Patent Description:
An aircraft propulsion system includes an inlet structure and a gas turbine engine. The inlet structure directs air into the gas turbine engine. Some known inlet structures include a variable airflow inlet area for tailoring a mass flow of the air entering the gas turbine engine. While these known inlet structures have various advantages, there is still room in the art for improvement. There is a need in the art therefore for an improved inlet assembly with a variable airflow inlet area.

<CIT> and <CIT> both disclose a prior art assembly according to the preamble of claim <NUM>.

According to an aspect, an assembly is provided for an aircraft propulsion system as set forth in claim <NUM>.

The following optional features may be applied to the above aspect:
The nacelle inlet structure may be configured as or otherwise include a scarfed inlet lip.

The first component of the assembly may be configured as or otherwise include the center body. The second component of the assembly may be configured as or otherwise include the scarfed inlet structure.

The scarfed inlet structure may have a leading edge. A first point on the leading edge may be axially displaced from a second point on the leading edge along the axis.

A plane of an inlet orifice to the inlet opening passage may be angularly offset from the axis.

The first surface is a tubular surface that radially tapers as center body extends in a forward direction along the axis.

The second surface is a tubular surface that radially tapers as center body extends in an aft direction along the axis.

The plateau surface may be aligned with an inlet lip of the scarfed inlet structure. The trailing edge of the plateau surface may have a scarfed configuration.

The scarfed inlet structure may include a leading edge. The center body trailing edge may be separated from the leading edge by a first axial distance along the axis when the first component of the assembly is in the first position. The trailing edge may be separated from the leading edge by a second axial distance along the axis when the first component of the assembly is in the second position. The second axial distance may be different than the first axial distance.

The plateau surface is a cylindrical surface with an axial length that changes as the cylindrical surface extends circumferentially about the axis.

The axis may be coaxial with a centerline of the aircraft propulsion system.

The axis may be offset from a centerline of the aircraft propulsion system.

The present disclosure may include any one or more of the individual features disclosed above and/or below alone or in any combination thereof, insofar as they fall within the scope of the claims.

<FIG> illustrates an aircraft propulsion system <NUM> for an aircraft such as, but not limited to, a commercial airliner or cargo plane. The aircraft propulsion system <NUM> includes a gas turbine engine <NUM> and a nacelle <NUM>.

The gas turbine engine <NUM> may be configured as a high-bypass turbofan engine. The gas turbine engine <NUM> of <FIG>, for example, includes a fan section <NUM>, a compressor section <NUM>, a combustor section <NUM> and a turbine section <NUM>. The compressor section <NUM> may include a low pressure compressor (LPC) section 27A and a high pressure compressor (HPC) section 27B. The turbine section <NUM> may include a high pressure turbine (HPT) section 29A and a low pressure turbine (LPT) section 29B.

The engine sections <NUM>-29B are arranged sequentially along an axial centerline <NUM> (e.g., a rotational axis) of the gas turbine engine <NUM> within an aircraft propulsion system housing <NUM>. This housing <NUM> includes an outer housing structure <NUM> and an inner housing structure <NUM>.

The outer housing structure <NUM> includes an outer case <NUM> (e.g., a fan case) and an outer structure <NUM> of the nacelle <NUM>; i.e., an outer nacelle structure. The outer case <NUM> houses at least the fan section <NUM>. The outer nacelle structure <NUM> houses and provides an aerodynamic cover for the outer case <NUM>. The outer nacelle structure <NUM> also covers a portion of an inner structure <NUM> of the nacelle <NUM>; i.e., an inner nacelle structure, which may also be referred to as an inner fixed structure (IFS). More particularly, the outer nacelle structure <NUM> axially overlaps and extends circumferentially about (e.g., completely around) the inner nacelle structure <NUM>. The outer nacelle structure <NUM> and the inner nacelle structure <NUM> thereby at least partially or completely form a bypass flow path <NUM> within the aircraft propulsion system <NUM>.

