Swept turbomachinery blade

Abstract

A swept turbomachinery blade for use in a cascade of such blades is disclosed. The blade ( 12 ) has an airfoil ( 22 ) uniquely swept so that an endwall shock ( 64 ) of limited radial extent and a passage shock ( 66 ) are coincident and a working medium ( 48 ) flowing through interblade passages ( 50 ) is subjected to a single coincident shock rather than the individual shocks. In one embodiment of the invention the forwardmost extremity of the airfoil defines an inner transition point ( 40 ) located at an inner transition radius r t -inner. The sweep angle of the airfoil is nondecreasing with increasing radius from the inner transition radius to an outer transition radius r t-outer , radially inward of the airfoil tip ( 26 ), and is nonincreasing with increasing radius between the outer transition radius and the airfoil tip.

We claim:  
     
       1. A turbomachinery blade for a turbine engine having a cascade of blades rotatable about a rotational axis so that each blade in the cascade has a leading neighbor and a trailing neighbor, and each blade cooperates with its neighbors to define flow passages for a working medium gas, the blade cascade being circumscribed by a case and under some operational conditions an endwall shock extends a limited distance radially inward from the case and also extends axially and circumferentially across the flow passages, and a passage shock also extends across the flow passages, the turbomachinery blade including an airfoil having a leading edge, a trailing edge, a root, a tip and an inner transition point located at an inner transition radius radially inward of the tip, the blade characterized in that at least a portion of the leading edge radially outward of the inner transition point is swept and a section of the airfoil radially coextensive with the endwall shock extending from the leading neighbor intercepts the endwall shock so that the endwall shock and the passage shock are coincident. 
     
     
       2. A turbomachinery blade for a turbine engine having a cascade of blades rotatable about a rotational axis so that each blade in the cascade has a leading neighbor and a trailing neighbor, and each blade cooperates with its neighbors to define flow passages for a working medium gas, the blade cascade being circumscribed by a case and under some operational conditions an endwall shock extends a limited distance radially inward from the case and also extends axially and circumferentially across the flow passages and a passage shock also extends across the flow passages, the turbomachinery blade including an airfoil having a leading edge, a trailing edge, a root, a tip located at a tip radius, an inner transition point located at an inner transition radius radially inward of the tip, and an outer transition point at an outer transition radius radially intermediate the inner transition radius and the tip radius, the blade having a tip region bounded by the outer transition radius and the tip radius, and an intermediate region bounded by the inner transition radius and the outer transition radius, the blade characterized in that the leading edge is swept in the intermediate region at a first sweep angle which is generally nondecreasing with increasing radius, and the leading edge is swept over at least a portion of the tip region at a second sweep angle which is generally nonincreasing with increasing radius so that the section of the airfoil radially coextensive with the endwall shock extending from the leading neighbor intercepts the endwall shock so that the endwall shock and the passage shock are coincident. 
     
     
       3. The turbomachinery blade of  claim 1  or  2  characterized in that the inner transition radius is coincident with the root at the leading edge of the blade. 
     
     
       4. A turbomachinery blade for a gas turbine engine fan comprising a plurality of blades mounted for rotation about a fan axis with neighboring blades forming passages for a working medium gas, wherein: 
       
         the blade has a configuration enabling the fan to rotate at speeds providing supersonic flow velocities over the blade in at least a portion of each passage causing the formation of a shock in the gas adjacent an inner wall of a case forming an outer boundary for the working medium gas flowing through the passages;  
       
       
         the blade has a leading edge with an inner region ending at an inward boundary of an intermediate region and a tip region beginning at an outward boundary of the intermediate region and extending to a tip end of the blade, the inner region being swept forward and the intermediate region being swept rearward at a sweep angle that does not decrease; and  
       
       
         the tip region is translated forward relative to a leading edge with the same sweep angle as the outward boundary of the intermediate region, to provide a sweep angle that causes the blade to intercept the shock. 
       
     
     
       5. The turbomachinery blade of  claim 4 , wherein throughout the tip region the sweep angle is less than the sweep angle at the outward boundary of the intermediate region. 
     
     
       6. The turbomachinery blade of  claim 5 , wherein the sweep angle decreases throughout the tip region. 
     
     
       7. The turbomachinery blade of  claim 6 , wherein the sweep angle increases throughout the intermediate region. 
     
     
       8. The turbomachinery blade of any one of  claims 4  to  7 , wherein the inner region extends between a root end of the blade and the inward boundary of the intermediate region, and the entire inner region is swept forward. 
     
     
       9. A blade for a gas turbine engine fan comprising a plurality of blades mounted for rotation within a case circumscribing the blades and forming an outer boundary for a working medium gas flowing through passages formed by neighboring blades, wherein: 
       
         the blade has a configuration enabling the fan to rotate at speeds providing supersonic flow velocities over the blade in at least a portion of each passage;  
       
       
         the blade has a leading edge with an inner region ending at an inward boundary of an intermediate region and a tip region beginning at an outward boundary of the intermediate region and extending to a tip end of the blade, the inner region being swept forward and the intermediate region being swept rearward at a sweep angle that does not decrease from the inward boundary of the intermediate region to the outward boundary of the intermediate region; and  
       
       
         throughout the tip region the sweep angle is less than the sweep angle at the outward boundary of the intermediate region. 
       
     
     
       10. The blade of  claim 9 , wherein the tip region is translated forward relative to a leading edge with the same sweep angle as the outward boundary of the intermediate region. 
     
     
       11. The blade of  claim 10 , wherein the inner region extends between a root end of the blade and the inward boundary of the intermediate region, and the entire inner region is swept forward. 
     
