Patent Publication Number: US-10787931-B2

Title: Method and structure of interdigitated turbine engine thermal management

Description:
FIELD 
     The present subject matter relates generally to gas turbine engine architecture. More particularly, the present subject matter relates to a method of thermal management and bearing arrangement for gas turbine engine turbine sections. 
     BACKGROUND 
     Gas turbine engines generally include a turbine section downstream of a combustion section that is rotatable with a compressor section to rotate and operate the gas turbine engine to generate power, such as propulsive thrust. General gas turbine engine design criteria often include conflicting criteria that must be balanced or compromised, including increasing fuel efficiency, operational efficiency, and/or power output while maintaining or reducing weight, part count, and/or packaging (i.e. axial and/or radial dimensions of the engine). 
     Conventional gas turbine engines generally include turbine sections defining a high pressure turbine in serial flow arrangement with an intermediate pressure turbine and/or low pressure turbine. The high pressure turbine includes an inlet or nozzle guide vane between the combustion section and the high pressure turbine rotor. Conventionally, combustion gases exiting the combustion section define a relatively low velocity compared to a velocity (e.g., along a circumferential or tangential direction) of the first rotating stage of the turbine, generally defined as the high pressure turbine rotor. Thus, conventionally, the nozzle guide vane serves to accelerate a flow of combustion gases exiting the combustion section to more closely match or exceed the high pressure turbine rotor speed along a tangential or circumferential direction. Such acceleration of flow using a nozzle guide vane to match or exceed high pressure turbine rotor speed is known to improve general engine operability and performance. 
     Thereafter, conventional gas turbine engine turbine sections generally include successive rows or stages of stationary and rotating airfoils, or vanes and blades, respectively. This conventional configuration generally conditions a flow of the combustion gases entering and exiting each stage of vanes and blades. However, conventional turbine sections, and especially stationary airfoils (i.e. vanes and nozzle guide vanes) require considerable quantities and masses of cooling air to mitigate damage due to hot combustion gases. For example, generally, nozzle guide vanes are designed to withstand a maximum combustion gas temperature along an annulus (i.e. hot spots), which may be significantly larger than an average combustion gas temperature along the annulus. Thus, conventional engines are designed to use significant quantities or masses of cooling air from a compressor section or unburned air from the combustion section to mitigate structural damage, wear, deterioration, and ultimately, maintenance and repair, of the nozzle guide vanes. However, this cooling air adversely affects overall engine efficiency, performance, fuel consumption, and/or operability by removing energy that could otherwise be used in combustion to drive the turbines, compressors, and fan. Still further, the nozzle guide vane is often a limiting component when determining maintenance and repair intervals for gas turbine engines, thereby limiting overall engine performance and efficiency. 
     A known solution to improve turbine section efficiency is to interdigitate the rotors of the turbine section (i.e. successive rows or stages of rotating airfoils or blades). For example, a known solution is to configure a turbine section, in serial flow arrangement from an upstream end to a downstream end along a longitudinal direction, with a nozzle guide vane, a high pressure turbine rotor, another turbine vane stage (i.e. stationary airfoils), and an intermediate pressure turbine interdigitated with a low pressure turbine. Another known solution is to configure a turbine section, in serial flow arrangement, with a nozzle guide vane, a high pressure turbine rotor, and various levels of interdigitated rotors thereafter, including low, intermediate, or high pressure turbine rotors. 
     However, despite various known solutions, there exists a need structures that enable further interdigitation of the turbine section toward the combustion section. Still further, there exists a need for methods of arranging and operating turbine section bearings that enable efficient use of compressed air for bearing operation and engine cooling. 
     BRIEF DESCRIPTION 
     Aspects and advantages of the invention will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the invention. 
     The present disclosure is directed to a method of turbine section thermal management for a gas turbine engine. The engine includes a first turbine bearing defining an outer air bearing disposed radially adjacent to a low speed turbine rotor hub of a low speed turbine rotor and an inner bearing disposed radially adjacent to a high pressure (HP) shaft coupled to a high speed turbine rotor, wherein a first manifold is in fluid communication from a pressure plenum of a combustion section to the first turbine bearing, and wherein a second manifold is in fluid communication from the first turbine bearing to a pressure regulating valve and an outer diameter secondary flowpath of the turbine section, and wherein a third manifold is in fluid communication from the pressure plenum of the combustion section to the pressure regulating valve. The method includes flowing air from the pressure plenum to the first turbine bearing; flowing air between the outer air bearing and the low speed turbine rotor hub; flowing air between the inner air bearing and the HP shaft; flowing air from the first turbine bearing to the pressure regulating valve via the second manifold; and flowing air from the first turbine bearing to the outer diameter secondary flowpath via the second manifold. 
     In one embodiment, flowing air to the outer diameter secondary flowpath includes operating the engine at a cruise power condition or higher. 
     In another embodiment, the method further includes flowing air from the outer diameter secondary flowpath to the core flowpath of a turbine section of the engine. 
     In various embodiments, the method further includes flowing air from the pressure regulating valve to a lower pressure sink. In one embodiment, flowing air from the pressure regulating valve includes flowing air to the core flowpath of the turbine section or to an approximately atmospheric pressure condition. In another embodiment, flowing air from the pressure regulating valve to a lower pressure sink includes operating the engine below a cruise power condition. 
     In one embodiment, flowing air through the outer diameter secondary flowpath includes opening a one way valve. 
     In another embodiment, opening a one way valve includes opening the one way valve at approximately a cruise power condition or higher of the engine. 
     In yet another embodiment, flowing air to the core flowpath of the turbine section includes flowing air proximate to the outer shroud and the outer shroud airfoils of the low speed turbine rotor. 
     In various embodiments, the method further includes providing a first supply of air from between the outer air bearing and the low speed turbine rotor hub; providing a second supply of air from between the inner air bearing and the HP shaft; and flowing the first supply of air and the second supply of air through the second manifold to the pressure regulating valve. In one embodiment, the method further includes monitoring a difference in air pressure from the second manifold and the third manifold. In another embodiment, monitoring a difference in air pressure from the second manifold and the third manifold includes monitoring a difference in pressure from the pressure plenum and pressure from the first supply of air and the second supply of air from the first turbine bearing. In yet another embodiment, flowing the first supply of air and the second supply of air through the second manifold includes collecting the first supply of air from an annular cavity of the outer air bearing and collecting the second supply of air from a first cavity defined between a compressor seal interface and the first turbine bearing and egressing the collected air through the second manifold to the pressure valve. 
     The present disclosure is further directed to a gas turbine engine defining a longitudinal direction, a radial direction extended from an axial centerline, and a circumferential direction. The gas turbine engine includes a compressor section, a combustion section, and a turbine section in serial flow arrangement along the longitudinal direction. The gas turbine engine includes a low speed turbine rotor comprising a hub extended along the longitudinal direction and radially within the combustion section; a high speed turbine rotor comprising a high pressure (HP) shaft coupling the high speed turbine rotor to a HP compressor in the compressor section, in which the HP shaft extends along the longitudinal direction and radially within the hub of the low speed turbine rotor; and a first turbine bearing disposed radially between the hub of the low speed turbine rotor and the HP shaft. The first turbine bearing defines an outer air bearing along an outer diameter of the first turbine bearing and adjacent to the hub of the low speed turbine rotor, and the first turbine bearing defines an inner air bearing along an inner diameter of the first turbine bearing and adjacent to the HP shaft. The gas turbine engine further includes a pressure regulating valve, a one way valve, and a second manifold extended from the first turbine bearing in fluid communication with the pressure regulating valve and the one way valve. The pressure regulating valve and the one way valve are in parallel arrangement relative to the second manifold. 
     In various embodiments, the turbine section further includes a turbine casing extended generally from the combustion section along the longitudinal direction and circumferentially encompassing the turbine section, in which the turbine section further defines an outer diameter (OD) secondary flowpath disposed generally between the outer shroud of the low speed turbine rotor and the turbine casing. In one embodiment, the engine further defines a turbine cooling supply manifold extended from the second manifold in fluid communication with the OD secondary flowpath. In another embodiment, the turbine cooling supply manifold defines the one way valve disposed along the turbine cooling supply manifold. 
     In various embodiments, the engine further includes a third manifold extended from the pressure plenum to the pressure regulating valve. The pressure plenum and the pressure regulating valve are in fluid communication via the third manifold. In one embodiment, the third manifold defines a reference pressure and/or motive pressure at the pressure regulating valve. 
