Patent Description:
<CIT> discloses a prior art turboshaft engine in accordance with the preamble of claim <NUM>, and a prior art method in accordance with the preamble of claim <NUM>.

<CIT> discloses a prior art gas turbine engine with a gear arrangement, <CIT> discloses a prior art gas turbine engine with a speed reduction device, and <CIT> discloses a brake for a turbine rotor. Relevant disclosures are also provided by "<NPL>, and "<NPL>.

According to a first aspect of the present invention, there is provided a turboshaft engine as set forth in claim <NUM>.

In an embodiment of the above, the speed change mechanism is an epicyclic gear train.

In a further embodiment of any of the above, the epicyclic gear train is a star gear system. The output is fixed to a ring gear and a carrier is fixed from rotation relative to an engine static structure with a gear ratio of about <NUM> to about <NUM>.

In a further embodiment of any of the above, the epicyclic gear train is a planet gear system. The output is fixed to a carrier and a ring gear is fixed from rotation relative to an engine static structure with a gear ratio of about <NUM> to <NUM>.

In a further embodiment of any of the above, the speed change mechanism is non-epicyclic and includes a gear ratio of about <NUM> to about <NUM>.

In a further embodiment of any of the above, the low pressure compressor includes a pressure ratio of greater than about <NUM> and less than about <NUM>. The high pressure compressor includes a pressure ratio of greater than about <NUM> and less than about <NUM>.

In a further embodiment of any of the above, a pressure ratio of the low pressure compressor is greater than about <NUM> and less than about <NUM>.

In a further embodiment of any of the above, the low pressure turbine includes at least one rotor stage and less than four rotor stages.

According to a further aspect of the present invention, there is provided a method as set forth in claim <NUM>.

In an embodiment of the above, the speed change mechanism is a star gear system. The output turboshaft is fixed to a ring gear and a carrier is fixed from rotation relative to an engine static structure with a gear ratio of about <NUM> to about <NUM>.

In a further embodiment of any of the above, the speed change mechanism is a planet gear system and the output turboshaft is fixed to carrier and a ring gear is fixed from rotation relative to an engine static structure with a gear ratio of about <NUM> to <NUM>.

The gas turbine engine may be a turboshaft engine as described in any of the above embodiments.

<FIG> schematically illustrates a gas turbine engine <NUM> according to a first non-limiting embodiment. The gas turbine engine <NUM> is disclosed herein as a two-spool turboshaft engine that generally incorporates an intake section <NUM>, a compressor section <NUM>, a combustor section <NUM>, and a turbine section <NUM>. The intake section <NUM> accepts air into an intake <NUM> and drives the air along a core flow path C into the compressor section <NUM> for compression and communication into the combustor section <NUM> 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>.

In one non-limiting embodiment, the low speed spool <NUM> and the high speed spool <NUM> are each supported by two separate bearing systems <NUM>. In another non-limiting embodiment (outside the wording of the claims), the low speed spool <NUM> and the high speed spool <NUM> are supported by a total of at least four bearing systems <NUM> and no more than ten bearing systems <NUM>. Furthermore, although the bearing systems <NUM> are depicted as ball bearings in the illustrated embodiment, other bearings, such as thrust bearings, roller bearings, journal bearings, or tapered bearings could be used to support the low speed spool <NUM> and the high speed spool <NUM>.

The low speed spool <NUM> generally includes an inner shaft <NUM> that interconnects a first (or low) pressure compressor <NUM> and a first (or low) pressure turbine <NUM>. The inner shaft <NUM> is connected to an output driveshaft <NUM> through a speed change mechanism, which in exemplary gas turbine engine <NUM> is illustrated as a geared architecture <NUM>, to turn the output driveshaft <NUM> at a lower rotational speed than the low speed spool <NUM>. The output driveshaft <NUM> is located on an axially forward end of the gas turbine engine <NUM> opposite the turbine section <NUM>. In another non-limiting embodiment, the output driveshaft <NUM> is located at an axially downstream end of the gas turbine engine <NUM>. In this disclosure, axial or axially is in relation to the axis A unless stated otherwise.

A combustor <NUM> is arranged in the exemplary gas turbine engine <NUM> between the high pressure compressor <NUM> and the high pressure turbine <NUM>. The output driveshaft <NUM> also rotates about the axis A. One of the bearing systems <NUM> may be located forward or aft of the high pressure turbine <NUM> such that one of the bearing systems <NUM> is associated with the mid-turbine frame <NUM>.

