Abstract:
A disclosed vane assembly includes a vane section formed of a plurality of circumferentially spaced fixed vanes. The vanes extend radially outward from an inner platform and hooked into case. The inner platform includes a mount rail extending radially inwardly from the inner platform. An air seal is attached to the inner platform of the vane section and includes a ring extending circumferentially about the axis. The disclosed air seal includes a plurality of tabs that receive lugs disposed on the mount rail. A ring nut is secured to the air seal and engaged to the mount rail for securing the vane section to the air seal.

Description:
BACKGROUND 
       [0001]    A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. 
         [0002]    Compressor and turbine sections include stages of rotating airfoils and stationary vanes. Radially inboard and outboard platforms and seals contain gas flow through the airfoils and vanes. Seals between rotating and static parts include edges that ride and abut static honeycomb elements. Moreover, cooling airflow is often directed through the static vanes to inner surfaces to provide an air pressure and/or flow that further contain the flow of hot gases between platforms of the airfoils and vanes. The structures required to define sealing interfaces and cooling air passages can be costly and complicate assembly. 
         [0003]    Accordingly, it is desirable to design and develop structures that reduce cost, simplify assembly while containing hot gas flow and defining desired cooling airflow passages. 
       SUMMARY 
       [0004]    A turbine section according to an exemplary embodiment of this disclosure, among other possible things includes first and second turbine rotors each carrying turbine blades for rotation about a central axis. The rotors each have at least one rotating seal at a radially inner location. A vane assembly includes a vane extending radially from a platform. An air seal is attached to the vane assembly, the air seal includes a ring extending circumferentially about the axis and a ring nut received on the air seal for securing the air seal to the vane assembly. 
         [0005]    In a further embodiment of the foregoing turbine section, the air seal includes mating features for circumferentially locating the air seal relative to the vane assembly. 
         [0006]    In a further embodiment of any of the foregoing turbine sections, includes a full ring seal disposed between a surface of the vane platform and the ring nut. 
         [0007]    In a further embodiment of any of the foregoing turbine sections, includes a lock ring engaged to the air seal and the ring nut for securing a relative position between the ring nut and the air seal. 
         [0008]    In a further embodiment of any of the foregoing turbine sections, includes a wire seal disposed between the ring nut and a surface of the air seal. 
         [0009]    In a further embodiment of any of the foregoing turbine sections, the platform includes a radially inward extending rim engaging a forward lip of the air seal. 
         [0010]    In a further embodiment of any of the foregoing turbine sections, the air seal includes a forward wall with openings for exhausting air flow. 
         [0011]    A vane assembly according to an exemplary embodiment of this disclosure, among other possible things includes a vane including an inner platform having a mount rail extending radially inwardly, an air seal attached to the inner platform of the vane section, the air seal includes a ring extending circumferentially about the axis including centering tabs receiving lugs disposed on the mount rail, and a ring nut received on the air seal and engaged to the mount rail for securing the air seal to the vane section. 
         [0012]    In a further embodiment of the foregoing vane assembly, includes a full ring seal disposed between a surface of the inner platform and the ring nut. 
         [0013]    In a further embodiment of any of the foregoing vane assemblies, the mount rail and the ring nut define a seal cavity and the full ring seal is disposed within the seal cavity. 
         [0014]    In a further embodiment of any of the foregoing vane assemblies, includes a lock ring engaged to the air seal and the ring nut for securing a relative position between the ring nut and the air seal. 
         [0015]    In a further embodiment of any of the foregoing vane assemblies, includes a wire seal disposed between the ring nut and a surface of the air seal. 
         [0016]    In a further embodiment of any of the foregoing vane assemblies, the inner platform includes a radially inward extending rim engaging a forward lip of the air seal. 
         [0017]    In a further embodiment of any of the foregoing vane assemblies, air seal includes a front wall with openings for exhausting cooling air flow. 
         [0018]    A method of assembling a vane assembly according to an exemplary embodiment of this disclosure, among other possible things includes defining a plurality of vanes circumferentially about an axis that extend from an inner platform, abutting a front hub of the inner platform against a lip of an air seal, and loading the front hub against the lip of the air seal with a ring nut threaded onto the air seal. 
