Power plants – Reaction motor – Interrelated reaction motors
Reexamination Certificate
2002-03-01
2004-05-11
Casaregola, Louis J. (Department: 3746)
Power plants
Reaction motor
Interrelated reaction motors
C060S268000
Reexamination Certificate
active
06732502
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
The invention relates to counter rotating aircraft gas turbine engines with counter rotating fans driven by counter rotating low pressure turbine rotors and, particularly, for such engines having high bypass and overall compressor ratios and low hub to tip ratios.
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. Some low pressure turbines have been designed with counter rotating turbines that power counter rotating fans and booster or low pressure compressors. U.S. Pat. Nos. 4,860,537, 5,307,622 and 4,790,133 disclose counter rotating turbines that power counter rotating fans and booster or low pressure compressors. Most of the thrust produced is generated by the fan.
Large modern commercial turbofan engines have higher operating efficiencies with higher bypass ratio configurations and larger transition ducts between low pressure and high pressure turbines. The frames, especially those located in the engine hot section, are complex and expensive. These engines feature high by pass ratio configurations that yield high propulsive efficiency and with large diameter fans rotating at low tip speeds that enable low noise and high fan efficiency with a corresponding fuel consumption reduction. The low speed of the fan rotor, which is beneficial to the fan, can have an adverse impact on the low pressure turbine configuration that benefits from higher rotational speeds that reduce aerodynamic loading and improve efficiency. These conflicting objectives requirements necessitate compromises in low pressure turbine and fan efficiencies, stage counts, and transition duct lengths between the core engine and low pressure turbine. These compromises lead to heavy and costly engine configurations. It is highly desirable to produce aircraft gas turbine engines with significantly lower levels of noise, weight, specific fuel consumption, and cost.
SUMMARY OF THE INVENTION
An aircraft gas turbine engine has a high pressure rotor including a high pressure turbine and a low pressure turbine having counter rotating low pressure inner and outer rotors located aft of the high pressure rotor. The low pressure inner and outer rotors include low pressure inner and outer shafts which are at least in part rotatably disposed co-axially with and radially inwardly of the high pressure rotor. A rotatable annular outer drum rotor is drivingly connected to a first fan blade row and a first booster by the low pressure inner shaft. A rotatable annular inner drum rotor is drivingly connected to a second fan blade row and a second booster by the low pressure outer shaft. The first and second boosters are axially located between the first and second fan blade rows.
A bypass duct is radially bounded by a fan casing and an annular radially inner bypass duct wall which surrounds the first and second boosters. A radially outer portion of the second fan blade row is radially disposed within the bypass duct. The engine has a fan inlet hub to tip radius ratio in a range between 0.20 and 0.35, a bypass ratio in a range of 5-15, an operational fan pressure ratio in a range of 1.4-2.5, and a sum of operational fan tip speeds of the first and second fan blade rows in a range of 1000 to 2500 feet per second. A high pressure compressor of the high pressure rotor is drivenly connected to the high pressure turbine by a high pressure shaft and the high pressure compressor is designed and operable to produce a compressor pressure ratio in a range of about 15-30 and overall pressure ratio in a range of about 40-65. The engine is designed such that the last stage of the booster and, in the exemplary embodiment, the second fan blade row are counter rotatable with respect to the high pressure compressor.
In the exemplary embodiment of the invention, the high pressure compressor includes between six and eight high pressure stages and about four variable vane stages. Less than four variable vane stages may be used. The first booster includes an integrally bladed annular first booster rotor section including a rotatable wall section from which axially spaced apart first booster blade rows extend radially inwardly. An outlet guide vane assembly is located directly aft of the low pressure turbine.
The invention also includes an aircraft gas turbine engine family having at least two different engine models or variations of the engine with substantially the same fan diameter. A first one of the engine models has a one stage high pressure turbine and a second one of the engine models has a two stage high pressure turbine.
Further embodiments of the invention include a second seal in sealing arrangement between forward ends of the low pressure turbine casing and the outer drum rotor, a third seal in sealing arrangement between the low pressure turbine casing and a final stage of the low pressure turbine blade rows which is bolted to an aft end of the outer drum rotor, and a first seal in sealing arrangement between the second fan and the fan frame. The seals are brush seals, however in other embodiments the seals may be non contacting seals or a combination of brush seals and non-contacting seals. The non-contacting seals may be aspirating seals or face seals.
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Brauer John C.
Dunbar Lawrence W.
Johnson James E.
Seda Jorge F.
Szucs Peter N.
Casaregola Louis J.
General Electric Company
Herkamp Nathan D.
Rosen Steven J.
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