Patent Description:
Gas turbine engines can include a fan section, a compressor section, a combustor section and a turbine section. The fan section includes a fan having fan blades for compressing a portion of incoming air to produce thrust and also for delivering a portion of air to the compressor 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 section and the fan.

Some fan sections may include guide vanes positioned in a bypass flow path downstream of the fan blades and a flow splitter. The guide vanes may direct the bypass airflow from the fan blades before being ejected from the bypass flow path.

<CIT> discloses a prior art engine with variable area exhaust including annular flow splitters spaced radially outwardly of a core engine and generally aft of a fan.

<CIT> discloses a prior art gas turbine engine with a stream diverter including first, second and third air ducts and an associated door.

<CIT> discloses a prior art convertible fan engine with a fan system having a front stage fan rotor, an aft fan rotor, a core, an annular inner bypass passage and an annular outer bypass passage.

<CIT> discloses a prior art gas turbine engine with a fan having blades with a camber distribution relative to a covered passage of the fan.

According to a first aspect of the present invention, there is provided a gas turbine engine as set forth in claim <NUM>. According to a further aspect of the present invention, there is provided a method of operation for a gas turbine engine as set forth in claim <NUM>. Further embodiments are provided as set forth in dependent claims <NUM>-<NUM> and <NUM>.

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

The fan <NUM> includes at least one row <NUM> of airfoils or blades <NUM> that are circumferentially distributed about, and are supported by, a rotatable hub <NUM>. The row <NUM> of blades <NUM> can extend in a generally spanwise or radial direction R outwardly from the hub <NUM>. The radial direction R can be perpendicular to the engine axis A. The hub <NUM> is rotatable in a clockwise or counterclockwise direction RR (<FIG>) about the engine axis A such that the blades <NUM> deliver airflow to the bypass flow path B and core flow path C in operation. In the illustrative example of <FIG>, the hub <NUM> is driven by the low pressure (or fan drive) turbine <NUM> through the geared architecture <NUM>. A spinner or nosecone <NUM> is supported relative to the hub <NUM> to provide an aerodynamic inner flow path into the fan section <NUM>.

The fan section <NUM> includes at least one row <NUM> of turning or fan exit guide vanes <NUM>. The guide vanes <NUM> are positioned in the bypass flow path B axially aft of the row <NUM> of blades <NUM> relative to the engine axis A. The row <NUM> of guide vanes <NUM> extend in the radial direction R across the bypass flow path B between the fan (e.g., first) case <NUM> and a core engine (e.g., second case) <NUM>. The fan case <NUM> includes a bypass (e.g., first) duct <NUM> surrounding the row <NUM> of blades <NUM> and the row <NUM> of guide vanes <NUM> to establish the bypass flow path B. The engine case <NUM> is dimensioned to extend about and along the engine axis A to at least partially surround and house the compressors <NUM>, <NUM>, combustor <NUM> and turbines <NUM>, <NUM>. The fan case <NUM> and engine case <NUM> establish a portion of the engine static structure <NUM>.

The engine case <NUM> includes a flow splitter <NUM> situated downstream and axially aft of the row <NUM> of blades <NUM> relative to the engine axis A. The flow splitter <NUM> can have a generally V-shaped cross sectional geometry and is arranged to divide airflow communicated from the row <NUM> of blades <NUM> between the bypass flow path B and a second duct <NUM> establishing the core flow path C. The second duct <NUM> can serve as an entrance or inlet to the compressor section <NUM> and is bounded by the flow splitter <NUM>.

Referring to <FIG>, with continuing reference to <FIG>, each of the blades <NUM> includes an airfoil body 62A that extends in the radial direction R from the hub <NUM> (<FIG>) between a root 62R coupled to the hub <NUM> and a tip 62T. Each airfoil body 62A extends axially in an axial or chordwise direction H between a leading edge 62LE and a trailing edge 62TE, and extends circumferentially in a thickness or circumferential direction T between a pressure side 62P and a suction side <NUM>. The directions R, H, T are perpendicular to each other. The chordwise direction H can be substantially parallel or transverse to the engine axis A. The radial direction R can have a major component that extends generally from or toward an axis of rotation of the blades <NUM>, which in the illustrated example coincides with the engine axis A. It should be understood that the radial direction R can include a minor component in the axial and/or circumferential directions H, T such that the blades <NUM> have a predefined amount of sweep and/or lean, for example.

The airfoil body 62A has an exterior airfoil surface 62E providing a contour that extends in the chordwise direction H between the leading and trailing edges 62LE, 62TE. The exterior airfoil surface 62E generates lift based upon its geometry and directs flow along the core flow path C and bypass flow path B. Each blade <NUM> can be constructed from a composite material, a metallic material such as steel, an aluminum or titanium alloy, or a combination of one or more of these. Abrasion-resistant coatings or other protective coatings may be applied to the blade <NUM>.

Referring to <FIG> and <FIG>, with continuing reference to <FIG> and <FIG>, each of the guide vanes <NUM> includes an airfoil body 68A that extends in the radial direction R between inner and outer surfaces 19A, 19B of the duct <NUM>, axially in the chordwise direction H between a leading edge 68LE and a trailing edge 68TE, and circumferentially in the thickness direction T between a pressure side 68P and a suction side <NUM>. The inner surfaces 19A of the first duct <NUM> can be provided by the engine case <NUM> at a location downstream of the flow splitter <NUM>, as illustrated in <FIG>.

The airfoil body 68A has an exterior airfoil surface 68E providing a contour that extends in the chordwise direction H between the leading and trailing edges 68LE, 68TE. The exterior airfoil surface 68E can be contoured to direct airflow F (<FIG>, <FIG>) compressed by the blades <NUM> through a gas path such as the bypass flow path B. The guide vanes <NUM> can be constructed from a composite material, a metallic material such as steel, an aluminum or titanium alloy, or a combination of one or more of these. The guide vanes <NUM> can serve as a structural component to transfer loads between the fan case <NUM> and the engine case <NUM> or another portion of the engine static structure <NUM>. The fan section <NUM> of <FIG> includes a single stage fan <NUM> having only one row <NUM> of blades <NUM> and only one row <NUM> of guide vanes <NUM>. In other examples, the fan section <NUM> has more than one row of blades <NUM> and/or guide vanes <NUM> configured with respect to any of the quantities disclosed herein.