     In another embodiment, the low speed turbine rotor and the high speed turbine rotor are together in alternating interdigitated arrangement along the longitudinal direction. 
     These and other features, aspects and advantages of the present invention will become better understood with reference to the following description and appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and, together with the description, serve to explain the principles of the invention. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       A full and enabling disclosure of the present invention, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended figures, in which: 
         FIG. 1  is a schematic cross sectional view of an exemplary gas turbine engine incorporating an exemplary embodiment of a turbine section according to an aspect of the present disclosure; 
         FIG. 2  is a schematic cross sectional view of an embodiment of a combustion section and turbine section of the engine shown in  FIG. 1  including an embodiment of a first turbine bearing supporting the turbine section; 
         FIG. 3  is a schematic cross sectional view of an embodiment of the engine including a first turbine bearing supporting the turbine section shown in  FIGS. 1-2 ; 
         FIG. 4  is a schematic cross sectional view of another embodiment of the engine including a first turbine bearing supporting the turbine section shown in  FIGS. 1-2 ; 
         FIG. 5  is a schematic cross sectional view of yet another embodiment of the engine including a first turbine bearing supporting the turbine section shown in  FIGS. 1-2 ; 
         FIG. 6  is a schematic cross sectional view of still another embodiment of the engine including a first turbine bearing supporting the turbine section shown in  FIGS. 1-2 ; 
         FIG. 7  is a schematic flowchart depicting flows through the engine including a first turbine bearing according to various embodiments shown in  FIGS. 1-7 ; 
         FIG. 8  is an exemplary flowchart outlining a method of operating the gas turbine engine shown and described in regard to  FIGS. 1-7 ; 
         FIG. 9  is an exemplary flowchart outlining a method of turbine section thermal management for the gas turbine engine shown and described in regard to  FIGS. 1-7 ; and 
         FIG. 10  is an exemplary flowchart outlining another method of turbine section thermal management for the gas turbine engine shown and described in regard to  FIGS. 1-7 . 
     
    
    
     Repeat use of reference characters in the present specification and drawings is intended to represent the same or analogous features or elements of the present invention. 
     DETAILED DESCRIPTION 
     Reference now will be made in detail to embodiments of the invention, one or more examples of which are illustrated in the drawings. Each example is provided by way of explanation of the invention, not limitation of the invention. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made in the present invention without departing from the scope or spirit of the invention. For instance, features illustrated or described as part of one embodiment can be used with another embodiment to yield a still further embodiment. Thus, it is intended that the present invention covers such modifications and variations as come within the scope of the appended claims and their equivalents. 
     As used herein, the terms “first”, “second”, and “third” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components. 
     The terms “upstream” and “downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction to which the fluid flows. 
     The terms “low”, “intermediate”, “high”, or their respective comparative degrees (e.g. −er, where applicable) each refer to relative speeds within an engine unless otherwise specified. For example, a “low turbine” or “low speed turbine” defines a rotational speed lower than a “high turbine” or “high speed turbine”. Alternatively, unless otherwise specified, the aforementioned terms may be understood in their superlative degree. For example, a “low turbine” may refer to the lowest rotational speed turbine within a turbine section, and a “high turbine” may refer to the highest rotational speed turbine within the turbine section. 
     Exemplary embodiments of a first turbine bearing for an interdigitated turbine section are generally provided as well as methods of operation and thermal management. The first turbine bearing defines an outer air bearing disposed adjacent to a low speed turbine rotor and an inner air bearing disposed adjacent to a high speed turbine rotor. The low speed turbine rotor generally defines a low speed turbine rotor interdigitated forward of and among rows of the high speed turbine rotor defining a high speed turbine rotor. In various embodiments, the first turbine bearing, and methods of operation thereof, include providing cooling to the high speed turbine rotor and/or an outer diameter (OD) secondary flowpath of the turbine section. 
     The embodiments of the gas turbine engine including the first turbine bearing, and methods of operation thereof, may enable efficient use of cooling air for operation of the first turbine bearing and turbine section, thereby improving overall gas turbine efficiency and/or reducing fuel consumption. The embodiments of the gas turbine engine with interdigitated turbine section shown and described herein may further enable additional stages of interdigitation of the low speed turbine rotor and the high speed turbine rotor, up to and including one or more stages of the low speed turbine rotor forward or upstream of the high speed turbine rotor. In various embodiments, the gas turbine engine with interdigitated turbine section may further reduce cooling air consumption, increase engine efficiency, performance, and/or operability, and/or reduce part quantities, weight, and/or packaging (i.e. axial and/or radial dimensions). Still further, the interdigitated turbine section may reduce a product of axial flow area and the square of the rotational speed (the product referred to as “AN 2 ”) while additionally reducing an average work factor per stage of the turbine section. 
     Referring now to the drawings,  FIG. 1  is a schematic cross sectional view of an exemplary gas turbine engine  10  (herein referred to as “engine  10 ”), shown as a high bypass turbofan engine, incorporating an exemplary embodiment of a turbine section  90  according to an aspect of the present disclosure. Although further described below with reference to a turbofan engine, the present disclosure is also applicable to turbomachinery in general, including propfan, turbojet, turboprop, and turboshaft gas turbine engines, including marine and industrial turbine engines and auxiliary power units. Still further, although described below as a three-spool gas turbine engine, the present disclosure is also applicable to two-spool gas turbine engines. As shown in  FIG. 1 , the engine  10  has a longitudinal or axial centerline axis  12  that extends there through for reference purposes. The engine  10  defines a longitudinal direction L, a radial direction R, and an upstream end  99  and a downstream end  98  along the longitudinal direction L. 
     In general, the engine  10  may include a substantially tubular outer casing  18  that defines an annular inlet  20 . The outer casing  18  encases or at least partially flows, in serial flow arrangement, a compressor section  21 , a combustion section  26 , and an interdigitated turbine section  90  (herein referred to as “turbine section  90 ”). Generally, the engine  10  defines, in serial flow arrangement from the upstream end  99  to the downstream end  98 , the fan assembly  14 , the compressor section  21 , the combustion section  26 , and the turbine section  90 . In the embodiment shown in  FIG. 1 , the compressor section  21  defines a high pressure (HP) compressor  24  and an intermediate pressure (IP) compressor  22 . In other embodiments, the fan assembly  14  may further include or define one or more stages of a plurality of fan blades  42  that are coupled to and extend outwardly in the radial direction R from a fan rotor  15  and/or a low speed shaft  36 . In various embodiments, multiple stages of the plurality of fan blades  42  coupled to the low speed shaft  36  may be referred to as a low pressure (LP) compressor. 
     An annular fan casing or nacelle  44  circumferentially surrounds at least a portion of the fan assembly  14  and/or at least a portion of the outer casing  18 . In one embodiment, the nacelle  44  may be supported relative to the outer casing  18  by a plurality of circumferentially-spaced outlet guide vanes or struts  46 . At least a portion of the nacelle  44  may extend over an outer portion (in radial direction R) of the outer casing  18  so as to define a bypass airflow passage  48  therebetween. 
     Referring now to  FIG. 2 , an exemplary embodiment of the turbine section  90  of the engine  10  is generally provided. The turbine section  90  includes a low speed turbine rotor  110  extended along the longitudinal direction L. The low speed turbine rotor  110  includes an inner shroud  112 , an outer shroud  114 , and at least one connecting airfoil  116  coupling the inner shroud  112  to the outer shroud  114 . The outer shroud  114  includes a plurality of outer shroud airfoils  118  extended inward along the radial direction R. In various embodiments, the inner shroud  112  may include a plurality of inner shroud airfoils  119  extended outward along the radial direction R. 
     The inner shroud  112  and the outer shroud  114  each extend generally along the longitudinal direction L. The inner shroud  112  and/or the outer shroud  114  may each extend at least partially in the radial direction R. In various embodiments, the inner shroud  112  extends from the connecting airfoil  116 . In one embodiment, the inner shroud  112  further extends toward the downstream end  98  along the longitudinal direction L. In still various embodiments, the outer shroud  114  extends from the connecting airfoil  116  toward the upstream end  99  along the longitudinal direction L toward the combustion section  26 . 
     As shown on  FIG. 4 , a turbine casing  155  extends generally extended from the combustion section  26  along the longitudinal direction L and circumferentially encompassing the turbine section  90 . The turbine casing  155  may define an outer diameter (OD) secondary flowpath  71  radially between the outer shroud  114  of the low speed turbine rotor  110  and the turbine casing  155 . The turbine casing  155  may include a plurality of orifices  156  through which a fluid may ingress or egress to the secondary flowpath  71  and/or core flowpath  70 . 