         [0019]    In a further embodiment of the foregoing method, includes the step of engaging a plurality of tabs on a lock ring with the ring nut to hold a position of the ring nut relative to the air seal. 
         [0020]    In a further embodiment of any of the foregoing methods, includes the sealing between the ring nut and a mount rail of the inner platform. 
         [0021]    In a further embodiment of any of the foregoing methods, includes defining an cooling air chamber between the air seal and the inner platform and exhausting cooling air flow from openings within the air seal. 
         [0022]    Although the different examples have the specific components shown in the illustrations, embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples. 
         [0023]    These and other features disclosed herein can be best understood from the following specification and drawings, the following of which is a brief description. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0024]      FIG. 1  is a schematic view of an example gas turbine engine. 
           [0025]      FIG. 2  is an enlarged cross-sectional view of a portion of the gas turbine engine. 
           [0026]      FIG. 3  is a sectional view of an example vane assembly. 
           [0027]      FIG. 4  is a sectional view of an example air seal. 
           [0028]      FIG. 5  is a cross-sectional view of an example lower platform. 
           [0029]      FIG. 6  is a perspective view of an example lock ring. 
           [0030]      FIG. 7  is a perspective view of an example ring nut. 
           [0031]      FIG. 8  is a schematic view of the example air seal including the lock ring. 
           [0032]      FIG. 9  is a front view of the example vane assembly. 
           [0033]      FIG. 10  is a rear view of the example vane assembly. 
       
    
    
     DETAILED DESCRIPTION 
       [0034]      FIG. 1  schematically illustrates an example gas turbine engine  20  that includes a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines might include an augmenter section (not shown) among other systems or features. The fan section  22  drives air along a bypass flow path B while the compressor section  24  draws air in along a core flow path C where air is compressed and communicated to a combustor section  26 . In the combustor section  26 , air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section  28  where energy is extracted and utilized to drive the fan section  22  and the compressor section  24 . 
         [0035]    Although the disclosed non-limiting embodiment depicts a turbofan gas turbine engine, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines; for example a turbine engine including a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section. 
         [0036]    The example engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided. 
         [0037]    The low speed spool  30  generally includes an inner shaft  40  that connects a fan  42  and a low pressure (or first) compressor section  44  to a low pressure (or second) turbine section  46 . The inner shaft  40  drives the fan  42  through a speed change device, such as a geared architecture  48 , to drive the fan  42  at a lower speed than the low speed spool  30 . The high-speed spool  32  includes an outer shaft  50  that interconnects a high pressure (or second) compressor section  52  and a high pressure (or first) turbine section  54 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via the bearing systems  38  about the engine central longitudinal axis A. 
         [0038]    A combustor  56  is arranged between the high pressure compressor  52  and the high pressure turbine  54 . In one example, the high pressure turbine  54  includes at least two stages to provide a double stage high pressure turbine  54 . In another example, the high pressure turbine  54  includes only a single stage. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine. 
         [0039]    The example low pressure turbine  46  has a pressure ratio that is greater than about  5 . The pressure ratio of the example low pressure turbine  46  is measured prior to an inlet of the low pressure turbine  46  as related to the pressure measured at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. 
         [0040]    A mid-turbine frame  58  of the engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The mid-turbine frame  58  further supports bearing systems  38  in the turbine section  28  as well as setting airflow entering the low pressure turbine  46 . 
         [0041]    The core airflow C is compressed by the low pressure compressor  44  then by the high pressure compressor  52  mixed with fuel and ignited in the combustor  56  to produce high speed exhaust gases that are then expanded through the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  58  includes vanes  60 , which are in the core airflow path and function as an inlet guide vane for the low pressure turbine  46 . Utilizing the vane  60  of the mid-turbine frame  58  as the inlet guide vane for low pressure turbine  46  decreases the length of the low pressure turbine  46  without increasing the axial length of the mid-turbine frame  58 . Reducing or eliminating the number of vanes in the low pressure turbine  46  shortens the axial length of the turbine section  28 . Thus, the compactness of the gas turbine engine  20  is increased and a higher power density may be achieved. 
