Endwall shape for use in turbomachinery

• US 6,669,445 B2
• Filed: 03/07/2002
• Issued: 12/30/2003
• Est. Priority Date: 03/07/2002
• Status: Expired due to Term

• Current Assignee: Raytheon Technologies Corporation d/b/a RTX (Rtx Corporation)
• Original Assignee: United Technologies Corporation (Rtx Corporation)
• Inventors: Staubach, J. Brent, Wagner, Joel H., Aggarwala, Andrew S.
• Primary Examiner(s): Look, Edward K.
• Assistant Examiner(s): MCALEENAN, JAMES M
• Application Number:
  US10/093,989
• Publication Number:
  US 20030170124A1
• Time in Patent Office: 663 Days
• Field of Search: 416/193.A, 416/248, 416/219.R, 416/220.R, 416/234, 416/239, 415/914, 415/181, 415/191, 415/192, 415/195, 415/210.1, 415/208.1
• US Class Current: 416/193.00A
• CPC Class Codes: F01D 5/142 of the blades of successive rotor or stator blade-rows
• CPC Class Codes: F01D 5/143 Contour of the outer or inner working fluid flow path wall, i.e. shroud or hub contour
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft
• CPC Class Codes: Y10S 415/914 Device to control boundary layer

Abstract

The present invention relates to an endwall shape for reducing shock strength on transonic turbomachinery airfoils which define at least one flow passage. The endwall shape includes a non-axisymmetric trough which extends from a leading portion of the at least one flow passage to a point near a trailing edge portion of the at least one flow passage.

Claims

1. A flow directing assembly for use in a rotary machine comprising:

a plurality of blades, each said blade having an airfoil and a platform;

a plurality of flow passages defined by said airfoils of said blades;

each of said flow passages having an inner endwall defined by platforms of adjacent ones of said blades;

said inner endwall of each said flow passage having means for minimizing shock effects within and aft of each of said flow passages;

said shock effect minimizing means comprising a non-axisymmetric trough which extends from a point within said flow passage to a point aft of said flow passage; and

said trough having an initial amplitude at a location adjacent a leading edge portion of said flow passage, a maximum amplitude adjacent a flow passage throat axial location, and a final amplitude at a downstream extent of said platform and wherein said initial and final amplitudes are less than said maximum amplitude.

2. A flow directing assembly according to claim 1, wherein each said trough is located in a portion of said platform associated with a respective blade.

3. A flow directing assembly according to claim 1, wherein each said airfoil has a suction side and each said trough has a curvature substantially identical to the curvature of a rear portion of said airfoil suction side.

4. A flow directing assembly according to claim 1, wherein said plurality of blades comprises a plurality of turbine blades.

5. A flow directing assembly according to claim 1, wherein said plurality of blades comprises a plurality of compressor blades.

6. An endwall shape for reducing shock strength on transonic turbomachinery airfoils having at least one flow passage defined by at least two airfoils comprising a non-axisymmetric trough extending from a leading edge portion of said at least one flow passage to a point near a trailing edge portion of said at least one flow passage and said non-axisymmetric trough has having a maximum amplitude near a passage throat axial location.

7. An endwall shape according to claim 6, wherein said trough has an initial amplitude at said leading edge portion and a final amplitude adjacent said point near said trailing edge portion and wherein both said initial amplitude and said final amplitude are less than said maximum amplitude.