Patent Publication Number: US-9845683-B2

Title: Gas turbine engine rotor blade

Description:
BACKGROUND 
     This disclosure relates to a gas turbine engine, and more particularly to a rotor blade for a gas turbine engine that provides improved aerodynamic performance. 
     Gas turbine engines typically include a compressor section, a combustor section and a turbine section. In general, 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 flow through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads. 
     Some gas turbine engines sections may utilize multiple stages to obtain the pressure levels necessary to achieve desired thermodynamic cycle goals. For example, the compressor and turbine sections of a gas turbine engine typically include alternating rows of moving airfoils (i.e., rotor blades) and stationary airfoils (i.e., stator vanes). Each stage consists of a row of rotor blades and a row of stator vanes. 
     One design feature of a rotor blade that can affect gas turbine engine performance is the airflow gap that extends between the tips of each rotor blade and a surrounding shroud assembly or engine casing. Airflow that escapes through these gaps can result in gas turbine engine performance losses. 
     SUMMARY 
     A rotor blade for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, an airfoil extending in span between a root region and a tip region and a tip portion extending at an angle from the tip region of the airfoil. 
     In a further non-limiting embodiment of the foregoing rotor blade, a span axis of the tip portion forms a dihedral angle relative to a span axis of the airfoil. 
     In a further non-limiting embodiment of either of the foregoing rotor blades, the dihedral angle is greater than or equal to 90° relative to the span axis of the airfoil. 
     In a further non-limiting embodiment of any of the foregoing rotor blades, the dihedral angle is less than or equal to 90° relative to the span axis of the airfoil. 
     In a further non-limiting embodiment of any of the foregoing rotor blades, the dihedral angle is between 45° and 135° degrees relative to the span axis of the airfoil. 
     In a further non-limiting embodiment of any of the foregoing rotor blades, the tip portion extends from a pressure side of the airfoil. 
     In a further non-limiting embodiment of any of the foregoing rotor blades, the tip portion extends in span between a root and a tip and extends in chord between a leading edge and a trailing edge, and the tip portion defines a plurality of cross-sectional slices that extend between the leading edge and the trailing edge along the span of the tip portion. 
     In a further non-limiting embodiment of any of the foregoing rotor blades, the tip portion is not tapered between the root and the tip of the tip portion. 
     In a further non-limiting embodiment of any of the foregoing rotor blades, the tip portion includes a converging taper between the root and the tip of the tip portion. 
     In a further non-limiting embodiment of any of the foregoing rotor blades, the tip portion includes a diverging taper between the root and the tip of the tip portion. 
     In a further non-limiting embodiment of any of the foregoing rotor blades, the tip portion forms a sweep angle that is defined between a chord axis and a span axis of the tip portion. 
     In a further non-limiting embodiment of any of the foregoing rotor blades, the tip portion includes an aft sweep. 
     In a further non-limiting embodiment of any of the foregoing rotor blades, the tip portion includes a forward sweep. 
     In a further non-limiting embodiment of any of the foregoing rotor blades, the tip portion defines a sweep angle and a dihedral angle that extend across an entire span of the tip portion. 
     In a further non-limiting embodiment of any of the foregoing rotor blades, a tip of the tip portion is rotated in a direction toward the root region. 
     In a further non-limiting embodiment of any of the foregoing rotor blades, a tip of the tip portion is rotated in a direction away from the root region. 
     A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a compressor section, a combustor section in fluid communication with the compressor section and a turbine section in fluid communication the combustor section. A plurality of rotor blades positioned within at least one of the compressor section and the turbine section, and each of the plurality of rotor blades includes an airfoil extending in span between a root region and a tip region and a tip portion extending at an angle from the tip region of the airfoil. 
     In a further non-limiting embodiment of the foregoing gas turbine engine, the plurality of rotor blades are at least partially radially surrounded by a shroud assembly. 
     In a further non-limiting embodiment of either of the foregoing gas turbine engines, the tip portion includes a dihedral angle and a sweep angle that extend across an entire span of the tip portion. 
     The various features and advantages of this disclosure will become apparent to those skilled in the art from the following detailed description. The drawings that accompany the detailed description can be briefly described as follows. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  illustrates a schematic, cross-sectional view of a gas turbine engine. 
         FIG. 2  illustrates a portion of a gas turbine engine. 