<FIG> schematically illustrate span positions of a blade <NUM> and guide vane <NUM>, respectively. Span positions are schematically illustrated from <NUM>% to <NUM>% in <NUM>% increments, for example, to define a plurality of sections 62C of the blade <NUM> and a plurality of sections 68C of the guide vane <NUM>. Each section 62C, 68C at a given span position is provided by a conical cut that corresponds to the shape of segments of the respective gas path, such as the bypass flowpath B or core flow path C, as shown by the large dashed lines.

In the case of a blade <NUM> with an integral platform, the <NUM>% span position (or zero span) corresponds to the generally radially innermost location where airfoil body 62A meets the fillet joining the airfoil body 62A to an adjacent platform <NUM> (see also <FIG>). In the case of a blade <NUM> without an integral platform, the <NUM>% span position corresponds to the generally radially innermost location where the discrete platform <NUM> meets the exterior blade surface 62E of the airfoil body 62A. A <NUM>% span position (or full span) corresponds to a section 62C of the blade <NUM> at the tip 62T. The <NUM>% position (or midspan) corresponds to a generally radial position halfway between the <NUM>% and <NUM>% span positions.

The <NUM>% span position of the vane <NUM> can correspond to the generally radially innermost location where the exterior surface 68E of the airfoil body 68A meets the inner surfaces 19A of the duct <NUM>. The <NUM>% span position can correspond to the generally radially outermost location where the exterior surface 68E of the airfoil body 68A meets the outer surfaces 19B of the duct <NUM>. The <NUM>% span position (or midspan) corresponds to a generally radial position halfway between the <NUM>% and <NUM>% span positions. As illustrated in <FIG> and <FIG>, each blade <NUM> is sectioned at a first generally radial position between the root 62R and the tip 62T, and each vane <NUM> is sectioned at a second generally radial position between the inner and outer surfaces 19A, 19B of the duct <NUM>. The first and second radial positions can be the same (e.g., both at <NUM>%, <NUM>% or <NUM>% span) or can differ (e.g., one at <NUM>% and the other at <NUM>% span).

Airfoil geometric shapes, stacking offsets, chord profiles, stagger angles, axial sweep and dihedral angles, tangential lean angles, bow, and/or other three-dimensional geometries, among other associated features, can be incorporated individually or collectively into the blades <NUM> and/or guide vanes <NUM> to improve various characteristics such as aerodynamic efficiency, structural integrity, and vibration mitigation.

Referring to <FIG>, the blades <NUM> and guide vanes <NUM> can be arranged at various orientations relative to the engine axis A. Each blade <NUM> establishes a chord dimension BCD. Each vane <NUM> establishes a chord dimension VCD. The chord dimension BCD/VCD is a length between the respective leading and trailing edges 62LE/68LE, 62TE/68TE. The chord dimension BCD/VCD may vary along the span of the airfoil body 62A/68A. The chord BCD/VCD forms a respective stagger angle α/β relative to the chordwise direction H or a plane parallel to the engine axis A.

The stagger angles α, β can vary along the airfoil span to define a twist. For example, the blades <NUM> and guide vanes <NUM> in <FIG> may be at a first span position, such as at <NUM>% span, and the blades <NUM> and guide vanes <NUM> in <FIG> may be at a second, different span position, such as at <NUM>% span. Referring to <FIG>, the stagger angle α of the blades <NUM> is larger at <NUM>% span than the stagger angle α of the blades <NUM> corresponding to the span position of <FIG>. In examples, the stagger angle α of the blade <NUM> is greater than or equal to about <NUM> degrees at <NUM>% span, or more narrowly less than or equal to about <NUM> degrees, such as between about <NUM> and <NUM> degrees at <NUM>% span. In examples, the stagger angle α is less than or equal to about <NUM> degrees at <NUM>% span, or more narrowly less than or equal to about <NUM> degrees, such as between about <NUM> and <NUM> degrees at <NUM>% span.

The stagger angle β of the vane <NUM> can be substantially the same from <NUM>% span to <NUM>% span or can differ. In examples, the stagger angle β is greater than or equal about <NUM> degrees at <NUM>% span, or more narrowly less than or equal to about <NUM> degrees, such as between about <NUM> and <NUM> degrees at <NUM>% span. In examples, the stagger angle β is less than or equal about <NUM> degrees at <NUM>% span, or more narrowly less than or equal to about <NUM> degrees, such as between about <NUM> and <NUM> degrees at <NUM>% span.

Each of the blades <NUM> can be cambered. A mean camber line <NUM> bisects the airfoil body 62A between the leading and trailing edges 62LE, 62TE of the blade <NUM>. A leading edge metal angle α1* at the leading edge 62LE and a trailing edge metal angle α2* at the trailing edge 62TE are established with respect to the mean camber line <NUM> and the chordwise direction H. The tip 62T of the blade <NUM> can rotate at a relatively faster speed, and the stagger of the blade <NUM> can be aligned with a relative flow direction F' to achieve an optimum incidence of the flow F relative to the leading edge metal angle α1* of the blade <NUM>. A camber κ1 of the blade <NUM> is equal to a difference between the leading edge metal angle α1* and the trailing edge metal angle α2*. In the illustrative example of <FIG>, the camber κ1 of the blade <NUM> can be substantially less at <NUM>% span (<FIG>) than the camber κ1 of the blade <NUM> at less than <NUM>% span (<FIG>). The camber κ1 of the airfoil body 62A contour that extends in the chordwise direction H between the leading and trailing edges 62LE, 62TE of the blade <NUM> can be less than or equal to about <NUM> degrees at <NUM>% span, or more narrowly less than or equal to about <NUM> degrees, such as between about <NUM> and <NUM> degrees at <NUM>% span. The camber κ1 can be greater than or equal to about <NUM> degrees at <NUM>% span, or more narrowly less than or equal to about <NUM> degrees, such as between about <NUM> and <NUM> degrees at <NUM>% span.

Each of the vanes <NUM> can be cambered. A mean camber line <NUM> bisects the airfoil body 68A between the leading and trailing edges 68LE, 68TE of the vane <NUM>. A leading edge metal angle β1* at the leading edge 68LE and a trailing edge metal angle β2* at the trailing edge 68TE are established with respect to the mean camber line <NUM> and the chordwise direction H. A camber κ2 of the vane <NUM> is equal to a difference between the leading edge metal angle β1* and the trailing edge metal angle β2*. The camber κ2 of the airfoil body 68A contour that extends in the chordwise direction H between the leading and trailing edges 68LE, 68TE of the vane <NUM> can be greater than or equal to about <NUM> degrees at <NUM>% span, or more narrowly less than or equal to about <NUM> degrees, such as between about <NUM> and <NUM> degrees at <NUM>% span. The camber κ2 can be less than or equal to about <NUM> degrees at <NUM>% span, or more narrowly less than or equal to about <NUM> degrees, such as between about <NUM> and <NUM> degrees at <NUM>% span. For the purposes of this disclosure, the terms "about" and "approximately" mean ±<NUM>% of the stated value unless otherwise indicated.