     Referring now to  FIG. 2 , the turbine section  90  further includes a high speed turbine rotor  120  and an intermediate speed turbine rotor  130  each disposed forward or upstream  99  of the one or more connecting airfoils  116  of the low speed turbine rotor  110 . The high speed turbine rotor  120  includes a plurality of high speed turbine airfoils  122  extended outward along the radial direction R. The intermediate speed turbine rotor  130  includes a plurality of intermediate speed turbine airfoils  132  extended outward along the radial direction R. The pluralities of high speed turbine airfoils  122  and intermediate speed turbine airfoils  132  are each disposed among the pluralities of outer shroud airfoils  118  of the low speed turbine rotor  110  along the longitudinal direction L. 
     In various embodiments, the low speed turbine rotor  110  defines a plurality of stages of rotating airfoils, such as the plurality of outer shroud airfoils  118  disposed along the longitudinal direction L, the one or more connecting airfoils  116 , and/or the plurality of inner shroud airfoils  119  disposed along the longitudinal direction L. In one embodiment, the low speed turbine rotor  110  defines at least one stage forward or upstream  99  of the high speed turbine rotor  120 . In another embodiment, the turbine section  90  defines a first stage of airfoils in which the first stage includes the plurality of outer shroud airfoils  118  of the low speed turbine rotor  110  forward or upstream  99  of each stage of the high speed turbine rotor  120 . 
     In still various embodiments, such as shown in  FIG. 2 , the engine  10  defines, in serial flow arrangement along the longitudinal direction L from the upstream end  99  to the downstream end  98 , the plurality of outer shroud airfoils  118  of the low speed turbine rotor  110 , the plurality of high speed turbine airfoils  122  of the high speed turbine rotor  120 , and the plurality of outer shroud airfoils  118  of the low speed turbine rotor  110 . In still various embodiments, additional iterations of interdigitation of the low speed turbine rotor  110  and the high speed turbine rotor  120  may be defined forward or upstream  99  of the connecting airfoils  116 . 
     In yet various embodiments, such as shown in  FIG. 2 , the engine  10  further defines two or more stages of the low speed turbine rotor  110  and one or more stages of the intermediate speed turbine rotor  130  in interdigitation along the longitudinal direction L downstream of the high speed turbine rotor  120 . For example, the engine  10  may define, in serial flow arrangement along the longitudinal direction L from the upstream end  99  to the downstream end  98 , the combustion section  26 , the first stage  101  of the low speed turbine rotor  110 , the high speed turbine rotor  120 , the low speed turbine rotor  110 , the intermediate speed turbine rotor  130 , and the low speed turbine rotor  110 . In one embodiment, the engine  10  may further define additional stages of interdigitation of the high speed turbine rotor  120  and/or the intermediate speed turbine rotor  130  in interdigitation (i.e. alternating arrangement along the longitudinal direction L) with the low speed turbine rotor  110 . 
     As another non-limiting example, as shown in  FIG. 2 , the engine  10  may further define the serial flow arrangement of the plurality of outer shroud airfoils  118 , the plurality of high speed turbine airfoils  122 , the plurality of outer shroud airfoils  118 , the plurality of intermediate speed turbine airfoils  132 , the plurality of outer shroud airfoils  118 , an additional plurality of intermediate speed turbine airfoils  132 , and the connecting airfoils  116 . It should be appreciated that although  FIG. 2  shows the high speed turbine rotor  120  as defining one stage, the high speed turbine rotor  120  may define generally one or more stages between the first stage  101  of the low speed turbine rotor  110  and the connecting airfoils  116  of the low speed turbine rotor  110 , and interdigitated therebetween along the longitudinal direction L. Similarly, it should be appreciated that although  FIG. 1  shows the intermediate speed turbine rotor  130  as defining two stages, the intermediate speed turbine rotor  130  may define generally one or more stages between the high speed turbine rotor  120  and the connecting airfoils  116  of the low speed turbine rotor  110 . 
     Referring now to  FIGS. 1-7 , in various embodiments, the low speed turbine rotor  110  is drivingly connected and rotatable with a low speed shaft  36  extended along the longitudinal direction L and generally concentric about the axial centerline  12 . In one embodiment, as shown in  FIG. 1 , the low speed shaft  36  is connected to the fan assembly  14 , of which is driven in rotation by the low speed turbine rotor  110  of the turbine section  90 . The low speed shaft  36  is connected to the fan rotor  15  of the fan assembly  14 . In various embodiments, the fan assembly  14  may define a plurality of stages of the plurality of fan blades  42 , of which further define the LPC. 
     In various embodiments, the intermediate speed turbine rotor  130  is drivingly connected and rotatable with an intermediate speed shaft  35  extended along the longitudinal direction L and generally concentric about the axial centerline  12 . In one embodiment, as shown in  FIG. 1 , the intermediate speed shaft  35  is connected to the IP compressor  22 , of which is driven in rotation by the intermediate speed turbine rotor  130  of the turbine section  90 . 
     Referring still to  FIGS. 1-7 , the high speed turbine rotor  120  of the turbine section  90  is drivingly connected and rotatable with a HP shaft  34  extended along the longitudinal direction L and generally concentric about the axial centerline  12 . The HP shaft  34  is connected to the HP compressor  24 , of which is driven in rotation by the high speed turbine rotor  120  of the turbine section  90 . 
     Referring to  FIG. 2 , the turbine section  90  further includes one or more turbine vanes  150 . The turbine vane  150  may define a plurality of stationary airfoils (i.e. vanes) in circumferential arrangement. In one embodiment, the turbine vane  150  is disposed between the pluralities of inner shroud airfoils  119  along the longitudinal direction L. In various embodiments, the turbine vane  150  is disposed downstream  98  of the connecting airfoil  116  of the low speed turbine rotor  110 . The turbine vane  150 , or pluralities thereof, interdigitated among the pluralities of inner shroud airfoils  119  may enable further conditioning of the combustion gases  86  and work or energy extraction from the low speed turbine rotor  110  via the plurality of inner shroud airfoils  119 . 
     During operation of the engine  10  as shown collectively in  FIGS. 1-7 , the high speed turbine rotor  120  rotates generally at a higher rotational speed than the intermediate speed turbine rotor  130 . The intermediate speed turbine rotor  130  rotates generally at a higher speed than the low speed turbine rotor  110 . During operation of the engine  10 , a volume of air as indicated schematically by arrows  74  enters the engine  10  through an associated inlet  76  of the nacelle and/or fan assembly  14 . As the air  74  passes across the fan blades  42 , a portion of the air as indicated schematically by arrows  78  is directed or routed into the bypass airflow passage  48  while another portion of the air as indicated schematically by arrows  80  is directed or through the fan assembly  14 . Air  80  is progressively compressed as it flows through the compressor section  21  toward the combustion section  26 . 
     The now compressed air, as indicated schematically by arrows  82 , flows into the combustion section  26  where a fuel  91  is introduced, mixed with at least a portion of the compressed air  82 , and ignited to form combustion gases  86 . The combustion gases  86  flow into the turbine section  90 , causing rotary members of the turbine section  90  to rotate and support operation of respectively coupled rotary members in the compressor section  21  and/or fan assembly  14 . 
     In various embodiments, the low speed turbine rotor  110 , and the low speed shaft  36  to which it is attached, rotates in a first direction along the circumferential direction. The high speed turbine rotor  120 , and the HP shaft  34  to which it is attached, rotates in a second direction opposite of the first direction along the circumferential direction. In one embodiment, the intermediate speed turbine rotor  130 , and the intermediate speed shaft  35  to which it is attached, rotates in the second direction in co-rotation with the high speed turbine rotor  120  and in counter-rotation with the low speed turbine rotor  110 . 
     It should further be understood that the first direction and the second direction as used and described herein are intended to denote directions relative to one another. Therefore, the first direction may refer to a clockwise rotation (viewed from downstream looking upstream) and the second direction may refer to a counter-clockwise rotation (viewed from downstream looking upstream). Alternatively, the first direction may refer to a counter-clockwise rotation (viewed from downstream looking upstream) and the second direction may refer to a clockwise rotation (viewed from downstream looking upstream). 
     Still further during an operation of the engine  10 , combustion gases  86  exiting the combustion section  26  define a generally low speed toward the downstream end  98  of the engine  10 . A low speed rotation (e.g. along a tangential or circumferential direction) of the first stage  101  of the low speed turbine rotor  110  accelerates a speed of the combustion gases  86 , such as in the tangential or circumferential direction, to approximately equal or greater than a speed of the high speed turbine rotor  120 . 
