Patent Publication Number: US-11021980-B2

Title: Gas turbine engine turbine vane ring arrangement

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This disclosure is a continuation of U.S. patent application Ser. No. 14/907,003, filed on Jan. 22, 2016, which is a National Stage Entry of PCT Application No. PCT/US2014/043110, filed on Jun. 19, 2014, which claims priority to U.S. Provisional Application No. 61/859,844, which was filed on Jul. 30, 2013. 
    
    
     BACKGROUND 
     This disclosure relates to a gas turbine engine vane arrangement, for example, in a turbine section. More particularly, the disclosure relates to a ring used to secure circumferentially arranged vanes to one another in, for example, a mid-turbine frame. 
     Gas turbine engines typically include a compressor section, a combustor section and a turbine section. During operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases are communicated through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads. 
     Both the compressor and turbine sections may include alternating series of rotating blades and stationary vanes that extend into the core flow path of the gas turbine engine. For example, in the turbine section, turbine blades rotate and extract energy from the hot combustion gases that are communicated along the core flow path of the gas turbine engine. The turbine vanes, which generally do not rotate, guide the airflow and prepare it for the next set of blades. 
     A mid-turbine frame is arranged axially between high and low turbine sections. One type of mid-turbine frame uses discrete vanes secured circumferentially to one another to provide an integral annular vane pack. The vane pack is reinforced using multiple rings secured to the vanes. An edge of the vane pack is disposed within a pocket of rotating blades of an adjacent turbine stage to provide a seal at the inner flow path. The reinforcement ring at this location is spaced from and outside of the pocket. 
     SUMMARY 
     In one exemplary embodiment, a vane pack for a gas turbine engine includes an annular arrangement of vanes. A ring is secured around the vanes and extends proud of an axial end of the vanes. 
     In a further embodiment of any of the above, the annular arrangement includes vane segments secured to one another circumferentially. 
     In a further embodiment of any of the above, the ring is secured to the vanes by mechanical elements. 
     In a further embodiment of any of the above, the mechanical elements include at least one of a braze, a weld and fasteners. 
     In a further embodiment of any of the above, the ring is secured to the vanes by an interference fit. 
     In a further embodiment of any of the above, the ring and the vanes include interlocking features that engage one another and are configured to prevent relative axial movement between the ring and the vanes. 
     In a further embodiment of any of the above, the ring is secured to an inner platform. 
     In a further embodiment of any of the above, the axial end is a leading edge. 
     In a further embodiment of any of the above, the ring provides an end configured to provide a seal with an adjacent rotating component. 
     In a further embodiment of any of the above, the end includes one of an annular pocket and an annular lip. 
     In another exemplary embodiment, a gas turbine engine includes a compressor section. A combustor is fluidly connected downstream from the compressor section. A turbine section is fluidly connected downstream from the combustor and includes high and low pressure turbine sections. A vane pack is arranged in one of the compressor or turbine sections. The vane pack includes a ring secured around an annular arrangement of vanes and extends proud of an axial end of the vanes to an end. The end interleaves with an adjacent rotating component to provide a seal. 
     In a further embodiment of any of the above, the vane pack is arranged in the turbine section. 
     In a further embodiment of any of the above, the rotating components include one of a pocket and a lip. The ring provides the other of the pocket and the lip. The lip is arranged in the pocket to provide the seal. 
     In a further embodiment of any of the above, the stage of rotating blades is provided by the high pressure turbine section. The vane pack provides a mid-turbine frame. 
     In a further embodiment of any of the above, the engine static structure supports a sealing ring that engages the reinforcement ring. 
     In a further embodiment of any of the above, the annular arrangement includes vane segments secured to one another circumferentially. 
     In a further embodiment of any of the above, the vanes are discrete from one another and hung from engine static structure. 
     In a further embodiment of any of the above, the reinforcement ring is secured to the vanes by at least one of a mechanical element and an interference fit. 
     In a further embodiment of any of the above, the reinforcement ring is secured to an inner platform. 
     In a further embodiment of any of the above, the axial end is a leading edge. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein: 
         FIG. 1  schematically illustrates a gas turbine engine embodiment. 
         FIG. 2  is an exploded perspective view of a mid-turbine frame vane pack. 
         FIG. 3  is a cross-sectional view of the mid-turbine frame vane pack arranged between the high and low turbine sections. 
         FIG. 4  is an enlarged view of a reinforcing ring of the vane pack arranged adjacent to rotating blades. 
         FIG. 5  is an enlarged view of another ring configuration adjacent to another blade. 
         FIG. 6  is an enlarged, broken view of another ring configuration secured to another vane arrangement. 
     
