Swept turbomachinery blade

Details Swept turbomachinery blade Swept turbomachinery blade

1999-06-30

2003-03-18

Verdier, Christopher (Department: 3745)

Rotary kinetic fluid motors or pumps

Means, disposition or arrangement for causing supersonic...

C416S238000, C416S242000

Reissue Patent

active

RE038040

ABSTRACT:

TECHNICAL FIELD
This invention relates to turbomachinery blades, and particularly to blades whose airfoils are swept to minimize the adverse effects of supersonic flow of a working medium over the airfoil surfaces.
BACKGROUND OF THE INVENTION
Gas turbine engines employ cascades of blades to exchange energy with a compressible working medium gas that flows axially through the engine. Each blade in the cascade has an attachment which engages a slot in a rotatable hub so that the blades extend radially outward from the hub. Each blade has a radially extending airfoil, and each airfoil cooperates with the airfoils of the neighboring blades to define a series of interblade flow passages through the cascade. The radially outer boundary of the flow passages is formed by a case which circumscribes the airfoil tips. The radially inner boundary of the passages is formed by abutting platforms which extend circumferentially from each blade.
During engine operation the hub, and therefore the blades attached thereto, rotate about a longitudinally extending rotational axis. The velocity of the working medium relative to the blades increases with increasing radius. Accordingly, it is not uncommon for the airfoil leading edges to be swept forward or swept back to mitigate the adverse aerodynamic effects associated with the compressibility of the working medium at high velocities.
One disadvantage of a swept blade results from pressure waves which extend along the span of each airfoil suction surface and reflect off the surrounding case. Because the airfoil is swept, both the incident waves and the reflected waves are oblique to the case. The reflected waves interact with the incident waves and coalesce into a planar aerodynamic shock which extends across the interblade flow channel between neighboring airfoils. These “endwall shocks” extend radially inward a limited distance from the case. In addition, the compressibility of the working medium causes a passage shock, which is unrelated to the above described endwall shock, to extend across the passage from the leading edge of each blade to the suction surface of the adjacent blade. As a result, the working medium gas flowing into the channels encounters multiple shocks and experiences unrecoverable losses in velocity and total pressure, both of which degrade the engine's efficiency. What is needed is a turbomachinery blade whose airfoil is swept to mitigate the effects of working medium compressibility while also avoiding the adverse influences of multiple shocks.
DISCLOSURE OF THE INVENTION
It is therefore an object of the invention to minimize the aerodynamic losses and efficiency degradation associated with endwall shocks by limiting the number of shocks in each interblade passage.
According to the invention, a blade for a blade cascade has an airfoil which is swept over at least a portion of its span, and the section of the airfoil radially coextensive with the endwall shock intercepts the endwall shock extending from the neighboring airfoil so that the endwall shock and the passage shock are coincident.
In one embodiment the axially forwardmost extremity of the airfoil's leading edge defines an inner transition point located at an inner transition radius radially inward of the airfoil tip. An outer transition point is located at an outer transition radius radially intermediate the inner transition radius and the airfoil tip. The outer transition radius and the tip bound a blade tip region while the inner and outer transition radii bound an intermediate region. The leading edge is swept at a first sweep angle in the intermediate region and is swept at a second sweep angle over at least a portion of the tip region. The first sweep angle is generally nondecreasing with increasing radius and the second sweep angle is generally non-increasing with increasing radius.
The invention has the advantage of limiting the number of shocks in each interblade passage so that engine efficiency is maximized.


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Leading edge sweep angle profiles of fan blades of Pratt & Whitney PW305 and PW306 gas turbine engines (no date).
Puterbaugh et al., “Design of a Rotor Incorporating Meridional Sweep and Circumferential Lean for Shock Loss Attenuation,” Feb. 1987, Contract AFWAL-TR-86-2013, Aero Propulsion Laboratories, Air Force Wright Aeronautical Laboratories, Wright-Paterson Air Force Base, Ohio.
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European Search Report, dated Feb. 25, 1998, in EP 774,567.
European Patent Office Official Action, dated Sep. 24, 1998, in EP 774,567.

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