         FIG. 3  illustrates an exemplary rotor blade that can be incorporated into a gas turbine engine. 
         FIGS. 4A, 4B and 4C  illustrate a tip portion of a rotor blade. 
         FIGS. 5A, 5B and 5C  illustrate various design characteristics that can be incorporated into a tip portion of a rotor blade. 
         FIGS. 6A, 6B and 6C  illustrate additional design characteristics of a rotor blade tip portion. 
         FIGS. 7A and 7B  illustrate other design features that can be incorporated into a tip portion of a rotor blade. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates a gas turbine engine  20 . The exemplary gas turbine engine  20  is a two-spool turbofan engine that generally incorporates 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 for features. The fan section  22  drives air along a bypass flow path B, while the compressor section  24  drives air along a core flow path C for compression and communication into the combustor section  26 . The hot combustion gases generated in the combustor section  26  are expanded through the turbine section  28 . 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 turbofan engines and these teachings could extend to other types of engines, including but not limited to, three-spool engine architectures. 
     The gas turbine engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine centerline longitudinal axis A. The low speed spool  30  and the high speed spool  32  may be mounted relative to an engine static structure  33  via several bearing systems  31 . It should be understood that other bearing systems  31  may alternatively or additionally be provided. 
     The low speed spool  30  generally includes an inner shaft  34  that interconnects a fan  36 , a low pressure compressor  38  and a low pressure turbine  39 . The inner shaft  34  can be connected to the fan  36  through a geared architecture  45  to drive the fan  36  at a lower speed than the low speed spool  30 . The high speed spool  32  includes an outer shaft  35  that interconnects a high pressure compressor  37  and a high pressure turbine  40 . In this embodiment, the inner shaft  34  and the outer shaft  35  are supported at various axial locations by bearing systems  31  positioned within the engine static structure  33 . 
     A combustor  42  is arranged between the high pressure compressor  37  and the high pressure turbine  40 . A mid-turbine frame  44  may be arranged generally between the high pressure turbine  40  and the low pressure turbine  39 . The mid-turbine frame  44  can support one or more bearing systems  31  of the turbine section  28 . The mid-turbine frame  44  may include one or more airfoils  46  that extend within the core flow path C. 
     The inner shaft  34  and the outer shaft  35  are concentric and rotate via the bearing systems  31  about the engine centerline longitudinal axis A, which is co-linear with their longitudinal axes. The core airflow is compressed by the low pressure compressor  38  and the high pressure compressor  37 , is mixed with fuel and burned in the combustor  42 , and is then expanded over the high pressure turbine  40  and the low pressure turbine  39 . The high pressure turbine  40  and the low pressure turbine  39  rotationally drive the respective high speed spool  32  and the low speed spool  30  in response to the expansion. 
     The pressure ratio of the low pressure turbine  39  can be pressure measured prior to the inlet of the low pressure turbine  39  as related to the pressure at the outlet of the low pressure turbine  39  and prior to an exhaust nozzle of the gas turbine engine  20 . In one non-limiting embodiment, the bypass ratio of the gas turbine engine  20  is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor  38 , and the low pressure turbine  39  has a pressure ratio that is greater than about five (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 disclosure is applicable to other gas turbine engines, including direct drive turbofans. 
     In one embodiment of the exemplary gas turbine engine  20 , a significant amount of thrust is provided by the bypass flow path B due to the high bypass ratio. The fan section  22  of the gas turbine engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. This flight condition, with the gas turbine engine  20  at its best fuel consumption, is also known as bucket cruise Thrust Specific Fuel Consumption (TSFC). TSFC is an industry standard parameter of fuel consumption per unit of thrust. 
     Fan Pressure Ratio is the pressure ratio across a blade of the fan section  22  without the use of a Fan Exit Guide Vane system. The low Fan Pressure Ratio according to one non-limiting embodiment of the example gas turbine engine  20  is less than 1.45. Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of [(Tram °R)/(518.7°R)]° 5 , where T represents the ambient temperature in degrees Rankine. The Low Corrected Fan Tip Speed according to one non-limiting embodiment of the example gas turbine engine  20  is less than about 1150 fps (351 m/s). 
     Each of the compressor section  24  and the turbine section  28  may include alternating rows of rotor assemblies and vane assemblies (shown schematically) that carry airfoils that extend into the core flow path C. For example, the rotor assemblies can carry a plurality of rotor blades  25 , while each vane assembly can carry a plurality of vanes  27  that extend into the core flow path C. 
