Cooled blades for a gas turbine engine

• US 5,700,131 A
• Filed: 08/24/1988
• Issued: 12/23/1997
• Est. Priority Date: 08/24/1988
• Status: Expired due to Term

• Current Assignee: United Technologies Corporation (Rtx Corporation)
• Original Assignee: United Technologies Corporation (Rtx Corporation)
• Inventors: McClelland, Robert J., Auxier, Thomas A., Hall, Kenneth B.
• Primary Examiner(s): Carone, Michael J.
• Assistant Examiner(s): MONTGOMERY, CHRISTOPHER KEITH
• Application Number:
  US07/236,093
• Time in Patent Office: 3,408 Days
• Field of Search: 416/92, 416/96 A, 416/97 R, 416/97 A
• US Class Current: 416/97.00R
• CPC Class Codes: F01D 5/186 Film cooling F01D5/187 takes precedence
• CPC Class Codes: F01D 5/187 Convection cooling
• CPC Class Codes: F05D 2260/201 by impingement of a fluid
• CPC Class Codes: F05D 2260/2212 by creating turbulence
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

An internally cooled turbine blade for a gas turbine engine is modified at the leading and trailing edges to include a dynamic cool air flowing radial passageway with an inlet at the root and a discharge at the tip feeding a plurality of radially spaced film cooling holes in the airfoil surface. Replenishment holes communicating with the serpentine passages radially spaced in the inner wall of the radial passage replenish the cooling air lost to the film cooling holes. The discharge orifice is sized to match the backflow margin to achieve a constant film hole coverage throughout the radial length. Trip strips may be employed to augment the pressure drop distribution

Claims

1. A turbine blade surface for a gas turbine engine having internal passages for flowing cooling air therein, the blade having an airfoil surface defining a root section, a leading edge section, a trailing edge section, a mid chord section, and a tip section, a plurality of radial internal passages defined by internal wall means formed adjacent said trailing edge section and leading edge section extending from said root section to said tip section defining a feed channel, a plurality of radially spaced film cooling holes in said airfoil surface communicating with said feed channel to flow a film of cooling air adjacent said airfoil surface, a plurality of replenishment holes spaced radially in said wall means for flowing cooling air from said mid cord section to said feed channel to replenish the cooling air in said feed channel that is otherwise lost in supplying cooling air to said film cooling holes, and means for communicating cooling air from said root section to discharge from an orifice in said airfoil surface at said tip section, and a source of cooling air for feeding cooling air to said root section.

2. A turbine blade as claimed in claim 1 wherein said orifice at said tip section is sized to minimize the flow of cooling air in said feed chamber to substantially provide a uniform film hole coverage of all the film holes being fed by said feed channel by matching the backflow margin for each radial position in said feed channel.

3. A turbine blade as claimed in claim 1 wherein said feed channel provides means for separating dirt from said cooling air and removing said separated dirt through said discharge orifice.

4. A turbine blade as in claim 1 including trip strips in said feed channel.

5. A turbine blade as in claim 4 wherein said trip strips are skewed relative to the direction of said fuel channel.