     By defining the low speed turbine rotor  110  as the first stage  101  of the turbine section  90  aft or downstream of the combustion section  26 , the engine  10  may obviate the need for a first turbine vane or nozzle guide vane to accelerate the combustion gases  86  forward or upstream of the high speed turbine rotor  120  defining a high speed turbine. As such, the engine  10  may reduce a quantity or mass of cooling air from the compressor section  21  and/or combustion section  26 , thereby increasing engine efficiency by enabling more energy (i.e. compressed air) to be used during combustion. Additionally, or alternatively, the turbine section  90  may reduce necessary cooling air and enable increased performance and/or operability of the compressor section  21 , including surge margin and/or efficiency, or decrease a required amount of work from the compressor section  21 , which may reduce axial dimensions or stages of the compressor section  21  and further reduce engine packaging, weight, and/or part count, and generally improve engine  10  performance. 
     Additionally, obviating the need for the first turbine vane or nozzle guide vane may enable the turbine section  90 , or more specifically, the first stage  101 , as a rotating stage, to be designed to an average combustion gas  86  temperature rather than designed to accommodate peak temperatures (i.e. high spots) along an annulus of the core flowpath  70  within the combustion section  26 . Therefore, as all of the plurality of outer shroud airfoils  118  of the first stage  101  are rotating, all of the plurality of outer shroud airfoils  118  may only transiently endure adverse effects of combustion hot spots rather than substantially sustained or constant exposure to a higher temperature from the combustion gases in contrast to other locations about the annulus of the core flowpath  70 . Still further, the turbine section  90  described herein may enable alternative design methods for the combustion section  26  due to a decreased adverse effect of combustion hot spots on the turbine section  90 . Therefore, the turbine section  90  may enable design of the combustion section  26  to further improve combustion stability, decrease emissions, increase operability across all or part of a flight envelope, increase altitude re-light performance, and/or decrease lean blowout (LBO). 
     Referring now to  FIG. 3 , a detailed schematic cross sectional view of the engine  10  including a first turbine bearing  200  supporting the low speed turbine rotor  110  and the high speed turbine rotor  120  is generally provided. The first turbine bearing  200  is disposed radially between the hub  105  of the low speed turbine rotor  110  and the HP shaft  34  coupled to the high speed turbine rotor  120 . The first turbine bearing  200  defines an outer air bearing  210  along an outer diameter  215  of the first turbine bearing  200  and adjacent to the hub  105  of the low speed turbine rotor  110 . The first turbine bearing  200  further defines an inner air bearing  220  along an inner diameter  225  and adjacent to the HP shaft  34  coupled to the high speed turbine rotor  120 . 
     In various embodiments, the first turbine bearing  200  includes an outer annular wall  211  at the outer diameter  215  of the outer air bearing  210  and an inner annular wall  221  at the inner diameter  225  of the inner air bearing  220 . Each of the outer and inner annular walls  211 ,  221  extend circumferentially around the axial centerline  12  of the engine  10  and are generally concentric to the axial centerline  12 . The outer annular wall  211  and inner annular wall  221  each extend at least partially along the longitudinal direction L. The first turbine bearing  200  further includes an outer plurality of orifices  213  arranged at least along the longitudinal direction L and circumferentially along the outer air bearing  210 . The outer plurality of orifices  213  is adjacent to the hub  105  of the low speed turbine rotor  110 . The first turbine bearing  200  still further includes an inner plurality of orifices  223  arranged at least along the longitudinal direction L and circumferentially along the inner air bearing  220 . The inner plurality of orifices  223  is adjacent to the HP shaft  34  coupled to the high speed turbine rotor  120 . 
     In various embodiments, the outer air bearing  210  and/or the inner air bearing  220  may be formed of a plurality of segments arranged substantially in circumferential arrangement about the axial centerline  12  of the engine  10  to together define an approximately 360 degree annular surface, such as the outer annular wall  211  and/or the inner annular wall  221 . 
     Referring to  FIG. 3 , the combustion section  26  includes a compressor/diffuser frame  25  generally surrounding a combustor assembly and defining a pressure plenum  84 . The compressor/diffuser frame  25  generally defines a pressure vessel defining the annular pressure plenum  84  downstream of an exit of the HP compressor  24  and generally upstream of and/or radially surrounding the combustion chamber  85  at which the compressed air  82  substantially defines a compressor exit pressure. The compressor/diffuser frame  25  generally includes an inner diffuser frame  27  generally annular and radially inward within the combustion section  26 . The inner diffuser frame  27  generally extends from the compressor section  21  toward the turbine section  90 , supporting the first turbine bearing  200  and a combustor assembly. The first turbine bearing  200  is coupled to and cantilevered from the inner diffuser frame  27  toward the turbine section  90 . 
     In various embodiments, the first turbine bearing  200  defines an at least partially annular groove  217  at the outer diameter  215  of the outer annular wall  211  of the outer air bearing  210 . The groove  217  is disposed radially adjacent to the hub  105  of the low speed turbine rotor  110 . In various embodiments, the groove  217  is defined proximate to an upstream end of the hub  105  of the low speed turbine rotor  110 . For example, the groove  217  may be defined on the outer air bearing  210  radially adjacent to within 50% of the longitudinal span of the hub  105  from the upstream end of the hub  105  (i.e., from the end of the hub  105  most proximate to the upstream end  99  of the engine  10 ). As another example, the groove  217  may be defined on the outer annular wall  211  of the first turbine bearing  200  adjacent along the radial direction R within 30% of the span of the hub  105  along the longitudinal direction L proximate to the upstream end  99  of the engine  10 . 
     In one embodiment, the engine  10  further defines a compressor seal interface  240  between the HP compressor  24  and the first turbine bearing  200 . Between the compressor seal interface  240  and the first turbine bearing  200 , the engine  10  defines a first cavity  245  extended at least partially circumferentially around the axial centerline  12  of the engine  10 . The engine  10  further defines a turbine seal interface  250  between the high speed turbine rotor  120  and the first turbine bearing. Between the turbine seal interface  250  and the first turbine bearing  200 , the engine  10  defines a second cavity  255  extended at least partially circumferentially around the axial centerline  12  of the engine  10 . The seal interfaces  240 ,  250  may each include a combination of a shroud and knife edge or labyrinth seal to define each seal interface  240 ,  250 . 
     The engine  10  may further include a first manifold  260  extended from the pressure plenum  84  to the first turbine bearing  200  and in fluid communication therebetween. The compressed air  82  from the compressor section  21  may flow from the pressure plenum  84  through the first manifold  260  into the first turbine bearing  200  as shown schematically by arrows  261 . The air  261  through the first turbine bearing  200  flows to and egresses through the first and second plurality of orifices  213 ,  223  at the outer air bearing  210  and the inner air bearing  220  of the first turbine bearing  200 . 
     The first turbine bearing  200  generally defines a hydrostatic air bearing through which the compressed air  261  egresses through the first and second plurality of orifices  213 ,  223  to define a separating fluid film between the rotating components. At the outer air bearing  210  of the first turbine bearing  200 , the compressed air  261  defines a fluid film between the outer annular wall  211  and the hub  105  of the low speed turbine rotor  110 . At the inner air bearing  220  of the first turbine bearing  200 , the compressed air  261  defines a fluid film between the inner annular wall  221  and the HP shaft  34  coupled to the high speed turbine rotor  120 . The pressurized air from the compressor section  21  and in the pressure plenum  84  at least partially defines or determines an amount of clearance or distance between the hub  105  of the low speed turbine rotor  110  and the first turbine bearing  200 , and between the HP shaft  34  of the high speed turbine rotor  120  and the first turbine bearing  200 . 
     Referring now to  FIG. 4 , the first turbine bearing  200  shown and described in regard to  FIG. 3  may further define a cooling orifice  230  axially adjacent to the high speed turbine rotor  120  through which a portion of the compressed air  261  within the first turbine bearing  200  may egress into the high speed turbine rotor  120  to provide thermal attenuation or cooling, such as shown schematically by arrows  231 . In various embodiments, the high speed turbine rotor  120  defines a turbine cooling circuit  125  extended therethrough at least partially along the longitudinal direction L and radial direction R. The turbine cooling circuit  125  is defined generally within the structure or walls of a rotor, hub, or disk of the high speed turbine rotor  120 . The turbine cooling circuit  125  may further be defined within the structure of the high speed turbine airfoils  122 . The turbine cooling circuit  125  may define one or more cooling circuit egresses  127  at the high speed turbine airfoils  122 , through which the compressed air  231  egresses the high speed turbine rotor  120  into the core flowpath  70  at the turbine section  90 . 
