Method for setting a gear ratio of a fan drive gear system of a gas turbine engine

A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a fan section including a fan rotatable about an axis and a speed reduction device in communication with the fan. The speed reduction device includes a star drive gear system with a star gear ratio of at least 1.5. A fan blade tip speed of the fan is less than 1400 fps.

BACKGROUND

This disclosure relates to a gas turbine engine, and more particularly to a method for setting a gear ratio of a fan drive gear system of a gas turbine engine.

A gas turbine engine may include a fan section, a compressor section, a combustor section, and a turbine section. Air entering the compressor section is compressed and delivered into the combustor section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. Among other variations, the compressor section can include low and high pressure compressors, and the turbine section can include low and high pressure turbines.

Typically, a high pressure turbine drives a high pressure compressor through an outer shaft to form a high spool, and a low pressure turbine drives a low pressure compressor through an inner shaft to form a low spool. The fan section may also be driven by the inner shaft. A direct drive gas turbine engine may include a fan section driven by the low spool such that a low pressure compressor, low pressure turbine, and fan section rotate at a common speed in a common direction.

A speed reduction device, which may be a fan drive gear system or other mechanism, may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section. This allows for an overall increase in propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the speed reduction device that drives the fan section at a reduced speed such that both the turbine section and the fan section can rotate at closer to optimal speeds.

Although gas turbine engines utilizing speed change mechanisms are generally known to be capable of improved propulsive efficiency relative to conventional engines, gas turbine engine manufacturers continue to seek further improvements to engine performance including improvements to thermal, transfer and propulsive efficiencies.

SUMMARY

A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a fan section including a fan rotatable about an axis and a speed reduction device in communication with the fan. The speed reduction device includes a star drive gear system with a star gear ratio of at least 1.5. A fan blade tip speed of the fan is less than 1400 fps.

In a further non-limiting embodiment of the foregoing gas turbine engine, the speed reduction device includes a star gear system gear ratio of at least 2.6.

In a further non-limiting embodiment of either of the foregoing gas turbine engines, the speed reduction device includes a system gear ratio less than or equal to 4.1.

In a further non-limiting embodiment of any of the foregoing gas turbine engines, a bypass ratio is included that is greater than about 6.0.

In a further non-limiting embodiment of any of the foregoing gas turbine engines, the bypass ratio is between about 11.0 and about 22.0.

In a further non-limiting embodiment of any of the foregoing gas turbine engines, the star system includes a sun gear, a plurality of star gears, a ring gear, and a carrier.

In a further non-limiting embodiment of any of the foregoing gas turbine engines, each of the plurality of star gears include at least one bearing.

In a further non-limiting embodiment of any of the foregoing gas turbine engines, the carrier is fixed from rotation.

In a further non-limiting embodiment of any of the foregoing gas turbine engines, a low pressure turbine is mechanically attached to the sun gear.

In a further non-limiting embodiment of any of the foregoing gas turbine engines, a fan section is mechanically attached to the ring gear.

In a further non-limiting embodiment of any of the foregoing gas turbine engines, an input of the speed reduction device is rotatable in a first direction and an output of the speed reduction device is rotatable in a second direction opposite to the first direction.

In a further non-limiting embodiment of any of the foregoing gas turbine engines, a low pressure turbine section is in communication with the speed reduction device, The low pressure turbine section includes at least three stages and no more than four stages.

In a further non-limiting embodiment of any of the foregoing gas turbine engines, the fan blade tip speed of the fan is greater than 1000 fps.

A method of improving performance of a gas turbine engine according to another exemplary aspect of the present disclosure includes, among other things, determining fan tip speed boundary conditions for at least one fan blade of a fan section and determining rotor boundary conditions for a rotor of a low pressure turbine, The stress level utilizes constraints in the rotor of the low pressure turbine and the at least one fan blade to determine if the rotary speed of the fan section and the low pressure turbine will meet a desired number of operating cycles.

In a further non-limiting embodiment of the foregoing method, a speed reduction device connects the fan section and the low pressure turbine and includes a star gear ratio of at least about 1.5 and no more than about 4.1.

In a further non-limiting embodiment of either of the foregoing methods, a fan pressure ratio is below 1.7.

In a further non-limiting embodiment of any of the foregoing methods, a fan pressure ratio is below 1.48.

In a further non-limiting embodiment of any of the foregoing methods, a bypass ratio is between about 11 and about 22.

In a further non-limiting embodiment of any of the foregoing methods, a fan blade tip speed of the at least one fan blade is less than 1400 fps.

In a further non-limiting embodiment of any of the foregoing methods, if a stress level in the rotor or the at least one fan blade is too high to meet a desired number of operating cycles, a gear ratio of a gear reduction device is lowered and the number of stages of the low pressure turbine is increased.

