Claims
- 1. In a gas turbine engine having a compressor assembly that includes a plurality of rotary compressor stages interconnected in torque transmitting relationship by rotary shafting such that the torque transmittal path between the compressor stages is disposed substantially radially outward of the axis of rotation of the compressor stages, a cooling airflow pickup system comprising:
- a collector arranged to collect pressurized, substantially nonswirling, cooling airflow from said compressor assembly at a first location radially outward of said shafting;
- an annular rotary member in said shafting having a plurality of straight, radially inwardly extending pickup holes arranged to receive the nonswirling airflow from said collector means, said pickup holes being circumferentially spaced about said annular member and being inclined reversely to the direction of rotation of said shafting at a preselected angle to a direct radial direction, such that the cooling airflow may pass radially inwardly through said pickup holes with minimal static pressure loss; and
- a vortex spoiler interconnected to rotate with said shafting, said spoiler defining a path extending radially inwardly from said pickup holes to a second location substantially closer to the axis of rotation than said first location, said spoiler including a plurality of circumferentially spaced blades smoothly curved reversely to the direction of rotation of said shafting and axially spanning said path to divide it into a plurality of minimally converging passages, the radially outer end of each of said blades being inclined in matching relation to said preselected angle of the pickup holes, and the radially inner end of each of said blades being in a substantially directly radial direction, such that cooling airflow may pass radially inwardly through said passages to said second location with minimal swirl increase and minimal static pressure loss.
- 2. In a gas turbine engine as set forth in claim 1, wherein at least one of said compressor stages comprises a unitary disc with a plurality of radially outwardly extending compressor blades integrally formed with said disc.
- 3. In a gas turbine engine as set forth in claim 2, wherein said compressor assembly further includes sets of stationary, deswirling diffuser vanes interspersed between said compressor stages, said collector arranged to receive to airflow from said compressor assembly as it discharges from a set of said stationary vanes.
- 4. In a gas turbine engine as set forth in claim 1, wherein said rotary member includes an axially extending cylindrical wall and a radial wall extending inwardly substantially perpendicularly to said axis of rotation to define a radial wall of said path, said pickup holes being disposed in said axially extending cylindrical wall.
- 5. In a gas turbine engine as set forth in claim 4, wherein said vortex spoiler includes a radial wall extending inwardly perpendicularly to said axis of rotation to define another radial wall of said path, said blades being integrally attached to said radial wall of the spoiler and extending axially across said path to said radial wall of the rotary member.
- 6. In a gas turbine engine as set forth in claim 5, wherein said spoiler further includes a plurality of bolts passing axially through said blades for firmly intersecuring said radial walls of the rotary member and spoiler.
- 7. In a gas turbine engine as set forth in claim 1, wherein said preselected angle is approximately 45 degrees.
- 8. In a gas turbine engine as set forth in claim 7, wherein the camber line of each of said blades is a circular arc.
- 9. A gas turbine engine, comprising:
- a compressor assembly including a plurality of rotary compressor stages;
- rotary shafting interconnecting said compressor stages in torque transmitting relationship such that the torque transmittal path between the compressor stages is disposed substantially radially outward of the axis of rotation of the compressor stages;
- a collector arranged to collect pressurized, substantially nonswirling, cooling airflow from said compressor assembly at a first location radially outward of said shafting;
- an annular rotary member in said shafting having a plurality of straight, radially inwardly extending pickup holes arranged to receive the nonswirling airflow from said collector means, said pickup holes being circumferentially spaced about said annular member and being inclined reversely to the direction of rotation of said shafting at a preselected angle to a direct radial direction; and
- a vortex spoiler interconnected to rotate with said shafting, said spoiler defining a path extending radially inwardly from said pickup holes to a second location substantially closer to the axis of rotation than said first location, said spoiler including a plurality of circumferentially spaced blades smoothly curved reversely to the direction of rotation of said shafting and axially spanning said path to divide it into a plurality of minimally converging passages, the radially outer end of each of said blades being inclined in matching relation to said preselected angle of the pickup holes, and the radially inner end of each of said blades being in a substantially directly radial direction.
- 10. In a gas turbine engine as set forth in claim 9 wherein at least one of said compressor stages comprises a unitary disc with a plurality of radially outwardly extending compressor blades integrally formed with said disc.
- 11. In a gas turbine engine as set forth in claim 9, wherein said compressor assembly further includes sets of stationary, deswirling diffuser vanes interspersed between said compressor stages, said collector arranged to receive to airflow from said compressor assembly as it discharges from a set of said stationary vanes.
- 12. In a gas turbine engine as set forth in claim 9, wherein said rotary member includes an axially extending cylindrical wall and a radial wall extending inwardly substantially perpendicular to said axis of rotation to define a radial wall of said path, said pickup holes being disposed in said axially extending cylindrical wall.
- 13. In a gas turbine engine as set forth in claim 12, wherein said vortex spoiler includes a radial wall extending inwardly perpendicularly to said axis of rotation to define another radial wall of said path, said blades being integrally attached to said radial wall of the spoiler and extending axially across said path to said radial wall of the rotary member.
- 14. In a gas turbine engine as set forth in claim 9, wherein said preselected angle is approximately 45 degrees.
- 15. In a gas turbine engine as set forth in claim 14, wherein the camber line of each of said blades is a circular arc.
CROSS REFERENCE TO RELATED APPLICATION
Priority is claimed to provisional application Serial No. 60/039,498, filed May 16, 1979 now abandoned.
US Referenced Citations (17)
Foreign Referenced Citations (3)
Number |
Date |
Country |
4214753 |
Nov 1993 |
DEX |
1355769 |
Nov 1987 |
RUX |
712051 |
Jul 1954 |
GBX |