FBO torque reducing feature in fan shaft

A fan shaft includes a first end, a second end, and an axis, the fan shaft and configured to be coupled to a gear assembly of a gas turbine engine at the first end and to a fan more proximal to the second end than the first end such that in response to being coupled, the fan shaft can transfer torque from the gear assembly to the fan. At least one axial portion of the fan shaft satisfies the relationshipwhere T represents peak torque during fan blade off, C represents a distance from a centerline of the gas turbine engine to an outer fiber of the fan shaft, J represents a polar moment of inertia of the fan shaft, and τ represents yield stress in shear of the fan shaft.

FIELD

The present disclosure relates generally to the design of gas turbine engine fan shafts and, more particularly, to fan shaft design features intended to reduce the torque load transmitted to the engine structure during a fan blade off event.

BACKGROUND

Gas turbine engines typically include a fan section, a compressor section, a combustor section and a turbine section. A fan blade may fracture and become liberated from the fan rotor while a gas turbine engine is running. This is a potentially hazardous event known as Fan Blade Off (FBO). Immediately following blade liberation, severe rub interaction between the remaining fan blades and the fan case occurs. The rub interaction produces a large torque load which is transmitted through the fan shaft.

SUMMARY

What is described is a fan shaft having a first end, a second end, and an axis, the fan shaft and configured to be coupled to a gear assembly of a gas turbine engine at the first end and to a fan more proximal to the second end than the first end such that in response to being coupled, the fan shaft can transfer torque from the gear assembly to the fan. At least one axial portion of the fan shaft satisfies the relationship

0.55≤T*CJ*τ≤0.95,
where T represents peak torque during fan blade off, C represents a distance from a centerline of the gas turbine engine to an outer fiber of the fan shaft, J represents a polar moment of inertia of the fan shaft, and τ represents yield stress in shear of the fan shaft.

In any of the foregoing fan shafts, the fan shaft includes a first bearing configured to rotatably couple the fan shaft to a static structure.

In any of the foregoing fan shafts, the at least one portion is positioned between the first bearing and the first end of the fan shaft.

In any of the foregoing fan shafts, the fan shaft includes a first bearing and a second bearing configured to rotatably couple the fan shaft to a static structure.

In any of the foregoing fan shafts, the at least one portion is positioned between the first bearing and the second bearing.

Any of the foregoing fan shafts may further include at least one rib positioned circumferentially about an inner diameter surface of the at least one portion.

Any of the foregoing fan shafts may further include an outer groove positioned circumferentially about an outer diameter surface of the at least one portion.

In any of the foregoing fan shafts, the at least one rib is axially aligned with the outer groove.

Any of the foregoing fan shafts may further include a first inner groove positioned axially forward and adjacent to the at least one rib, the first inner groove having a forward surface extending radially outward as it approaches the at least one rib, as well as a second inner groove positioned axially aft and adjacent to the at least one rib, the second inner groove having an aft surface extending radially outward as it approaches the at least one rib.

Also described is a fan section of a gas turbine engine. The fan section includes a fan configured to rotate about a centerline. The fan section also includes a fan shaft configured to transfer torque to the fan and including a first bearing, a second bearing aft of the first bearing, and at least one rib positioned on an inner diameter surface of the fan shaft, the first bearing and the second bearing being configured to rotatably couple the fan shaft to a static structure.

In any of the foregoing fan sections, the at least one rib is positioned axially between the first bearing and the second bearing.

Any of the foregoing fan sections may also include an outer groove positioned circumferentially about an outer diameter surface of the fan shaft such that the at least one rib is axially aligned with the outer groove.

In any of the foregoing fan sections, at least a portion of the fan shaft between the first bearing and the second bearing satisfies the relationship

0.55≤T*CJ*τ≤0.95,
where T represents peak torque during fan blade off, C represents a distance from a centerline of the gas turbine engine to an outer fiber of the fan shaft, J represents a polar moment of inertia of the fan shaft, and τ represents yield stress in shear of the fan shaft.