     
       12. The blade of  claim 11 , wherein: 
       
         the intermediate region sweep angle increases throughout the intermediate region; and  
       
       
         the tip region sweep angle decreases throughout the tip region. 
       
     
     
       13. The blade of  claim 10 , wherein the tip region sweep angle decreases throughout the tip region. 
     
     
       14. The blade of  claim 13 , wherein the intermediate region sweep angle increases throughout the intermediate region. 
     
     
       15. The blade of  claim 9 , wherein the tip region maintains a rearward sweep throughout the tip region. 
     
     
       16. A gas turbine engine fan, comprising a plurality of blades mounted for rotation within a case circumscribing the blades and forming an outer boundary for a working medium gas flowing through passages formed by neighboring blades, wherein: 
       
         each blade has a configuration enabling the fan to rotate at speeds providing supersonic working medium gas velocities over the blade at least in the vicinity of the passage proximate to the case;  
       
       
         each blade has a leading edge with an inner region ending at an inward boundary of a swept intermediate region and a swept tip region beginning at an outward boundary of the intermediate region and extending to a tip end of the blade, the inner region of each blade being swept forward and the intermediate region of each blade being swept rearward at a sweep angle that does not decrease from the inward boundary of the intermediate region to the outward boundary of the intermediate region; and  
       
       
         throughout the tip region the sweep angle of each blade is less than the sweep angle at the outward boundary of the intermediate region. 
       
     
     
       17. The gas turbine engine fan of  claim 16 , wherein the tip region is translated forward relative to a leading edge with the same sweep angle as the outward boundary of the intermediate region. 
     
     
       18. The gas turbine engine fan of  claim 17 , wherein: 
       
         the intermediate region sweep angle of each blade increases throughout the intermediate region; and  
       
       
         the tip region sweep angle of each blade decreases throughout the tip region. 
       
     
     
       19. The gas turbine engine fan of  claim 18 , wherein the inner region of the leading edge of each blade begins at a root end of the blade and extends to the inward boundary of the intermediate region, and the entire inner region of each blade is swept forward. 
     
     
       20. A gas turbine engine fan comprising a plurality of identical blades, each blade being mounted for rotation within a case circumscribing the blades and having an inner wall forming an outer boundary for a working medium gas flowing through passages formed by neighboring blades, wherein: 
       
         each blade has a configuration enabling the fan to rotate at speeds providing supersonic working medium gas velocities over the blade in the vicinity of the passages proximate to the case;  
       
       
         each blade has a leading edge with an inner region, an intermediate region and a tip region, the inner region extending to an inward boundary of the intermediate region, and the tip region extending from an outward boundary of the intermediate region to a tip end of the blade; and  
       
       
         the inner region is swept forward, the intermediate region is swept rearward at a sweep angle that does not decrease, and the tip region is translated forward relative to a leading edge with the same sweep angle as the outward boundary of the intermediate region. 
       
     
     
       21. The gas turbine engine fan of  claim 20 , wherein the tip region maintains a rearward sweep throughout the tip region. 
     
     
       22. The gas turbine engine fan of  claim 20 , wherein: 
       
         the intermediate region sweep angle of each blade increases throughout the intermediate region; and  
       
       
         the tip region of each blade is swept at a sweep angle that decreases throughout the tip region. 
       
     
     
       23. The gas turbine engine fan of  claim 20 , wherein the inner wall of the case is perpendicular to pressure waves that extend spanwise of the blades as they rotate, the waves being incident to the case wall in a region of the blades. 
     
     
       24. The gas turbine engine fan of  claim 20 , wherein a projection of the tip end of each blade onto a radial plane is parallel to the inner wall of the casing in longitudinal cross- section.   
     
     
       25. The gas turbine engine fan of  claim 20 , wherein the inner region of the leading edge of each blade begins at a root end of the blade, and the entire inner region of each blade is swept forward. 
     
     
       26. A blade for a gas turbine engine rotatable within a case at speeds providing supersonic flow over at least a portion of the blade, wherein the blade has a leading edge with a forward swept inner region, the inner region ending at a rearward swept middle region having a sweep angle that does not decrease throughout the middle region, the middle region ending at a tip region that is translated forward relative to a leading edge with the same sweep angle as the end of the middle region. 
     
     
       27. The blade of  claim 26 , wherein the tip region maintains a rearward sweep throughout the tip region. 
     
     
       28. The blade of  claim 26 , wherein the inner region extends from a blade root to the middle region and the leading edge is swept forward throughout the inner region. 
     
     
       29. The blade of  claim 28 , wherein the sweep angle of the middle region increases throughout the middle region. 
     
     
       30. The blade of  claim 29 , wherein throughout the tip region the sweep angle is less than the sweep angle at the end of the middle region. 
     
     
       31. The blade of  claim 30 , wherein the sweep angle of the tip region decreases from the end of the middle region to a tip end of the blade. 
     
     
       32. A blade for a gas turbine engine rotatable within a case at speeds providing supersonic flow over at least a portion of the blade, wherein the blade has a leading edge with a forward swept middle region having a sweep angle that does not decrease throughout the middle region and ending at a tip region that is translated rearward relative to a leading edge with the same sweep angle as the end of the middle region. 
     
     
       33. The blade of  claim 32 , wherein the tip region maintains a forward sweep throughout the tip region. 
     
     
       34. The blade of  claim 32 , wherein the leading edge has a rear swept inner region. 
     
     
       35. The blade of  claim 34 , wherein the sweep angle of the middle region increases throughout the middle region. 
     
     
       36. The blade of  claim 35 , wherein throughout the tip region the sweep angle is less than the sweep angle at the end of the middle region. 
     
     
       37. The blade of  claim 36 , wherein the sweep angle of the tip region decreases from the end of the middle region to a tip end of the blade.