The inner housing structure <NUM> includes an inner case <NUM> (e.g., a core case) and the inner nacelle structure <NUM>. The inner case <NUM> houses one or more of the engine sections 27A-29B, which engine sections 27A-29B may be collectively referred to as an engine core. The inner nacelle structure <NUM> houses and provides an aerodynamic cover for the inner case <NUM>.

Each of the engine sections <NUM>, 27A, 27B, 29A and 29B includes a bladed rotor <NUM>-<NUM>. The fan rotor <NUM> and the LPC rotor <NUM> are connected to and driven by the LPT rotor <NUM> through a low speed shaft <NUM>. The HPC rotor <NUM> is connected to and driven by the HPT rotor <NUM> through a high speed shaft <NUM>. The shafts <NUM> and <NUM> are rotatably supported by a plurality of bearings (not shown). Each of these bearings is connected to the aircraft propulsion system housing <NUM> (e.g., the inner case <NUM>) by at least one stationary structure such as, for example, an annular support strut.

During operation, air enters the aircraft propulsion system <NUM> through an inlet structure <NUM> of the outer nacelle structure <NUM>; i.e., a nacelle inlet structure. This air is directed through a duct <NUM> (e.g., a fan duct in the fan section <NUM>) and into a core flow path <NUM> and the bypass flow path <NUM>. The core flow path <NUM> extends axially along the axial centerline <NUM> within the aircraft propulsion system <NUM>, through the engine sections 27A-29B, to a core nozzle outlet, where the core flow path <NUM> is radially within the inner case <NUM>. The bypass flow path <NUM> extends axially along the axial centerline <NUM> within the aircraft propulsion system <NUM> to a bypass nozzle outlet, where the bypass flow path <NUM> is radially between the outer nacelle structure <NUM> and the inner nacelle structure <NUM>. The air within the core flow path <NUM> may be referred to as "core air". The air within the bypass flow path <NUM> may be referred to as "bypass air".

The core air is compressed by the compressor rotors <NUM> and <NUM> and directed into a combustion chamber of a combustor in the combustor section <NUM>. Fuel is injected into the combustion chamber and mixed with the compressed core air to provide a fuel-air mixture. The rotation of the LPT rotor <NUM> also drives rotation of the fan rotor <NUM>, which propels bypass air through and out of the bypass flow path <NUM>. The propulsion of the bypass air may account for a majority of thrust generated by the turbine engine <NUM>. The aircraft propulsion system <NUM> of the present disclosure, however, is not limited to the exemplary gas turbine engine configuration described above.

Optimal mass flow requirements of the air entering the aircraft propulsion system <NUM> through the nacelle inlet structure <NUM> may change depending upon one or more parameters. These parameters may include, but are not limited to, modes of operation, aircraft maneuvers and operating conditions. For example, where the aircraft flies at supersonic speeds, the nacelle inlet structure <NUM> may be configured to direct a first mass flow of the air into the aircraft propulsion system <NUM>. When the aircraft flies at subsonic speeds, the nacelle inlet structure <NUM> may be configured to direct a second mass flow of the air into the aircraft propulsion system <NUM>, where the second mass flow is greater than the first mass flow.

To accommodate changing mass flows, the aircraft propulsion system <NUM> of <FIG> includes a variable area inlet assembly <NUM>. The variable area inlet assembly <NUM> of <FIG> and <FIG> includes at least the nacelle inlet structure <NUM> and a center body <NUM>; e.g., an inlet cone or an inlet spike. At least these variable area inlet assembly components <NUM> and <NUM> collectively form an annular inlet duct, which inlet duct forms an inlet passage <NUM> into and within the aircraft propulsion system <NUM>. Referring to <FIG>, the inlet duct and its inlet passage <NUM> extend axially along the axial centerline <NUM> from an inlet orifice <NUM> of the inlet passage <NUM> to the fan duct <NUM>. The inlet passage <NUM> of <FIG> includes a metering portion <NUM>; e.g., a choke point. The term "metering portion" may describe a portion of the inlet passage <NUM> with the smallest cross-sectional flow area.

