초록 ▼

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 subjec

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

대표청구항 ▼

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 casca

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, which fan comprises a plurality of blades mounted for rotation about a fan axis with neighboring blades forming passages for a working medium gas, wherein the blade comprises: 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; andthe tip region is translated forward, which translation includes a circumferential component, 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. A turbomachinery blade for a gas turbine engine fan, which fan comprises a plurality of blades mounted for rotation about a fan axis with adjacent blades forming passages for a working medium gas, the blade comprising: 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;wherein the inner region is swept forward and the intermediate region is swept rearward; andthe tip region is configured to mitigate the adverse aerodynamic effects of the compressibility of the working medium gas moving at supersonic speed in a passage in the vicinity of a case surrounding the fan, which adverse aerodynamic effects include a shock in the gas adjacent the case;wherein the tip region configuration is such that a portion of the tip region is disposed to intercept the shock;wherein the configuration of the tip region is such that at least a portion of the tip region has a forward sweep angle, and the portion of the tip region having a forward sweep angle includes an axial component, or a circumferential component, or a combination thereof; andwherein the configuration of the tip region is such that at least a first portion of the tip region is angled upstream of a second portion of the tip region, which second and first portions are contiguous with one another. 6. The turbomachinery blade of claim 5, wherein the portion of the tip region that is angled upstream includes an axial upstream component, or a circumferential upstream component, or a combination thereof. 7. A fan stage of a ducted fan gas turbine engine that is rotatable about an axis of rotation and defines a downstream direction along the axis of rotation, comprising: a fan casing that defines an inner duct wall having a fan rotor region, the inner duct wall of the fan casing at the fan rotor region being convergent;a hub disposed concentrically relative to the fan casing;a fan rotor that includes multiple swept fan blades, the swept fan blades being spaced apart around the hub and being capable of rotating at speeds providing supersonic working medium gas velocities over the blades to cause a shock in the gas adjacent the inner duct wall, each of the multiple swept fan blades having:a tip profile that corresponds to the inner duct wall of the fan casing;a leading edge that defines a variable sweep angle in a direction perpendicular to the axis of rotation, the leading edge including: an inner region adjacent the hub, the inner region defining a forward sweep angle;an intermediate region between the inner region and the fan casing, the intermediate region defining a rearward sweep angle; andan outer region between the intermediate region and the fan casing, the outer region being translated forward relative to a leading edge with the same sweep angle as an outward boundary of the intermediate region to provide a sweep angle that causes the blade to intercept the shock.