Patent Description:
Airfoils are present in many aerodynamic applications including, but not limited to, turbines of gas turbine engines. These turbine airfoils each have a root, a tip, pressure and suction surfaces that extend from root to tip and leading and trailing edges at leading and trailing sides of the pressure and suction surfaces. In a turbine, the turbine airfoils can aerodynamically interact with high temperature and high pressure fluids to cause a rotor to rotate.

Turbine airfoils and other similar features are held in place within gas turbine engines by the root being provided with a dovetail shape (i.e., the blade dovetail) that fits within a complementarily shaped slot in a hub platform (i.e., the fan hub slot). During operations of the gas turbine engine, as the rotor rotates at increasingly higher speeds, the blade dovetail begins to mechanically deform due to centrifugal forces and to engage with pressure surfaces of the fan hub slot. It has been seen that these engagements are often provided such that the fan hub slot and the blade dovetail both experience stress concentrations at their respective ends due to higher loads being transferred at these locations.

Accordingly, a need exists for a blade dovetail or a fan hub slot that distributes loads and thus reduces stress concentrations.

<CIT> discloses a turbine rotor blade, the turbine rotor blade comprising a blade root with a flute bordering one of the two end sides of the root. The document discloses the technical features of the preamble of claim <NUM>.

According to an aspect of the invention, an aerodynamic element assembly is provided according to claim <NUM>.

The aerodynamic element may include a turbine blade.

The slot and the dovetail section may be curved.

The slot and the dovetail section may be straight.

A cross-sectional shape of the slot includes inwardly facing pressure surfaces and a cross-sectional shape of the dovetail section includes outwardly facing pressure surfaces.

The ends of each side of the slot at the opposite ends thereof may be flared outwardly by about <NUM> inches (<NUM>).

With the dovetail section assuming the initial configuration, ends of each side of the dovetail section at opposite ends thereof may be shaved inwardly and thereby configured to be spaced apart from corresponding ends of pressure surfaces of the slot. The ends of each side of the dovetail section at the opposite ends thereof may be shaved inwardly by about <NUM> inches (<NUM>).

According to another aspect of the invention, a gas turbine engine is provided aaccording to claim <NUM>.

According to another aspect of the invention, a method of assembling an aerodynamic element assembly of a gas turbine engine is provided according to claim <NUM>.

The following descriptions are by way of example only and should not be considered limiting in any way.

The fan section <NUM> drives air along a bypass flow path B in a bypass duct, while the compressor section <NUM> drives air along a core flow path C for compression and communication into the combustor section <NUM> and then expansion through the turbine section <NUM>.

The exemplary gas turbine engine <NUM> generally includes a low speed spool <NUM> and a high speed spool <NUM> mounted for rotation about an engine central longitudinal axis A relative to an engine static structure <NUM> via several bearing systems <NUM>.

A combustor <NUM> is arranged in the gas turbine engine <NUM> between the high pressure compressor <NUM> and the high pressure turbine <NUM>. The engine static structure <NUM> is arranged generally between the high pressure turbine <NUM> and the low pressure turbine <NUM>. The engine static structure <NUM> further supports the bearing systems <NUM> in the turbine section <NUM>.

The core airflow is compressed by the low pressure compressor <NUM> and then the high pressure compressor <NUM>, is mixed and burned with fuel in the combustor <NUM> and is then expanded over the high pressure turbine <NUM> and the low pressure turbine <NUM>. The high and low pressure turbines <NUM> and <NUM> rotationally drive the low speed spool <NUM> and the high speed spool <NUM>, respectively, in response to the expansion. For example, geared architecture <NUM> may be located aft of the combustor section <NUM> or even aft of the turbine section <NUM>, and the fan section <NUM> may be positioned forward or aft of the location of geared architecture <NUM>.

The gas turbine engine <NUM> in one example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine <NUM> bypass ratio is greater than about six (<NUM>), with an example embodiment being greater than about ten (<NUM>), the geared architecture <NUM> is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about <NUM> and the low pressure turbine <NUM> has a pressure ratio that is greater than about five. In one disclosed embodiment, the gas turbine engine <NUM> bypass ratio is greater than about ten (<NUM>:<NUM>), the fan diameter is significantly larger than that of the low pressure compressor <NUM>, and the low pressure turbine <NUM> has a pressure ratio that is greater than about five (<NUM>:<NUM>). The geared architecture <NUM> may be an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about <NUM>:<NUM>. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans.

The fan section <NUM> of the gas turbine engine <NUM> is designed for a particular flight condition--typically cruise at about <NUM> Mach and about <NUM>,<NUM> feet (<NUM>,<NUM> meters).

