270 . At least one of the individual seal segments  302  includes an anti-rotation tab  218  that interfaces with a stop  222  on the rim  284  of the disk  268 . It should be appreciated that various interfaces may be alternatively or additionally provided on one, or multiple, seal segments  302 . 
         [0031]    Referring to  FIGS. 2 and 3 , each seal segment  302  generally includes a first radial span  330 , a second radial span  332 , a first axial span  334 , a second axial span  336 , and a seal span  338  with a plurality of knife edges  340  that engage the abradable annular seal  296 . The first axial span  334 , the second axial span  336 , and the seal span  338  extend generally axially between the first radial span  330  and the second radial span  332  which extend generally radially with respect to the engine axis  120 . It should be appreciated that the first axial span  334 , the second axial span  336 , and the seal span  338  may include a generally arcuate configuration to facilitate resistance to interstage loads from the first and second rotor assemblies  260 ,  262 . 
         [0032]    The first radial span  330  and the second radial span  332  include a respective radial support  342 ,  344  that engage the respective pilot diameter  214 ,  216 . The first radial span  330  and the second radial span  332  also include a respective axial supports  346 ,  348  that maintain axial spacing of the segmented interstage seal assembly  300 . The axial supports  346 ,  348  include a circumferential groove  350 ,  352  to support wire seals  354  that seal with the associated disks  268 ,  270 . The first radial span  330  and the second radial span  332  may further comprise respective pressure seals  312 ,  314  extending from the junctions between the radial spans  330 ,  332  and the seal span  338 . The pressure seals  312 ,  314  may contact disks  268 ,  270 . The pressure seals  312 ,  314 , axial supports  346 ,  348 , and radial supports  342 ,  344  may limit the transfer of hot gas from the core gas path into the first cavity C 1 . 
         [0033]    The seal span  338  and the second radial span  332  include a multiple of apertures  362 ,  364  that receive and direct a cooling flow from a passage  366  through the vanes  290  into the second stage blade  266 . The plurality of apertures  364  in the second radial span  332  may thereby operate similar to a TOBI. 
         [0034]    The first radial span  330 , the second radial span  332 , the first axial span  334 , and the second axial span  336  may form a torque box  370  to minimize bending and resist a crushing load between the first and second stage rotor assemblies  260 ,  262 . In other words, shear loads may be transferred to all the sides to improve structural efficiency. The torque box  370  stiffens each seal segment  302  against axial deflection and maintains the disks  268 ,  270  at a uniform axial displacement. In various embodiments, any number of internal beams with apertures may be located between the axial supports  346 ,  348  to further strengthen the torque box  370 . 
         [0035]    The seal segments  302 , in combination with the cooling air flow path configurations described herein, may thus provide thermally isolated turbine disks. Although described primarily with respect to a two-stage HPT, the present disclosure may be consistent with HPTs with any number of stages, including a single stage turbine. 
         [0036]    Referring to  FIG. 4 , a flowchart of a process  400  for cooling a turbine section of a gas turbine engine is illustrated according to various embodiments. Cooling air may be supplied to the turbine section (step  410 ). In various embodiments, the cooling air may be supplied by a TOBI. A first portion of the cooling air may be directed into a rotor blade of a first stage rotor assembly (step  420 ). In various embodiments, an annular seal upstream of the first stage rotor assembly may divide the cooling air into the first portion and a second portion. The second portion of the cooling air may be directed into a cavity between a disk of the first stage rotor assembly and a disk of a second stage rotor assembly (step  430 ). In various embodiments, the TOBI may be a single source of cooling air for the first portion of the cooling air and the second portion of the cooling air. The second portion of the cooling air may be directed through an orifice in a disk arm of the first stage rotor assembly. The first portion of the cooling air and the second portion of the cooling air may thermally isolate the disk of the first stage rotor assembly. 
         [0037]    Benefits, other advantages, and solutions to problems have been described herein with regard to specific embodiments. Furthermore, the connecting lines shown in the various figures contained herein are intended to represent exemplary functional relationships and/or physical couplings between the various elements. It should be noted that many alternative or additional functional relationships or physical connections may be present in a practical system. However, the benefits, advantages, solutions to problems, and any elements that may cause any benefit, advantage, or solution to occur or become more pronounced are not to be construed as critical, required, or essential features or elements of the disclosure. The scope of the disclosure is accordingly to be limited by nothing other than the appended claims, in which reference to an element in the singular is not intended to mean “one and only one” unless explicitly so stated, but rather “one or more.” Moreover, where a phrase similar to “at least one of A, B, or C” is used in the claims, it is intended that the phrase be interpreted to mean that A alone may be present in an embodiment, B alone may be present in an embodiment, C alone may be present in an embodiment, or that any combination of the elements A, B and C may be present in a single embodiment; for example, A and B, A and C, B and C, or A and B and C. Different cross-hatching is used throughout the figures to denote different parts but not necessarily to denote the same or different materials. 
         [0038]    Systems, methods and apparatus are provided herein. In the detailed description herein, references to “one embodiment”, “an embodiment”, “various embodiments”, etc., indicate that the embodiment described may include a particular feature, structure, or characteristic, but every embodiment may not necessarily include the particular feature, structure, or characteristic. Moreover, such phrases are not necessarily referring to the same embodiment. Further, when a particular feature, structure, or characteristic is described in connection with an embodiment, it is submitted that it is within the knowledge of one skilled in the art to affect such feature, structure, or characteristic in connection with other embodiments whether or not explicitly described. After reading the description, it will be apparent to one skilled in the relevant art(s) how to implement the disclosure in alternative embodiments. 
         [0039]    Furthermore, no element, component, or method step in the present disclosure is intended to be dedicated to the public regardless of whether the element, component, or method step is explicitly recited in the claims. No claim element herein is to be construed under the provisions of 35 U.S.C. 112(f) unless the element is expressly recited using the phrase “means for.” As used herein, the terms “comprises”, “comprising”, or any other variation thereof, are intended to cover a non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements does not include only those elements but may include other elements not expressly listed or inherent to such process, method, article, or apparatus.