<FIG> illustrates a gas turbine engine <NUM> according to another example. In this disclosure, like reference numerals designate like elements where appropriate and reference numerals with the addition of one-hundred or multiples thereof designate modified elements that are understood to incorporate the same features and benefits of the corresponding original elements. The engine <NUM> includes a fan section <NUM> including a fan (or rotor) <NUM>, a compressor section <NUM>, a combustor section <NUM> and a turbine section <NUM>. The compressor, combustor and turbine sections <NUM>, <NUM>, <NUM> establish respective portions of the core flow path C. The turbine section <NUM> drives the fan section <NUM> and the compressor section <NUM>. The compressor section <NUM> includes a low pressure compressor <NUM> and a high pressure compressor <NUM>. The turbine section <NUM> includes a low pressure (or fan drive) turbine <NUM> and a high pressure turbine <NUM>.

The fan <NUM> includes at least one row <NUM> of blades <NUM> carried by and mounted to a rotatable hub <NUM>. In examples, the blades <NUM> are incorporated into a fixed-pitch arrangement in which the blades <NUM> have a fixed stagger angle α. In other examples, the blades <NUM> are incorporated into a variable-pitch arrangement in which the stagger angle α varies in response to rotation of the blades <NUM> about a respective blade axis that extends between the blade root 162R and tip 162T.

The engine <NUM> can include a geared architecture <NUM> that drives the fan <NUM> at a different speed than a speed of the fan drive turbine <NUM> in operation. In examples, the geared architecture <NUM> is an epicyclic gear system such as a planetary gear system having a fixed ring gear and a rotatable carrier, or such as a star gear system having a fixed carrier and a rotatable ring gear. The geared architecture <NUM> can establish a gear reduction ratio of greater than about <NUM> or <NUM>, or more narrowly less than about <NUM> or <NUM>, for example. In other examples, the geared architecture <NUM> is omitted such that the fan drive turbine <NUM> directly drives the fan <NUM> at a common speed and in a common direction.

The fan section <NUM> includes a fan (e.g., first) case <NUM> including a bypass (e.g., first) duct <NUM> surrounding the row <NUM> of blades <NUM> to establish a bypass (e.g., first) flow path B. The engine <NUM> includes an engine (e.g., second) case <NUM> that establishes a core (e.g., second) flow path C.

The engine <NUM> can include a nacelle <NUM> and core cowling (or housing) <NUM>. The nacelle <NUM> is mechanically attached or otherwise secured to the fan case <NUM> and establishes an engine inlet <NUM>. The engine inlet 123I conveys airflow F to the fan <NUM> in operation. The core cowling <NUM> is mechanically attached or otherwise secured to the engine case <NUM>. The core cowling <NUM> at least partially surrounds the engine case <NUM>. The bypass flow path B terminates at a bypass exit (or nozzle) <NUM> established by a trailing edge of the nacelle <NUM>.

At least one row <NUM> of turning or fan exit guide vanes <NUM> are situated in the bypass duct <NUM>. The guide vanes <NUM> extend in a radial direction R across the bypass flow path B. Each guide vane <NUM> extends in the radial direction R between a <NUM>% span position along an inner surface 119A and a <NUM>% span position along an outer surface 119B of the bypass duct <NUM>.

An airfoil body 168A of the guide vane <NUM> extends along a longitudinal axis LA, as illustrated in <FIG>. A projection of the longitudinal axis LA intersects the engine axis A to establish a vane angle ϕ. The vane angle ϕ can be approximately <NUM> degrees such that the guide vane is substantially perpendicular to the engine axis A, as illustrated by the guide vane <NUM> of <FIG>. In examples, the vane angle ϕ is less <NUM> degrees such that the guide vane has a component that extends aftwards in the axial direction from <NUM>% to <NUM>% span with respect to the engine axis A, as illustrated by the guide vane <NUM> of <FIG>. The vane angle ϕ can be greater than or equal to <NUM> degrees, or more narrowly between <NUM> and <NUM> degrees, for example.

Various quantities of blades <NUM> and guide vanes <NUM> can be incorporated into the engine <NUM>. The row <NUM> of blades <NUM> establishes a blade quantity (BQ). The row <NUM> of guide vanes <NUM> establishes a vane quantity (VQ), which can be the same or can differ from the blade quantity (BQ). The blade quantity (BQ) can be less than <NUM> or <NUM> blades, or more narrowly at least <NUM> but not more than <NUM> blades, such as <NUM> or <NUM> blades, for example. The vane quantity (VQ) can be no more than <NUM> guide vanes, or more narrowly at least <NUM> guide vanes, for example. In examples, the ratio of VQ/BQ is between <NUM> and <NUM>, or more narrowly between <NUM> and <NUM>, which may reduce noise due to flow interaction between the blades <NUM> and guide vanes <NUM>. For the purposes of this disclosure, the term "between" is inclusive of the stated value(s) unless otherwise stated.

Various stage counts can be incorporated in the compressors <NUM>, <NUM> and turbines <NUM>, <NUM>. For example, the low pressure compressor <NUM> can include at least <NUM> but no more than <NUM> compressor stages, such as <NUM> compressor stages as illustrated by <FIG> and <FIG>. In examples, the low pressure turbine <NUM> includes at least <NUM> but no more than <NUM> turbine stages, such as <NUM> turbine stages as illustrated by <FIG>, <FIG> turbine stages, or <NUM> turbine stages as illustrated by <FIG>. The high pressure turbine <NUM> can include a single stage, or can include two stages as illustrated by <FIG> and <FIG>, for example. A ratio between the blade quantity (BQ) and a quantity of the stages of the low pressure turbine <NUM> can be between about <NUM> and about <NUM>, or more narrowly less than or equal to about <NUM>, such as about <NUM>. The example low pressure turbine <NUM> provides the driving power to rotate the fan <NUM> and therefore the relationship between the number of turbine stages of the low pressure turbine <NUM> and the blade quantity (BQ) disclose an example gas turbine engine <NUM> with increased power transfer efficiency.

