Patent Application
20100247286
This invention relates to gas turbine component cooling techniques and, more specifically, to a manner of feeding cooling air to film cooling holes in turbine components with seal slots.
Gas turbine engines operate at elevated temperatures, and film cooling is widely used to protect components from the harsh high-temperature environment. Maintaining metal temperatures for gas turbine components within material limits has been addressed by many different techniques such as film cooling, impingement cooling, low conductivity coatings and heat augmentation devices such as turbulators, ribs, pin fin banks, etc.
Film cooling is widely used in connection with gas turbine first-stage components and to a lower extent in subsequent stages. Standard practice among the industry is to feed these film cooling holes from existing cavities built into the component. This severely limits flexibility with respect to drilling holes at locations not aligned with the cavities. As a result, the designer oftentimes cannot place film cooling at locations of high level temperatures, or has to orient the cooling holes at angles that reduce the impact of the film cooling. Competitors have addressed this issue in the past by machining dedicated chambers and serpentine passages into the component. These features are only manufactured for the purpose of feeding these holes, and add extra manufacturing cost to the component.
Specific examples in the prior art include cooling holes fed from cavities cast into the turbine sidewalls as exemplified by U.S. Pat. No. 5,344,283. Other approaches for casting dedicated chambers into the sidewalls with the intent of feeding film cooling holes are disclosed in U.S. Pat. Nos. 6,254,333 and 6,210,111. A cavity formed by seal plates in a cold side of a stage one turbine nozzle is disclosed in U.S. Pat. No. 5,417,545. A concept for machining multiple cooling holes such that they feed from the same aperture in a cold side cavity is disclosed in U.S. Pat. No. 5,062,768. The assignee of this invention presents a concept for pressurizing a seal slot with air from cooling cavities for the purpose of cooling the seal itself in U.S. Pat. No. 6,340,285.
In a first exemplary but non-limiting aspect, the present invention relates to a cooling arrangement for a turbine component having a slot along an edge thereof, the slot having a closed end formed with at least one cooling cavity, and at least one cooling passageway extending between the cavity and an external surface of the turbine component.
In another aspect, the invention relates to a cooling arrangement for a first component of a turbine having a seal slot formed in a forward face of the component, the seal slot extending about a generally rectangular opening in said forward face and opening in a direction toward a second turbine component and adapted to receive a flange portion of a seal extending between the first component and the second component; the slot having a closed aft end formed with at least one cooling cavity provided with at least one cooling passage extending between the cavity and an external surface of the first component, and wherein said at least one cooling passage extends at an acute angle relative to a rotor axis of the turbine.
In still another aspect, the invention relates to a method of film cooling a turbine component formed with at least one seal slot adapted to receive a seal element, the method comprising (a) forming one or more cavities at a closed end of the seal slot; (b) forming one or more cooling passages in each of the one or more cavities, the one or more cooling passages extending between the one or more cavities and a surface of the turbine component to be cooled.
The invention will now be described in detail in connection with the drawings identified below.
With reference initially to
In accordance with a nonlimiting implementation of the invention, an aft or rearward wall of the seal slot 26 is formed to provide one or more cooling cavities 28 as best seen in
In a second exemplary but non-limiting embodiment, (also shown in
As shown in
If film cooling during such transient conditions is not regarded as critical, it would be of little or no consequence if the leg 22 of the seal 18 partially or completely blocks the flow of cooling air into the film cooling cavities 28. On the other hand, if cooling is viewed as critical even under transient conditions, one or more radial (or other) grooves 42 may be formed in the forward edge or face of the first stage nozzle 14 to insure cooling air to flow into the seal slot 26 and into the cooling cavities 28 (or 36), noting that there is some clearance between the seal leg 24 itself and the seal slot 26.
The above-described arrangements provide easy access for drilling the cooling holes or passages and allow the designer to locate those cooling holes or passages at locations where existing cavities otherwise do not provide access. In addition, by angling the cooling passages 28 as shown, the path itself has a greater length, thereby enhancing conduction cooling within the nozzle, while at the same time, enhancing cooling air film formation along the surface of the nozzle. Thus, the arrangements provide a way to apply more efficient film cooling air so as to reduce flow requirements and leakages, while increasing component life and improving engine performance.
It will also be appreciated that the cooling configurations described above are also readily employed in any stationary seal slots within the hot gas flow path of the turbine.
While the invention has been described in connection with what is presently considered to be the most practical and preferred embodiment, it is to be understood that the invention is not to be limited to the disclosed embodiment, but on the contrary, is intended to cover various modifications and equivalent arrangements included within the spirit and scope of the appended claims.
