Patent Description:
Gas turbine engines are known and typically include a fan delivering air into a compressor. The air is compressed and delivered into a combustor where it is mixed with fuel and ignited. Products of this combustion pass downstream over turbine rotors driving them to rotate. The turbine rotors, in turn, drive compressor and fan rotors.

Historically, the fan rotor rotated at the same speed as a turbine rotor. More recently, it has been proposed to include a gear reduction between a fan driving turbine and the fan rotor. With this change, the fan rotor may increase in diameter and rotate at slower speeds. However, the inclusion of the gear reduction raises packaging challenges.

<NPL>) discloses a prior art gas turbine engine. A prior art gas turbine engine, having the features of the preamble of claim <NUM>, is disclosed in <CIT>,.

According to the present invention, a gas turbine engine is provided according to claim <NUM>.

In another embodiment according to the previous embodiment, the gear reduction is a star gear reduction.

In another embodiment according to any of the previous embodiments, the gear reduction is equal to about <NUM>.

In another embodiment according to any of the previous embodiments, a bypass ratio may be defined as a volume of air delivered by the fan rotor into a bypass duct compared to the volume of air delivered into the compressor, and wherein the bypass ratio is greater than about <NUM>.

In a further example, the engine <NUM> bypass ratio is greater than about six (<NUM>:<NUM>), with an example embodiment being greater than about ten (<NUM>:<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> (<NUM>: <NUM>) and the low pressure turbine <NUM> has a pressure ratio that is greater than about five (<NUM>:<NUM>). In one disclosed embodiment, the 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 epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about <NUM> (<NUM>:<NUM>).

"Low corrected fan tip speed" is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R) / (<NUM> °R)]<NUM> (where °R = K × <NUM>/<NUM>).

<FIG> shows an alternative embodiment. <FIG> shows an embodiment <NUM>, wherein there is a fan drive turbine <NUM> driving a shaft <NUM> to in turn drive a fan rotor <NUM>. A gear reduction <NUM> may be positioned between the fan drive turbine <NUM> and the fan rotor <NUM>. This gear reduction <NUM> may be structured and operate like the gear reduction disclosed above. A compressor rotor <NUM> is driven by an intermediate pressure turbine <NUM>, and a second stage compressor rotor <NUM> is driven by a turbine rotor <NUM>. A combustion section <NUM> is positioned intermediate the compressor rotor <NUM> and the turbine section <NUM>.

<FIG> shows an engine <NUM> which may be a relatively small diameter engine. Fan blades <NUM> extend from a hub <NUM>. A nacelle <NUM> defines a bypass duct. A radially outer tip <NUM> of the fan blades <NUM> at an inlet end is spaced from a radially inner inlet end <NUM> of the hub <NUM>. A first radius r<NUM> can be defined between the centerline A and the point <NUM>. A second radius r<NUM> is defined between centerline A and point <NUM>. It is desirable to decrease a ratio of r<NUM>:r<NUM>. However, there are limitations on how small the ratio can be made. The ratio of r<NUM>:r<NUM> is greater than or equal to about <NUM> and less than or equal to about <NUM>.

A point <NUM> is a radially outermost or "highest" point on a curved or contoured surface leading into the compressor section <NUM>. Note, this structure could also be included in an engine as disclosed in <FIG>. A point <NUM> is defined immediately upstream of the first blade row <NUM> of the compressor section <NUM>. As can be seen, a surface <NUM> extends between points <NUM> and <NUM>. A gear reduction <NUM> is positioned intermediate the points <NUM> and <NUM>. As can be appreciated from this figure, the gear reduction includes multiple components which must be packaged radially inwardly of the surface <NUM>. In addition, the gear reduction <NUM> is positioned intermediate the fan hub <NUM> and the point <NUM>.

<FIG> shows the gear reduction <NUM>. An input shaft <NUM> is driven by the fan drive turbine <NUM>/<NUM> and, in turn, drives a sun gear <NUM>. A carrier <NUM> mounts the sun gear <NUM> and a plurality of star gears <NUM>. Star gears <NUM> are mounted on journals <NUM> which are fixed within the carrier <NUM>. As known, the sun gear <NUM> rotates and, in turn, rotates the star gears <NUM>, which then cause a ring gear <NUM> to rotate. Ring gear <NUM> drives a flexible shaft <NUM> which, in turn, drives the fan hub <NUM>.