       FIG. 2  schematically illustrates a portion  100  of a gas turbine engine, such as the gas turbine engine  20  of  FIG. 1 . The portion  100  may be representative of a section of either the compressor section  24  or the turbine section  28  of the gas turbine engine  20 . The portion  100  includes a plurality of stages that each include alternating rows of rotor blades  25  and stator vanes  27 . Although two stages are illustrated by  FIG. 2 , it should be understood that the portion  100  could include a greater or fewer number of stages. 
     The rotor blades  25  rotate about the engine centerline longitudinal axis A in a known manner to either create or extract energy (in the form of pressure) from the core airflow that is communicated through the gas turbine engine  20  along the core flow path C. The stator vanes  27  convert the velocity of airflow into pressure, and turn the airflow in a desired direction to prepare the airflow for the next set of rotor blades  25 . 
     The rotor blades  25  are at least partially radially surrounded by a shroud assembly  50  (i.e., an outer casing of the engine static structure  33  of  FIG. 1 ). A gap  52  can extend between each rotor blade  25  and the shroud assembly  50  to provide clearance for accommodating the rotation of the rotor blades  30 . 
       FIG. 3  illustrates an exemplary rotor blade  25  that can be incorporated into a gas turbine engine. For example, one or more rotor blades of the compressor section  24  and/or the turbine section  28  of the gas turbine engine  20  may include a design similar to the exemplary rotor blade  25 . The teachings of this disclosure could also extend to other portions of a gas turbine engine  20 . The rotor blade  25  can include one or more design characteristics that provide improved aerodynamic performance, thereby improving gas turbine engine performance. 
     In this exemplary embodiment, the rotor blade  25  includes an airfoil  56  that axially extends in chord between a leading edge portion  60  and a trailing edge portion  62 . The airfoil  56  also extends in span across a span axis SA between a root region  64  and a tip region  54 . The airfoil  56  may also circumferentially extend between a pressure side  66  and a suction side  68 . 
     A tip portion  58  may extend from the airfoil  56  of the rotor blade  25 . In one embodiment, the tip portion  58  extends from the tip region  54  at an angle relative to the airfoil  56 . In this embodiment, the tip portion  58  extends from the pressure side  66  of the airfoil  56 . That is, the tip portion  58  only extends from a single side of the airfoil  56 . The tip portion  58  may extend from the airfoil  56  such that it is parallel to the shroud assembly  50 , which radially surrounds the rotor blade  25 . 
     Although not shown in  FIG. 3 , the rotor blade  25  may also include platform and root portions for attaching the rotor blade  25  to a rotor disk (see feature  39  of  FIG. 2 , for example). 
       FIGS. 4A, 4B and 4C  illustrate the tip portion  58  of the rotor blade  25  of  FIG. 3 . The tip portion  58  can form a dihedral angle α relative to a span axis SA of the airfoil  56 . 
     In one embodiment, the tip portion  58  forms a dihedral angle α 1  that is 90° relative to the span axis SA (see  FIG. 4A ). In other words, the tip portion  58  can extend across a span axis SA-T that is perpendicular to the span axis SA of the airfoil  56 . In another embodiment, the tip portion  58  forms a dihedral angle α 2  is less than 90° relative to the span axis SA (see  FIG. 4B ). The tip portion  58  could also form a dihedral angle α 3  that is greater than 90° relative to the span axis SA (see  FIG. 4C ). In yet another embodiment, the dihedral angle is between 45° and 135° relative to the span axis SA of the airfoil  56 . 
       FIGS. 5A, 5B and 5C  illustrate possible variations in the chord length over the span of a tip portion  58  of a rotor blade  25 . The tip portion  58  extends in span between a root  70  (near the airfoil  56 ) and a tip  72  (spaced from the airfoil  56 ) and extends in chord between a leading edge  74  and a trailing edge  76 . A plurality of cross-sectional chord slices CL extend between the leading edge  74  and the trailing edge  76  across the span between the root  70  and tip  72 . 