In any of the foregoing fan sections, the fan shaft is configured to receive torque from a gear assembly and the gear assembly is configured to receive torque from a turbine section of the gas turbine engine.

Any of the foregoing fan sections may also include a first inner groove positioned axially forward and adjacent to the at least one rib, the first inner groove having a forward surface extending radially outward as it approaches the at least one rib, as well as a second inner groove positioned axially aft and adjacent to the at least one rib, the second inner groove having an aft surface extending radially outward as it approaches the at least one rib.

Also described is a gas turbine engine having an axis. The gas turbine engine includes a static structure, a low pressure turbine section, and an inner shaft coupled to the low pressure turbine section and configured to receive torque from the low pressure turbine section. The gas turbine engine also includes a gear assembly coupled to the inner shaft. The gas turbine engine also includes a fan shaft coupled to and configured to receive torque from the gear assembly, including a first bearing and a second bearing configured to allow rotation of the fan shaft relative to the static structure, the fan shaft having at least one portion that satisfies the relationship

0.55≤T*CJ*τ≤0.95.
T represents peak torque during fan blade off, C represents a distance from a centerline of the gas turbine engine to an outer fiber of the fan shaft, J represents a polar moment of inertia of the fan shaft, and τ represents yield stress in shear of the fan shaft The fan shaft also includes a fan coupled to the fan shaft and configured to rotate in response to torque being received by the fan shaft.

In any of the foregoing gas turbine engines, the at least one portion is positioned axially between the first bearing and the second bearing.

In any of the foregoing gas turbine engines, the fan shaft further includes at least one rib extending circumferentially about an inner diameter surface of the at least one portion.

In any of the foregoing gas turbine engines, the fan shaft further includes an outer groove extending circumferentially about an outer diameter surface of the at least one portion.

In any of the foregoing gas turbine engines, the fan shaft further includes a first inner groove positioned axially forward and adjacent to the at least one rib, the first inner groove having a forward surface extending radially outward as it approaches the at least one rib, as well as a second inner groove positioned axially aft and adjacent to the at least one rib, the second inner groove having an aft surface extending radially outward as it approaches the at least one rib.

DETAILED DESCRIPTION

With reference toFIG. 1, a gas turbine engine20is provided. An A-R-C axis illustrated in each of the figures illustrates the axial (A), radial (R) and circumferential (C) directions. As used herein, “aft” refers to the direction associated with the tail (e.g., the back end) of an aircraft, or generally, to the direction of exhaust of the gas turbine engine. As used herein, “forward” refers to the direction associated with the nose (e.g., the front end) of an aircraft, or generally, to the direction of flight or motion. As utilized herein, radially inward refers to the negative R direction and radially outward refers to the R direction.

Gas turbine engine20may be a two-spool turbofan that generally incorporates a fan section22, a compressor section24, a combustor section26and a turbine section28. Alternative engines include an augmentor section among other systems or features. In operation, fan section22drives air along a bypass flow-path B while compressor section24drives air along a core flow-path C for compression and communication into combustor section26then expansion through turbine section28. Although depicted as a turbofan gas turbine engine20herein, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures.

Gas turbine engine20generally comprise a low speed spool30and a high speed spool32mounted for rotation about an engine central longitudinal axis A-A′ relative to an engine static structure36via several bearing systems38,38-1, and38-2. It should be understood that various bearing systems38at various locations may alternatively or additionally be provided, including for example, bearing system38, bearing system38-1, and bearing system38-2.