As will be described below, stress concentrations at respective ends of fan hub slots and blade dovetails (and similar features of gas turbine engines, such as the gas turbine engine <NUM> of <FIG>) are reduced by at least one of the ends of the fan hub slots being flared outwardly starting from a predefined distance from the ends and pressure surfaces of the ends of the blade dovetail being shaved starting from a predefined distance from the ends. Thus, when the respective ends of the fan hub slots and the blade dovetails start to engage as the blade dovetail deforms due to increasing rotational speed of the gas turbine engine, load transfer is distributed along extended lengths of the blade dovetails and stress concentrations at the respective ends are reduced.

With reference to <FIG>, an aerodynamic element assembly <NUM> is provided and includes a hub <NUM> and an aerodynamic element <NUM>. The hub <NUM> includes a body <NUM> that is formed to define a slot <NUM>. The slot <NUM> has a cross-sectional shape that includes radially inwardly facing pressure surfaces <NUM> on either side of the slot <NUM>. The aerodynamic element <NUM> includes a dovetail section <NUM> and an airfoil section <NUM> (see <FIG>). The dovetail section <NUM> is receivable in the slot <NUM> and includes radially outwardly facing pressure surfaces <NUM> on either side of the dovetail section <NUM> to register and engage with the pressure surfaces <NUM> of the slot <NUM>. The airfoil section <NUM> extends radially from the dovetail section <NUM> and is configured to aerodynamically interact with a working fluid to drive rotations of the hub <NUM> and the aerodynamic element <NUM> around a rotational axis (i.e., the engine central longitudinal axis A of <FIG>).

During operations of the aerodynamic element assembly <NUM>, the dovetail section <NUM> is deformable from an initial configuration, in which the dovetail section assumes an undeformed condition, to a deformed configuration. With the dovetail section <NUM> assuming the initial configuration, at least one or both of ends <NUM> (see <FIG>) of the slot <NUM> are flared and ends <NUM> (see <FIG>) of the dovetail section <NUM> are shaved. That is, with the dovetail section <NUM> assuming the initial configuration and with the dovetail section <NUM> being received in the slot <NUM>, ends <NUM> of the pressure surfaces <NUM> of the slot <NUM> can be flared outwardly from corresponding portions of the pressure surfaces <NUM> of the dovetail section <NUM> so as to be spaced apart from the corresponding portions of the pressure surfaces <NUM> of the dovetail section <NUM>. Conversely, with the dovetail section <NUM> assuming the initial configuration and with the dovetail section <NUM> being received in the slot <NUM>, ends <NUM> of the pressure surfaces <NUM> of the dovetail section <NUM> can be shaved inwardly from corresponding portions of the pressure surfaces <NUM> of the slot <NUM> so as to be spaced apart from the corresponding portions of the pressure surfaces <NUM> of the slot <NUM>.

In addition, it is to be understood that, with the dovetail section <NUM> assuming the initial configuration and with the dovetail section <NUM> being received in the slot <NUM>, the ends <NUM> of the pressure surfaces <NUM> of the slot <NUM> and the ends <NUM> of the pressure surfaces <NUM> of the dovetail section <NUM> can be flared outwardly and shaved inwardly, respectively, so as to be spaced apart from one another.

In accordance with embodiments, the aerodynamic assembly <NUM> of <FIG> can be provided at various locations throughout a gas turgine engine, such as a hub/dovetail interface of the gas turbine engine <NUM> of <FIG>, as well as other types of engines. As such, as shown in <FIG> and <FIG> in particular, the aerodynamic element <NUM> can include or be provided as a turbine blade <NUM> and the hub <NUM> can include or be provided as a fan hub <NUM> of the turbine section <NUM> (described above) and/or the aerodynamic element <NUM> can include or be provided as a blade <NUM> and the hub <NUM> can include or be provided as a bladed disk <NUM> of the compressor section <NUM> (described above).

In any case, with reference to <FIG>, both the slot <NUM> and the dovetail section <NUM> can be curved (see <FIG>) or both the slot <NUM> and the dovetail section <NUM> can be straight (see <FIG>).

For those cases in which the ends <NUM> of the pressure surfaces <NUM> of the slot <NUM> are flared outwardly from the corresponding portions of the pressure surfaces <NUM> of the dovetail section <NUM>, the ends <NUM> of the pressure surfaces <NUM> at each side of the slot <NUM> at each of the opposite ends of the slot <NUM> can be flared outwardly by about <NUM> inches (<NUM>) (machining tolerances are about <NUM> inches (<NUM>)). The outward flaring extends from the opposite ends of the slot <NUM> inwardly by about <NUM> inch (<NUM>).