         [0042]    The disclosed gas turbine engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine  20  includes a bypass ratio greater than about six (6), with an example embodiment being greater than about ten (10). The example geared architecture  48  is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about 2.3. 
         [0043]    In one disclosed embodiment, the gas turbine engine  20  includes a bypass ratio greater than about ten (10:1) and the fan diameter is significantly larger than an outer diameter of the low pressure compressor  44 . It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines. 
         [0044]    A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section  22  of the engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft., with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of pound-mass (lbm) of fuel per hour being burned divided by pound-force (lbf) of thrust the engine produces at that minimum point. 
         [0045]    “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.50. In another non-limiting embodiment the low fan pressure ratio is less than about 1.45. 
         [0046]    “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/518.7) 0.5 ]. The “Low corrected fan tip speed”, as disclosed herein according to one non-limiting embodiment, is less than about 1150 ft/second. 
         [0047]    The example gas turbine engine includes the fan  42  that comprises in one non-limiting embodiment less than about 26 fan blades. In another non-limiting embodiment, the fan section  22  includes less than about 20 fan blades. Moreover, in one disclosed embodiment the low pressure turbine  46  includes no more than about 6 turbine rotors schematically indicated at  34 . In another non-limiting example embodiment the low pressure turbine  46  includes about 3 turbine rotors. A ratio between the number of fan blades  42  and the number of low pressure turbine rotors is between about 3.3 and about 8.6. The example low pressure turbine  46  provides the driving power to rotate the fan section  22  and therefore the relationship between the number of turbine rotors  34  in the low pressure turbine  46  and the number of blades  42  in the fan section  22  disclose an example gas turbine engine  20  with increased power transfer efficiency. 
         [0048]    Referring to  FIG. 2  with continued reference to  FIG. 1 , the example the high pressure turbine  54  includes first and second rotors  62 ,  64 , and corresponding first and second airfoils  74  and  76  that rotate with the first and second rotors  62 ,  64 . Vane assembly  66  is disposed between rotors  62  and  64 . The vane assembly  66  is fixed relative the rotation of the rotors  62  and  64  and includes vane  68  extending between an upper platform  70  and a lower platform  72 . Leakage of hot gases through the turbine section  54  is undesirable and therefore features are provided to maintain gas flow between upper and lower platforms  70 ,  72 . 
         [0049]    Each of the airfoils  74  and  76  include upper and lower platforms and outer static shrouds that define the gas flow path. Each of the rotors  62 ,  64  include knife edge seals  78 , and  80  that engage a honeycomb portion  82  that is fixed to the static vane assembly  66 . The knife edges  78  correspond with the honeycomb  82  to seal and contain gas flow within the defined gas path through the high pressure turbine  54 . 
         [0050]    Cooling air indicated by arrows  25  is injected into a space between the fixed vane assembly  66  and the rotor  62 . The cooling air in this space provides an increased pressure that aids in maintaining gas within the desired flowpath and preventing gas from flowing between the vanes and rotating airfoil  74 ,  76 . 
         [0051]    Cooling airflow is shown by the arrow  15  and flows from an outer portion of the turbine case  55  down through openings (not shown) through the vane  68  into a chamber  108  defined below the lower platform  72  of the vane assembly  66 . The chamber  108  includes a plurality of openings  102  ( FIG. 3 ) to allow cooling air  25  to flow forward into the gap between the rotor  62  and the fixed stator assembly  66 . 
         [0052]    Referring to  FIG. 3  with continued reference to  FIG. 2 , the example vane assembly  66  includes an integral one piece ring air seal  84  that receives cooling air that flows through the vanes  68  into the chamber  108 . The air seal  84  is one continuous uninterrupted structure from a wall  98  to the aft most edge  95 . The air seal  84  is attached to and mounted to the lower platform  72 . The air seal  84  extends about the entire circumference of the lower platform  72  and about the axis A. 
         [0053]    The example air seal  84  includes the forward wall  98  that defines a front lip  100  that engages a vane rim  110  that creates a forward seal for defining the cooling air chamber  108 . The forward wall  98  includes a plurality of openings  102  that eject cooling air  25  into the forward gap between the rotor  62  and the vane assembly  66 . 
