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.sub.t -inner. The sweep angle of the airfoil is nondecreasing with increasing radius from the inner transition radius to an outer transition radius r.sub.t-outer, radially inward of the airfoil tip (26), and is nonincreasing with increasing radius between the outer transition radius and the airfoil tip.

Abstract: A compressor or fan blade (112) for a gas turbine engine includes a suction surface midspan shroud (122) which is cambered so that the shroud trailing edge (132) is substantially aligned with the direction of the air flow in the vicinity of the shroud trailing edge (132). The cambered portion originates at the chordwise location where a shock wave (60) attached to the leading edge of the neighboring blade crosses the shroud (136). The cambered portion gradually transitions to an uncambered profile at a circumferential location intermediate the suction surface (110) of the blade (112) and the pressure surface shroud of the neighboring blade.

Abstract: A flow directing assembly 24, 24a for a rotary machine 10 is disclosed. Various construction details are developed to decrease aerodynamic losses and promote the ease of design. In one particular embodiment, a concave region 58 of an end wall is employed to create convex wall regions 54, 56 upstream and downstream of the concave region.