The engine <NUM> includes a core bypass (or third) duct <NUM> established between the engine case <NUM> and core cowling <NUM>. The core bypass duct <NUM> establishes an intermediate (or third) flow path I. The core bypass duct <NUM> can be a separate and distinct structure or can be established by a volume of the core engine bay between the engine core and core cowling <NUM>. The core bypass duct <NUM> can be coupled to an annular core vent port (or nozzle) <NUM> established along an outer periphery of the core cowling <NUM>. In the illustrated example of <FIG>, the guide vanes <NUM> are an axially aftmost row of fan exit guide vanes in the bypass flow path B between the row <NUM> of blades <NUM> and the bypass exit <NUM> relative to the engine axis A. In other examples, one or more rows of vanes are arranged in the bypass flow path B axially forward and/or aft of the row <NUM> of guide vanes <NUM>. The hub <NUM> is rotatable about the engine axis A such that the row <NUM> of blades <NUM> deliver or convey airflow to the bypass, intermediate and core flow paths B, I, C in operation.

The engine <NUM> includes a first flow splitter <NUM> and a second flow splitter <NUM> arranged in a cascade. The flow splitters <NUM>, <NUM> are situated downstream and axially aft of the row <NUM> of blades <NUM> relative to the engine axis A. The first flow splitter <NUM> can be incorporated into or mechanically attached to the core cowling <NUM>. The second flow splitter <NUM> can be incorporated into or mechanically attached to the engine case <NUM>. The first flow splitter <NUM> establishes a portion of the bypass duct <NUM> and a portion of the second duct <NUM>. The second flow splitter <NUM> establishes a portion of the second duct <NUM> and a portion of an entrance (e.g., fourth) duct <NUM>. The entrance duct <NUM> can establish an entrance or inlet to the compressor section <NUM>.

The first flow splitter <NUM> is arranged to divide airflow conveyed by the row <NUM> of blades <NUM> between the bypass flow path B and the second duct <NUM>. The second flow splitter <NUM> is radially inboard of the first flow splitter <NUM> relative to the engine axis A. The second duct <NUM> branches between the core flow path C and the intermediate flow path I at the second flow splitter <NUM>. The second flow splitter <NUM> is arranged to divide airflow conveyed to the second duct <NUM> between the intermediate flow path I and the core flow path C along the entrance duct <NUM>. The flow splitters <NUM>, <NUM> can have a generally V-shaped cross sectional geometry dimensioned to divide the airflow.

A bypass port <NUM> is established along an outer periphery 125P of the core cowling <NUM>. The bypass port <NUM> interconnects the intermediate flow path I and bypass flow path B at a position downstream of the row <NUM> of guide vanes <NUM> relative to a general direction of airflow through the bypass flow path B such that airflow conveyed by the intermediate flow path I through the bypass port <NUM> bypasses or otherwise is not communicated across the guide vanes <NUM> in operation. The bypass port <NUM> can interconnect the intermediate flow path I and bypass flow path B at a position axially aft of the <NUM>% span position of the row <NUM> of guide vanes <NUM> relative to the engine axis A. In the illustrated example of <FIG>, the vent port <NUM> is established downstream and axially aft of both the bypass port <NUM> and the bypass exit <NUM> relative to the engine axis A. In other examples, the vent port <NUM> is established upstream and axially forward of the bypass exit <NUM> such that airflow in the bypass flow path B mixes in the bypass duct <NUM> with airflow conveyed by the vent port <NUM> prior to exiting the bypass exit <NUM>.

One or more louvers <NUM> can be situated in the bypass port <NUM>. The louvers <NUM> can have a generally airfoil-shaped cross sectional geometry and are arranged to direct airflow from the intermediate flow path I into the bypass flow path B in a direction generally downstream of the guide vanes <NUM> and toward the bypass exit <NUM>.

The engine <NUM> can establish an adaptive fan flow arrangement in which a relative distribution of airflow conveyed by the fan <NUM> to the flow paths B, I and/or C varies during one or more modes. One or more blocker doors <NUM> can be moved between opened and closed positions to selectively communicate airflow through the bypass port <NUM> (open and closed positions indicated in dashed lines at <NUM>, <NUM>' of <FIG> for illustrative purposes).

In <FIG>, the engine <NUM> includes one or more blocker doors <NUM> situated in the core bypass duct <NUM> to modulate airflow through the intermediate flow path I. The blocker doors <NUM> are situated in the intermediate flow path I downstream of the bypass port <NUM>. The blocker doors <NUM> are moveable between a closed position (<FIG>) and an open position (<FIG>) to selectively communicate airflow in the intermediate flow path I downstream to the vent port <NUM>. The blocker doors <NUM> at least partially or substantially block airflow through the intermediate flow path I in the closed position. In the illustrative example of <FIG>, the blocker doors <NUM> are pivotable flaps having a pivot point established along the core bypass duct <NUM>.

Referring to <FIG>, with continuing reference to <FIG>, the nacelle <NUM> can include at least one thrust reverser <NUM> and/or a variable area fan nozzle (VAFN) <NUM> for adjusting various characteristics of the bypass flow. <FIG> illustrates the thrust reverser <NUM> and the variable area fan nozzle <NUM> in stowed positions. <FIG> illustrates the thrust reverser <NUM> in the stowed position and the variable area fan nozzle <NUM> in a deployed position. <FIG> illustrates the thrust reverser <NUM> in a deployed position and the variable area fan nozzle <NUM> in the stowed position. The nacelle <NUM> includes a stationary portion <NUM>-<NUM> mounted to the fan case <NUM>, a first moveable portion <NUM>-<NUM> and a second movable portion <NUM>-<NUM>. The moveable portions <NUM>-<NUM>, <NUM>-<NUM> are moveable relative to the stationary portion <NUM>-<NUM> and relative to each other.

The thrust reverser <NUM> is operable to convey airflow in the bypass flow path B for producing reverse thrust such as during approach and/or landing conditions of an aircraft associated with the engine <NUM>. The thrust reverser <NUM> includes a thrust reverser body 180B, which is configured with the first moveable portion <NUM>-<NUM> of the nacelle <NUM>. The thrust reverser body 180B may have a generally tubular geometry. The thrust reverser <NUM> can include one or more blocker doors 180D, one or more actuators 180A, and/or one or more cascades 180C of turning vanes 180V arranged circumferentially about the engine axis A.