     In various embodiments, the cooling orifice  230  of the first turbine bearing  200  defines a nozzle to increase a speed of and/or direct a flow of air  231  to the cooling circuit  125  of the high speed turbine rotor  120 . For example, the cooling orifice  230  may define a plurality of vanes defining a pressure side, a suction side, a leading edge, and a trailing edge, across which the compressed air  231  accelerates a flow of the air  231  along the longitudinal direction L and/or along the circumferential direction to approximately match the rotational speed of the high speed turbine rotor  120 . In various embodiments, the plurality of vanes defined within the cooling orifice  230  may further accelerate the flow of air  231  along the same direction of rotation as the high speed turbine rotor  120  (i.e., co-directional or co-rotational). For example, in the embodiments described herein, the plurality of vanes of the cooling orifice  230  may accelerate the air  231  generally along the second direction along the circumferential direction generally co-directional to the high speed turbine rotor  120  and opposite of the low speed turbine rotor  110 . However, in various embodiments, the cooling orifice  230  may accelerate the flow of air along the first direction. 
     Referring now to  FIGS. 3-4 , as the compressed air  261  pressurizes the space between the outer air bearing  210  and the hub  105 , a portion of the compressed air, shown schematically as  262 , may egress into the core flowpath  70 . Furthermore, as the compressed air  261  pressurizes the space between the first turbine bearing  200  and the hub  105  of the low speed turbine rotor  110 , the compressed air  261  is collected in an annular cavity  218  defined by the groove  217  at the outer annular wall  211 . Furthermore, as the compressed air  261  pressurizes the space between the first turbine bearing  200  and the HP shaft  34  coupled to the high speed turbine rotor  120 , the compressed air  261  is collected in the first cavity  245  defined between the compressor seal interface  240  and the first turbine bearing  200 . In various embodiments, the compressed air  261  may further be collected in the second cavity  255  defined between the turbine seal interface  250  and the first turbine bearing  200 . The air collected in each cavity  218 ,  245 ,  255  may then be collected within the first turbine bearing  200  and egressed through a second manifold  270  as shown schematically by arrows  271 . The second manifold  270  is extended from the first turbine bearing  200  to a pressure regulating valve  300  and in fluid communication therebetween. 
     In various embodiments, the second manifold  270  is in fluid communication with the annular cavity  218  and the first cavity  245 . The second manifold  270  may define a flow and pressure of the compressed air  271  therethrough defined by the air from the annular cavity and the first cavity. In another embodiment, the second manifold  270  is further in fluid communication with the second cavity  255 . The second manifold  270  may define a flow and pressure of the compressed air  271  therethrough defined by the air from the annular cavity, the first cavity, and the second cavity. 
     In various embodiments, the engine  10  further includes a third manifold  280  extended from the pressure plenum  84  of the combustion section  26  to the pressure regulating valve  300  and in fluid communication therebetween. In one embodiment, the third manifold  280  communicates a reference pressure or motive pressure to the pressure regulating valve  300 . The compressed air  82  in the pressure plenum  84  pressurizes the third manifold  280  as shown schematically by arrows  281 . 
     During operation of the engine  10  shown in  FIGS. 1-4 , the compressed air  281  from the pressure plenum generally provides a reference pressure to the pressure regulating valve  300  that may generally reflect or correspond to the pressure in pressure plenum  84 . The pressure regulating valve  300  generally regulates a difference in pressure between the third manifold  280  and the compressed air  281  within it and the second manifold  270  and the compressed air  271  egressing the first turbine bearing  200 . As such, the pressure regulating valve  300  regulates the pressure of air between the hub  105  of the low speed turbine rotor  110  and the outer air bearing  210  as well as the pressure of air between the inner air bearing  220  and the HP shaft  34  coupled to the high speed turbine rotor  120 . 
     Referring now to  FIG. 5 , the engine  10  may be configured substantially similarly to the embodiments shown and described in regard to  FIGS. 1-4 . In  FIG. 5 , the engine  10  may further define an OD turbine cooling supply manifold  275  extended from the second manifold  270  and coupled to the OD secondary flowpath  71  of the turbine section  90 . In one embodiment, a one way valve  310  is coupled to the turbine cooling supply manifold  275  to enable a flow of the compressed air  271  to the OD secondary flowpath  71  of the turbine section  90 . In various embodiments, the one way valve  310  may be referred to as a check valve or non-return valve, enabling a single direction of flow along the second manifold  270  from the first turbine bearing  200 . 
     During operation of the engine  10  shown in  FIG. 5 , when the pressure in the pressure plenum  84  and, subsequently, the third manifold  280 , is low, the pressure regulating valve  300  opens and sends the compressed air, as shown schematically by arrows  299 , to a lower pressure sink. In one embodiment, the lower pressure sink includes the core flowpath  70  toward a downstream end  98  of the turbine section  90 . In another embodiment, the lower pressure sink includes an ambient area radially outside of the engine  10  (e.g., at approximately ambient or atmospheric pressure condition, such as at approximately 1 bar, or less at altitudes above sea level). When the pressure regulating valve  300  opens and sends the compressed air  299  to the lower pressure sink, the one way valve  310  generally closes. Conversely, when pressure in the pressure plenum  84  is high, the one way valve  310  generally opens and enables a flow of the compressed air, as shown schematically by arrows  276 , into the turbine section  90 . The compressed air  276  provides cooling to the interdigitated low speed turbine rotor  110 , such as the outer shroud  114  and one or more of the plurality of outer shroud airfoils  118  extended therefrom. The compressed air  276  may additionally, or alternatively, provide buffering at one or more seals within the OD secondary flowpath  71  to restrict a flow of air from higher pressure cavities (e.g., from the combustion section  26  or generally from higher pressures upstream of others within the OD secondary flowpath  71  of the turbine section  90 ). 
     When included in an aircraft, the engine  10  may define the pressure in the pressure plenum  84  as low or high based, at least in part, on a landing and take-off (LTO) cycle of the aircraft, such as, but not limited to, those defined by standards, regulations, or requirements from the Federal Aviation Administration or the European Aviation Safety Agency. For example, the LTO as known in the art generally includes taxi, take-off, climb, approach, landing, and taxi. The LTO may further include cruise following take-off and descent following cruise and prior to approach. Generally, the engine  10  defines a high pressure at the pressure plenum  84  at a cruise condition or higher, such as including climb and take-off. The engine  10  defines a low pressure at the pressure plenum  84  below a cruise condition, such as at approach, landing, and taxi. Although various embodiments of the engine  10  may define quantitatively different high and low pressures at the pressure plenum  84 , it should be understood that various embodiments of the engine  10  may define qualitatively similar pressures relative to the LTO. 
     It should be appreciated that although the engine  10  may define the aforementioned LTO during normal operation, deviations in the LTO may define a high pressure in the engine  10  at lower altitudes, such as, but not limited to, altitudes commensurate to approach conditions or lower. For example, the engine  10  may define a high pressure approximately similar to the cruise condition or higher when an aircraft increases power to increase altitude or speed from an approach condition. As used herein, cruise condition includes those abnormal operating conditions that may define pressures, flows, and/or temperatures within the engine  10  generally similar to cruise condition. Still further, although the engine  10  is discussed in regard to within an aircraft, it should be appreciated that similar conditions may be defined in other apparatuses, including, but not limited to, power generation (e.g., industrial, auxiliary power unit, etc.), marine, locomotive, etc. 
     The first turbine bearing  200  supports the low speed turbine rotor  110  inward of the plurality of outer shroud airfoils  118  at the first stage  101 . For example, the first turbine bearing  200  may support the overhung or cantilevered low speed turbine rotor  110  generally forward or upstream  99  of the high speed turbine rotor  120 . Still further, the first turbine bearing  200  supports the high speed turbine rotor  120  and the HP shaft  34  coupled to the HP compressor  24 . 
     The arrangement of the first bearing  200  may provide support toward the upstream end  99  of the low speed turbine rotor  110  to be interdigitated forward and/or among the high speed turbine rotor  120 . Furthermore, the first bearing  200  provides support toward the upstream end  99  of the low speed turbine rotor  110  that limits an overhanging or cantilevered weight of the low speed turbine rotor  110  from the connecting airfoil  116  upstream toward the combustion section  26 . Still further, the first bearing  200  provides support toward the upstream end  99  of the low speed turbine rotor  110  that provides balance to the inner shroud  112  and the plurality of inner shroud airfoils  119  extended therefrom toward the downstream end  98  of the turbine section  90 . In various embodiments, the axially extended hub  105  of the low speed turbine rotor  110  may further define one or more balance planes. The balance plane may define features to which weight may be added to or removed from the low speed turbine rotor  110  to aid rotor balance and operation. 
