Abstract:
A method of assembling a blade array includes the step of inserting a blade having a platform into a rotor. The blade includes a pocket radially beneath the platform that includes an interference feature. The blade corresponds to one of first and second blades that are scaled versions of one another. The method includes the step of selecting a damper seal, and inserting the damper seal into the pocket. The damper seal corresponds to one of first and second damper seals. The first damper seal cooperates with the interference feature thereby permitting the first damper seal to fully seat within the pocket. The second damper seal is obstructed by the interference feature thereby preventing the second damper seal from fully seating within the pocket.

Description:
BACKGROUND 
       [0001]    This disclosure relates to damper pocket seals for blades used in a turbine blade array, for example. In particular, the disclosure relates to mistake proofing the installation of the damper pocket seals into the blades. 
         [0002]    Damper seals are used to prevent leakage between circumferentially adjacent blade platforms within a stage of a gas turbine engine. The damper seals are arranged in adjacent pockets to block a circumferential gap between the adjacent platforms. Additionally, the damper seals minimize undesired movement between the adjacent blades. 
         [0003]    One type of gas turbine engine may include a core that is a scaled version of another gas turbine engine core. The scaled cores provide different thrust, but rely upon generally the same engine design. As a result, the blades between the scaled versions may have a virtually identical shape such that they are indistinguishable from one another without careful measurement. Typically, the blades and damper seals are provided in kits for a given gas turbine engine such that it is difficult to interchange parts between scaled cores during a maintenance or overhaul procedure. Nonetheless, it still may be possible to insert the damper seal from one engine into its scaled counterpart engine. If such a mistake occurs, the damper seal may fall out during engine operation. 
       SUMMARY 
       [0004]    In one exemplary embodiment, a method of assembling a blade array includes the step of inserting a blade having a platform into a rotor. The blade includes a pocket radially beneath the platform that includes an interference feature. The blade corresponds to one of first and second blades that are scaled versions of one another. The method includes the step of selecting a damper seal, and inserting the damper seal into the pocket. The damper seal corresponds to one of first and second damper seals. The first damper seal cooperates with the interference feature thereby permitting the first damper seal to fully seat within the pocket. The second damper seal is obstructed by the interference feature thereby preventing the second damper seal from fully seating within the pocket. 
         [0005]    In a further embodiment of any of the above, the first and second blades each include an airfoil and a root that are substantially the same shape as one another. 
         [0006]    In a further embodiment of any of the above, the first and second blades each include a platform that are substantially the same, excluding the interference feature. 
         [0007]    In a further embodiment of any of the above, the first blade has a scale factor of 1.1 or less compared to the second blade. 
         [0008]    In a further embodiment of any of the above, the first blade has a scale factor of about 1.04 compared to the second blade. 
         [0009]    In a further embodiment of any of the above, the first and second blades are configured to be used for the same stage of different gas turbine engines that have scaled cores relative to one another. 
         [0010]    In a further embodiment of any of the above, the first and second blades are turbine blades. 
         [0011]    In a further embodiment of any of the above, the method includes the step of inserting another blade into the rotor adjacent to the other blade. The damper seal is inserted into adjacent pockets of the adjacent blades to seal a circumferential gap between adjacent platforms of the adjacent blades. 
         [0012]    In a further embodiment of any of the above, the damper seal includes a generally C-shaped wall having forward and aft ends abutting an inner surface of the pocket. 
         [0013]    In a further embodiment of any of the above, the pocket includes an aft side, and the damper seal includes forward and aft ends. The aft side provides the interference feature such that the aft end is obstructed by the aft side of the pocket. 
         [0014]    In a further embodiment of any of the above, the wall includes a lateral tab. The interference feature corresponds to a protrusion extending into the pocket. The lateral tab of the second damper seal is obstructed by the protrusion. 