    
    
     DETAILED DESCRIPTION 
       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 temperature 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 . 
     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 with or without 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. 
     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 X 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. 
     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 first) 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 second) 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 X. 
     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. 
     The example low pressure turbine  46  has a pressure ratio that is greater than about five (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. 
     A mid-turbine frame  57  of the engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The mid-turbine frame  57  further supports bearing systems  38  in the turbine section  28  as well as setting airflow entering the low pressure turbine  46 . 
     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  57  includes vanes  59 , which are in the core airflow path and function as an inlet guide vane for the low pressure turbine  46 . Utilizing the vane  59  of the mid-turbine frame  57  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  57 . 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. 
     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. 
     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. 
     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. 
     “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. 
     “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)] 0.5 . The “Low corrected fan tip speed”, as disclosed herein according to one non-limiting embodiment, is less than about 1150 ft/second. 
     An exploded view of a vane pack  60  is illustrated in  FIG. 2 . The vane pack  60  provides a gas path portion of the mid-turbine frame  57  in one example gas turbine engine. The vane pack may be provided in other sections of the engine  20 , such as the compressor section and other areas of the turbine section. In one example, the vane pack  60  is provided by multiple vane segments  62  circumferentially arranged and secured with respect to one another to provide an annular structure. Each vane  62  includes an inner and outer platform  64 ,  66  joined to one another by the vane airfoil  59 . 
     In one example, the vanes  62  are constructed from a nickel alloy and brazed to one another. Forward inner and outer diameter rings  68 ,  70  and aft inner and outer diameter rings  72 ,  74  are secured to the vane segments  62  for structural reinforcement. In one example, the rings  68 ,  70 ,  72 ,  74  are secured to the vane segments  62  by brazing. 
     Although multiple discrete circumferential vane segments are shown in  FIG. 2 , it should be understood that a cast and/or machined structure may provide clusters of vanes or all of the vanes and associated inner and outer platforms in a single, unitary annular configuration. 
     In one example, the vane airfoils  59  provide a hollow cavity  76  that accommodate oil lines, structural members, wires, bleed air conduits or other elements that may be passed from the outer portion of the engine static structure  36  to an inner portion. 
     Referring to  FIGS. 2 and 3 , the vanes  62  includes a boss  78  that receives a bushing  79 . A pin  80  is secured to the engine static structure  36  and received by the bushing  79  to locate the vane pack  60  with respect to the engine static structure  36 . Engine static structure  36  supports one of the bearings  38  mounted to the high pressure turbine shaft  32 . 
     First, second, third and fourth sealing rings  82 ,  84 ,  86 ,  88  are supported by the engine static structure  36  and respectively engage the forward inner and outer diameter ring  68 ,  70  and the aft inner and outer diameter ring  72 ,  74  to seal the flow path gases within the core flow path C from other components. 
     As shown in  FIGS. 3 and 4 , the high pressure turbine section  54  includes an aft stage blade  90 , which includes a pocket  94 . The forward inner diameter ring  68  includes an end  100  secured around the vanes  60  that extends proud of an axial end of the vanes, in the example the leading edge  99  of the inner platform  64 . The end  100  provides an annular lip that is arranged at least partially within the pocket  94  and radially beneath the blade platform  96 . The forward inner diameter ring  68  is secured to the main segments  62  at an interface  98  by brazing, for example, if one or more of the vane segments  62  begins to separate from the forward inner diameter ring  68 , the vane segments  62  will not physically interfere with the rotation of the aft stage blade  90 . 
     The low pressure turbine section  46  includes a forward stage blade  92 . In the example, the aft inner diameter ring  72  does not extend beyond the vane segment  62  as does the forward inner diameter ring  68 , since there is more clearance between the vane segments  62  and the forward stage blade  92 . However, an end of the forward outer diameter ring  70  and aft inner and outer diameter rings  72 ,  74  may extend axially beyond the vane segments  62  if desired where running clearances are tighter. 
     In the example shown in  FIG. 5 , the blade  190  includes a platform  196  having a lip received in an annular pocket  194  provided by the end  200  of the ring  168 , which is secured to the vane  162 . Thus, it should be understood that the platform and end may include any geometry suitable for providing a seal between the blade and vane. 
     Referring to  FIG. 6 , discrete single vanes or cluster of vanes is shown at  290  and is supported or hung relative to the engine static structure  36  by an attachment feature, such as a hook  291 . The vane segment  262  and ring  268  include complementary shaped interlocking features to prevent the ring  268  from migrating axially toward the blade. In the example, one of the interlocking features is a groove  269  and the other of the interlocking features is a tab  271 . In another example, the interlocking features may be provided by conical surfaces that provide a wedge-like interface. The interlocking features may obviate the need for any additional mechanical securing elements, such as brazing and/or fasteners. 
     Although example embodiments have been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of the claims. For that and other reasons, the following claims should be studied to determine their true scope and content.