The thrust reverser body 180B can include at least one recess 180R that houses the cascades 180C and the actuators 180A when the thrust reverser <NUM> is in the stowed position. The cascades 180C are dimensioned to span across a reversal port 180P when the thrust reverser <NUM> is in the deployed position, as illustrated by <FIG>. The cascade 180C selectively communicates airflow from the bypass duct <NUM> outwardly through the reversal port 180P in the deployed position. The cascade 180C can extend axially aft of both the row <NUM> of guide vanes <NUM> and the bypass port <NUM> with respect to the engine axis A, and the row <NUM> of guide vanes <NUM> can be situated axially forward of the blocker doors 180D relative to the engine axis A, as illustrated by <FIG>.

Each blocker door 180D is pivotally connected to the thrust reverser body 180B. The bypass port <NUM> can be established axially forward of the blocker doors 180D with respect to the engine axis A, as illustrated in <FIG>. The actuators 180A are operable to axially translate the thrust reverser body 180B between the stowed and deployed positions. As the thrust reverser body 180B translates aftwards, the blocker doors 180B pivot radially inward into the bypass duct <NUM> and divert at least some or substantially all of the bypass airflow F as flow F1 through the cascades 180C to provide the reverse engine thrust, as illustrated in <FIG>. In other examples, the cascades 180C are configured to translate axially with a respective thrust reverser body 180B. The thrust reverser body 180B and/or cascades 180C can include one or more circumferential segments that synchronously or independently translate or otherwise move between deployed and stowed positions.

The variable area fan nozzle <NUM> includes a nozzle body 182B and one or more actuators 182A. The nozzle body 182B is configured with the second movable portion <NUM>-<NUM> of the nacelle <NUM>, and is arranged radially within and may nest with the thrust reverser body 180B. The nozzle body 182B may have a generally tubular geometry that extends about an axially contoured outer periphery 125P of the core cowling <NUM>. The actuators 182A are operable to axially translate the nozzle body 182B between the stowed position of <FIG> and the deployed position of <FIG>. As the nozzle body 182B translates aftwards, a radial distance RD of the bypass exit <NUM> between a trailing edge or aft end 123TE of the nacelle <NUM> and the outer periphery 125P of the core cowling <NUM> established at the stowed position may change (e.g., increase) to a radial distance RD' and thereby change (e.g., increase) a flow area of the bypass exit <NUM>. In this manner, the variable area fan nozzle <NUM> may adjust a pressure drop or ratio across the bypass flowpath B defined by the bypass duct <NUM> by changing the flow area of the bypass exit <NUM>.

The variable area fan nozzle <NUM> can define or otherwise include at least one auxiliary port 182P to vary the bypass flow. The auxiliary port 182P can be established in the nozzle body 182B, as illustrated by <FIG>. The auxiliary port 182P is moveable relative to a trailing edge portion 123P of the first moveable portion <NUM>-<NUM> of the nacelle <NUM> to selectively block and open the auxiliary port 182P. The auxiliary port 182P is opened as the nozzle body 182B translates axially aftwards relative to the engine axis A. A flow area through the auxiliary port 182P augments the flow area of the bypass exit <NUM>, thereby increasing an effective flow area of the variable area fan nozzle <NUM>. The variable area fan nozzle <NUM> therefore may adjust a pressure drop or ratio across the bypass flowpath B defined by the bypass duct <NUM> while translating the nozzle body 182B over a relatively smaller axial distance. In the illustrative example of <FIG>, the variable area fan nozzle <NUM> is omitted.

Referring to <FIG>, with continuing reference to <FIG>, the blades <NUM> can be dimensioned to provide a relatively radially compact engine arrangement. Each of the blades <NUM> extends in the radial direction R between a <NUM>% span position at the hub <NUM> and a <NUM>% span position at a tip 162T. A tip radius RT of the blades <NUM> is established between a radially outermost portion of the tip 162T and the engine axis A. A forwardmost portion of a leading edge 162LE of each the blades <NUM> is arranged along a first reference plane REF1. The first reference plane REF1 is perpendicular to the engine axis A. A hub radius RH is established along the first reference plane REF1 between an outer periphery of the hub <NUM> and the engine axis A. The tip radius RT can be between <NUM> and <NUM> inches (<NUM> and <NUM>,<NUM>), or more narrowly between <NUM> and <NUM> inches (<NUM>,<NUM> and <NUM>,<NUM>), for example.

Each of the blades <NUM> extends in the radial direction R between the <NUM>% span position and the <NUM>% span position to establish a geometric hub-to-tip ratio (RH:RT). The hub-to-tip ratio (RH:RT) can be less than or equal to about <NUM>, or more narrowly greater than or equal to about <NUM>, measured relative to the forwardmost portion of the leading edge 162LE of the blades <NUM>. In examples, the hub-to-tip ratio (RH:RT) is between <NUM>-<NUM>, or more narrowly between <NUM>-<NUM>. In examples, the hub-to-tip ratio (RH:RT) is established such that a maximum value of the fan pressure ratio across the blade <NUM> alone is less than or about equal to about <NUM> measured at cruise at <NUM> Mach and <NUM>,<NUM> feet (<NUM>,<NUM>). In further examples, a maximum value of the fan pressure ratio is greater than or equal to about <NUM>, or more narrowly less than or equal to about <NUM>. The fan pressure ratios and hub-to-tip ratios disclosed herein can be utilized alone or in combination to establish relatively high bypass turbo fan engine arrangements.

The flow splitters <NUM>, <NUM> are arranged at respective positions relative to each other and the row <NUM> of blades <NUM> to establish a predetermined distribution of airflow through the bypass, intermediate and core flow paths B, I, C in operation. A forwardmost edge 170F of the first flow splitter <NUM> is situated in the radial and axial directions R, H at a first splitter position P1 relative to the engine axis A. The first splitter position P1 is established along a second reference plane REF2. A forwardmost edge 172F of the second flow splitter <NUM> is situated in the radial and axial directions R, H at a second splitter position P2 relative to the engine axis A. The second splitter position P2 is established along a third reference plane REF3. The reference planes REF2, REF3 are perpendicular to the engine axis A.

The reference planes REF2, REF3 are established at different positions relative to the engine axis A to establish a cascade arrangement. The forwardmost edge 172F of the second flow splitter <NUM> is axially aft of the forwardmost edge 170F of the first flow splitter <NUM> with respect to the engine axis A. The forwardmost edges 170F, 172F of the flow splitters <NUM>, <NUM> are axially forward of the leading edges 168LE of the row <NUM> of guide vanes <NUM> at the <NUM>% span position with respect to the engine axis A. The forwardmost edges 170F, 172F can be dimensioned to extend axially forward of the geared architecture <NUM> relative to the engine axis A, as illustrated by <FIG>, which can establish a relatively compact core engine arrangement.

