초록

An airfoil for use as rotor blades in compressors for turbomachines, such as gas turbine engines. The airfoil includes increased forward sweep and forward dihedral effective to reduce losses generated by interaction of tip clearance flow, secondary flows and passage shocks.

대표청구항 ▼

1. A compressor airfoil for pressurizing air inside a surrounding casing, said airfoil comprising: laterally opposite pressure and suction sides joined together at chordally opposite leading and trailing edges and extending in span from a root to a tip;a radially inner portion and a radially outer p

1. A compressor airfoil for pressurizing air inside a surrounding casing, said airfoil comprising: laterally opposite pressure and suction sides joined together at chordally opposite leading and trailing edges and extending in span from a root to a tip;a radially inner portion and a radially outer portion of said airfoil defined along said span;a leading edge aerodynamic sweep defined relative to a stream surface of flow passing said airfoil;a leading edge aerodynamic dihedral defined relative to said stream surface; anda ratio of said leading edge sweep to said leading edge dihedral being in a range effective to reduce losses generated by interaction of tip clearance flow, secondary flows and passage shocks, said ratio being between about 1:1 to about 3:1 along said radially outer portion to said airfoil. 2. The airfoil of claim 1, wherein said radially outer portion of said airfoil is located in a range of about 70% to about 100% span from said root. 3. The airfoil of claim 2, wherein said leading edge sweep and said leading edge dihedral both increase in a forward direction along said radially outer portion of said airfoil progressing in a radially outward direction. 4. The airfoil of claim 3, including a transition portion between about 50% to about 70% span from said root, wherein radial sections of said airfoil have centers-of-gravity that are offset an increasing amount in an aft circumferential direction, opposite to the direction of blade rotation, from a location adjacent to said root to said transition portion and that are offset an increasing amount in a forward circumferential direction, in the direction of blade rotation, from said transition portion to said tip. 5. The airfoil of claim 4, wherein said radial sections of said airfoil have centers-of-gravity that are offset an increasing amount in an axially aft direction from said root to said transition portion and that are offset an increasing amount in an axially forward direction from said transition portion to said tip. 6. The airfoil of claim 3, wherein a chord distribution defined by a variation in the chord of the airfoil with increasing span comprises a non-linear increasing chord distribution. 7. The airfoil of claim 2, wherein said ratio of leading edge sweep to leading edge dihedral is substantially equal to 2:1 along said radially outer portion of said airfoil. 8. The airfoil of claim 2, wherein said sweep is within a range of about 10° to about 45° forward sweep along said radially outer portion. 9. The airfoil of claim 2, wherein said dihedral is within a range of about 5° to about 22.5° forward dihedral along at least a portion of said radially outer portion. 10. The airfoil of claim 9, wherein said forward sweep and said forward dihedral both increase monotonically along said radially outer portion of said airfoil to a location adjacent to said tip. 11. A compressor airfoil for pressurizing air inside a surrounding casing, said airfoil comprising: laterally opposite pressure and suction sides joined together at chordally opposite leading and trailing edges and extending in span from a root to a tip;a radially inner portion and a radially outer portion of said airfoil defined along said span, said radially outer portion being located in a range of about 70% to about 100% span from said root;a leading edge aerodynamic sweep defined relative to a stream surface of a flow passing said airfoil;a leading edge aerodynamic dihedral defined relative to said stream surface;said leading edge sweep and said leading edge dihedral both increasing monotonically in a forward direction along said radially outer portion of said airfoil progressing in a radially outward direction;a transition portion located between said radially inner portion and said radially outer portion, said transition portion being located in a range of about 50% to about 70% span from said root;radial sections of said airfoil defining centers-of-gravity wherein said centers-of-gravity are offset an increasing amount in an aft circumferential direction, opposite to the direction of blade rotation from a location adjacent to said root to said transition portion and are offset an increasing amount in a forward direction, in the direction of blade rotation, from said transition portion to said tip, and said centers-of-gravity are offset an increasing amount in an axially aft direction from said root to said transition portion and are offset an increasing amount in an axially forward direction from said transition portion to said tip; anda ratio of said leading edge sweep to said leading edge dihedral being in a range effective to reduce losses generated by interaction of tip clearance flow, secondary flows and passage shocks, said ratio being between about 1:1 to about 3:1 along said radially outer portion to said airfoil. 12. The airfoil of claim 11, wherein a chord distribution defined by a variation in the chord of the airfoil with increasing span comprises a non-linear increasing chord distribution. 13. The airfoil of claim 11, wherein said ratio of leading edge sweep to leading edge dihedral is substantially equal to 2:1 along said radially outer portion of said airfoil. 14. The airfoil of claim 11, wherein said sweep is within a range of about 10° to about 45° forward sweep along said radially outer portion. 15. The airfoil of claim 14, wherein said dihedral is within a range of about 5° to about 22.5° forward dihedral along at least a portion of said radially outer portion. 16. A compressor blade for a gas turbine engine, said compressor blade having an airfoil comprising: laterally opposite pressure and suction sides joined together at chordally opposite leading and trailing edges and extending in span from a root to a tip;a radially inner portion and a radially outer portion of said airfoil defined along said span;a leading edge aerodynamic sweep defined relative to a stream surface of a flow passing said airfoil;a leading edge aerodynamic dihedral defined relative to said stream surface; andwherein said leading edge aerodynamic sweep and dihedral of said radially outer portion is designed so that tip losses generated by the interaction of tip clearance, secondary flows and passage shocks are reduced, said leading edge aerodynamic sweep and said leading edge aerodynamic dihedral are defined substantially in accordance with the values of LE Sweep and LE Dihedral, respectively, set forth at locations identified by span locations, N, 11-17 in Table 1. 17. The compressor blade of claim 16, wherein, along the length of said span, a chord, an axial center-of-gravity and a tangential center-of-gravity are defined substantially in accordance with the values of Chord, x-cg and y-cg, respectively, set forth at locations identified by span locations, N, 1-17 in Table 1. 18. A compressor blade for a gas turbine engine, said compressor blade having an airfoil comprising: laterally opposite pressure and suction sides joined together at chordally opposite leading and trailing edges and extending in span from a root to a tip;a radially inner portion and a radially outer portion of said airfoil defined along said span;a leading edge aerodynamic sweep defined relative to a stream surface of a flow passing said airfoil;a leading edge aerodynamic dihedral defined relative to said stream surface; andwherein said leading edge aerodynamic sweep and dihedral of said radially outer portion is designed so that tip losses generated by the interaction of tip clearance, secondary flows and passage shocks are reduced, said leading edge aerodynamic sweep and said leading edge aerodynamic dihedral are defined substantially in accordance with the values of LE Sweep and LE Dihedral, respectively, set forth at locations identified by span locations, N, 11-17 in Table 2. 19. The compressor blade of claim 18, wherein, along the length of said span, a chord, an axial center-of-gravity and a tangential center-of-gravity are defined substantially in accordance with the values of Chord, x-cg and y-cg, respectively, set forth at locations identified by span locations, N, 1-17 in Table 2.