Abstract:
A blade has an airfoil extending radially outwardly of a dovetail. The dovetail has edges that will be at circumferential sides of the blade when the blade is mounted within a rotor. A bottom surface of the dovetail will be radially inward when the rotor blade is mounted in a rotor, and is formed such that a circumferentially central portion of the bottom surface is radially thicker than are circumferential edges. A fan and a gas turbine engine are also described.

Description:
This application is a United States National Phase of PCT Application No. PCT/US2013/031868 filed on Mar. 15, 2013 which claims priority to U.S. Provisional Application Ser. No. 61/703,498, filed on Sep. 20, 2012. 
    
    
     BACKGROUND 
     This disclosure relates to dovetail geometry for fan blades used in individually bladed rotors of gas turbine engines. 
     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. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines. 
     The high pressure turbine drives the high pressure compressor through an outer shaft to form a high spool, and the low pressure turbine drives the low pressure compressor through an inner shaft to form a low spool. The fan section may also be driven by the low inner shaft. A direct drive gas turbine engine includes a fan section driven by the low spool such that the low pressure compressor, low pressure turbine and fan section rotate at a common speed in a common direction. 
     A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a reduced speed such that both the turbine section and the fan section can rotate at closer to optimal speeds. 
     Individual fan blades are mounted within a hub or rotor driven by the gear assembly. The configuration and geometry of the fan blades balance propulsive efficiency with durability and fatigue requirements. 
     A dovetail of each fan blade is received in a correspondingly shaped slot in the fan rotor. The dovetail provides a bearing surface which reacts against a load surface of the slot. It is desirable to provide a small dovetail for weight savings. However, the dovetail is sized to provide sufficient strength to retain the fan blades in the fan rotor throughout engine operation and during a variety of conditions. Several prior art dovetails have provided a ratio of neck width to a vertical bearing surface height of 1.19, 1.35 and 1.49. A prior art ratio of fan radius to vertical bearing surface height of 45, 47 and 56 has been provided, and a prior art ratio of dovetail width to vertical bearing surface height of 2.31, 2.93 and 2.99 has been provided. 
     Although geared architecture have improved propulsive efficiency, turbine engine manufacturers continue to seek further improvements to engine performance including improvements to thermal, transfer and propulsive efficiencies. 
     SUMMARY 
     In one exemplary embodiment, a fan blade for a gas turbine engine includes an airfoil that defines a pressure and suction side. A neck supports the airfoil and interconnects to a dovetail by opposing radii. The dovetail includes a bearing surface tangent to one of the radii at a tangent point. The dovetail has a dovetail width and a vertical bearing surface height from the tangent point to a bottom of the dovetail opposite the neck. A first ratio of the dovetail width to the vertical bearing surface height is in a range of substantially between 2.32 to 2.90. 
     In a further embodiment of the above, the first ratio is in a range of 2.35 to 2.70. 
     In a further embodiment of any of the above, the first ratio is about 2.40. 
     In a further embodiment of any of the above, the neck includes a neck width. A second ratio of the neck width to the vertical bearing surface height is in a range of 0.95 to 1.18. 
     In a further embodiment of any of the above, the second ratio is in a range of about 1.10. 
     In a further embodiment of any of the above, the vertical bearing height is in the range of 0.675 inch to 0.775 inch. 
     In another exemplary embodiment, a fan section of a gas turbine engine includes a fan rotor that has a slot and a load surface. An airfoil defines a pressure and suction side. A neck supports the airfoil and interconnects to a dovetail by opposing radii. The dovetail is received in the slot. The dovetail includes a bearing surface tangent to one of the radii at a tangent point. The dovetail has a dovetail width and a vertical bearing surface height from the tangent point to a bottom of the dovetail opposite the neck. The first ratio of the dovetail width to the vertical bearing surface height is in a range of substantially between 2.32 to 2.90. 
     In a further embodiment of any of the above, the bearing surface is configured to react against the load surface during engine operation. A spacer is provided between the bottom and the slot. 
     In a further embodiment of any of the above, the neck includes a neck width. A second ratio of the neck width to the vertical bearing surface height is in a range of 0.95 to 1.18. 
     In a further embodiment of any of the above, a fan section of a gas turbine engine includes a fan radius. A third ratio of the fan radius to the vertical bearing surface height is in a range of 30 to 44. 
     In a further embodiment of any of the above, the third ratio is about 39. 
     In a further embodiment of any of the above, the vertical bearing height is in the range of 0.675 inch to 0.775 inch. 
     In one exemplary embodiment, a gas turbine engine includes a fan section that has a plurality of fan blades that are mounted to a fan rotor that is rotatable about an axis. A gas turbine engine also includes a compressor section and a combustor that are in fluid communication with one another. A turbine section is in fluid communication with the combustor. A fan drive gear system module is coupled to the turbine section for rotating the fan about the axis. A neck supports the fan blades and is interconnected to a dovetail by opposing radii. The dovetail is received by the fan rotor. The dovetail includes a bearing surface tangent to one of the radii at a tangent point. The dovetail has a dovetail width and a vertical bearing surface height from the tangent point to a bottom of the dovetail opposite the neck. A first ratio of the dovetail width to the vertical bearing surface height is in a range of substantially between 2.32 to 2.90. 
     In a further embodiment of any of the above, the neck includes a neck width. A second ratio of the neck width to the vertical bearing surface height is in a range of 1.18 to 0.95. 
     In a further embodiment of any of the above, a gas turbine engine includes a fan radius. A third ratio of the fan radius to the vertical bearing surface height is in a range of 30 to 44. 
    