     Referring now to  FIG. 6 , another exemplary embodiment of the engine  10  including the first turbine bearing  200  supporting the low speed turbine rotor  110  and the high speed turbine rotor  120  is generally provided. The exemplary embodiment provided in  FIG. 6  may be configured substantially similarly as embodiments shown and described in regard to  FIGS. 1-5 . However, the embodiment of the engine  10  in  FIG. 6  further includes a fourth manifold  290  extended from the pressure plenum  84  and coupled to the first turbine bearing  200 . The fourth manifold  290  provides fluid communication of flow and pressure of air from the pressure plenum  84  to the first turbine bearing  200 , such as shown schematically by arrows  291 , through the first turbine bearing  200  and egressing through the cooling orifice  230 . The fourth manifold  290  may collect air from the pressure plenum  84  independently of the first manifold  260  and the compressed air  261  pressurized therethrough to the outer air bearing  210  and the inner air bearing  220 . As such, the fourth manifold  290  may define orifices, volumes, or areas, including, but not limited to, restrictor or orifice plates, variable geometry, valves, etc. that may induce a different pressure of compressed air  291  through first turbine bearing  200  through the cooling orifice  230  and egressing into the turbine cooling conduit  125  of the high speed turbine rotor  120  (as shown schematically by arrows  231 ) than that of the compressed air  261 . 
     In various embodiments, the fourth manifold  290  may extend to a separate external compressed air source (e.g., outside of the pressure plenum  84  or outside of the combustion section  26 ) to induce a flow and/or pressure independent of the flow or pressure through the pressure plenum  84 . The flow and/or pressure of air  291  through the fourth manifold  290  may be actively or passively controlled separately from the air  261  through the first manifold  260  into the first turbine bearing  200 . Active control may include one or more of a valve or actuation to open or restrict air  291  to the turbine cooling circuit  125 . Passive control may include one or more of a metering or restrictor plate, orifice, or nozzle, conduit, or manifold defining an area of volume sized to provide a desired minimum or maximum flow and/or pressure at one or more engine operating conditions. 
     Referring now to  FIG. 7 , an exemplary flowchart depicting flows of air and combustion gases through the engine  10  is generally provided.  FIG. 7  depicts flows as shown and described in regard to  FIGS. 1-6 , and various embodiments thereof. However, it should be appreciated that the flowchart may include or omit any one or several embodiments as discussed in regard to each of  FIGS. 1-6  or embodiments thereof.  FIG. 7  shows an initial flow of air  80  entering the core flowpath  70  of the engine  10 . The air  80  enters the IP compressor  22  (or, in two-spool embodiments, the LP compressor) and the HP compressor  24  to become compressed air  82  entering the combustion section  26 . A portion of the compressed air  82  egresses from the pressure plenum  84  and enters into the first turbine bearing  200 , as shown schematically at arrows  261 . As shown and discussed in regard to  FIGS. 1-5 , the compressed air  261  egresses through the outer plurality of orifices  213  of the outer air bearing  210  and the inner plurality of orifices  223  of the inner air bearing  220 , each of which generate a gap or clearance in support of the low speed turbine rotor  110  and the high speed turbine rotor  120 , respectively. 
     As shown and discussed in regard to  FIGS. 1-6 , the compressed air  261 , after egressing from the outer air bearing  210  and inner air bearing  220 , is collected in one or more cavities, such as the annular cavity  218 , the first cavity  245 , and/or the second cavity  255  ( FIGS. 1-6 ), and egressed from the first turbine bearing  200 , as shown schematically by arrows  271 , to the pressure regulating valve  300 . 
     Referring still to  FIG. 7 , a portion of compressed air  82  may be egressed from the pressure plenum  84  directly to the pressure regulating valve  300 , such as shown schematically by arrows  281 . The compressed air  281  provides a reference pressure or motive pressure at the pressure regulating valve  300  in comparison to the input pressure from the compressed air  271  egressing the first turbine bearing  200  and a differential from the air  299  egressing the pressure regulating valve  300  to a lower pressure sink (e.g., such as the core flowpath  70  downstream of the turbine section  90 , or an ambient condition). 
     In various embodiments, such as shown in  FIGS. 6-7 , compressed air  291  may enter the first turbine bearing  200  and egress into high speed turbine rotor  120 , as shown schematically by arrows  231 , independently of the compressed air  261  at least partially going through the outer air bearing  210  and inner air bearing  220 . 
     In still various embodiments, at least a portion of air may egress from the first turbine bearing  200  to the core flowpath at or upstream of the turbine section  90 , as shown schematically by arrows  262 . 
     Referring still to  FIG. 7 , a portion of air may egress from the pressure regulating valve  300  into the OD secondary flowpath  71  of the turbine section  90 , such as shown schematically by arrows  276 . The compressed air  276  may provide cooling to the low speed turbine rotor  110  interdigitated among the high speed turbine rotor  120 . In various embodiments, the one way valve  310  is disposed upstream of the OD secondary flowpath  71  and parallel to the flow of compressed air  271  going to the pressure regulating valve  300 . The one way valve  310  may restrict a flow of the air  276  toward the OD secondary flowpath  71  and disable flow toward the pressure regulating valve  300 . 
     Referring now to  FIG. 8 , a flowchart outlining an exemplary method of operating a gas turbine engine  800  is generally provided (hereinafter referred to as “method  800 ”). The method  800  generally provides an operation of an engine such as that shown in  FIGS. 1-7  and described in regard thereto. The engine  10  generally includes a first turbine bearing  200  defining an outer air bearing  210  radially adjacent to a low speed turbine rotor  110  and an inner air bearing  210  radially adjacent to a high speed turbine rotor  120  and an HP shaft  34  coupled thereto. The engine  10  includes a first manifold  260  extended to the first turbine bearing  200  and in fluid communication with a pressure plenum  84  of a combustion section  26  of the engine  10 . A second manifold  270  is coupled to the first turbine bearing  200 . The second manifold  270  provides fluid communication from the first turbine bearing  200  to the pressure regulating valve  300  or to a lower pressure sink (e.g., an ambient or atmospheric condition outside of the engine, or a lower pressure, such as in the core flowpath proximate to the downstream end of the engine  10 ). The method  800  outlined in  FIG. 8  and described herein may provide operation and support of the interdigitated turbine section  90 , such as of the low speed turbine rotor  110  and the high speed turbine rotor  120 . 
     The method  800  includes at  810  flowing air from the pressure plenum to the first turbine bearing; at  820  flowing air between the outer air bearing and the low speed turbine rotor hub; at  830  flowing air between the inner air bearing and the HP shaft; and at  840  flowing air from the first turbine bearing to the pressure regulating valve via the second manifold. Although  FIG. 8  depicts steps performed in a particular order for the purposes of illustration and discussion, those of ordinary skill in the art, using the disclosures provided herein, will understand that various steps of any of the methods disclosed herein may be adapted, modified, rearranged, omitted, or expanded in various ways without deviating from the scope of the present disclosure. 
     In various embodiments of the method  800  outlined in  FIG. 8  and described herein, in conjunction with the embodiments of the engine  10  shown in  FIGS. 1-7 , the step at  810  may include flowing air  261  from the pressure plenum  84  to the first turbine bearing  200  including the outer turbine bearing  210  and the inner turbine bearing  220 . At  820 , flowing air  261  between the outer air bearing  220  and the hub  105  of the low speed turbine rotor  110  includes defining a pressure of the air  261  between the outer air bearing  220  and the hub  105  such that defines a non-contacting gap or clearance between the first turbine bearing  200  and the hub  105  of the low speed turbine rotor  110 . In one embodiment at  830 , flowing air between the inner air bearing  220  and the HP shaft  34  defines a pressure of the air  261  such that defines a non-contacting gap or clearance between the first turbine bearing  200  and the HP shaft  34  coupled to the high speed turbine rotor  120 . In an embodiment at  840 , flowing air  261  from the first turbine bearing  200  to the pressure regulating valve  200  via the second manifold  270  defines a flow and pressure of the air  271  as shown and described in regard to  FIGS. 3-7 . 