         [0054]    A ring nut  86  engages threads  104  ( FIG. 4 ) of the air seal  84  to hold the lower platform  72  of the vane assembly  66  between the front lip  100  and a shoulder  120  of the ring nut  86 . The ring nut  86  includes a cavity  122  that corresponds with a slot or groove  125  disposed on the lower platform  72  to define an annular cavity for seal  92 . In this example, the seal  92  comprises a W-shaped seal that biases outward against surfaces of the ring nut  86  and the lower platform  72 . 
         [0055]    The lower platform  72  includes the mount rail  112  that defines the annular groove  125  that corresponds with the cavity  122  defined in the ring nut  86 . The seal  92  is an annular seal that extends about the circumference of the lower platform  72  to provide the desired seal. A second seal  90  is disposed within a groove  106  that is defined in the air seal  84  and a forward surface of the locking nut  86 . In this example, the second seal  90  includes a circular cross-section such as an O-ring or wire seal that is compressed sufficiently to provide the desired sealing features. The combination of the first seal  92  and the second seal  90  provides for the containment of cooling air flow that flows into the cooling chamber  108  defined between the lower platform  72  and the air seal  84 . The first seal  92  and the second seal  90  are fabricated from a seal material including properties compatible with the pressures and temperatures encountered in the high pressure turbine  54 . 
         [0056]    Referring to  FIGS. 4 ,  5 ,  6 ,  7  and  8  with continued reference to  FIG. 3 , the example ring nut  86  includes slots  116  disposed at equally spaced intervals about the circumference of the locking nut  86 . Locking ring segments  88  includes openings  124  that receive tabs  96  of the air seal  84  to fix the locking ring segments  88  relative to the air seal  84 . The locking ring segments  88  includes tabs  126  that bend upward into the slots  116  once the locking nut  86  is tightened to a desired torque valve. The tabs  126  disposed within the slots  116  of the nut  86  prevent rotation of the nut  86  away from the desired locked position. The example locking nut  86  includes threads  118  that correspond with the threads  104  provided on the air seal  84 . 
         [0057]    The example lower platform  72  includes the forward vane rim  110  and the mounting rail  112 . The mounting rail  112  is disposed approximately midway between a fore and aft edges of the lower platform  72 . The example mounting rail  112  abuts the shoulder  120  of the locking ring  86  to bias the vane rim  110  into engagement with the front lip  100  of the air seal  84 . The interface between the front lip  100  and the vane rim  110  provides the sealing required to contain cooling airflow in the chamber  108 . 
         [0058]    The air seal  84  includes a plurality of tabs  94  disposed about the circumference of the air seal  84 . The example tabs  94  are evenly spaced, however, the tabs  94  cold be spaced in any manner about the air seal  84 . A space between the tabs  94  receives lugs  114  on the mounting rail  112  of the lower platform  72 . The lugs  114  received within the space between tabs  94  prevent rotation and maintain a relative circumferential position between the lower platform  72  and the example air seal  84 . As appreciated, although only a few lugs  114  are illustrated, a plurality of lugs  114  are spaced at intervals about the circumference of the mounting rail  112  and are received between tabs  94  within the example air seal  84 . 
         [0059]    Referring to  FIGS. 9 and 10  with continued reference to  FIG. 3 , the example vane assembly  66  includes a plurality of vanes  68  between the upper platform  70  and a lower platform  72 . The lower platform  72  is mounted to the air seal  84  such that cooling airflow can be channeled through the various vanes  68  to the chamber  108  ( FIG. 3 ) defined between the lower platform  72  and the air seal  84 . The example air seal  84  is a continuous ring about the axis A and eliminates complications caused by multiple pieces or segmented structures. 
         [0060]    Accordingly, the example air seal  84  provides a continual seal engagement with the lower platform  72  to provide the desired cooling passages and support the honeycomb structure  82  that engages seal knife edges  78 ,  80  on the rotors  62 ,  64 . The single piece annular locking nut  86  is locked in place by a single, or multiple, segmented lock ring(s)  88  to provide the desired sealing function and connection to the lower platform  72 . 
         [0061]    Although an example embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this disclosure. For that reason, the following claims should be studied to determine the scope and content of this disclosure.