         [0015]    In a further embodiment of any of the above, the first damper seal includes a first tab having a first forward edge spaced a first distance from a first forward end. The second damper seal includes a second tab having a second forward edge spaced a second distance from a second forward end. The first and second distances are different than one another. 
         [0016]    In a further embodiment of any of the above, the lateral tab extends radially inwardly from the wall. 
         [0017]    In a further embodiment of any of the above, the damper seal is a stamped steel, and the blade is a nickel alloy. 
         [0018]    In another exemplary embodiment, a blade array includes a rotor, and a blade is supported in the rotor. The blade includes a platform and a pocket arranged radially beneath the platform that includes an interference feature. A correct damper seal is arranged in the pocket and cooperates with the interference feature thereby permitting the correct damper seal to fully seat within the pocket. The interference feature is configured to obstruct an incorrect damper seal thereby preventing the incorrect damper seal from fully seating within the pocket. 
         [0019]    In a further embodiment of any of the above, the correct and incorrect damper seals include a generally C-shaped wall having forward and aft ends abutting an inner surface of the pocket. 
         [0020]    In a further embodiment of any of the above, the pocket includes an aft side, and the correct and incorrect damper seal include forward and aft ends. The aft side provides the interference feature such that the aft end of the incorrect damper seal is obstructed by the aft side of the pocket. 
         [0021]    In a further embodiment of any of the above, the wall of each correct and incorrect damper seal includes a lateral tab. The interference feature corresponds to a protrusion extending into the pocket, and the lateral tab of the incorrect damper seal is obstructed by the protrusion. 
         [0022]    In a further embodiment of any of the above, the correct damper seal includes a first tab having a first forward edge spaced a first distance from a first forward end. The incorrect damper seal includes a second tab having a second forward edge spaced a second distance from a second forward end, and the first and second distances different than one another. 
         [0023]    In a further embodiment of any of the above, the lateral tab extends radially inwardly from the wall. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0024]    The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein: 
           [0025]      FIG. 1  schematically illustrates a gas turbine engine embodiment. 
           [0026]      FIG. 2  is a cross-sectional view through a high pressure turbine section. 
           [0027]      FIG. 3  is a schematic view of adjacent blades having a damper seal installed into adjacent pockets. 
           [0028]      FIG. 4A  illustrates a first version of a first stage turbine blade with a correct damper seal. 
           [0029]      FIG. 4B  illustrates the turbine blade of  FIG. 4A  with an incorrect damper seal. 
           [0030]      FIG. 4C  illustrates a second version of a first stage turbine blade with a correct damper seal. 
           [0031]      FIG. 4D  illustrates the turbine blade of  FIG. 4C  with an incorrect damper seal. 
       
    
    
     DETAILED DESCRIPTION 
       [0032]      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 . 
         [0033]    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. 
         [0034]    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. 
         [0035]    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. 
         [0036]    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. 
         [0037]    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. 
         [0038]    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 . 
         [0039]    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. 
         [0040]    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. 
         [0041]    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. 
         [0042]    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. 
         [0043]    “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. 
         [0044]    “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. 
         [0045]    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. 
         [0046]    Referring to  FIG. 2 , a cross-sectional view through a high pressure turbine section  54  is illustrated. In the example high pressure turbine section  54 , first and second fixed vane arrays  60 ,  62  are axially spaced apart from one another. A first stage array of turbine blades  64  is arranged axially between the first and second fixed vane arrays  60 ,  62 . A second stage array of turbine blades  66  is arranged aft of the second fixed vane array  62 . The first and second stage arrays of turbine blades  64 ,  66 , which are constructed from a nickel alloy, are arranged within a core flow path C and connected to a spool  32 . 
         [0047]    A root  74  of the turbine blade  64  is mounted to the rotor disk  68 . The root  74  supports a platform  76  from which an airfoil extends  78 . The airfoil  78 , which includes leading and trailing edges  82 ,  84 , provides the tip  80  arranged adjacent to a blade outer air seal  70  mounted to a turbine case  72 . A platform  58  of the second fixed vane array  62  is arranged in an overlapping relationship with the turbine blades  64 ,  66 . 
