Patent Publication Number: US-11377967-B2

Title: Pre-formed faceted turbine blade damper seal

Description:
BACKGROUND 
     This disclosure relates to a gas turbine engine, and more particularly to a pre-formed damper seal that is used in a gas turbine engine. 
     A gas turbine engine includes a plurality of turbine blades each received in a slot of a turbine disk. The turbine blades are exposed to aerodynamic forces that can result in vibratory stresses. A damper can be located under platforms of adjacent turbine blades to reduce the vibratory response and provide frictional damping between the turbine blades. The damper slides on an underside of the platforms. The damper is made of a material that is dissimilar from the material of the turbine blades. When the vibratory motions of adjacent turbine blades oppose each other (that is, occur out of phase), the damper slides to absorb the energy of vibration. It is usually a stiff slug of metal with rigid features to provide consistent contact with each side of the platform. 
     Additionally, the turbine blades are exposed to hot gasses. An air cavity between a turbine disk and a gas path of a turbine blade may be pressurized with cooling air to protect the turbine disk from high temperatures. A separate seal is often located near the platform to control the leakage of the cooling air into the hot gasses, improving engine performance and fuel efficiency. 
     During assembly of the high pressure turbine rotor, a damper or damper seal sits loosely between neighboring blades. In order for the damper to reach design intent and reach maximum effectiveness, it requires a break-in period to conform to the blade under-platform geometry. This is achieved during the initial engine start-up and operation acceptance testing, where under heat and centrifugal loading, the damper begins to deform and take the shape of the blade under-platform geometry which increases the damping effectiveness and seals the mate-face gap. 
     Accordingly, it is desire to provide a damper or damper seal that reduces the required break in period. 
     BRIEF DESCRIPTION 
     Disclosed is a damper seal for a turbine blade of a gas turbine engine, the damper seal having: an upper portion; a first downwardly curved portion; and a second downwardly curved portion, the first downwardly curved portion and the second downwardly curved portion extend from opposing end regions of the upper portion, the upper portion having a length extending between the opposing end regions of the upper portion and a width transverse to the length, wherein the upper portion is curved along the entire width as it extends along the length. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the width of the upper portion has a constant radius profile running along the entire width as it extends along the length. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the first downwardly curved portion includes a first tab and a second tab each extending in opposing directions with respect to the first downwardly curved portion, and a third tab that extends from the first tab and the second tab of the first downwardly curved portion in the same general direction as the first downwardly curved portion; and the second downwardly curved portion includes a first tab and a second tab each extending in opposing directions with respect to the second downwardly curved portion, and a third tab that extends from the first tab and the second tab of the second downwardly curved portion in the same general direction as the second downwardly curved portion. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, a height of the second downwardly curved portion relative to the upper portion is longer than a height of the first downwardly curved portion relative to the upper portion. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the damper seal is formed from stamped sheet metal. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the damper seal further includes a mistake proofing tab extending from the third tab of the first downwardly curved portion and a mistake proofing opening located in the third tab of the second downwardly curved portion. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the width of the upper portion has a constant radius profile running along the entire width as it extends along the length. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, a height of the second downwardly curved portion relative to the upper portion is longer than a height of the first downwardly curved portion relative to the upper portion. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the damper seal is formed from stamped sheet metal. 
     Also disclosed is a turbine disk of a gas turbine engine having a plurality of turbine blades each of the plurality of turbine blades being secured to the turbine disk, at least one of the plurality of turbine blades having: a root; a platform located between the root and an airfoil of the blade, wherein the platforms of adjacent blades of the disk define a cavity; and a damper seal received in the cavity the damper seal having: an upper portion; a first downwardly curved portion; and a second downwardly curved portion, the first downwardly curved portion and the second downwardly curved portion extend from opposing end regions of the upper portion, the upper portion having a length extending between the opposing end regions of the upper portion and a width transverse to the length, wherein the upper portion is curved along the entire width as it extends along the length, the upper portion being position to cover a mate face gap between platforms of adjacent turbine blades of the disk. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the width of the upper portion has a constant radius profile running along the entire width as it extends along the length. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the first downwardly curved portion includes a first tab and a second tab each extending in opposing directions with respect to the first downwardly curved portion, and a third tab that extends from the first tab and the second tab of the first downwardly curved portion in the same general direction as the first downwardly curved portion; and the second downwardly curved portion includes a first tab and a second tab each extending in opposing directions with respect to the second downwardly curved portion, and a third tab that extends from the first tab and the second tab of the second downwardly curved portion in the same general direction as the second downwardly curved portion. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, a height of the second downwardly curved portion relative to the upper portion is longer than a height of the first downwardly curved portion relative to the upper portion. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the damper seal is formed from stamped sheet metal. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the damper seal further comprises a mistake proofing tab extending from the third tab of the first downwardly curved portion and a mistake proofing opening located in the third tab of the second downwardly curved portion. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the width of the upper portion has a constant radius profile running along the entire width as it extends along the length. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, a height of the second downwardly curved portion relative to the upper portion is longer than a height of the first downwardly curved portion relative to the upper portion. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the turbine disk is a first stage of a high pressure turbine. 