The variable area inlet assembly <NUM> is configured to provide a variable airflow inlet area. More particularly, at least the variable area inlet assembly components <NUM> and <NUM> are configured to provide the metering portion <NUM> of the inlet passage <NUM> with a variable cross-sectional area. Referring to <FIG> for example, during a first (e.g., supersonic) mode of operation, at least one of the variable area inlet assembly components <NUM> and <NUM> is configured in / moves (e.g., rotates) to a first position; e.g., a fully closed position. In this first position, the metering portion <NUM> has a first area 74A; e.g., a cross-sectional area in the plane of <FIG>. Referring to <FIG>, during a second (e.g., subsonic) mode of operation, the at least one of the variable area inlet assembly components <NUM> and <NUM> is configured in / moves (e.g., rotates) to a second position; e.g., a fully open position. In this second position, the metering portion <NUM> has a second area 74B; e.g., a cross-sectional area in the plane of <FIG>. The second area 74B is different (e.g., greater) than the first area 74A.

Referring to <FIG>, the nacelle inlet structure <NUM> is disposed at a forward, upstream end of the nacelle <NUM>. The nacelle inlet structure <NUM> may be configured as a stationary and/or scarfed inlet structure. The nacelle inlet structure <NUM> of <FIG> includes a tubular inner barrel <NUM>, a tubular outer barrel <NUM> and an annular scarfed inlet lip <NUM>.

The inner barrel <NUM> extends circumferentially about (e.g., completely around) an axial centerline <NUM> (e.g., an axis) of the variable area inlet assembly <NUM>, which centerline <NUM> may be coaxial with the axial centerline <NUM>. The inner barrel <NUM> extends axially along the axial centerline <NUM>, <NUM> between a first (e.g., forward, upstream) end <NUM> of the inner barrel <NUM> and a second (e.g., aft, downstream) end <NUM> of the inner barrel <NUM>. The inner barrel second end <NUM> of <FIG> is connected to a (e.g., forward, upstream) end of the outer case <NUM>. The inner barrel <NUM> may be configured to attenuate noise generated during aircraft propulsion system operation and, more particularly for example, noise generated by rotation of the fan rotor <NUM>. The inner barrel <NUM>, for example, may include at least one tubular noise attenuating acoustic panel or a circumferential array of arcuate noise attenuating acoustic panels arranged around the axial centerline. The present disclosure, however, is not limited to such an acoustic inner barrel configuration.

The outer barrel <NUM> extends circumferentially about (e.g., completely around) the axial centerline <NUM>, <NUM>. The outer barrel <NUM> extends axially along the axial centerline <NUM>, <NUM> between a first (e.g., forward, upstream) end <NUM> of the outer barrel <NUM> and a second (e.g., aft, downstream) end <NUM> of the outer barrel <NUM>. The outer barrel second end <NUM> of <FIG> is disposed next to respective (e.g., forward, upstream) ends of a pair of fan cowls (one visible in <FIG>) of the outer nacelle structure <NUM>.

Referring to <FIG>, the inlet lip <NUM> forms a leading edge <NUM> of the nacelle <NUM> as well as an outer peripheral boundary of the inlet orifice <NUM> and a forward, upstream portion of the inlet passage <NUM>. A flat plane <NUM> defined by at least three points along (or an entirety of) the nacelle leading edge <NUM> (see plane of <FIG>) is angularly offset from the axial centerline <NUM>, <NUM> by an angle <NUM>; e.g., an acute angle. With this arrangement, a first point <NUM> (e.g., one of the at least three points) on the nacelle leading edge <NUM> is axially displaced from a second point <NUM> (e.g., another one of the at least three points) on the nacelle leading edge <NUM> by a non-zero axial distance <NUM>. The first point <NUM> of <FIG> may be a forwardmost, upstream-most point along the nacelle leading edge <NUM> located at, for example, a gravitational top of the aircraft propulsion system <NUM>. The second point <NUM> of <FIG> may be an aftmost, downstream-most point along the nacelle leading edge <NUM> and/or diametrically opposed to the first point <NUM> located at, for example, a gravitational bottom of the aircraft propulsion system <NUM>. The present disclosure, however, is not limited to such an exemplary inlet lip configuration. For example, in other embodiments, the first point <NUM> and/or the second point <NUM> may be respectively located along opposing sides of the aircraft propulsion system <NUM>. Furthermore, while the first point <NUM> may be the only forwardmost, upstream-most point along the nacelle leading edge <NUM> and/or the second point <NUM> may be the only aftmost, downstream-most point along the nacelle leading edge <NUM>, the present disclosure is not limited to such an exemplary configuration. The nacelle leading edge <NUM>, for example, may alternatively have an undulating (e.g., wavy) geometry.

