Film cooling hole construction in gas turbine moving-vanes

• US 6,474,947 B1
• Filed: 04/18/2001
• Issued: 11/05/2002
• Est. Priority Date: 03/13/1998
• Status: Expired due to Term

• Current Assignee: Mitsubishi Heavy Industries, Ltd.
• Original Assignee: Mitsubishi Heavy Industries, Ltd.
• Inventors: Yuri, Masanori
• Primary Examiner(s): Ryznic, John E.
• Application Number:
  US09/807,748
• Time in Patent Office: 566 Days
• Field of Search: 416/97.R, 415/115, 415/116, 415/117
• 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 2260/22141 using fins or ribs

Abstract

The interior of a gas turbine moving blade is sectioned by a rib into cooling passage portions. Cooling air enters one of the cooling passage portions and turns to flow into the other cooling passage portion. A stagnation area occurs in an end corner portion of the one cooling passage portion, but a cooling hole is provided so that air flow comes outside of the blade through the cooling hole, and thus cooling air flow occurs thereat. Also, a separation area occurs in a tip end portion of the rib due to separation of air flow, but another cooling hole is provided so that air flow comes outside of the blade and cooling air flow occurs thereat. Further, in a gas turbine moving blade having turbulators provided in multi-stages on a cooling passage inner wall, a film cooling hole structure for eliminating separation of cooling air flow between the turbulators is also provided.

Claims

1. A film cooling hole structure for a gas turbine blade having an interior which is sectioned by a rib into a cooling air flow upstream side passage and a cooling air flow downstream side passage that communicate with each other to define a serpentine cooling passage, comprising:

a film cooling hole in an end comer portion of the cooling air flow upstream side passage; and

a film cooling hole in the cooling air flow downstream side passage and adjacent to a tip portion of the rib,

such that when the gas turbine blade is to be cooled, cooling air is conveyed into the serpentine cooling passage and exits from said film cooling hole in the end comer portion of the cooling air flow upstream side passage and also from said film cooling hole in the cooling air flow downstream side passage adjacent to the tip portion of the rib.

2. The film cooling hole structure according to claim 1, wherein turbulators are arranged in multi-stages on an inner wall of the serpentine cooling passage so as to cross the cooling air when conveyed into the serpentine cooling passage, further comprising:

another film cooling hole between every two immediately adjacent turbulators, such that when the cooling air is conveyed into the serpentine cooling passage to cool the gas turbine blade the cooling air also exits from said another film cooling hole between every two immediately adjacent turbulators,

wherein a width of each of the turbulators is e, and a distance between a downstream side of each of the turbulators and a center of a respective said another film cooling hole is d, with said respective said another film cooling hole being positioned such that the following relationship results.