Referring to <FIG>, with continuing reference to <FIG>, a first annulus area A1 is established by, and between, inner and outer diameter surfaces 119A, 119B of the bypass duct <NUM> at the forwardmost edge 170F of the first flow splitter <NUM>. A second annulus area A2 is established by, and between, inner and outer diameter surfaces 121A, 121B of the second duct <NUM> at the forwardmost edge 170F of the first flow splitter <NUM>. A third annulus area A3 is established by, and between, inner and outer diameter surfaces 133A, 133B of the entrance duct <NUM> at the second flow splitter <NUM>. The annulus areas A1, A2, A3 extend about the engine axis A to establish a portion of the respective flow paths B, I, C.

The annulus areas A1, A2, A3 can be dimensioned to establish a predetermined distribution of airflow to the flow paths B, I, C in operation. A first geometric bypass area ratio (A1:A2) associated with the first flow splitter <NUM> is defined as the first annulus area A1 divided by the second annulus area A2. A second geometric bypass area ratio (A1:A3) associated with the second flow splitter <NUM> is defined as the first annulus area A1 divided by the third annulus area A3. The annulus areas A1, A2, A3 can be dimensioned to establish a predetermined value of the first bypass area ratio (A1: A2) and a predetermined value of the second bypass area ratio (A1:A3). In examples, the first bypass area ratio (A1:A2) is greater than or equal to about <NUM> to establish a relatively high bypass turbo fan arrangement. In examples, the first bypass area ratio (A1:A2) is less than or equal to about <NUM>, or more narrowly less than or equal to about <NUM>. The second bypass area ratio (A1:A3) is greater than the first bypass area ratio (A1:A2). For example, a value of the second bypass area ratio (A1:A3) can be between about <NUM>-<NUM>% of the first bypass area ratio (A1:A2), such as approximately twice or <NUM>% of the first bypass area ratio (A1:A2). In examples, the second bypass area ratio (A1:A3) is greater than or equal to about <NUM>, or more narrowly greater than or equal to about <NUM>.

Referring back to <FIG>, the forwardmost edges 170F, 172F of the flow splitters <NUM>, <NUM> can be dimensioned relative to a span of the blades <NUM> to establish the predetermined distribution of airflow through the bypass, intermediate and core flow paths B, I, C in operation. In examples, the first splitter position P1 can be radially aligned with or radially inward of a <NUM>% span position of the blades <NUM>, or more narrowly can be radially aligned with or radially outward of a <NUM>% span position of the blades <NUM> relative to the radial direction R, such as between about <NUM>-<NUM>% span. The second splitter position P2 can be radially aligned with or radially outward of a <NUM>% span position of the blades <NUM>, or more narrowly can be radially aligned with or radially inward of a <NUM>% span position of the blades <NUM> relative to the radial direction R, such as between about <NUM>-<NUM>% span. The splitter positions P1, P2 can be established at a radial distance of at least <NUM>% span apart from each other relative to the blades <NUM> such that a relatively larger amount of airflow is communicated to the intermediate flow path I and bypasses the guide vanes <NUM>. In examples, the radial distance between the splitter positions P1, P2 is between about <NUM>-<NUM>% span, relative to the blades <NUM>. The blades <NUM> can have any of the stagger angles α and/or camber κ values disclosed herein at the respective spanwise positions of the first splitter position P1 and/or second splitter position P2 to convey airflow in a predetermined direction towards the splitters <NUM>, <NUM>, which can improve efficiency.

The guide vanes <NUM> and flow splitters <NUM>, <NUM> can be arranged relative to the fan blades <NUM> to establish a relatively axially compact envelope. The leading edges 162LE of the blades <NUM> at the <NUM>% span position are established along a fourth reference plane REF4. The leading edges 168LE of the guide vanes <NUM> at the <NUM>% span position are established along a fifth reference plane REF5. The reference planes REF4, REF5 are perpendicular to the engine axis A. A first axial length L1 is established between the reference planes REF4, REF5. In examples, a ratio (LLRT) of the first axial length L1 divided by the tip radius RT is greater than or equal to about <NUM>, which can reduce fan noise propagation through the bypass flow path B. The ratio (LLRT) can be less than or equal to about <NUM>, which can establish a relatively axially compact fan section. A second axial length L2 is established between the first and second splitter positions P1, P2 along the reference planes REF2, REF3 relative to the engine axis A. In examples, a ratio (L2:L1) of the second axial length L2 divided by the first axial length L1 is less than or equal to about <NUM>, or more narrowly between about <NUM> and about <NUM>, which can further establish a relatively axially compact engine.

<FIG> illustrates a plot of bypass area ratio relative to airfoil span position, such as span positions of the blades <NUM>. <FIG> illustrates selected portions of <FIG>. Example profiles or curves C1-C6 are shown. The curves C1-C6 correspond to fan or rotor arrangements having hub-to-tip ratios (RH:RT) of <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, respectively. The fan <NUM> can be dimensioned to establish any of the hub-to-tip ratios (RH:RT) and corresponding values along the curves C1-C6.

Each curve C1-C6 can be established as function of annulus area associated with respective bypass area ratios. Splitter position values for each % span position along the respective curve C1-C6 can be defined in one of the following ways: <MAT> <MAT> where (MR) is the radial splitter position of the respective splitter <NUM>, <NUM>, (RH) is the hub radius and (RT) is the tip radius as previously defined, and (λ) is the hub-to-tip ratio (RH:RT). Since the annulus areas established by the splitters are non-linear, the corresponding curves C1-C6 are also non-linear.