     In one embodiment, the method  800  further includes at  822  providing a first supply of air from between the outer air bearing and the low speed turbine rotor hub to the second manifold; at  824  providing a second supply of air from between the inner air bearing and the HP shaft to the second manifold; and at  825  flowing the first supply of air and the second supply of air through the second manifold to a lower pressure sink. For example, the method  800  at  822  may include providing air  271  from the annular cavity  218  defined by the groove  217  in the outer air bearing  210 , and egressing the air  271  through the first turbine bearing  200  to the second manifold  270 . The method  800  at  824  may include providing air  271  from the first cavity  245  through the first turbine bearing  200  to the second manifold  270 . The method  800  at  824  may further include providing air  271  from the second cavity  255  through the first turbine bearing  200  to the second manifold  270 . The method  800  at  825  may include collecting the air  271  from the annular cavity  218  and the first cavity  245  such that the pressure of the air  271  is normalized, and egressed through the second manifold  270 , as shown schematically by arrows  271  in  FIGS. 3-7 . The method  800  at  825  may further include collecting the air  271  from the annular cavity  218 , the first cavity  245 , and the second cavity  255 , normalizing the pressure, and egressing the air  271  through the second manifold  270 . 
     In one embodiment, the step at  825  includes flowing the first supply of air and the second supply of air through the second manifold  270  to the pressure regulating valve  300 . In still various embodiments, the step at  825  may include flowing the first supply of air and the second supply of air through the second manifold to a lower pressure sink. For example, the lower pressure sink may be an ambient or atmospheric pressure condition outside of the engine  10  or toward a downstream end  98  of the core flowpath  70  of the turbine section  90 . In various examples, the lower pressure sink generally defines a pressure less than within the pressure plenum  84 , the first turbine bearing  200 , or the second manifold  270 . In one embodiment, the lower pressure sink is approximately 1 bar or less. 
     In one embodiment of the method  800 , the method  800  further includes at  850  flowing air from the pressure plenum of the combustion section to the pressure regulating valve via the third manifold. For example, the method  800 , such as implemented in conjunction with the engine  10  shown in  FIGS. 1-7 , may include flowing or communicating to the pressure regulating valve  300  the pressure within the pressure plenum  84  via the third manifold  280 , such as shown via the compressed air  281  from the pressure plenum  84  to the pressure regulating valve  300 . 
     In another embodiment, the method  800  further includes at  860  flowing air from between the outer air bearing and the low speed turbine rotor hub to the core flowpath of the engine at the turbine section. For example, referring to  FIGS. 1-7 , the method  800  at  860  includes directing a flow of air  262  from the outer plurality of orifices  213  of the outer air bearing  210  to the core flowpath  70  at the turbine section  90  of the engine  10 . 
     Referring now to  FIG. 9 , a flowchart outlining an exemplary method of operating a gas turbine engine  900  is generally provided (hereinafter referred to as “method  900 ”). The method  900  generally provides an operation of an engine such as that shown in  FIGS. 1-7  and described in regard thereto. The method  900  may further include one or more steps outlined and described in regard to method  800  outlined and described in regard to  FIG. 8 . The engine  10  generally includes a first turbine bearing  200  defining an outer air bearing  210  radially adjacent to a low speed turbine rotor  110  and an inner air bearing  210  radially adjacent to a high speed turbine rotor  120  and an HP shaft  34  coupled thereto. The engine  10  includes a first manifold  260  extended to the first turbine bearing  200  and in fluid communication with a pressure plenum  84  of a combustion section  26  of the engine  10 . A second manifold  270  is coupled to the first turbine bearing  200 . A third manifold  280  is coupled to the combustion section  26  and provides fluid communication from the pressure plenum  84  to a pressure regulating valve  300 . The second manifold  270  provides fluid communication from the first turbine bearing  200  to the pressure regulating valve  300 . The method  900  outlined in  FIG. 9  and described herein may provide operation and support of the interdigitated turbine section  90 , such as of the low speed turbine rotor  110  and the high speed turbine rotor  120 , while also providing thermal attenuation at the high speed turbine rotor  120 . 
     The method  900  may include at  910  flowing air from the pressure plenum to the first turbine bearing; at  920  flowing air between the outer air bearing and the low speed turbine rotor hub; at  930  flowing air between the inner air bearing and the HP shaft; at  940  flowing air from the first turbine bearing to the pressure regulating valve via the second manifold; and at  945  flowing air from the first turbine bearing through a turbine cooling circuit of the high speed turbine rotor. The steps at  910 ,  920 ,  930 , and  940 , and various embodiments thereof, may be undertaken as described in regard to the steps at  810 ,  820 ,  830 , and  840 , respectively, and in various embodiments thereof. 
     In various embodiments, such as in regard to  FIGS. 4-6 , the step at  945  includes egressing the flow of air  231  from the first turbine bearing  200  through the cooling orifice  230  into the turbine cooling conduit  125  of the high speed turbine rotor  120 . For example, in one embodiment, in reference to the engine  10  shown and described in regard to  FIGS. 4-5 , the first turbine bearing  200  defines a conduit there through to provide compressed air  261  from the first manifold  260  and pressure plenum  84  and egressing through the cooling orifice  230  adjacent to the turbine cooling circuit  125  of the high speed turbine rotor  120 , as depicted schematically by arrows  231 . 
     In one embodiment, the method  900  includes at  950  flowing air from the pressure plenum of the combustion section to the pressure regulating valve via the third manifold, such as described in regard to step  850  of method  800  and shown and described in regard to  FIGS. 1-7 . 
     In another embodiment, the method  900  includes at  960  flowing air from between the outer air bearing and the low speed turbine rotor hub to the core flowpath of the turbine section, such as described in regard to step  860  of method  800  and shown and described in regard to  FIGS. 1-8 . 
     In still another embodiment, flowing air to the turbine cooling circuit  125  of the high speed turbine rotor  120  includes at  970  modulating or adjusting a flow and/or pressure of the compressed air  231  egressing the first turbine bearing  200  into the turbine cooling circuit  125 . For example, referring to  FIGS. 6-7 , modulating or adjusting a flow and/or pressure of compressed air  231  may include flowing air from the fourth manifold  290  independently of the first manifold  260 . 
     In still another embodiment, the method  900  may further include at  980  monitoring a difference in air pressure from the second manifold and air pressure from the third manifold. For example, monitoring the difference in air pressure includes a difference between the compressed air  271  in the second manifold  270  and the compressed air  281  in the third manifold  280 . The compressed air  281 , or more specifically, the pressure of the compressed air  281  provides to the pressure regulating valve  300  a reference pressure or muscle air for operating the pressure regulating valve  300  to maintain a desired pressure of the compressed air  261  at the first turbine bearing  200  for maintaining a desired gap or clearance between the first turbine bearing  200  and each of the hub  105  and HP shaft  34 . 
     In various embodiments, the method  900  may include at  922  providing a first supply of air from between the outer air bearing and the low speed turbine rotor hub; at  924  providing a second supply of air from between the inner air bearing and the HP shaft; and at  926  flowing the first supply of air and the second supply of air through the second manifold to the pressure regulating valve. The steps at  922  and  924  may be performed substantially similarly as described in regard to steps  822  and  824  of method  800 . 
     The method  900  at  926  may include collecting the air  271  from the annular cavity  218  and the first cavity  245  such that the pressure of the air  271  is normalized, and egressed through the second manifold  270 , as shown schematically by arrows  271  in  FIGS. 3-7 . The method  900  at  926  may further include collecting the air  271  from the annular cavity  218 , the first cavity  245 , and the second cavity  255 , normalizing the pressure, and egressing the air  271  through the second manifold  270 . For example, in one embodiment, the step at  926  includes flowing the first supply of air and the second supply of air through the second manifold  270  to the pressure regulating valve  300 . 
     Referring now to  FIG. 10 , a flowchart outlining an exemplary method of thermal management a gas turbine engine  1000  is generally provided (hereinafter referred to as “method  1000 ”). The method  1000  may provide thermal attenuation of the low speed turbine rotor  110  of the interdigitated turbine section  90  such as shown and described in regard to  FIGS. 1-7 . The method  1000  may further include one or more steps of the method  800  and/or method  900  and outlined and described in regard to  FIGS. 8-9 . 
     The method  1000  may include at  1010  flowing air from the pressure plenum to the first turbine bearing; at  1020  flowing air between the outer air bearing and the low speed turbine rotor hub; at  1030  flowing air between the inner air bearing and the HP shaft; and at  1040  flowing air from the first turbine bearing to the pressure regulating valve via the second manifold. The steps at  1010 ,  1020 ,  1030 , and  1040 , and various embodiments thereof, may be undertaken as described in regard to the steps at  810 ,  820 ,  830 , and  840 , respectively, and in various embodiments thereof. 