Due to the environment in which the gas turbine engine <NUM> may be operating, there is a need to separate particles, such as sand, dirt, or other debris, from the core flow path C entering the gas turbine engine <NUM>. Particles entering the intake <NUM> traveling through the core flow path C are separated into a particle stream P that enters a particle separator <NUM> on a radially outer side of the core flow path C. The particle stream P is formed due to the geometry of the intake <NUM>. The intake <NUM> includes a component in the radially outer direction upstream of a portion with a component in a radially inward direction. This change in direction forces the particles against a radially outer surface of the intake <NUM> and into the particle separator <NUM> while the majority of the air is able to continue into the low pressure compressor <NUM> along the core flow path C. In this disclosure, radial or radially is in relation to the axis A unless stated otherwise.

The core flow path C is compressed by the low pressure compressor <NUM> then the high pressure compressor <NUM>, mixed and burned with fuel in the combustor <NUM>, then expanded over the high pressure turbine <NUM> and low pressure turbine <NUM>. It will be appreciated that each of the positions of the intake section <NUM>, compressor section <NUM>, combustor section <NUM>, turbine section <NUM>, and geared architecture <NUM> may be varied. For example, geared architecture <NUM> may be located aft of combustor section <NUM> or even aft of turbine section <NUM>.

The gas turbine engine <NUM> is a zero bypass engine, such that the gas turbine engine <NUM> includes a bypass ratio of zero because the gas turbine engine <NUM> includes the core flow path C without having a bypass duct forming a flow path surrounding the gas turbine engine <NUM>.

According to one non-limiting embodiment, the geared architecture <NUM> is an epicyclic gear train, such as a star gear system or a planet gear system, with a gear reduction ratio of greater than about <NUM> and less than about <NUM>. The output rotational speed of the epicyclic gear train would be fixed relative to the rotational speed of the low speed spool <NUM> such that a rotational speed of the output driveshaft <NUM> would vary with the rotational speed of the low speed spool <NUM> by a fixed gear ratio in the epicyclic gear train.

As shown in the non-limiting embodiments of <FIG>, the geared architecture <NUM> may be a star gear system with a gear ratio of about <NUM> to about <NUM>. The star gear system includes a sun gear <NUM> mechanically attached to the inner shaft <NUM> with a sun gear flexible coupling <NUM> and a plurality of star gears <NUM> surrounding the sun gear <NUM> supported by a carrier <NUM>. The carrier <NUM> is fixed from rotation relative to the engine static structure <NUM> with a carrier flexible coupling <NUM>. A ring gear <NUM> is located radially outward from the carrier <NUM> and the star gears <NUM>. The ring gear <NUM> is attached to the output driveshaft <NUM>, which is supported by drive shaft bearings <NUM>, such as roller or ball bearings. The sun gear flexible coupling <NUM> and the carrier flexible coupling <NUM> provide flexibility into the star gear system to accommodate for any misalignment during operation. Because the geared architecture <NUM> is a star gear system, the inner shaft <NUM> and the output driveshaft <NUM>, rotate in opposite rotational directions.

In another non-limiting embodiment shown in <FIG>, the geared architecture <NUM> may be a planet gear system with a gear ratio of <NUM> to about <NUM>. The planet gear system is similar to the star gear system of <FIG> and <FIG> except where described below or shown in <FIG>. The planet gear system includes the sun gear <NUM> mechanically attached to the inner shaft <NUM> with the sun gear flexible coupling <NUM> and planet gears <NUM> surrounding the sun gear <NUM>. The planet gears <NUM> are supported by the carrier <NUM>. The carrier <NUM> is allowed to rotate relative to the engine static structure <NUM>. The carrier <NUM> drives the output driveshaft <NUM>. The ring gear <NUM> is located radially outward from the carrier <NUM> and the planet gears <NUM> and is fixed from rotation relative to the engine static structure <NUM> with a ring gear flexible coupling <NUM>. The sun gear flexible coupling <NUM> and the ring gear flexible coupling <NUM> provide flexibility into the planet gear system to accommodate for any misalignment during operation. Because the geared architecture <NUM> is a planet gear system, the inner shaft <NUM> and the output driveshaft <NUM>, rotate in the same rotational direction.

