Passive clearance system for turbine blades

• US 5,282,721 A
• Filed: 01/04/1993
• Issued: 02/01/1994
• Est. Priority Date: 09/30/1991
• Status: Expired due to Term

• Current Assignee: United Technologies Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Original Assignee: United Technologies Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Inventors: Kildea, Robert J.
• Primary Examiner(s): Look, Edward K.
• Assistant Examiner(s): LARSON, JAMES
• Application Number:
  US08/000,529
• Time in Patent Office: 393 Days
• Field of Search: 416/90 R, 416/92, 416/97 R, 416/97 A, 416/228, 416/235, 416/236 R, 416/237, 415/115, 415/173.1
• US Class Current: 416/97.00R
• CPC Class Codes: F01D 11/10 using sealing fluid, e.g. steam
• CPC Class Codes: F01D 5/20 Specially-shaped blade tips to seal space between tips and stator F01D5/225 takes precedence

Abstract

Spent cooling air from the internal passage(s) in a turbine blade of a gas turbine engine are judiciously located to inject air at the pressure side of the blade in proximity to or in the gap between the tip of the blade and outer air seal to reduce leakage in the gap and improve engine performance. In one embodiment, a projection at the tip of the blade on the pressure side is utilized in combination with an angled discharge passageway and in another embodiment a pair of discharge angled passages, one interconnecting an internal passage on the pressure side and the other connecting an internal passage on the suction side of the blade is utilized. A projection at the tip adjacent the pressure side may also be incorporated in the second embodiment.

Claims

1. A turbine blade for a gas turbine engine having at least one internal cooling passage internally of the turbine blade, said turbine blade being driven by engine fluid working medium and having an airfoil section including an outer wall defining a pressure surface, a suction surface, a tip and a lower end of said internal cooling passage for leading cooling air from said lower end to flow through said internal passage and discharge through a plurality of spaced holes formed in said outer wall, said turbine blade being rotatably mounted in said engine such that said tip is spaced adjacent an annular member which defines therewith a gap, the improvement comprising a projection having an outer surface extending axially from the tip of the airfoil in the same plane of the airfoil'"'"'s tip surface and an inclined inner surface extending from the pressure surface to the end of said outer surface, and an angled hole disposed substantially parallel to said inclined inner surface extending through the outer wall on the pressure surface connecting said internal passage for leading spent cooling air externally of said blade in proximity to said projection at a location spaced from said tip such that the flow of spent cooling air flows along said inner surface of said projection toward said tip whereby said flow combines with said fluid working medium flowing radially adjacent said pressure surface so that the tendency of said engine fluid working medium from migrating into said gap is reduced and the performance of said turbine blade is enhanced.

2. A turbine blade as claimed in claim 1 wherein the angle of said hole is selected to inject the flow from said blade at an angle substantially equal to between 35 and 60 degrees relative to the center line of said engine.

3. A turbine blade for a gas turbine engine having a plurality of internal cooling passageways internally of the turbine blade, said turbine blade being driven by the fluid working medium of said engine having an airfoil section including an outer wall defining a pressure surface, a suction surface, a tip and a lower extremity, at least one of said plurality of internal cooling passageways being adjacent said pressure surface and another of said passageways being adjacent said suction surface, said turbine blade being rotatably mounted in said engine such that said tip is spaced adjacent an annular member which defines therewith a gap, the improvement comprising a projection having an outer surface extending axially from the tip of the airfoil in the same plane as the airfoil tip'"'"'s surface and an inclined inner surface extending from the pressure surface to the end of said outer surface, a first acutely angled passageway substantially parallel to said inclined inner surface extending through the outer wall on the pressure surface connecting said one of said internal passageways adjacent said pressure surface for leading spent cooling air externally of said blade on said pressure side in proximity to said tip and a second acutely angled passageway extending through said outer wall on said tip surface connecting said one of said passageways adjacent said suction surface for leading spent cooling air to said gap in proximity to the pressure surface, such that the flow of spent cooling air from said first angled passageway flows along the pressure surface toward said tip and said gap and together with said flow from said second acutely angled passageway obstructs the flow in said gap to minimize the tendency of the fluid working medium to flow into said gap and enhance the performance of said turbine blade.

4. A turbine blade as claimed in claim 3 wherein the angle of first and second said angled passageways is selected to inject the flow from said blade at an angle substantially equal to between 35 and 60 degrees relative to the surface of the tip of the airfoil.

5. A turbine blade as claimed in claim 4 wherein the second angle passageway is staggered relative to said first angle passageway.