The inlet lip <NUM> of <FIG> has a cupped (e.g., a generally U-shaped or V-shaped) side sectional geometry. The inlet lip <NUM> and its cupped side sectional geometry extend circumferentially about (e.g., completely around) the axial centerline <NUM>, <NUM>; see also <FIG>. The inlet lip <NUM> of <FIG>, for example, includes axially overlapping inner and outer lip portions <NUM> and <NUM>. The inner lip portion <NUM> is connected to and may be integral with the outer lip portion <NUM> at and along the nacelle leading edge <NUM>. An (e.g., aft, downstream) end of the inner lip portion <NUM> is axially adjacent and/or connected to the inner barrel first end <NUM>. An (e.g., aft, downstream) end of the outer lip portion <NUM> is axially adjacent and/or connected to the outer barrel first end <NUM>.

Referring to <FIG>, the center body <NUM> may be configured as a double tapered center body. The center body <NUM> of <FIG>, for example, has a double tapered exterior skin that extends circumferentially about (e.g., completely around) an axial centerline <NUM> (e.g., a rotational axis) of the center body <NUM>, which centerline <NUM> may be coaxial with the axial centerline <NUM>, <NUM> (see also <FIG>). The center body <NUM> and its exterior surface extend axially along the axial centerline <NUM>, <NUM>, <NUM> from a first (e.g., forward, upstream) end <NUM> of the center body <NUM> to a second (e.g., aft, downstream) end <NUM> of the center body <NUM>. The exterior skin of <FIG> includes a first (e.g., forward, upstream) surface <NUM>, a second (e.g., aft, downstream) surface <NUM> and a (e.g., intermediate) plateau surface <NUM>.

The first surface <NUM> may be a tapered and/or tubular surface. The first surface <NUM> of <FIG>, for example, is configured as a substantially conical surface. This first surface <NUM> extends circumferentially about (e.g., completely around) the axial centerline <NUM>, <NUM>, <NUM>. The first surface <NUM> extends axially along the axial centerline <NUM>, <NUM>, <NUM> from a leading end tip (e.g., forward, upstream point) of the center body <NUM> at the center body first end <NUM> to a (e.g., annular) leading edge <NUM> of the plateau surface <NUM>. The first surface <NUM> flares (e.g., continuously or intermittently) radially away from the axial centerline <NUM>, <NUM>, <NUM> as the center body <NUM> and its surface <NUM> extend axially in an aft, downstream direction along the axial centerline <NUM>, <NUM>, <NUM> from the leading end tip towards (e.g., to) the plateau surface <NUM>. The first surface <NUM> thereby tapers (e.g., continuously or intermittently) radially towards the axial centerline <NUM>, <NUM>, <NUM> as the center body <NUM> and its surface <NUM> extend axially in a forward, upstream direction along the axial centerline <NUM>, <NUM>, <NUM> from the plateau surface <NUM> towards (e.g., to) the leading end tip.