     Also disclosed is a method of damping vibrations between adjoining blades of a gas turbine engine, the method including the steps of: locating a damper seal adjacent to a mate face gap defined by adjacent platforms of blades secured to a disk of the gas turbine engine, the damper seal comprising an upper portion; a first downwardly curved portion; and a second downwardly curved portion, the first downwardly curved portion and the second downwardly curved portion extend from opposing end regions of the upper portion, the upper portion having a length extending between the opposing end regions of the upper portion and a width transverse to the length, wherein the upper portion is curved along the entire width as it extends along the length. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the width of the upper portion has a constant radius profile running along the entire width as it extends along the length. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The following descriptions should not be considered limiting in any way. With reference to the accompanying drawings, like elements are numbered alike: 
         FIG. 1  is a schematic, partial cross-sectional view of a gas turbine engine in accordance with this disclosure; 
         FIG. 2  is a portion of a turbine section of the engine illustrated in  FIG. 1 ; 
         FIG. 3  illustrates a turbine blade secured to a turbine disk; 
         FIG. 4A  illustrates a bottom perspective view of the turbine blade of  FIG. 3 ; 
         FIG. 4B  illustrates a retention nub of the turbine blade the taken along section A-A of  FIG. 4A ; 
         FIG. 5  is a top (partial cross-sectional view) illustrating a damper seal installed between two adjacent turbine blades; 
         FIG. 6  is a cross-sectional side view along lines  6 - 6  of  FIG. 5 ; 
         FIG. 7  is a perspective view of a damper seal in accordance with an embodiment of the present disclosure; 
         FIG. 8  is a top plan view of a damper seal in accordance with an embodiment of the present disclosure; 
         FIG. 9  is a side view of a damper seal in accordance with an embodiment of the present disclosure; 
         FIG. 10  is a partial perspective view illustrating the damper seal secured to a turbine blade; 
         FIG. 11  is a side view illustrating the damper seal secured to a turbine blade; 
         FIG. 12  is a view along lines  12 - 12  of  FIG. 11  when a damper seal is secured to a pair of turbine blades; 
         FIG. 13  is top plan view of a damper seal without a curved upper portion and illustrating initial line contacts of the damper seal with the platforms of adjacent turbine blades; 
         FIG. 14  is top plan view of a damper seal in accordance with an embodiment of the present disclosure and with a curved upper portion, illustrating initial line contacts of the damper seal with the platforms of adjacent turbine blades; 
         FIG. 15  is a superimposed side view illustrating two turbine blades one with a damper seal not pre-formed in accordance with an embodiment of the present disclosure (no curved upper portion) and one with a damper seal preformed in accordance with an embodiment of the present disclosure (curved upper portion); 
         FIG. 15A  is a view along lines  15 A- 15 A of  FIG. 15  when the damper seals are secured to a pair of turbine blades; 
         FIG. 15B  is a view along lines  15 B- 15 B of  FIG. 15  when the damper seals are secured to a pair of turbine blades; 
         FIG. 15C  is a view along lines  15 C- 15 C of  FIG. 15  when the damper seals are secured to a pair of turbine blades; 
         FIG. 15D  is a view along lines  15 D- 15 D of  FIG. 15  when the damper seals are secured to a pair of turbine blades; 
         FIG. 15E  is a view along lines  15 E- 15 E of  FIG. 15  when the damper seals are secured to a pair of turbine blades; and 
         FIG. 16  is an enlarged view of  FIG. 15D . 
     
    
    
     DETAILED DESCRIPTION 
     A detailed description of one or more embodiments of the disclosed apparatus and method are presented herein by way of exemplification and not limitation with reference to the FIGS. Reference is made to U.S. Pat. No. 9,810,075 the contents of which are incorporated herein by reference thereto. 
       FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as a two-spool turbofan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines might include other systems or features. The fan section  22  drives air along a bypass flow path B in a bypass duct, while the compressor section  24  drives air along a core flow path C for compression and communication into the combustor section  26  then expansion through the turbine section  28 . Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. 
     Although depicted as a turbofan gas turbine engine in the disclosed non-limiting embodiment, 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 or geared turbofan architectures. 
     The fan section  22  drives air along a bypass flowpath B while the compressor section  24  drives air along a core flowpath C for compression and communication into the combustor section  26  then expansion through the turbine section  28 . 
     The 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. 
     The low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a low pressure compressor  44  and a low pressure turbine  46 . The inner shaft  40  is connected to the fan  42  through 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 compressor  52  and a high pressure turbine  54 . 
     As shown in  FIG. 2 , the high pressure turbine  54  includes a first stage  70  and a second stage  72 . The first stage  70  includes a static vane  66 A and plurality of turbine blades  68 A. The second stage  72  includes a static vane  66 B and a plurality of turbine blades  68 B. 
     A combustor  56  is arranged between the high pressure compressor  52  and the high pressure turbine  54 . 
     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 . 
     The inner shaft  40  and the outer shaft  50  are concentric and rotate via bearing systems  38  about the engine central longitudinal axis A, which is collinear with their longitudinal axes. 
     The core airflow C is compressed by the low pressure compressor  44 , then the high pressure compressor  52 , mixed and burned with fuel in the combustor  56 , then expanded over the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  58  includes airfoils  60  which are in the core airflow path. The turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. 
     The engine  20  is in one example a high-bypass geared aircraft engine. In a further example, the engine  20  bypass ratio is greater than about six (6:1) with an example embodiment being greater than ten (10:1). The geared architecture  48  is an epicyclic gear train (such as a planetary gear system or other gear system) with a gear reduction ratio of greater than about 2.3 (2.3:1). The low pressure turbine  46  has a pressure ratio that is greater than about five (5:1). The low pressure turbine  46  pressure ratio is pressure measured prior to inlet of low pressure turbine  46  as related to the pressure at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. 
     In one disclosed embodiment, the engine  20  bypass ratio is greater than about ten (10:1), and the fan diameter is significantly larger than that of the low pressure compressor  44 . The low pressure turbine  46  has a pressure ratio that is greater than about five (5:1). The geared architecture  48  may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.5 (2.5:1). It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans. 
     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 (11,000 meters). The flight condition of 0.8 Mach and 35,000 feet (11,000 meters), with the engine at its best fuel consumption, also known as bucket cruise Thrust Specific Fuel Consumption (“TSFC”). TSFC is the industry standard parameter of lbm of fuel being burned divided by 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 1.45. 
     “Low corrected fan tip speed” is the actual fan tip speed in feet per second 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 feet per second (350.5 meters per second). 
       FIG. 2  illustrates the turbine section  28 . The turbine section  28  includes turbine discs  61  that each rotate about the axis A. In the first stage  70  of the high pressure turbine  54 , a plurality of turbine blades  68 A are mounted on a turbine disk  61 . In the second stage  72  of the high pressure turbine  54 , a plurality of turbine blades  68 B are mounted on another turbine disk  61 . 
       FIG. 3  illustrates a perspective view of a turbine blade  68 A partially installed in a turbine disk  61 . In one example, the turbine blades  68 A are made of a nickel alloy. The turbine disk  61  includes a plurality of slots  74  separated by turbine disk lugs  76 . The slot may be in the shape of a dovetail, a fir tree shape, or some other configuration. The turbine blade  68 A includes a root  78  that is received in one of the plurality of turbine disk slots  74  of the turbine disk  61 , a platform  80  including retention shelves  82  and buttresses  93 , and an airfoil  84 . The platform  80  has a length L. The airfoil  84  has a leading edge  86  and a trailing edge  88 . A neck cavity  90  is defined between the platform  80  and the retention shelf  82 . A buttress  93  is also located in the neck cavity  90  and under the platform  80  of each turbine blade  68 A. The buttress  93  is a support structure that connects the platform  80  to the retention shelf  82 . Although  FIG. 3  illustrates a single turbine blade  68 A a plurality of turbine blades are secured to the turbine disk  61 . For convenience, only a portion of the turbine disk  61  is illustrated. 