With this construction, with the dovetail section <NUM> assuming the initial configuration and with the dovetail section <NUM> being received in the slot <NUM>, the ends <NUM> of the pressure surfaces <NUM> at each side of the slot <NUM> and at the opposite ends of the slot <NUM> are at least initially spaced from the corresponding portions of the pressure surfaces <NUM> of the dovetail section <NUM>. In the exemplary case of the aerodynamic element <NUM> including or being provided as the turbine blade <NUM> and the hub <NUM> including or being provided as the fan hub <NUM> of the turbine section <NUM>, as the rotations of the hub <NUM> and the aerodynamic element <NUM> speed up, the dovetail section <NUM> mechanically deforms toward assuming the deformed condition. At this point, outward flared pressure surfaces <NUM> of the slot <NUM> register and engage with the complementary pressure surfaces <NUM> of the dovetail section <NUM> but the contact area between the pressure surfaces <NUM> and <NUM> is extended along lengths of the slot <NUM>. In this way, transferred loads are distributed more evenly along the slot <NUM> and stress concentrations at the ends of the slot <NUM> are reduced.

For those cases in which the ends <NUM> of the pressure surfaces <NUM> of the dovetail section <NUM> are shaved inwardly from the corresponding portions of the pressure surfaces <NUM> of the slot <NUM>, the ends <NUM> of the pressure surfaces <NUM> at each side of the dovetail section <NUM> at each of the opposite ends of the slot dovetail section <NUM> can be shaved inwardly by about <NUM> inches (<NUM>) (machining tolerances are about <NUM> inches (<NUM>)). The inward shaving extends from the opposite ends of the dovetail section <NUM> inwardly by about <NUM> inch (<NUM>).

With this construction, with the dovetail section <NUM> assuming the initial configuration and with the dovetail section <NUM> being received in the slot <NUM>, the ends <NUM> of the pressure surfaces <NUM> at each side of the dovetail section <NUM> and at the opposite ends of the dovetail section <NUM> are at least initially spaced from the corresponding portions of the pressure surfaces <NUM> of the slot <NUM>. In the exemplary case of the aerodynamic element <NUM> including or being provided as the turbine blade <NUM> and the hub <NUM> including or being provided as the fan hub <NUM> of the turbine section <NUM>, as the rotations of the hub <NUM> and the aerodynamic element <NUM> speed up, the dovetail section <NUM> mechanically deforms toward assuming the deformed condition. At this point, inward shaved pressure surfaces <NUM> of the dovetail section <NUM> register and engage with the complementary pressure surfaces <NUM> of the slot <NUM> but the contact area between the pressure surfaces <NUM> and <NUM> is extended along lengths of the slot <NUM>. In this way, transferred loads are distributed more evenly along the dovetail section <NUM> and stress concentrations at the ends of the dovetail section <NUM> are reduced.

With reference to <FIG>, a method of assembling an aerodynamic element assembly of a gas turbine engine as described above is provided and includes determining an extent of stress concentrations between pressure surfaces of a dovetail section of an aerodynamic element and pressure surfaces of a slot defined in a hub (<NUM>) and at least one of flaring ends of the pressure surfaces of the slot outwardly and shaving ends of the pressure surfaces of the dovetail section inwardly in accordance with a determined extent of the stress concentrations (<NUM>).

Benefits of the features described herein are the provision of one of flared fan hub slots and shaved blade dovetails that come into contact during blade dovetail deformation such that load transfer is distributed along extended lengths of the blade dovetails and stress concentrations at the respective ends of the fan hub slots and the blade dovetails are reduced.

Claim 1:
An aerodynamic element assembly (<NUM>) for a gas turbine engine, comprising:
a hub (<NUM>) defining a slot (<NUM>); and
an aerodynamic element (<NUM>) comprising a dovetail section (<NUM>) receivable in the slot (<NUM>) and an airfoil section (<NUM>) configured to aerodynamically interact with fluid to drive hub (<NUM>) and aerodynamic element rotations around a rotational axis,
the dovetail section (<NUM>) being deformable during operational conditions from an initial configuration to a deformed configuration,
characterized in that:
with the dovetail section (<NUM>) assuming the initial configuration and with the dovetail section being received in the slot (<NUM>): axial ends (<NUM>) of each side of the pressure surfaces (<NUM>) of the slot (<NUM>) at opposite ends thereof are flared outwardly and thereby configured to be spaced apart from corresponding ends (<NUM>) of pressure surfaces (<NUM>) of the dovetail section (<NUM>).