Patent Application
20100192595
This invention relates generally to gas turbine engines, and more specifically to gas turbine engine assemblies and methods of assembling the same.
At least some known gas turbine engines include a fan, a core engine, and a power turbine. The core engine includes at least one compressor, a combustor, a high-pressure turbine and a low-pressure turbine coupled together in a serial flow relationship. More specifically, the compressor and high-pressure turbine are coupled through a shaft to define a high-pressure rotor assembly. Air entering the core engine is mixed with fuel and ignited to form a high energy gas stream. The high energy gas stream flows through the high-pressure turbine to rotatably drive the high-pressure turbine such that the shaft, in turn, rotatably drives the compressor.
The gas stream expands as it flows through the low-pressure turbine positioned aft of the high-pressure turbine. The low-pressure turbine includes a rotor assembly having a fan coupled to a drive shaft. The low-pressure turbine rotatably drives the fan through the drive shaft.
Modern commercial turbofans tend toward higher bypass ratios to improve efficiency. For acoustic and fan efficiency reasons, it is desirable to reduce fan RPM or tip speed. However, a lower RPM increases low-pressure turbine loading, diameter and/or stage count. A fan directly driven by the low-pressure turbine limits the choice in fan speed because a slight reduction in fan speed for improved performance results in poorer performance in the low-pressure turbine.
In one aspect, a turbine engine assembly including, a core gas turbine engine, a first low-pressure turbine section in serial flow communication with the core gas turbine engine, the first low-pressure turbine section configured to rotate in a first rotational direction, a first gear assembly coupled to the first low-pressure turbine section, a second low-pressure turbine section coupled to the gear assembly, the second low-pressure turbine section configured to rotate in a second rotational direction, and a fan assembly coupled to the second low-pressure turbine section.
In another aspect, a method for assembling a gas turbine engine including, coupling a first section of a low pressure turbine downstream of a core gas turbine engine, wherein the first section of the low pressure turbine rotates in a first direction, coupling a first gear assembly to the first stage of the low pressure turbine, coupling a second section of the low pressure turbine to the first gear assembly, wherein the second section of the low pressure turbine rotates in a second direction, and coupling a single stage fan assembly to the second section of the low pressure turbine.
In a further aspect, a gas turbine engine assembly including, a core gas turbine engine, a first low-pressure turbine section in serial flow communication with the core gas turbine engine, a second low-pressure turbine section in serial flow communication with the core gas turbine engine aft of the first low-pressure turbine section, a fan assembly coupled to the second low-pressure turbine section, a booster compressor coupled to the first low-pressure turbine section, and a gear assembly coupled between said booster compressor and said fan assembly.
Core gas turbine engine 12 includes an outer casing 20 that defines an annular core engine inlet 22. Casing 20 surrounds a low-pressure booster compressor 24. Single-stage fan assembly 16 increases the pressure of incoming air to a first pressure level and directs a portion of the incoming air to the low-pressure booster compressor 24. Low-pressure booster compressor 24 receives air from the single-stage fan assembly 16 and facilitates increasing the pressure to a higher, second pressure level. In one embodiment, core gas turbine engine 12 is a core CFM56 gas turbine engine available from General Electric Aircraft Engines, Cincinnati, Ohio. In another embodiment a high pressure core is used and the booster is not.
In some embodiments, a high-pressure, multi-stage, axial-flow compressor 26 receives pressurized air from booster compressor 24 and further increases the pressure of the air to a third, higher pressure level. The high-pressure air is channeled to a combustor 28 and is mixed with fuel. The fuel-air mixture is ignited to raise the temperature and energy level of the pressurized air. The high energy combustion products flow to a first or high-pressure turbine 30 for driving compressor 26 through a first rotatable drive shaft 32, and then to second or low-pressure turbine 14. After driving low-pressure turbine 14, the combustion products leave turbine engine assembly 10 through an exhaust nozzle (not shown) to provide propulsive jet thrust.
In one embodiment, booster compressor 24 includes a plurality of rows of rotor blades 70 that are coupled to a respective rotor disk 72. Booster compressor 24 is positioned aft of an inlet guide vane assembly 74 and is coupled to a drive shaft 34 such that booster compressor 24 rotates at a rotational speed that is substantially equal to a rotational speed of fan assembly 16. Although booster compressor 24 is shown as having only three rows of rotor blades 70, booster compressor 24 may have any suitable number and/or rows of rotor blades 70, such as a single row of rotor blades 70 or a plurality of rows of rotor blades 70 that are interdigitated with a plurality of rows of guide vanes 76. In one embodiment, guide vanes 76 are fixedly or securely coupled to a booster case 78. In an alternative embodiment, rotor blades 70 are rotatably coupled to rotor disk 72 such that guide vanes 76 are movable during engine operation to facilitate varying a quantity of air channeled through booster compressor 24. In another alternative embodiment, turbine engine assembly 10 does not include booster compressor 24.
Low-pressure turbine 14 includes two sections, a first section 80 and a second section 82. Although first section 80 is shown with one stage and second section 82 is shown with two stages each section may have multiple or single stages in other embodiments. First section 80 is coupled to a first intermediate drive shaft 84 and rotates in a first direction with a first rotational speed while second section 82 is coupled to a second intermediate drive shaft 86 and rotates in a second direction with a second rotational speed. Both first and second intermediate drive shafts 84 and 86 are coupled to second rotatable drive shaft 34 through a gear assembly 88. In the exemplary embodiment gear assembly 88 is a planetary (star type) reversing and speed reducing gear assembly. In other embodiments, gear assembly 88 may be any other type of gear assembly.
Second rotatable drive shaft 34 drives fan assembly 16. Fan assembly 16 is configured to rotate about longitudinal axis 11 in a second rotational direction, includes at least one row of rotor blades 60, and is positioned within a fan case 64. Rotor blades 60 are coupled to a rotor disk 66.
Exemplary embodiments of a gas turbine engine assembly and methods of assembly the gas turbine engine assembly are described above in detail. The assembly and method are not limited to the specific embodiments described herein, but rather, components of the assembly and/or steps of the method may be utilized independently and separately from other components and/or steps described herein. Further, the described assembly components and/or the method steps can also be defined in, or used in combination with, other assemblies and/or methods, and are not limited to practice with only the assembly and/or method as described herein.
This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to make and use the invention. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.