The plot of curves C1-C6 can serve as a nomograph utilized to select radial locations of the first and second positions P1, P2 of the flow splitters <NUM>, <NUM> with respect to hub-to-tip (RH:RT) ratio and/or bypass area ratio (A1:A2), (A1:A3), or vice versa. For example, referring to <FIG>, the fan <NUM> can establish a hub-to-tip (RH:RT) ratio of <NUM> corresponding to the curve C1. The radial location of the first and second splitter positions P1, P2 of the flow splitters <NUM>, <NUM> correspond to points P1(<NUM>), P2(<NUM>) selected along the curve C1. The points P1(<NUM>), P2(<NUM>) can be selected with respect to a predetermined bypass area ratio. For example, point P1(<NUM>) of the first flow splitter <NUM> can be established in the radial direction R at about a <NUM>% span position of the blades <NUM>. Point P1(<NUM>) corresponds to a first bypass area ratio (A1:A2) of about <NUM> and a mass average fan pressure ratio of about <NUM>. The radial location of the second flow splitter <NUM> can be determined by assuming analytically that the second flow splitter <NUM> is the first flow splitter <NUM> (e.g., a single flow splitter arrangement). The point P2(<NUM>) of the second flow splitter <NUM> can be established in the radial direction R at about a <NUM>% span position of the blades <NUM>. Point P2(<NUM>) corresponds to a second bypass area ratio (A1:A3) of about <NUM> and a mass average fan pressure ratio of about <NUM>. For the purposes of this disclosure, mass average fan pressure ratio is defined as an average pressure ratio across the blades <NUM> alone measured at cruise at <NUM> Mach and <NUM>,<NUM> feet between the <NUM>% span position and the corresponding span position of the respective splitter <NUM>, <NUM>. The mass across the blades <NUM> associated with the mass average fan pressure ratio is approximately equal to the amount of the mass captured by the respective splitter <NUM>, <NUM>.

As another example, the fan <NUM> can establish a hub-to-tip (RH:RT) ratio of <NUM> corresponding to the curve C6. The radial location of the first and second splitter positions P1, P2 of the flow splitters <NUM>, <NUM> correspond to points P1(<NUM>), P2(<NUM>) selected along the curve C6. Point P1(<NUM>) of the first flow splitter <NUM> can be established in the radial direction R at about a <NUM>% span position of the blades <NUM>. Point P1(<NUM>) corresponds to a first bypass area ratio (A1:A2) of about <NUM> and a mass average fan pressure ratio of about <NUM>. Point P2(<NUM>) of the second flow splitter <NUM> can be established in the radial direction R at about a <NUM>% span position of the blades <NUM>. Point P2(<NUM>) corresponds to a second bypass area ratio (A1:A3) of about <NUM> and a mass average fan pressure ratio of about <NUM>.

The flow splitters <NUM>, <NUM> and other portions of the fan section <NUM> can be dimensioned with respect to any other values along the curves C1-C6. For example, points P1'(<NUM>), P2'(<NUM>) can be established along the curve C1 at about <NUM>% and about <NUM>% span corresponding to first and second bypass area ratios (ALA2), (ALAS) of about <NUM> and <NUM>, respectively. As another example, points P1'(<NUM>), P2'(<NUM>) can be established along the curve C6 at about <NUM>% and about <NUM>% span corresponding to first and second bypass area ratios (A1:A2), (A1:A3) of about <NUM> and <NUM>, respectively.

As illustrated by the curves C1-C6 of <FIG>, for the same bypass area ratio, a lower hub-to-tip ratio (RH:RT) can result in the splitter positions P1, P2 of the splitters <NUM>, <NUM> being relatively further away from the <NUM>% span position of the blades <NUM>. The splitters <NUM>, <NUM> can be dimensioned with respect to relatively lower % span positions to establish relatively larger bypass area ratios, as illustrated by the curves C1-C6 of <FIG>.

<FIG> illustrates a plot of fan pressure ratio relative to airfoil span position, such as span positions of the blades <NUM>. <FIG> illustrates selected portions of the plot of <FIG>. Curve PR1 corresponds to discrete fan pressure ratios at each span position. Lines LP1(<NUM>), LP2(<NUM>) correspond to the splitter positions P1(<NUM>), P2(<NUM>) of the splitters <NUM>, <NUM> associated with a hub-to-tip ratio (RH:RT) equal to <NUM>. Lines MA1(<NUM>), MA2(<NUM>) are the mass averaged fan pressure ratios corresponding to the respective splitter positions P1(<NUM>), P2(<NUM>). <FIG> depicts a curve PR1(INT) representing a spanwise integration of the discrete fan pressure ratios of curve PR1. An intersection of the lines LP1(<NUM>), LP2(<NUM>) corresponding to the points P1(<NUM>), P2(<NUM>) and the curve PR1(INT) establish a position of the respective lines MA1(<NUM>), MA2(<NUM>) relative to the y-axis. In the illustrative example of <FIG>, lines MA1(<NUM>), MA2(<NUM>) correspond to mass average fan pressure ratios of <NUM> and <NUM>, respectively, which correspond to values previously discussed with respect to the points P1(<NUM>), P2(<NUM>) along the curve C1 of <FIG>.

<FIG> illustrates another plot of fan pressure ratio relative to airfoil span position, such as span positions of the blades <NUM>. Curve PR2 corresponds to discrete fan pressure ratios at each span position of the blade <NUM>. Lines LP1(<NUM>), LP2(<NUM>) correspond to splitter positions P1(<NUM>), P2(<NUM>) of the splitters <NUM>, <NUM> associated with a hub-to-tip ratio (RH:RT) equal to <NUM>. In the illustrative example of <FIG>, lines MA1(<NUM>), MA2(<NUM>) correspond to mass average fan pressure ratios of <NUM> and <NUM>, respectively, which correspond to values previously discussed with respect to the points P1(<NUM>), P2(<NUM>) along the curve C6 of <FIG>.

The engine <NUM> can include a controller CONT (shown in dashed lines in <FIG> for illustrative purposes). The controller CONT is operable to cause the blocker doors <NUM>, <NUM>, thrust reverser <NUM> and variable area fan nozzle <NUM> to move between the various disclosed positions or modes. The controller CONT can include one or more computing devices, each having one or more of a computer processor, memory, storage means, network device and input and/or output devices and/or interfaces according to some example. The memory may, for example, include UVPROM, EEPROM, FLASH, RAM, ROM, DVD, CD, a hard drive, or other computer readable medium which may store data and/or the algorithms corresponding to the various functions of this disclosure. In other examples, the controller CONT is an analog or electromechanical device configured to provide the disclosed functions of this disclosure. The controller CONT can be a portion of a full-authority digital electronic control FADEC or an electronic engine control (EEC), another system, or a stand-alone system located within the aircraft remote from the engine <NUM>.

<FIG> illustrates a plurality of discrete modes M1-M4 established by the position of the blocker door(s) <NUM> and position of the blocker door(s) <NUM> across the bypass port <NUM>. The controller CONT can be programmed with logic or otherwise configured to cause the engine <NUM> to execute each of the modes M1-M4. During modes M1 and M3, the blocker doors <NUM> are stowed to establish an open position to at least partially communicate airflow through vent port <NUM>. During modes M2 and M4, the blocker doors <NUM> are deployed to establish a closed position to at least partially block airflow from being communicated to the vent port <NUM>. During modes M1 and M2, the bypass port <NUM> is open or uncovered to at least partially communicate airflow from the intermediate flow path I to the bypass flow path B. During modes M3 and M4, the bypass port <NUM> is closed or covered to at least partially block airflow being communicated from the intermediate flow path I to the bypass flow path B. The engine <NUM> can be operated in the first mode M1 during takeoff, for example, which may be associated with relatively lower engine inlet airflow such that the fan <NUM> communicates relatively more airflow which may improve fan stability.