In a further non-limiting embodiment of any of the foregoing methods, if a stress level in the rotor or the at least one fan blade is too high to meet a desired number of operating cycles, a gear ratio of a gear reduction device is lowered and an annular area of the low pressure turbine is increased.

DETAILED DESCRIPTION

FIG. 1schematically illustrates a gas turbine engine20. The exemplary gas turbine engine20is a two-spool turbofan engine that generally incorporates a fan section22, a compressor section24, a combustor section26and a turbine section28. Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section22drives air along a bypass flow path B, while the compressor section24drives air along a core flow path C for compression and communication into the combustor section26. The hot combustion gases generated in the combustor section26are expanded through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to two-spool turbofan engines and these teachings could extend to other types of engines, including but not limited to, three-spool engine architectures.

The exemplary gas turbine engine20generally includes a low speed spool30and a high speed spool32mounted for rotation about an engine centerline longitudinal axis A. The low speed spool30and the high speed spool32may be mounted relative to an engine static structure33via several bearing systems31. It should be understood that other bearing systems31may alternatively or additionally be provided, and the location of bearing systems31may be varied as appropriate to the application.

The low speed spool30generally includes an inner shaft34that interconnects a fan36, a low pressure compressor38and a low pressure turbine39. The inner shaft34can be connected to the fan36through a speed change mechanism, which in exemplary gas turbine engine20is illustrated as a geared architecture45, such as a fan drive gear system50(seeFIGS. 2 and 3). The speed change mechanism drives the fan36at a lower speed than the low speed spool30. The high speed spool32includes an outer shaft35that interconnects a high pressure compressor37and a high pressure turbine40. In this embodiment, the inner shaft34and the outer shaft35are supported at various axial locations by bearing systems31positioned within the engine static structure33.

A combustor42is arranged in exemplary gas turbine20between the high pressure compressor37and the high pressure turbine40. A mid-turbine frame44may be arranged generally between the high pressure turbine40and the low pressure turbine39. The mid-turbine frame44can support one or more bearing systems31of the turbine section28. The mid-turbine frame44may include one or more airfoils46that extend within the core flow path C. It will be appreciated that each of the positions of the fan section22, compressor section24, combustor section26, turbine section28, and fan drive gear system50may be varied. For example, gear system50may be located aft of combustor section26or even aft of turbine section28, and fan section22may be positioned forward or aft of the location of gear system50.

The inner shaft34and the outer shaft35are concentric and rotate via the bearing systems31about the engine centerline longitudinal axis A, which is co-linear with their longitudinal axes. The core airflow is compressed by the low pressure compressor38and the high pressure compressor37, is mixed with fuel and burned in the combustor42, and is then expanded over the high pressure turbine40and the low pressure turbine39. The high pressure turbine40and the low pressure turbine39rotationally drive the respective high speed spool32and the low speed spool30in response to the expansion.

In a non-limiting embodiment, the gas turbine engine20is a high-bypass geared aircraft engine. In a further example, the gas turbine engine20bypass ratio is greater than about six (6:1). The geared architecture45can include an epicyclic gear train, such as a planetary gear system, a star gear system, or other gear system. The geared architecture45enables operation of the low speed spool30at higher speeds, which can enable an increase in the operational efficiency of the low pressure compressor38and low pressure turbine39, and render increased pressure in a fewer number of stages.

The pressure ratio of the low pressure turbine39can be pressure measured prior to the inlet of the low pressure turbine39as related to the pressure at the outlet of the low pressure turbine39and prior to an exhaust nozzle of the gas turbine engine20. In one non-limiting embodiment, the bypass ratio of the gas turbine engine20is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor38, and the low pressure turbine39has a pressure ratio that is greater than about five (5:1). In another non-limiting embodiment, the bypass ratio is greater than 11 and less than 22, or greater than 13 and less than 20. It should be understood, however, that the above parameters are only exemplary of a geared architecture engine or other engine using a speed change mechanism, and that the present disclosure is applicable to other gas turbine engines, including direct drive turbofans. In one non-limiting embodiment, the low pressure turbine39includes at least one stage and no more than eight stages, or at least three stages and no more than six stages. In another non-limiting embodiment, the low pressure turbine39includes at least three stages and no more than four stages.

In this embodiment of the exemplary gas turbine engine20, a significant amount of thrust is provided by the bypass flow path B due to the high bypass ratio. The fan section22of the gas turbine engine20is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. This flight condition, with the gas turbine engine20at its best fuel consumption, is also known as bucket cruise Thrust Specific Fuel Consumption (TSFC). TSFC is an industry standard parameter of fuel consumption per unit of thrust.