    
     
       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. 2A  is a schematic view of an example prior art fan blade. 
         FIG. 2B  is an end view of the fan blade shown in  FIG. 2A  received in a fan rotor. 
         FIG. 3  is a schematic end view of an example fan blade dovetail according to the disclosure. 
     
    
    
     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 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 . 
     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. 
     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. 
     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 A. 
     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 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  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 . 
     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. 
     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. 
     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. 
     A fan blade  42  is illustrated in  FIG. 2A  having an airfoil  64  extending radially outwardly from a dovetail  62 . A leading edge  66  and a trailing edge  68  define the forward and rear limits of the airfoil  64 . The airfoil  64  has a suction side  70  and a pressure side  72 . The dovetail  62  includes ends  74  and  76  that are found at ends of the overall blade  42  associated with the leading edge  66  and trailing edge  68 . 
     As shown in  FIG. 2B , a fan rotor  78  receives the dovetail  62  in slot  80 , shown schematically, to mount the fan blade  42  with the airfoil  64  extending radially outwardly. A spacer  81  is provided within the slot  80  between the fan rotor  78  and the dovetail  62 . The rotor  78  carries the fan blade  42  as the rotor is rotationally driven. 
     An aluminum hybrid metallic hollow fan blade includes dovetails to secure the blade into the disk or rotor  78 . It is desired to minimize stress for low cycle and high cycle fatigue in addition to improve impact capability under bird strike and fan blade out loads. The disclosed tall dovetail and its desired ratios ( FIG. 3 ) provide more stiffness and stress reduction than a typical dovetail design ( FIG. 2 ). The disclosed dovetail proposal reduces stresses and maximizes stiffness for aluminum and applies to any hybrid metallic or singular metallic fan blade (aluminum, titanium, etc.). 
     The disclosed dovetail design geometry proposes to minimize stress in the dovetail  84  and reduce weight in the heavier (since material is denser) disk or rotor  78 . The current dovetail design has a short depth, or vertical bearing surface height  82 , as shown in  FIG. 3 . 
     Referring to  FIG. 3 , the slot  80  has a radial slot height  86 . A neck  89  supports the airfoil  88  interconnected to the dovetail  84  by opposing radii  98 . The dovetail  84  includes a bearing surface  90  that is configured to react against a load surface  110  of the slot  80  during engine operation. The prior art dovetail  62  includes a stress area  65  provided by a plane  97  that is normal to a tangent point  96  of the bearing surface  90  to the radius  98 . The stress area  92  of the example dovetail  94  is provided by a plane  99  that is normal to a tangent point  100  of the bearing surface  90  to the radius  98 . As can be appreciated from  FIG. 3 , the stress area  92  is significantly larger than the stress area  97 . Since the dovetail  84  is larger than the prior art dovetail  62 , a smaller spacer  87  is used between a bottom of the dovetail  84  and the slot  84  for the same size slot. 
