Swept turbomachinery blade

a turbomachinery and blade technology, applied in the field of turbomachinery blades, can solve the problems of pressure waves which extend along the span and degrade and achieve the effect of optimizing the efficiency of the engin

Inactive Publication Date: 2003-03-18
UNITED TECH CORP
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  • Summary
  • Abstract
  • Description
  • Claims
  • Application Information

AI Technical Summary

Benefits of technology

The invention has the advantage of limiting the number of shocks

Problems solved by technology

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.
As a result, the worki

Method used

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  • Swept turbomachinery blade
  • Swept turbomachinery blade
  • Swept turbomachinery blade

Examples

Experimental program
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Embodiment Construction

Referring to FIGS. 1-3, the forward end of a gas turbine engine includes a fan section 10 having a cascade of fan blades 12. Each blade has an attachment 14 for attaching the blade to a disk or hub 16 which is rotatable about a longitudinally extending rotational axis 18. Each blade also has a circumferentially extending platform 20 radially outward of the attachment. When installed in an engine, the platforms of neighboring blades in the cascade abut each other to form the cascade's inner flowpath boundary. An airfoil 22 extending radially outward from each platform has a root 24, a tip 26, a leading edge 28, a trailing edge 30, a pressure surface 32 and a suction surface 34. The axially forwardmost extremity of the leading edge defines an inner transition point 40 at an inner transition radius r.sub.t -inner, radially inward of the tip. The blade cascade is circumscribed by a case 42 which forms the cascade's outer flowpath boundary. The case includes a rubstrip 46 which partially...

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PUM

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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 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.

Description

TECHNICAL FIELDThis 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 INVENTIONGas 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 ...

Claims

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Application Information

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IPC IPC(8): F01D5/16F01D5/14F04D29/38F04D29/32F02K3/06F04D21/00
CPCF01D5/16F01D5/141F04D21/00F04D29/324F04D29/384F04D29/386F05D2220/327F05D2240/302F05D2250/70F05D2250/711F05D2250/712F05D2250/713
Inventor SPEAR, DAVID A.KANTOR, DENNIS N.BIEDERMAN, BRUCE P.OROSA, JOHN A.
Owner UNITED TECH CORP
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