     The method  1000  further includes at  1047  flowing air from the first turbine bearing to the outer diameter (OD) secondary flowpath via the second manifold. For example, the method  1000  at  1047  may include flowing compressed air  271  from the second manifold  270  through the turbine cooling supply manifold  275  coupled to the outer turbine casing  155  and in fluid communication with the OD secondary flowpath  71 , such as shown schematically by arrows  276  in  FIG. 5 . 
     In one embodiment of the method  1000  at  1047 , flowing air  276  through the OD secondary flowpath  71  may include opening a one way valve (e.g., one way valve  310 ), such as by defining a pressure of the air  271  in the second manifold  270  that overcomes the one way valve  310  at the turbine cooling supply manifold  275 . In various embodiments, the method  1000  at  1047  may include operating the engine  10  at a cruise power condition or higher of the LTO. 
     In one embodiment, the method  1000  may further includes at  1049  flowing air from the OD secondary flowpath to the core flowpath. For example, in regard to  FIG. 5 , compressed air, shown schematically by arrows  276 , flows from the turbine cooling supply manifold  275  through the OD secondary flowpath  71  into the core flowpath  70 . More specifically, the compressed air  276  flows proximate to the outer shroud  114  and the outer shroud airfoils  118  of the low speed turbine rotor  110  to provide cooling. In various embodiments, the compressed air  276  may further flow toward the downstream end  98  within the OD secondary flowpath  71  to further provide cooling along the low speed turbine  110 . 
     In another embodiment, the method  1000  may further include at  1090  flowing air from the pressure regulating valve to a lower pressure sink. For example, referring to  FIG. 5 , the method  1000  at  1090  may include flowing air  299  from the pressure regulating valve  300  to the core flowpath  70  at the downstream end  98  of the turbine section  90  (e.g., proximate to the inner shroud airfoils  119  of the low speed turbine rotor  110 ). As another example, the method  1000  at  1090  may include flowing air  299  to an ambient condition outside of the core flowpath  70 . In various embodiments, the method  1000  at  1090  includes operating the engine  10  below a cruise power condition (e.g., approach, landing, taxi, or generally at low pressure/low altitude conditions). 
     In various embodiments, the method  1000  may further includes steps shown and described in regard to method  800  and method  900 , such as, the steps  822 ,  824 ,  926 ,  945 ,  850 ,  860 ,  970 , or  980  commensurate to steps  1022 ,  1024 ,  1026 ,  1045 ,  1050 ,  1060 ,  1070 , or  1080  herein. 
     The turbine section  90  shown and described herein may improve upon existing turbine sections by providing improved fuel efficiency, operational efficiency, and/or power output while maintaining or reducing weight, part count, and/or packaging. The plurality of outer shroud airfoils  118  of the low speed turbine rotor  110  interdigitated among the plurality of high speed turbine airfoils  122  of the high speed turbine rotor  120  may reduce packaging and reduce part count by removing stages of stationary airfoils between each rotating component. Additionally, the turbine section  90  may provide efficiency benefits comparable to a reduction gearbox without adding weight or size (e.g. axial length) to the engine  10 . The low speed turbine rotor  110 , as the first stage downstream of the combustion section  26 , may further improve engine efficiency by reducing cooling air appropriated away from producing combustion gases  86 , thereby allowing more energy from the compressor section  21  to be used in combustion and operation of the engine  10 . Furthermore, removing the nozzle guide vane between the combustion section  26  and the low speed turbine rotor  110  of the turbine section  90  may reduce or eliminate design constraints related to hot spots in the combustion gases along the annulus of the core flowpath  70 . 
     The various embodiments of the turbine section  90  generally shown and described herein may be constructed as individual blades installed into drums or hubs, or integrally bladed rotors (IBRs) or bladed disks, or combinations thereof. The blades, hubs, or bladed disks may be formed of ceramic matrix composite (CMC) materials and/or metals appropriate for gas turbine engine hot sections, such as, but not limited to, nickel-based alloys, cobalt-based alloys, iron-based alloys, or titanium-based alloys, each of which may include, but are not limited to, chromium, cobalt, tungsten, tantalum, molybdenum, and/or rhenium. The turbine section  90 , or portions or combinations of portions thereof, including the inner shroud  112 , the outer shroud  114 , the connecting airfoil(s)  116 , the plurality of outer shroud airfoils  118 , and/or the plurality of inner shroud airfoils  119 , may be formed using additive manufacturing or 3D printing, or casting, forging, machining, or castings formed of 3D printed molds, or combinations thereof. The turbine section  90 , or portions thereof, such as stages of the rotating components  110 ,  120 ,  130 , the outer shroud  114 , the inner shroud  112 , and/or various shrouds, seals, and other details may be mechanically joined using fasteners, such as nuts, bolts, screws, pins, or rivets, or using joining methods, such as welding, brazing, bonding, friction or diffusion bonding, etc., or combinations of fasteners and/or joining methods. Still further, it should be understood that the low speed turbine rotor  110 , including the inner and/or outer shroud  112 ,  114 , may incorporate features that allow for differential expansion. Such features include, but are not limited to, aforementioned methods of manufacture, various shrouds, seals, materials, and/or combinations thereof. 
     The systems and methods shown in  FIGS. 1-10  and described herein may decrease fuel consumption, increase operability, increase engine performance and/or power output while maintaining or reducing weight, part count, and/or packaging (e.g. radial and/or axial dimensions). The systems provided herein may allow for increased high bypass ratios and/or overall pressure ratios over existing gas turbine engine configurations, such as turbofans, while maintaining or reducing packaging relative to other gas turbine engines of similar power output. The systems described herein may contribute to improved bypass ratio and/or overall pressure ratio and thereby increase overall gas turbine engine efficiency. The systems provided herein may increase overall gas turbine engine efficiency by reducing or eliminating stationary airfoils that require cooling air (e.g. nozzle guide vane). 
     Still further, the systems and methods shown in  FIGS. 1-10  and described herein may reduce a product of a flow area and the square of the rotational speed (the product herein referred to as “AN 2 ”) of the gas turbine engine. For example, engine  10  shown and described in regard to  FIGS. 1-10  may generally reduce AN 2  relative to a conventional geared turbofan configuration. Generally, lowering the AN 2 , such as by reducing the rotational speed and/or the flow area, increases the required average stage work factor (i.e. the average required loading on each stage of rotating airfoils). However, the systems described herein may lower the AN 2  while also lowering the average stage work factor and maintaining axial length of the turbine section  90  (compared to engines of similar thrust output and packaging) by interdigitating the low speed turbine rotor  110  among the one or more stages of the high speed turbine rotor  120  and the intermediate speed turbine  130  while also defining a non-digitated turbine structure (i.e. the inner shroud  112  and the plurality of inner shroud airfoils  119 ) toward the downstream end  98  of the turbine section  90 . Therefore, the low speed turbine rotor  110  may increase the quantity of rotating stages of airfoils while reducing the average stage work factor, and therefore the AN 2 , while mitigating increases in axial length to produce a similar AN 2  value. The low speed turbine rotor  110  may further reduce the AN 2  while additionally reducing the overall quantity of airfoils, rotating and stationary, in the turbine section  90  relative to turbine sections of gas turbine engines of similar power output and/or packaging. 
     Furthermore, the systems and methods shown in  FIGS. 1-10  and described herein may further improve engine efficiency, reduce airfoil quantity, reduce engine weight, and/or alleviate combustion section design constraints by interdigitating the low speed turbine rotor  110  forward or upstream  99  of the high speed turbine rotor  120 . For example, defining the first stage of the low speed turbine rotor  110  as immediately downstream  98  of the combustion section  26 , without a first turbine vane or nozzle guide vane therebetween, as well as defining the low speed turbine rotor  110  in counter-rotation with the high speed turbine rotor  120 , may reduce effects of overall combustion hot spots on the first stage of the low speed turbine rotor  110  in contrast to a stationary, first turbine vane or nozzle guide vane. As such, the turbine section  90  and engine  10  described herein may remove constraints to combustion section  26  design by de-emphasizing hot spots, or combustion pattern factor, in favor of other design criteria, such as decreasing emissions, improving lean blow-out (LBO) and/or altitude re-light, improving overall operability across part or all of an operating envelope, or increasing the operating envelope. 
     This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.