GAS TURBINE DISK CAVITY INGESTION INHIBITOR

• US 20020159880A1
• Filed: 04/26/2001
• Published: 10/31/2002
• Est. Priority Date: 04/26/2001
• Status: Active Grant

• Current Assignee: Honeywell International Inc.
• Original Assignee: Honeywell International Inc.
• Inventors: Wright, Carl D., Okpara, Nnawuihe A., Morris, Mark C., Howe, William J., Freiberg, Douglas P.
• Granted Patent: US 6,481,959 B1
• Time in Patent Office: Days
• US Class Current: 415/1
• CPC Class Codes: F01D 11/001 for sealing space between stator blade and rotor
• CPC Class Codes: F01D 25/12 Cooling
• CPC Class Codes: F01D 5/082 on the side of the rotor disc
• 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: F01D 5/145 Means for influencing boundary layers or secondary circulations for compressors F04D29/68
• CPC Class Codes: F05D 2220/50 for auxiliary power units (APU's)
• CPC Class Codes: F05D 2240/81 Cooled platforms
• CPC Class Codes: F05D 2260/202 by film cooling
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

A supplemental air cooling system for use in gas turbine engines to inhibit the ingestion of hot flow path gases into circumferential locations of turbine disk cavities is provided. The supplemental air cooling is provided through a simple set of cooling air holes located on each side of the turbine nozzle airfoil trailing edges, and proximately placed to be below the turbine nozzle structural element flow discouragers. Turbine disk cavity cooling purge air entering the disk cavity through the cooling air holes produces dynamic pressure cooling air jets which force the incoming hot ingestion air to turn circumferentially and go back out in the flow path before it enters the turbine disk cavity. The result is a decrease in hot gas ingestion, a reduction in disk rotor and static structural metal temperatures, a reduction in the amount of required cooling air flow, and enhanced performance of the gas turbine engine by virtue of improved specific fuel consumption.

Claims

1. A supplemental air cooling system for use in gas turbine engines to inhibit ingestion of hot gas flow into certain circumferential locations of turbine disk cavities, said air cooling system comprising:

a plurality of ingestion inhibiting dynamic jet orifices located on an underside of turbine nozzle trailing edges;

said orifices penetrating through turbine nozzle hot gas flow path discouraging structural elements;

said orifices being optionally located on each side of each stator blade airfoil trailing edge, or directly underneath each stator blade airfoil trailing edge;

the centerline of each said orifices directed to a point within the turbine disk cavity;

said point being located about midway between the stationary turbine nozzle and the rotating turbine blade hot flow path gas discouraging structural elements;

a flow of gas turbine engine cooling air traversing through the plurality of ingestion inhibiting dynamic jet orifices and flowing into the turbine disk cavity;

whereby the flow of cooling air traversing through the plurality of ingestion inhibiting dynamic jet orifices and into the turbine disk cavity forces incoming hot flow gases to turn circumferentially and go back out into the path of hot gas flow before it can cause an adverse increase in internal component temperatures.

2. The supplemental air cooling system of claim 1, wherein the general direction of the centerline of each ingestion inhibiting dynamic jet cooling air orifice is determined by extending an imaginary line in the airfoil trailing edge flow direction from the turbine nozzle stator blade trailing edge airfoil to a point where it intersects the trailing edge of the turbine nozzle flow discouraging structure.

3. The supplemental air cooling system of claim 1, wherein the ingestion inhibiting dynamic jet cooling air orifices are positioned circumferentially to approximately encompass a length which is about equivalent to between ¼

and ½

of the turbine stator airfoil pitch.

4. The supplemental air cooling system of claim 1, wherein the general direction of the centerline of each ingestion inhibiting dynamic jet cooling air orifice has a radial component to direct the airflow for improved turbine nozzle trailing edge cooling.

5. The supplemental air cooling system of claim 1, wherein between one to eight ingestion inhibiting dynamic jet cooling air orifices are optionally placed adjacent to each side of each stator airfoil trailing edge, or directly underneath each stator airfoil trailing edge.

6. The supplemental air cooling system of claim 1, wherein the ingestion inhibiting dynamic jet orifice diameter is about 0.010 to 0.030 inches.

7. A method of conveying cooling air for use in gas turbine engines to inhibit the ingestion of hot flow path gases into certain circumferential locations of turbine disk cavities, said air conveyance method comprising:

conveying said cooling air flow to a point within the turbine disk cavity located about midway between the stationary turbine nozzle and the rotating turbine blade hot flow path gas discouraging structural elements;

directing said cooling air into the turbine disk cavity forcing incoming hot flow gases to turn circumferentially and return to the path of hot gas flow before it can cause an adverse increase in internal component temperatures.