     The prior art stress area  65  has a vertical bearing surface height  82  from the tangent point  96  to the bottom of the dovetail  62 . The vertical bearing surface height  104  from the tangent point  100  to the bottom of the dovetail  84  is significantly larger. The dovetail width  108  is the same between the prior art dovetail  62  and the disclosed dovetail  84  for the same size slot. 
     The dovetail  84  has a first ratio of the dovetail width  108  to the vertical bearing surface height  104  in the range of 2.32 to 2.90. In one example, the first ratio is in a range of 2.35 to 2.70. In another example, the first ratio is about 2.40, or 2.40+/−0.05. 
     The neck  89  includes a neck width  102 . A second ratio of the neck width  102  to the vertical bearing surface height  104  is in the range of 0.95 to 1.18. In another example, the second ratio is in a range of about 1.10, or 1.10+/−0.05. The vertical bearing height is in the range of 0.675 in (17.15 mm) to 0.775 in (19.69 mm). 
     The engine  20  has a fan radius R, as shown in  FIG. 1 . A third ratio of the fan radius to the vertical bearing surface height  104  is in the range of 30 to 44. In one example, the third ratio is about 39, or 39+/−5. 
     The tall dovetail  84  increases stiffness and increases the life of the part by decreasing stresses. The tall dovetail  84  could also be made shorter in the radial direction with an optimized increase in radial height resulting in less system weight. The bearing surface  90  experiences high stress levels due to tooth bending of the dovetail  84 . Reducing tooth bending provides a longer life part as well as reduced system weight. The disclosed dovetail includes an increased shear area  92  that reduces shear strains during fan blade loss. 
     The disclosed shear area  92  is greater than a shear area  65  of a prior dovetail  62 . In one disclosed example, the shear area  92  is between about 20% to 30% greater than the shear area  65  of prior dovetail  62 . In another disclosed example, the shear area is approximately 25% greater than previous dovetails represented by the prior art dovetail  62 . The shear area  92  is related to the increased radial height  106 . In one non-limiting dimensional embodiment, the height is between 0.95 inches and 1.20 inches. Moreover, in another non-limiting dimensional embodiment the height is increased between 0.1 inches and 0.2 inches. 
     This invention allows the dovetail to provide maximized blade stiffness and reduces system weight since are a lighter weight blade material is replaced with heaver disk material. The increased thickness in the tall dovetail  84  provides additional area for fan blade loss. This is a benefit on an aluminum hybrid metallic hollow blades and it applicable to other materials. 
     The disclosed fan blade tall dovetail  84  provides a radially tall dovetail that result in lower strains during a bird/blade impact and a fan blade out situations. This results in no or reduced blade material being liberated. The fan blade is a low cycle fatigue part. The limiting location is on the bearing surface, which experiences high stress levels due to tooth bending in the dovetail. Eliminating this tooth bending provides for a longer lasting part as well as a reduced weight part. Prior art dovetail fillet run out has the potential to fail in shear during a fan blade out event. The extra shear area provided by the disclosed tall dovetail  84  increases capability by reducing shear stresses during such an event. In one example the disclosed tall dovetail  84  includes a reduction in stress of between 10% and 30% over previous dovetail configurations. In another disclosed embodiment, the reduction in stress is between about 12% and 26%. 
     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.