The engine <NUM> can be operated in the first mode M1 and/or the second mode M2 during cruise, for example. The engine <NUM> can be operated in the second mode M2 during a beginning of climb condition to increase fan pressure ratio and generate relatively greater thrust. The thrust reverser <NUM> can be deployed during the second mode M2, for example, to increase an amount of airflow being communicated to the bypass flow path to the thrust reverser <NUM> for generating reverse thrust. The engine <NUM> can be operated in the third mode M3 during a middle of climb condition (i.e., between beginning climb and top of climb). The engine <NUM> can be operated in the fourth mode M4 during a top of climb condition to increase airflow to the core flow path C and bypass flow path B to maximize or otherwise increase thrust. The modes M1-M4 are exemplary and it should be appreciated the blocker door(s) <NUM> and bypass port <NUM> can be configured to establish fewer or more than four modes.

The controller CONT can be coupled to one or more sensors to determine any of the conditions disclosed herein, such as temperature, pressure and/or speed sensors operable to measure various conditions or states of the engine <NUM>. Temperature and/or pressure sensors can be situated along the bypass, intermediate and/or core flow paths B, I, C, for example. Speed sensors can be situated adjacent to the rotors to determine a speed of the fan rotor, low spool, and/or high spool, for example. The controller CONT can utilize other data and information to determine any of the conditions disclosed herein, including aircraft velocity, altitude and throttle position. One would understand how to situate the sensors and program or otherwise configure the controller CONT with logic to obtain data and other information from the sensors and other systems in view of the teachings disclosed herein.

The engine <NUM> can be operated as follows. Referring to <FIG>, the fan <NUM> is driven by the fan drive turbine <NUM> such that the blades <NUM> deliver airflow to the bypass, intermediate and core flow paths B, I, C. Driving the fan <NUM> can include driving the hub <NUM> of the fan <NUM> through the geared architecture <NUM> at a different speed than a speed of the fan drive turbine <NUM>. The engine <NUM> can be operated in any of the modes disclosed herein, including the modes M1-M4 of <FIG>. Airflow in the intermediate flow path I can be ejected from the bypass port <NUM> into the bypass flow path B to bypass the guide vanes <NUM>. One or more blocker doors <NUM> can be selectively moved between open and closed positions to at least partially block airflow from the intermediate flow path I being communicated downstream to the core bypass duct <NUM> and vent port <NUM>. One or more of the blocker doors <NUM> can be moved between open and closed positions to selectively communicate airflow through the bypass port <NUM> (<NUM> shown in dashed lines in <FIG> for illustrative purposes).

Referring to <FIG>, an operating line of the fan <NUM> at an operating condition such as takeoff can be shifted relatively further away from the stall line by upflowing the fan <NUM> in response to selectively communicating airflow to the vent port <NUM>. One or more of the blocker doors <NUM> can be selectively moved to the open position to communicate airflow in the intermediate flow path I downstream to the vent port <NUM>. Airflow communicated to the core bypass duct <NUM> can be ejected from the vent port <NUM> downstream of the bypass port <NUM> and/or bypass exit <NUM>.

The disclosed splitter arrangements can improve aerodynamic loading and efficiency of the guide vanes, including lowering inlet Mach number to the guide vanes and improving aerodynamic diffusion in the guide vanes. The disclosed splitter arrangements can improve blade and guide vane performance near <NUM>% span, which can be utilized to achieve a relatively more efficient spanwise pressure ratio across the entire span of the blade and the entire span of the guide vane. Engines made with the disclosed architecture, and including arrangements as set forth in this application, and with modifications coming from the scope of the claims in this application, thus provide very high efficient operation, relatively high stall margins, and are compact and lightweight relative to their thrust capability. Two-spool and three-spool direct drive engine architectures can also benefit from the teachings herein.

Claim 1:
A gas turbine engine (<NUM>) comprising:
a fan section (<NUM>) including a fan having a row of blades (<NUM>) extending in a radial direction (R) between a <NUM>% span position at a hub (<NUM>) and a <NUM>% span position at a tip (162T), wherein the hub (<NUM>) is rotatable about an engine longitudinal axis (A) such that the row of blades (<NUM>) deliver flow to a bypass flow path (B), an intermediate flow path (I) and a core flow path (C);
a compressor section (<NUM>) establishing the core flow path (C);
a turbine section (<NUM>) that drives the fan section (<NUM>) and the compressor section (<NUM>);
a fan case (<NUM>) including a bypass duct (<NUM>) surrounding the row of blades (<NUM>) to establish the bypass flow path (B);
a housing including a first flow splitter (<NUM>) that divides flow between the bypass flow path (B) and a second duct (<NUM>);
a row of guide vanes (<NUM>) in the bypass duct (<NUM>) that extend in the radial direction (R) across the bypass flow path (B);
an engine case (<NUM>) including a second flow splitter (<NUM>) radially inboard of the first flow splitter (<NUM>) and that divides flow from the second duct (<NUM>) between the intermediate flow path (I) and the core flow path (C);
a bypass port (<NUM>) that interconnects the intermediate and bypass flow paths (I, B) at a position downstream of the row of guide vanes (<NUM>); and
one or more blocker doors (<NUM>) situated in the intermediate flow path (I) downstream of the bypass port (<NUM>), wherein the one or more blocker doors (<NUM>) are moveable to selectively communicate flow in the intermediate flow path (I) to a vent port (<NUM>);
wherein each of the blades (<NUM>) extends in the radial direction (R) between the <NUM>% span position and <NUM>% span position to establish a hub-to-tip ratio, and the hub-to-tip ratio is less than or equal to <NUM> measured relative to a forwardmost portion of a leading edge (162LE) of the blades (<NUM>), and wherein a forwardmost edge of the second flow splitter (<NUM>) is axially aft of a forwardmost edge of the first flow splitter (<NUM>) with respect to the engine longitudinal axis (A), and the forwardmost edges of the first and second flow splitters (<NUM>, <NUM>) are axially forward of the row of guide vanes (<NUM>) with respect to the engine longitudinal axis (A).