Power plants – Reaction motor – Interrelated reaction motors
Reexamination Certificate
2001-11-29
2004-03-23
Casaregola, Louis J. (Department: 3746)
Power plants
Reaction motor
Interrelated reaction motors
C060S797000
Reexamination Certificate
active
06708482
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
The invention relates to aircraft gas turbine engines and, particularly, for such engines having frames that support the rotors in bearings and are used to mount the engines to the aircraft.
2. Description of Related Art
A gas turbine engine of the turbofan type generally includes a forward fan and booster compressor, a middle core engine, and an aft low pressure power turbine. The core engine includes a high pressure compressor, a combustor and a high pressure turbine in a serial flow relationship. The high pressure compressor and high pressure turbine of the core engine are interconnected by a high pressure shaft. The high pressure compressor, turbine, and shaft essentially form the high pressure rotor. The high pressure compressor is rotatably driven to compress air entering the core engine to a relatively high pressure. This high pressure air is then mixed with fuel in the combustor and ignited to form a high energy gas stream. The gas stream flows aft and passes through the high pressure turbine, rotatably driving it and the high pressure shaft which, in turn, rotatably drives the compressor.
The gas stream leaving the high pressure turbine is expanded through a second or low pressure turbine. The low pressure turbine rotatably drives the fan and booster compressor via a low pressure shaft, all of which form the low pressure rotor. The low pressure shaft extends through the high pressure rotor. Most of the thrust produced is generated by the fan. Engine frames are used to support and carry the bearings which, in turn, rotatably support the rotors. Conventional turbofan engines have a fan frame, a mid-frame, and an aft turbine frame. Bearing supporting frames are heavy and add weight, length, and cost to the engine.
Large modern commercial turbofan engines have higher operating efficiencies with higher by pass ratio configurations, larger transition ducts between low pressure and high pressure turbines. The frames, especially those located in the engine hot section, are complex and expensive. Other mid-size turbofan engines eliminate one frame by providing HP rotor support through a differential bearing arrangement in which the high pressure rotor rides on the low pressure rotor with an inter-shaft or differential bearing between them. New commercial engine designs are incorporating counter-rotating rotors for improved turbine efficiency. Counter-rotating rotors can have a detrimental impact on high pressure ratio components clearances especially in the hot section which rely on tight clearance control to provide fuel efficiency benefits.
Consequently, a need exists for an alternative bearing support assembly which will avoid the above mentioned drawbacks and reduce, engine, length, weight and cost and tip improve clearance performance.
SUMMARY OF THE INVENTION
An aircraft gas turbine engine turbine frame includes a first structural ring, a second structural ring disposed co-axially with and radially spaced inwardly of the first structural ring about a centerline axis, and a plurality of circumferentially spaced apart struts extending radially between the first and second structural rings. Forward and aft sump members having forward and aft central bores are fixedly joined to forward and aft portions of the turbine frame, respectively. A frame connecting means for connecting the engine to an aircraft is disposed on the first structural ring. The forward and aft central bores may be cylindrical and the frame connecting means may include at least one U-shaped clevis.
One embodiment of the invention is a gas turbine engine assembly wherein the frame is an inter-turbine frame axially located between first and second turbines of first and second rotors, respectively. The first turbine is located forward of the second turbine and the second rotor includes a second shaft which is at least in part rotatably disposed co-axially with and radially inwardly of the first rotor. The second rotor is supported by a respective aftwardmost second turbine frame bearing mounted in the aft central bore of the aft sump member and the first rotor is partly supported by a respective first turbine frame bearing mounted in the forward central bore of the forward sump member. An axial center of gravity of the second turbine passes though or very near the second turbine frame bearing. In a more particular embodiment of the invention, the second turbine includes a turbine disk assembly having axially adjacent rotor disks interconnected by structural disk forward and aft spacer arms, respectively. The turbine disk assembly is connected to the second shaft at or near the axial center of gravity. A conical shaft extension may be used to drivingly connect the turbine disk assembly to the second shaft. The conical shaft extension is connected to the turbine disk assembly at or near the axial center of gravity. The rotor disks have hubs connected to rims by webs extending radially outwardly from the hubs, each of the rotor disks supports a row of blades supported in the disk rim.
The aft sump member may have a first radius as measured from the engine centerline axis that is substantially greater than a second radius of the forward sump members. The first radius may be in a range of 150 to 250 percent larger than the second radius.
The present invention replaces a turbine rear frame with an outer guide vane assembly that results in cost and weight reduction benefits by using the turbine transition duct spacing to incorporate an inter-turbine frame to rotatably support both HP and LP rotors. Improved clearance performance results from LP shaft critical speed being disengaged from the HP rotor speed influence. Mounting the low pressure turbine bearing between turbines improves clearance performance because the low pressure turbine bearing diameter is increased resulting in increased stiffness of the low pressure turbine support. Increasing the low pressure turbine bearing diameter also results in reduction of the length of the low pressure turbine shaft LP shaft cone.
REFERENCES:
patent: 3844115 (1974-10-01), Freid
patent: 4558564 (1985-12-01), Bouiller et al.
patent: 4790133 (1988-12-01), Stuart
patent: 4860537 (1989-08-01), Taylor
patent: 4965994 (1990-10-01), Ciokajlo et al.
patent: 4969325 (1990-11-01), Adamson et al.
patent: 4976102 (1990-12-01), Taylor
patent: 5160251 (1992-11-01), Ciokajlo
patent: 5232339 (1993-08-01), Plemmons et al.
patent: 5307622 (1994-05-01), Ciokajlo
patent: 5361580 (1994-11-01), Ciokajlo et al.
patent: 5443229 (1995-08-01), O'Brien et al.
patent: 5452575 (1995-09-01), Freid
patent: 5634767 (1997-06-01), Dawson
patent: 5813214 (1998-09-01), Moniz et al.
patent: 5860275 (1999-01-01), Newton et al.
patent: 6185925 (2001-02-01), Proctor et al.
Andes William S.
Casaregola Louis J.
Rosen Steven J.
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