     Hot gasses flow along a hot gas flow path E. The neck cavity  90  between adjacent turbine blades  68 A is pressurized with a flow of cooling air F to protect the turbine discs  61  from the hot gasses in the hot gas flow path E. 
       FIG. 4A  illustrates a lower perspective view of a turbine blade  68 A to be located in the first stage  70  of the high pressure turbine  54 , for example. The neck cavity  90  includes a retention nub  92  located on a lower surface  91  of the platform  80 . 
       FIG. 4B  illustrates a cross-sectional view of the retention nub  92  taken along section  4 B- 4 B of  FIG. 4A . The retention nub  92  includes a first surface  94  and a second surface  96 . An angle J defined between the first surface  94  and a horizontal plane is approximately 30 to 60 degrees. An angle K defined between the second surface  96  and the horizontal plane is approximately 45 to 85 degrees. 
       FIGS. 5 and 6  illustrate a damper seal  98  installed between adjacent turbine blades  68 A 1  and  68 A 2 . The damper seal  98  is located in a neck cavity  90  of the turbine blades  68 A 1  and  68 A 2 . The damper seal  98  is located in an under-platform pocket  97  depicted by the dashed lines in  FIG. 5 . The damper seal  98  is located under the platforms  80  and above the retention shelves  82  of the adjacent blades  68 A 1  and  68 A 2  and spans a space or mate face gap  100  between a leading edge  99  and a trailing edge  101  of the platforms  80  of the turbine blades  68 A 1  and  68 A 2 . The retention nub  92  of the turbine blade  68 A 2  is received in an opening  120  of the damper seal  98 . 
     By employing a damper seal  98  that combines the features of a damper and a seal into a single component, the number of parts and the weight is reduced. Additionally, the assembly process is simplified by requiring only one component to be installed between adjacent turbine blades  68 A. 
     The damper seal  98  imposes a normal load on the turbine blades  68 A. The resulting frictional force created by the normal load produces damping, reducing a vibratory response. The damper seal  98  prevents the cooling air F from leaking from the neck cavity  90  of the turbine blades  68 A and into the hot gas flow path E along arrows G (shown in  FIG. 3 ). 
       FIG. 6  illustrates a side view of the turbine blade  68 A with the damper seal  98  installed in the neck cavity  90 . The retention nub  92  of the turbine blade  68 A is received in the opening  120  of the damper seal  98 . 
     In the past and during assembly of the high pressure turbine rotor, the damper seal  98  sits loosely between neighboring blades. In order for the damper seal  98  to reach its design intent and reach its maximum effectiveness, a break-in period is typically required to conform to the damper seal  98  to the blade under-platform geometry. In the past, this is achieved during the initial engine start-up and operation acceptance testing, where the damper seal  98  is subject to heat from the main gas path flow (arrows  122 ), which is applied to the damper seal  98  through conductive paths (arrows  124 ) of the blade  68 A. In addition, centrifugal loading in the direction of arrow  126  is also applied to the damper seal  98 . As such, the damper seal  98  moves radially outward and begins to deform and take the shape of the blade under-platform geometry which increases the damping effectiveness and seals the mate-face gap  100 . 
     In accordance with an embodiment of the present disclosure, a damper seal  98  is provided that reduces the aforementioned break-in period and allows the damper seal  98  to reach its effectiveness quicker. 
     Referring now to  FIGS. 7-9 , a damper seal  98  in accordance with the present disclosure is illustrated. The damper seal  98  spans the space or mate face gap  100  (as shown in  FIG. 5 ) between platforms  80  of adjacent turbine blades  68 A in the first stage  70  of the high pressure turbine  54  to provide both damping and sealing and prevent the leakage of the cooling air F. The damper seal  98  imposes a normal load on the adjacent turbine blades  68 A due to centrifugal force. The resulting frictional force created by the normal load produces damping to reduce a vibratory response. The damper seal  98  prevents the cooling air F in the neck cavity  90  from leaking into the hot flow gas path E along arrows G (shown in  FIG. 3 ). 
     In one non-limiting embodiment, the damper seal  98  is formed from stamped sheet metal. The damper seal  98  can also be formed by direct metal laser sintering. Other manufacturing methods are possible. 