       FIG. 5A  illustrates one possible configuration that can be embodied by the tip portion  58 . In this embodiment, the tip portion  58  is not tapered between the root  70  and the tip  72 . In other words, a chord CL 1  that extends through the root  70  (between the leading edge  74  and the trailing edge  76 ) is the same length as a chord CL 2  that extends through the tip  72  (between the leading edge  74  and the trailing edge  76 ). 
     In another embodiment, the tip portion  58  includes a converging taper between the root  70  and the tip  72 . In other words, as shown in  FIG. 5B , a chord CL 1  that extends through the root  70  can include greater length than a chord CL 2  that extends through the tip  72 . A converging taper such as illustrated by  FIG. 5B  defines taper angles β 1 , β 2  relative to reference axes A 1 , A 2  that extend axially through a leading edge  75  and a trailing edge  77  of the root  70 . The taper angles β 1 , β 2  may be the same or different angles. In this configuration, the leading edge  74  of the tip portion  58  extends toward the trailing edge  76  of the tip portion  58  and the trailing edge  76  extends toward the leading edge  74  to define the converging taper. 
       FIG. 5C  illustrates a tip portion  58  having a diverging taper between the root  70  and the tip  72 . The diverging taper establishes a larger chord CL 2  at the tip  72  as compared to a chord CL 1  that extends through the root  70 . The diverging taper illustrated by  FIG. 5C  defines taper angles β 1 , β 2  relative to reference axes A 1 , A 2  that extend axially from the leading edge  75  and trailing edge  77  of the root  70 . In this configuration, the leading edge  74  of the tip portion  58  extends away from the trailing edge  76  and the trailing edge  76  extends away from the leading edge  74  to define the diverging taper. The taper angles β 1 , β 2  may be the same or different angles. 
       FIGS. 6A, 6B and 6C  illustrate additional design features that can be incorporated into a tip portion  58  of a rotor blade  25 . For example, the tip portion  58  can also form a sweep angle μ. The sweep angle β 1 , β 2  is defined between a chord axis CL 1  and a span axis SP 1  of the tip portion  58 . In one non-limiting embodiment, the span axis SP 1  intersects the chord axis CL 1  at 25% of the length of the chord axis CL 1  between the leading edge  74  and the trailing edge  76 . 
     The tip portion  58  can include no sweep (see  FIG. 6A ), an aft sweep (see  FIG. 6B ) or a forward sweep (see  FIG. 6C ). The aft sweep extends in a downstream direction DD relative to the airfoil  56  (i.e., toward the trailing edge  62 ). A forward sweep extends in an upstream direction UD relative to the airfoil  56  (i.e., toward the leading edge  60 ). 
       FIGS. 7A and 7B  illustrate additional characteristics that can be designed into the tip portion  58  of a rotor blade  25 . The tip portion  58  may include an airfoil tip rotation. As shown in  FIG. 7A , the tip  72  of the tip portion  58  may be rotated by an angle Δ 1  toward the root region  64  (see  FIG. 3 ) of the airfoil  56 . Alternatively, as shown in  FIG. 7B , the tip  72  of the tip portion  58  can be rotated by an angle Δ 2  in a direction away from the root region  64 . In other words, the tip  72  of the tip portion  58  can include a nose down or a nose up configuration. 
     Although the design characteristics described above and illustrated in  FIGS. 4, 5A, 5B, 5C, 6A, 6B, 6C, 7A and 7B  of this application are shown individually, it should be understood that any given tip portion of a rotor blade can include any combination of these design configurations. For example, one exemplary rotor blade can include a tip portion having a dihedral angle that is greater than 90°, a converging taper, no sweep and a nose down configured tip. In another configuration, a tip portion of a rotor blade can include a normal dihedral angle, a diverging taper, forward sweep and no tip rotation. It should be understood that the specific design characteristics for any given rotor blade can vary depending upon design specific parameters, including but not limited to, the aerodynamic and performance requirements of a gas turbine engine. 
     Although the different non-limiting embodiments are illustrated as having specific components, the embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from any of the non-limiting embodiments in combination with features or components from any of the other non-limiting embodiments. 
     It should be understood that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be understood that although a particular component arrangement is disclosed and illustrated in these exemplary embodiments, other arrangements could also benefit from the teachings of this disclosure. 
     The foregoing description shall be interpreted as illustrative and not in any limiting sense. A worker of ordinary skill in the art would understand that certain modifications could come within the scope of this disclosure. For these reasons, the following claims should be studied to determine the true scope and content of this disclosure.