Low speed spool30generally includes an inner shaft40that interconnects a fan42, a low pressure (or first) compressor section44and a low pressure (or first) turbine section46. Inner shaft40is connected to fan42through a geared architecture48that can drive fan42at a lower speed than low speed spool30. Fan42may be enclosed by a fan case90radially outward from fan42. Geared architecture48includes a gear assembly60enclosed within a gear housing62. Gear assembly60couples inner shaft40to a fan shaft70that is coupled to a rotating fan structure, including fan42. Fan shaft70is coupled to engine static structure36by bearing system which, in various embodiments, includes a first bearing202and a second bearing204that is aft of first bearing202. High speed spool32includes an outer shaft50that interconnects a high pressure (or second) compressor section52and high pressure (or second) turbine section54. A combustor56is located between high pressure compressor52and high pressure turbine54. A mid-turbine frame57of engine static structure36is located generally between high pressure turbine54and low pressure turbine46. Mid-turbine frame57supports one or more bearing systems38in turbine section28. Inner shaft40and outer shaft50are concentric and rotate via bearing systems38about the engine central longitudinal axis A-A′, which is collinear with their longitudinal axes. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine.

The core airflow C is compressed by low pressure compressor section44then high pressure compressor52, mixed and burned with fuel in combustor56, then expanded over high pressure turbine54and low pressure turbine46. Mid-turbine frame57includes airfoils59which are in the core airflow path. Turbines46,54rotationally drive the respective low speed spool30and high speed spool32in response to the expansion.

Gas turbine engine20is a high-bypass ratio geared aircraft engine. The bypass ratio of gas turbine engine20may be greater than about six (6). The bypass ratio of gas turbine engine20may also be greater than ten (10:1). Geared architecture48may be an epicyclic gear train, such as a star gear system (sun gear in meshing engagement with a plurality of star gears supported by a carrier and in meshing engagement with a ring gear) or other gear system. Geared architecture48may have a gear reduction ratio of greater than about 2.3 and low pressure turbine46may have a pressure ratio that is greater than about five (5). The diameter of fan42may be significantly larger than that of the low pressure compressor section44, and the low pressure turbine46may have a pressure ratio that is greater than about five (5:1). The pressure ratio of low pressure turbine46is measured prior to inlet of low pressure turbine46as related to the pressure at the outlet of low pressure turbine46. It should be understood, however, that the above parameters are exemplary of various embodiments of a suitable geared architecture engine and that the present disclosure contemplates other turbine engines including direct drive turbofans.

The next generation turbofan engines are designed for higher efficiency and use higher pressure ratios and higher temperatures in high pressure compressor52than are conventionally experienced. These higher operating temperatures and pressure ratios create operating environments that cause thermal loads that are higher than the thermal loads conventionally experienced, which may shorten the operational life of current components.

With reference now toFIG. 2, a portion of fan section22is illustrated. Fan shaft70has a first end280that is coupled to gear assembly60and a second end282. A fan blade42A of fan42is coupled to fan shaft70near second end282. Stated differently, fan blade42A is more proximal second end282than first end280. Fan shaft70is coupled to static structure36via one or more bearings, which may be included in bearing system38-1as a first bearing202and a second bearing204that is positioned aft of first bearing202.

Fan shaft70may have a cylindrical shape and be positioned circumferentially about a centerline261. In response to input torque being received at gear assembly60, gear assembly60causes fan shaft70to rotate about centerline261. Fan blade42A is coupled to fan shaft70such that fan42(and thus fan blade42A) rotates about centerline261at the same angular velocity as fan shaft70. In various embodiments, fan shaft70is also coupled to a spinner cap200that may rotate about centerline261along with fan shaft70and fan blade42A.

With reference now toFIGS. 1 and 2, gas turbine engine20may be subject to a fan blade off (FBO) event. During FBO, while gas turbine engine20is operating such that fan42is rotating about centerline261, an event causes fan blade42A becomes liberated from fan42and, thus, fan shaft70. This results in a rotating centrifugal load that causes fan42to skew to a side of fan42opposing fan blade42A. Other fan blades of fan42may drag on fan case90, resulting in deceleration of fan shaft70. Much of the resulting force is absorbed by fan shaft70near bearing system38-1.