     The damper seal  98  has an upper portion  130 . A first downwardly curved portion  132  and a second downwardly curved portion  134  that extend from opposing end regions of the upper portion  130 . In one example, relative to the upper portion  130  of the damper seal  98 , a height H 2  of the second downwardly curved portion  134  is longer than a height H 1  of the first downwardly curved portion  132 . 
     An end region of the first downwardly curved portion  132  includes a first tab  136  and a second tab  138  that each extend in opposing directions with respect to the first downwardly curved portion  132 . A third tab  140  extends from tabs  136  and  138  and also extends in the same general direction as the first downwardly curved portion  132 . The third tab  140  provides sealing to the neck cavity  90  and prevents the passage of the cooling air F into the hot gas flow path E. 
     An end region of the second downwardly curved portion  134  includes a first tab  142  and a second tab  144 . A third tab  146  extends from tabs  142  and  144  and also extends in the same general direction as the second downwardly curved portion  134 . The third tab  146  provides sealing to the neck cavity  90  and prevents the passage of the cooling air F into the hot gas flow path E. 
     Tabs  136 ,  138 ,  142  and  144  prevent rocking of the damper seal  98  when it is between platforms  80  of adjacent turbine blades  68 A. 
     In accordance with an embodiment of the present disclosure, the upper portion  130  of the damper seal  98  is substantially curved in the direction of arrows  148 . As such, the upper portion  130  is generally curved along its width W. In one embodiment, the upper portion  130  is curved along its entire width W. As illustrated herein the width W extends in the same directions as tabs  136 ,  138 ,  142  and  144 . In other words, the width W of the upper portion  130  is transverse to the length L of the upper portion or the length L of the upper portion extends along a major axis of the upper portion  130  and the width W extends along a minor axis of the upper portion  130 . 
     In one non-limiting exemplary embodiment, the damper seal shape of the upper portion  130  or an outboard mating surface of the upper portion  130  that contacts the under-side of the blade platforms will have a constant radius profile running from leading to trailing ends of the underside of the blade/platform until transitioning to the first downwardly curved portion  132  and the second downwardly curved portion  134  which include the tabs  136 ,  138 ,  140 ,  142 ,  144 ,  146 . 
       FIG. 10  is a partial perspective view illustrating the damper seal  98  secured to a turbine blade  68 A. 
     Referring now to at least  FIGS. 11-16  differences between a damper seal  150  without a curved upper portion  130  and a damper seal  98  with a curved upper portion  130  in accordance with the present disclosure is illustrated. 
     In  FIG. 11  is a side view of a turbine blade  68 A with a damper seal is illustrated. In  FIG. 13  a top plan view of the damper seal  150  without a curved upper portion  130  is illustrated.  FIG. 13  illustrates initial lines of contact  152  of the damper seal  150  with an underside  154  of platforms  80  of adjacent turbine blades  68 A prior to the aforementioned break-in period. 
     In contrast and in  FIG. 14 , a top plan view of the damper seal  98  without a curved upper portion  130  is illustrated.  FIG. 14  also illustrates initial lines of contact  152  of the damper seal  150  with an underside  154  of platforms  80  of adjacent turbine blades  68 A prior to the aforementioned break-in period. 
     As clearly illustrated, the initial lines of contact  152  of the damper seal  98  are much closer to each other than the initial lines of contact  152  of the damper seal  150 . Also illustrated in  FIGS. 13 and 14  is the location of the mate face gap  100  on the upper portion  130  of damper seals  98  and  150  when they are initially located between a pair of turbine blades  68 A prior to the aforementioned break-in period. This location is illustrated by pair of lines  156 . Also and as illustrated in  FIGS. 13 and 14 , the initial lines of contact  152  of the damper seal  98  are much closer to the mate face gap  100 . 
     Referring now to  FIG. 12 , a view along lines  12 - 12  of  FIG. 11  is illustrated when the damper seal is located underneath the platforms  80  of adjacent turbine blades  68 A prior to the aforementioned break-in period. In  FIG. 12 , the locations of both damper seal  98  with a curved upper portion  130  and damper seal  150  without a curved upper portion  130  are superimposed on each other. As clearly illustrated, the damper seal  98  with the curved upper portion  130  pre-conformed to the contours of the underside  154  of the platforms  80  of the turbine blades  68 A will have a greater surface area in direct contact with the underside  154 . 
       FIG. 15  is a side view illustrating two turbine blades superimposed on each other, one with a damper seal  150  (not pre-formed in accordance with an embodiment of the present disclosure) and one with a damper seal  98  (preformed in accordance with an embodiment of the present disclosure). 
