Power plants – Combustion products used as motive fluid – Combustion products generator
Reexamination Certificate
1999-02-12
2001-02-13
Thorpe, Timothy S. (Department: 3746)
Power plants
Combustion products used as motive fluid
Combustion products generator
C415S177000, C415S178000
Reexamination Certificate
active
06185925
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
This invention relates generally to external cooling of turbine casings and flanges and, more particularly, to external cooling of low pressure turbine frames.
2. Discussion of the Background Art
Gas turbine engines, such as the General Electric CFM56-5B and -7B with dual annular combustors (DAC) engines, have been designed to operate both efficiently and with low amounts of pollution emissions. As a consequence, such engines have historically operated at very high EGT (Exhaust Gas Temperature) levels under ground idle and low power conditions. The turbine rear frame (TRF) and various nacelle components may experience temperatures that are high enough such that the mechanical properties of the materials are reduced and no longer acceptable from a fatigue and/or from an ultimate strength standpoint. To reduce these high temperatures, the present invention incorporates an external cooling manifold assembly capable of supplying sufficient cooling to this region of the engine such that temperature operating limits will be met or exceeded at all points of the flight mission.
Low pressure turbine (LPT) active clearance control systems which use an external manifold system to impinge fan discharge air on the LPT case have long been used for the purposes of maintaining desirable tip clearances between rotating turbine blades and respective surrounding shrouds, see U.S. Pat. No. 4,019,320, entitled “External Gas Turbine Engine Cooling For Clearance Control” as an example. Also long known in the art is external cooling of the engine case for the purposes of thermal control. Such cooling heretofore has been by flowing fan air over the casing during part or the entire engine operation.
The problem is particularly more acute for the rear turbine frame which typically supports the bearing assembly which supports the low pressure rotor and which contains the lugs through which pins are disposed to mount the engine to an aircraft pylon.
SUMMARY OF THE INVENTION
Briefly, in accordance with one aspect of the present invention, an apparatus for cooling a gas turbine engine annular turbine frame includes a first plenum box having an inlet for receiving cooling air and an annular first manifold tube in fluid communication with a first outlet of and extending in a first circumferential direction away from the plenum box. The first manifold tube is circumscribed radially outward of and around a centerline and has axially spaced apart first arcuate sections circumscribed about and perpendicular to the centerline. Each of the first arcuate sections of the first manifold tube has a corresponding portion of a first plurality of impingement cooling holes disposed through the first manifold tube and aimed generally radially inward. In one embodiment of the present invention, the first plurality of impingement cooling holes includes at least one set of impingement cooling holes disposed along a first set of arcs wherein each of the first arcs lay on one of the first arcuate sections. At least one additional set of impingement cooling holes may be disposed along at least one additional set of arcs such that each of the additional arcs lay on the one of the first arcuate sections axially spaced apart from the first arc. Preferably, the at least one set of impingement cooling holes are aimed towards a first circumference on the frame and the at least one additional set of impingement cooling holes are aimed towards a second circumference on the frame. A cooling air bleed is preferably connected to the plenum box with a selectable on and off control valve therebetween. In another particular embodiment of the invention, the cooling air bleed is in fluid communication with at least one stage of a compressor of the engine.
In another embodiment, an annular second manifold tube in fluid communication with a second outlet of the plenum box extends in a second circumferential direction away from the plenum box and is circumscribed radially outward of and around the centerline and axially spaced apart from and aft of the first manifold tube. The second manifold tube has axially spaced apart second arcuate circumscribed about and perpendicular to the centerline and each of the second arcuate sections of the second manifold tube has a corresponding portion of a second plurality of impingement cooling holes disposed through the second manifold tube and aimed generally radially inward. First and second portions of the first and second pluralities of impingement cooling holes are aimed at first and second axially spaced apart circles respectively around the frame and the first and second portions of the first and second pluralities of impingement cooling holes are in at least two of the first and second arcuate sections, respectively. Preferably, a counter flow means is used to flow cooling air in the first and second manifold tubes in opposite first and second circumferential directions, respectively. Preferably, the first and second manifold tubes are nearly 360 degrees in circumference and extend circumferentially from first and second outlets in the plenum box to first and second manifold tube ends, respectively, near the plenum box. The counter flow means may include the first and second outlets being circumferentially spaced apart in circumferentially opposite sides of the plenum box.
ADVANTAGES OF THE INVENTION
The present invention increases the effective usefulness of high operating temperature combustors by providing an effective cooling system and apparatus for cooling a turbine rear frame of a gas turbine engine. The cooling manifold system of the present invention may be an optional feature of the engine and does not have to be used on all engine models. It provides sufficient cooling capability to operate the engine within the limits imposed by various customers and even under hot day, fully deteriorated, fully loaded conditions.
The cooling system and its manifold arrangement of the present invention has a minimal impact on other engine systems. The fuel burn impact is negligible and does not negatively effect the engine's fuel efficiency. The invention has little impact on the weight of the engine. In summary, the turbine rear frame cooling apparatus of the present invention provides a light-weight fuel efficient design for aircraft gas turbine engines which is particularly useful for such engines with constructed high temperature highly fuel efficient combustors which provides an overall increase in the fuel efficiency of the entire engine.
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Epstein Michael J.
Farley Scott E.
Fraser Gordon G.
Koch Richard F.
Proctor Robert
General Electric Company
Herkamp Nathan D.
Hess Andrew C.
Thorpe Timothy S.
Torrente David J.
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