In traditional gas turbine engines, fan shafts are designed to allow a relatively small amount of torsional yielding during FBO. However, fan shaft70is designed to allow a predetermined amount of torsional yielding. The torsional yielding of fan shaft70safely dissipates energy and thus reduces the potentially destructive torque load transmitted to the rest of gas turbine engine20. In particular, fan shaft70may be designed to allow torsional yielding between first bearing202and second bearing204. For example, a one percent (1%) increase in torsional yielding can reduce peak torque transmitted through the fan shaft during FBO by as much as 30%.

In order to achieve a desirable amount of torsional yielding of fan shaft70, the portion of fan shaft70designed to allow the torsional yielding (such as the portion between first bearing202and second bearing204) should satisfy equation 1 below. Satisfaction of equation 1 has an additional benefit of reducing maneuver bending loads transmitted from an overhung fan to gear assembly60.

In equation 1, τ represents yield stress in shear (material strength) and is a material property. Thus, the value of τ is based on selection of material of fan shaft70and/or heat treating of the material of fan shaft70. T represents peak torque during FBO. Thus, the value of T is based on the architecture of gas turbine engine20as well as any external loads applied, such as the charge. C represents the distance from centerline261to the outer fiber of fan shaft70, as shown by radius266. Thus, the value of C is based on the geometry of fan shaft70. J represents the polar moment of inertia of fan shaft70. The polar moment of inertia of fan shaft70is determined based on radius262from centerline261to the inner edge of fan shaft70and radius264which from centerline261to the outer metal of fan shaft70. Thus, the value of J is also based on the geometry of fan shaft70. Accordingly, the material of fan shaft70, the architecture of gas turbine engine20, external loads applied to gas turbine engine20, and/or the geometry of fan shaft70may be adjusted in order for fan shaft70to satisfy equation 1.

With reference toFIG. 3, a fan shaft370may be coupled to a gear assembly360and one or more fan blades of a fan342. InFIG. 3, a bearing system338includes a first bearing302but does not include a second bearing. Thus, it is desirable for equation 1 to be satisfied by at least one portion of fan shaft370between gear assembly360and first bearing302. For example, it may be desirable for fan shaft370to satisfy equation 1 between first bearing302and a point halfway, or one third of the way, or two thirds of the way between first bearing and gear assembly360.

With reference now toFIGS. 2 and 4, a portion438of the fan shaft70may be positioned between first bearing202and second bearing204, may extend from first bearing202to second bearing204, may be positioned between first bearing202and gear assembly60, may extend between first bearing202and gear assembly60, and/or may extend between first bearing202and a point halfway between, one third of the way to, or two thirds of the way from first bearing202and gear assembly60.

With reference directed toFIG. 4, portion438of fan shaft70may have an axial distance420. In various embodiments, portion438may satisfy equation 1 at every point over axial distance420and/or may on average satisfy equation 1 over portion438. For example, some areas of fan shaft70over axial distance420may not satisfy equation 1 and some may satisfy equation 1, but the average of fan shaft70over axial distance420may satisfy equation 1.

Portion438may have an outer diameter surface424and an inner diameter surface422radially inward from outer diameter surface424. Portion438may include one or more ribs including a rib426and a rib428. Rib426and rib428may be positioned on inner diameter surface422and may extend circumferentially about inner diameter surface422. The ribs may reduce the likelihood of fan shaft70buckling during FBO.

A first inner groove430may be positioned axially forward of rib426and a second inner groove434may be positioned axially aft of rib428. A third inner groove432may be positioned between rib426and rib428. First inner groove430may have a forward surface486that extends radially outward as it approaches rib426and second inner groove434may have an aft surface488that extends radially outward as it approaches rib428.

Portion438may have an outer groove445positioned on outer diameter surface424. Outer groove445may extend circumferentially about outer diameter surface424. Outer groove445may have an axial length that is at least half of axial distance420. In various embodiments, outer groove445is axially aligned with rib426and rib428and, in various embodiments, may be axially aligned with first inner groove430and second inner groove434. In various embodiments, portion438satisfies equation 1 at all points, or on average, between an axially forward end of forward surface486and an axially aft end of aft surface488.