         [0048]    The engine  20  includes a core section that is a scaled version of another engine core section. That is, two engines of different sizes and thrusts generally share the same design such that the core components from one engine are scaled versions of the other engine core components. A first blade of a first core and a second blade of a second core each include an airfoil and a root that are substantially the same shape as one another, although the blades may have slightly different cooling features. However, the differences in cooling features may not be visible or may be subtle. As a result, the turbine blades for the same stages of the cores have a substantially identical shape or external contour. This makes it difficult to discern one core&#39;s components from the other core&#39;s components. In one example, the first blade has a scale factor of 1.1 or less compared to the second blade such that there is a 10% or less size difference between the different blades. In another example, the first blade has a scale factor of about 1.04 compared to the second blade such that there is only about a 4% size difference between the different blades. 
         [0049]    During maintenance or overhaul of an engine, a blade array, shown in  FIG. 3 , is assembled by inserting a blade  64  into a rotor  68  ( FIG. 2 ). Another blade  64  is inserted into the rotor adjacent to the other blade  64  to provide an arrangement shown in  FIG. 3 . The blades  64  each include laterally spaced pressure and suction side pockets  86 ,  88  radially beneath the platform  76 . A circumferential gap  90  is provided circumferentially between the adjacent platforms  76 . A damper seal  92 , which may be stamped steel, is inserted into adjacent pockets  86 ,  88  of the adjacent blades  64  to seal the circumferential gap  90 . 
         [0050]    Like the scaled blades, the damper seals for the same stage of different cores may look alike and be of substantially the same shape. To prevent the incorrect damper seal from being used with the wrong turbine blades, an interference feature, such as protrusion  106 , may be provided in one or both of the pockets  86 ,  88 . The correct damper seal for a given blade cooperates with the interference feature to permit the first damper seal to fully seat within the pocket. The incorrect damper seal is obstructed by the interference feature to prevent the second damper seal from fully seating within the pocket. In this manner, the interference feature ensures that only the correct damper seal can be used for a particular blade, which is shaped substantially the same as a scaled version of that blade. 
         [0051]    Referring to  FIGS. 4A-4C , the damper seal  92  is correct for the blade  64 , and the damper seal  192  is correct for the blade  164 . The damper seals  92 ,  192  includes a generally C-shaped wall  98 ,  198 , respectively. Referring to  FIG. 4A , the wall  98  includes a forward end  100  received in a forward recess  96  of the pocket  86 . An aft end  102  engages an inner surface  94  of the pocket  86  at an aft side  97 . A tab  104  extends laterally and radially inward from the wall  98 . The tab  104  includes forward and aft edges  105 ,  107 . The forward edge  105  is spaced a first distance D 1  from the forward end  100 . The position of the protrusion  106  accommodates the tab  104  to permit the damper seal  92  to fully seat within the pocket  86 . 
         [0052]    Referring to  FIG. 4B , the protrusion  106  prevents installation of the smaller damper seal  192 . The tab  204  includes forward and aft edges  205 ,  207 . The forward edge  205  is spaced a second distance D 2  from the forward end  200 , which is different than the first distance D 1 . Thus, in this example, the placement of the tab  104 ,  204  ensures the proper damper seal is used with the proper blade. 
         [0053]    Referring to  FIG. 4C , the blade  164  includes a platform  176  on root  174  that supports an airfoil  178 . The forward end  200  of the damper seal  192  is received in the forward recess  196 . The aft edge  207  is positioned forward of the protrusion  206 , which accommodates the tab  204  to permit the damper seal  192  to fully seat within the pocket  186 . As shown in  FIG. 4D , the tab  104  is obstructed by the protrusion  206 , preventing the damper seal  92  from being fully seated within the pocket  186 . 
         [0054]    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 reason, the following claims should be studied to determine their true scope and content.