Fan Pressure Ratio is the pressure ratio across a blade of the fan section22without the use of a Fan Exit Guide Vane system. The low Fan Pressure Ratio according to one non-limiting embodiment of the example gas turbine engine20is less than 1.45. In another non-limiting embodiment of the example gas turbine engine20, the Fan Pressure Ratio is less than 1.38 and greater than 1.25. In another non-limiting embodiment, the fan pressure ratio is less than 1.48. In another non-limiting embodiment, the fan pressure ratio is less than 1.52. In another non-limiting embodiment, the fan pressure ratio is less than 1.7. Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)]0.5, where T represents the ambient temperature in degrees Rankine. The Low Corrected Fan Tip Speed according to one non-limiting embodiment of the example gas turbine engine20is less than about 1150 fps (351 m/s). The Low Corrected Fan Tip Speed according to another non-limiting embodiment of the example gas turbine engine20is less than about 1400 fps (427 m/s). The Low Corrected Fan Tip Speed according to another non-limiting embodiment of the example gas turbine engine20is greater than about 1000 fps (305 m/s).

FIG. 2schematically illustrates the low speed spool30of the gas turbine engine20. The low speed spool30includes the fan36, the low pressure compressor38, and the low pressure turbine39. The inner shaft34interconnects the fan36, the low pressure compressor38, and the low pressure turbine39. The inner shaft34is connected to the fan36through the fan drive gear system50. In this embodiment, the fan drive gear system50provides for counter-rotation of the low pressure turbine39and the fan36. For example, the fan36rotates in a first direction D1, whereas the low pressure turbine39rotates in a second direction D2 that is opposite of the first direction D1.

FIG. 3illustrates one example embodiment of the fan drive gear system50incorporated into the gas turbine engine20to provide for counter-rotation of the fan36and the low pressure turbine39. In this embodiment, the fan drive gear system50includes a star gear system with a sun gear52, a ring gear54disposed about the sun gear52, and a plurality of star gears56having journal bearings57positioned between the sun gear52and the ring gear54. A fixed carrier58carries and is attached to each of the star gears56. In this embodiment, the fixed carrier58does not rotate and is connected to a grounded structure55of the gas turbine engine20.

The sun gear52receives an input from the low pressure turbine39(seeFIG. 2) and rotates in the first direction D1 thereby turning the plurality of star gears56in a second direction D2 that is opposite of the first direction D1. Movement of the plurality of star gears56is transmitted to the ring gear54which rotates in the second direction D2 opposite from the first direction D1 of the sun gear52. The ring gear54is connected to the fan36for rotating the fan36(seeFIG. 2) in the second direction D2.

A star system gear ratio of the fan drive gear system50is determined by measuring a diameter of the ring gear54and dividing that diameter by a diameter of the sun gear52. In one embodiment, the star system gear ratio of the geared architecture45is between 1.5 and 4.1. In another embodiment, the system gear ratio of the fan drive gear system50is between 2.6 and 4.1. When the star system gear ratio is below 1.5, the sun gear52is relatively much larger than the star gears56. This size differential reduces the load the star gears56are capable of carrying because of the reduction in size of the star gear journal bearings57. When the star system gear ratio is above 4.1, the sun gear52may be much smaller than the star gears56. This size differential increases the size of the star gear56journal bearings57but reduces the load the sun gear52is capable of carrying because of its reduced size and number of teeth. Alternatively, roller bearings could be used in place of journal bearings57.

Improving performance of the gas turbine engine20begins by determining fan tip speed boundary conditions for at least one fan blade of the fan36to define the speed of the tip of the fan blade. The maximum fan diameter is determined based on the projected fuel burn derived from balancing engine efficiency, mass of air through the bypass flow path B, and engine weight increase due to the size of the fan blades.

Boundary conditions are then determined for the rotor of each stage of the low pressure turbine39to define the speed of the rotor tip and to define the size of the rotor and the number of stages in the low pressure turbine39based on the efficiency of low pressure turbine39and the low pressure compressor38.

Constraints regarding stress levels in the rotor and the fan blade are utilized to determine if the rotary speed of the fan36and the low pressure turbine39will meet a desired number of operating life cycles. If the stress levels in the rotor or the fan blade are too high, the gear ratio of the fan drive gear system50can be lowered and the number of stages of the low pressure turbine39or annular area of the low pressure turbine39can be increased.

Although the different non-limiting embodiments are illustrated as having specific components, the embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from any of the non-limiting embodiments in combination with features or components from any of the other non-limiting embodiments.

It should be understood that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be understood that although a particular component arrangement is disclosed and illustrated in these exemplary embodiments, other arrangements could also benefit from the teachings of this disclosure.