The gear reduction <NUM> is a star gear reduction and has a gear ratio of greater than or equal to about <NUM>. In one embodiment, the gear reduction was <NUM>.

In order to package the gear reduction <NUM> within a relatively small space, a diameter Di of the gear reduction <NUM> is desirably reduced. To achieve this reduction, the journal bearings <NUM> and the ring gear <NUM> are made to be relatively axially long or extend for a relatively great distance l<NUM> measured along the axis A.

An inlet to surface <NUM> leads into the compressor. The surface <NUM> curves from a radially outermost point <NUM> radially inwardly to a point <NUM> leading into the first compressor blade row <NUM>. The gear reduction <NUM> is positioned between the radially outermost point <NUM> and point <NUM>. Point <NUM> is radially inward of a radially outermost point of ring gear <NUM>.

In this manner, an engine can be designed which has a smaller diameter than in the past. Thus, surface <NUM> can move inwardly to result in this small diameter. Also, surface <NUM> can be designed for best operation of the engine rather than being constrained by the need to package gear reduction <NUM>.

A volume of the ring gear <NUM> plus the carrier <NUM> is greater than equal to <NUM> inches<NUM> (<NUM><NUM>) and less than or equal to <NUM> inches<NUM> (<NUM><NUM>). The length L<NUM> may be greater than or equal to about <NUM> inches and less than or equal to about <NUM> inches.

A ratio of the length Li to the diameter Di may be greater than or equal to about <NUM> and less than or equal to about <NUM>.

The disclosed engine is particularly useful in lower thrust ranges. Engines having greater than or equal to <NUM>,<NUM> lbs (<NUM> kN) of thrust and less than or equal to about <NUM>,<NUM> lbs (<NUM> kN) of thrust benefit from this design.

In exemplary engines, the bypass ratio may be greater than <NUM> and, in one embodiment, may be greater than or equal to about <NUM>.

Claim 1:
A gas turbine engine (<NUM>) having, in operation, greater than or equal to <NUM>,<NUM> lbs (<NUM> kN) of thrust and less than or equal to about <NUM>,<NUM> lbs (<NUM> kN) of thrust, comprising:
a fan rotor having a hub (<NUM>) and a plurality of fan blades (<NUM>) extending radially outwardly of said hub (<NUM>);
a compressor positioned downstream of the fan rotor, the compressor having a first compressor blade row (<NUM>) defined along a rotational axis (A) of said fan rotor; and
a gear reduction (<NUM>) positioned axially between said first compressor blade row (<NUM>) and said fan rotor, said gear reduction (<NUM>) including a ring gear (<NUM>) and a carrier (<NUM>), said carrier (<NUM>) having an axial length (L<NUM>) and said ring gear (<NUM>) having an outer diameter (D<NUM>), and said gear reduction (<NUM>) is connected to drive said hub to rotate (<NUM>),
wherein a volume is defined for said carrier (<NUM>) and said ring gear (<NUM>), and said volume being designed to be greater than or equal to <NUM> inches<NUM> (<NUM><NUM>) and less than or equal to about <NUM> inches<NUM> (<NUM><NUM>),
the hub (<NUM>) has a radius (r<NUM>) defined at an inlet point of said hub (<NUM>), said fan blades (<NUM>) have a radius (r<NUM>), and a ratio of said hub radius (r<NUM>) to said fan blade radius (r<NUM>) is less than or equal to <NUM> and greater than or equal to <NUM>,
a gear ratio of said gear reduction (<NUM>) is greater than or equal to <NUM>, and
an inlet into said compressor extends radially, inwardly along a surface (<NUM>) from a radially outermost point (<NUM>) to a point (<NUM>) leading into said first compressor blade row (<NUM>), said gear reduction (<NUM>) is positioned between said radially outermost point (<NUM>) and said point (<NUM>) leading into said first compressor blade row (<NUM>), and said point (<NUM>) leading into said first compressor blade row (<NUM>) is radially inward of said ring gear (<NUM>), characterised in that:
a ratio of said axial length (L<NUM>) to said outer diameter (D<NUM>) is greater than or equal to <NUM> and less than or equal to <NUM>.