Alternatively, the geared architecture <NUM> could be a non-epicyclic gear system including helical, spur, or bevel gears to create a gear reduction ratio of greater than about <NUM> and less than about <NUM>. The output rotational speed of the non-epicyclic gear system would be fixed relative to the rotational speed of the low speed spool <NUM> such that a rotational speed of the output driveshaft <NUM> would vary with the rotational speed of the low speed spool <NUM> by a fixed gear ratio in the non-epicyclic gear system.

In the illustrated non-limiting embodiment shown in <FIG>, the low pressure compressor <NUM> includes an array of inlet guide vanes <NUM> directing air from the intake <NUM> in the intake section <NUM> past multiple rotor stages <NUM> each including an array of rotor blades <NUM>. The rotor stages <NUM> are separated by stators <NUM> each including an array of vanes <NUM>. The vanes <NUM> could be variable pitch or fixed from rotating about an axis. In the illustrated non-limiting embodiment, the low pressure compressor <NUM> includes three rotor stages <NUM> and three stators <NUM> and includes a pressure ratio between about <NUM> and about <NUM>. In this disclosure, about equates to within ten (<NUM>) percent of the stated value unless stated otherwise.

The high pressure compressor <NUM> includes an array of inlet guide vanes <NUM> axially upstream of a first axial compressor stage <NUM>. The first axial compressor stage <NUM> includes an axial stage rotor <NUM> having an array of rotor blades <NUM>. A centrifugal compressor stage <NUM> is located downstream and separated from the axial compressor stage <NUM> by an array of vanes <NUM>. The centrifugal compressor stage <NUM> includes an array of blades <NUM> that direct compressed air downstream and radially outward and toward the combustor section <NUM>. The high pressure compressor <NUM> generates a pressure ratio between about <NUM> and about <NUM>. This allows the overall pressure ratio of the compressor section <NUM> to be greater than about <NUM> and less than about <NUM>. However, the overall pressure ratio of the compressor section <NUM> could reach <NUM>.

The high pressure turbine <NUM> includes an array of inlet guide vanes <NUM> that direct the core flow path C past a single rotor stage <NUM> having an array of rotor blades <NUM> upstream of the airfoils <NUM> on the mid-turbine frame <NUM>.

Furthermore, in the illustrated non-limiting embodiment, the low pressure turbine <NUM> includes three rotor stages <NUM> each including an array of rotor blades <NUM>. Each of the rotor stages <NUM> are separated by stators <NUM> having an array of vanes <NUM>. The vanes <NUM> could be variable vanes or fixed from rotation about an axis. In another non-limiting embodiment, the low pressure turbine <NUM> includes at least one rotor stage <NUM> and less than four rotor stages <NUM>. An outlet vane <NUM> is located downstream of the low pressure turbine <NUM> and directs the core flow path C out of an exhaust nozzle <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 invention is applicable to other gas turbine engines.

During operation of the gas turbine engine <NUM>, the high speed spool <NUM> rotates at a maximum rotation speed of about <NUM>,<NUM> rpms to about <NUM>,<NUM> rpms while the low speed spool operates a rotational speed of about <NUM>,<NUM> rpms. Because the rotational speed of about <NUM>,<NUM> rpms of the low speed spool is generally much higher than is desired during operation, the input to the geared architecture <NUM> is coupled to the low speed spool <NUM> to reduce the rotation speed of the low speed spool by a ratio of about <NUM> to about <NUM> at an output of the geared architecture <NUM> to drive the output driveshaft <NUM>.

Claim 1:
A turboshaft engine (<NUM>) comprising:
a high speed spool (<NUM>) including an outer shaft (<NUM>) connecting a high pressure compressor (<NUM>) with a high pressure turbine (<NUM>); and
a low speed spool (<NUM>) including an inner shaft (<NUM>) connecting a low pressure compressor (<NUM>) with a low pressure turbine (<NUM>);
a speed change mechanism (<NUM>) including an input in communication with the low speed spool (<NUM>) and a fixed gear ratio; and
an output turboshaft (<NUM>) in communication with an output of the speed change mechanism (<NUM>),
characterised in that:
the turboshaft engine (<NUM>) comprises a bypass ratio of zero and does not have a bypass duct forming a flow path surrounding the gas turbine engine (<NUM>); and
the output turboshaft (<NUM>), the inner shaft (<NUM>), and the outer shaft (<NUM>) rotate about an engine central longitudinal axis (A), wherein the low speed spool (<NUM>) is supported by no more than two bearing systems (<NUM>) and the high speed spool (<NUM>) is supported by no more than two bearing systems (<NUM>).