       FIG. 15A  is a view along lines  15 A- 15 A of  FIG. 15  when the damper seals  98 ,  150  are secured to a pair of turbine blades  68 A.  FIG. 15B  is a view along lines  15 B- 15 B of  FIG. 15  when the damper seals  98 ,  150  are secured to a pair of turbine blades  68 A.  FIG. 15C  is a view along lines  15 C- 15 C of  FIG. 15  when the damper seals  98 ,  150  are secured to a pair of turbine blades  68 A.  FIG. 15D  is a view along lines  15 D- 15 D of  FIG. 15  when the damper seals  98 ,  150  are secured to a pair of turbine blades  68 A.  FIG. 15E  is a view along lines  15 E- 15 E of  FIG. 15  when the damper seals  98 ,  150  are secured to a pair of turbine blades  68 A.  FIGS. 15A-15E  clearly illustrate that a greater surface area of upper portion  130  of damper seal  98  contacts the underside  154  than the upper portion  130  of damper seal  150 . 
       FIG. 16  is an enlarged view of  FIG. 15D . As clearly illustrated, the initial lines of contact  152  for damper seal  98  in comparison to damper seal  150  are moved towards the damper seal center or mate face gap center  100 . This results in an increased stiffness of the damper seal. Reduction in the distance L between the initial points of contact  152  of the damper seal  150  and the initial points of contact  152  of the damper seal  98  helps with this increased stiffness of the damper seal. 
     By providing a damper seal  98  with a curved upper portion or curved central portion  130  and as discussed above, this reduces break-in period requirements, which achieves early damper seal effectiveness, and thus reduces overall engine testing time. As such and in order to reduce an overall initial engine testing time, a pre-formed damper seal with a curved upper portion is needed. 
     In contrast to the flat outboard surface or upper portion  130  provided in damper seal  150 , the radial profile of the damper seal  98  shifts the initial contact zones on both blades towards the center of the platform gap or mate face gap. As such, this radial profile or curved upper portion allows the damper  98  to conform to geometry quickly as it can rotate tangentially (relative to the rotor axis) to accommodate the total tolerance stack of the assembled hardware (e.g., adjacent blades  68 A). 
     Ensuring better initial contact between the damper seal and the neighboring blades  68 A as well as the ability to quickly center with the tolerance stack range of the assembly achieves a reduction in engine break-in period requirements and thus, reduces overall engine testing time. 
     Referring now to at least  FIGS. 8, 9 and 14 , the damper seal  98  may comprise a mistake proofing tab  170  extending from the third tab  140  of the first downwardly curved portion  132  and a mistake proofing opening or hole  172  located in the third tab  146  of the second downwardly curved portion  134 . Mistake proofing tab  170  and mistake proofing opening or hole  172  will help ensure that the damper seal if properly located in between adjacent turbine blades  68 A as tab  170  and/or opening  172  will prevent proper insertion of the damper seal between adjacent blades  68 A by for example having tab  170  engage a feature of the turbine blades  68 A and/or a protrusion being received within opening or hole  172 . Although a mistake proofing tab  170  and a mistake proofing opening or hole  172  are illustrated in at least  FIGS. 8, 9 and 14 , it is contemplated that the damper seal  98  can be made without mistake proofing tab  170  and mistake proofing opening or hole  172 . In other words, at least one embodiment of the present application does not have or require the mistake proofing tab  170  and/or the mistake proofing opening or hole  172 . 
     The term “about” is intended to include the degree of error associated with measurement of the particular quantity based upon the equipment available at the time of filing the application. For example, “about” can include a range of ±8% or 5%, or 2% of a given value. 
     The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the present disclosure. As used herein, the singular forms “a”, “an” and “the” are intended to include the plural forms as well, unless the context clearly indicates otherwise. It will be further understood that the terms “comprises” and/or “comprising,” when used in this specification, specify the presence of stated features, integers, steps, operations, elements, and/or components, but do not preclude the presence or addition of one or more other features, integers, steps, operations, element components, and/or groups thereof. 
     While the present disclosure has been described with reference to an exemplary embodiment or embodiments, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the present disclosure. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the present disclosure without departing from the essential scope thereof. Therefore, it is intended that the present disclosure not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this present disclosure, but that the present disclosure will include all embodiments falling within the scope of the claims.