The second surface <NUM> may be a tapered and/or tubular surface. The second surface <NUM> of <FIG>, for example, is configured as a substantially conical surface. This second surface <NUM> extends circumferentially about (e.g., completely around) the axial centerline <NUM>, <NUM>, <NUM>. The second surface <NUM> extends axially along the axial centerline <NUM>, <NUM>, <NUM> from a trailing edge of the center body <NUM> at the center body second end <NUM> to a (e.g., annular) trailing edge <NUM> of the plateau surface <NUM>. The second surface <NUM> tapers (e.g., continuously or intermittently) radially towards the axial centerline <NUM>, <NUM>, <NUM> as the center body <NUM> and its surface <NUM> extend axially in the aft, downstream direction along the axial centerline <NUM>, <NUM>, <NUM> from the plateau surface <NUM> towards (e.g., to) the center body trailing edge. The second surface <NUM> thereby flares (e.g., continuously or intermittently) radially away from the axial centerline <NUM>, <NUM>, <NUM> as the center body <NUM> and its surface <NUM> extend axially in the forward, upstream direction along the axial centerline <NUM>, <NUM>, <NUM> from the center body trailing edge towards (e.g., to) the plateau surface <NUM>.

The plateau surface <NUM> may be an apex surface of the center body <NUM>. The plateau surface <NUM> of <FIG>, for example, is configured as a substantially cylindrical surface. This plateau surface <NUM> extends circumferentially about (e.g., completely around) the axial centerline <NUM>, <NUM>, <NUM>. The plateau surface <NUM> extends axially along the axial centerline <NUM>, <NUM>, <NUM> from the plateau surface leading edge <NUM> to the plateau surface trailing edge <NUM>. A flat plane <NUM> defined by at least three points along (or an entirety of) the plateau surface leading edge <NUM> is angularly offset from the axial centerline <NUM>, <NUM>, <NUM> by a first angle <NUM>; e.g., a right angle. A flat plane <NUM> defined by at least three points along (or an entirety of) the plateau surface trailing edge <NUM> is angularly offset from the axial centerline <NUM>, <NUM>, <NUM> by a second angle <NUM>; e.g., an acute angle. The second angle <NUM> of <FIG> is different (e.g., less) than the first angle <NUM>. The plateau surface trailing edge plane <NUM> of <FIG> is thereby angularly offset from the plateau surface leading edge plane <NUM>. The plateau surface trailing edge plane <NUM> may (or may not) be parallel with the inlet structure leading edge plane <NUM> (see <FIG>). Thus, the plateau surface trailing edge plane <NUM> may have a scarfed configuration.

A first point <NUM> (e.g., one of the at least three points) on the plateau surface trailing edge <NUM> is axially displaced from a second point <NUM> (e.g., another one of the at least three points) on the plateau surface trailing edge <NUM> by a non-zero axial distance <NUM>. The first point <NUM> of <FIG> may be a forwardmost, upstream-most point along the trailing edge <NUM> located at, for example, the gravitational top of the aircraft propulsion system <NUM> during the first (e.g., supersonic) mode of operation. The second point <NUM> of <FIG> may be an aftmost, downstream-most point along the trailing edge <NUM> and/or diametrically opposed to the first point <NUM> located at, for example, the gravitational bottom of the aircraft propulsion system <NUM> during the first (e.g., supersonic) mode of operation. The present disclosure, however, is not limited to such an exemplary inlet lip configuration. For example, in other embodiments, the first point <NUM> and/or the second point <NUM> may be respectively located along opposing sides of the center body <NUM> during the first (e.g., supersonic) mode of operation. Furthermore, while the first point <NUM> may be the only forwardmost, upstream-most point along the plateau surface trailing edge <NUM> and/or the second point <NUM> may be the only aftmost, downstream-most point along the plateau surface trailing edge <NUM>, the present disclosure is not limited to such an exemplary configuration. The plateau surface leading edge <NUM>, for example, may alternatively have an undulating (e.g., wavy) geometry.

The plateau surface <NUM> has an axial length <NUM>. This plateau surface axial length <NUM> (e.g., continuously or intermittently) changes as the plateau surface <NUM> extends circumferentially about (e.g., completely around) the axial centerline <NUM>, <NUM>, <NUM>. The plateau surface axial length <NUM> may have a first (e.g., maximum) value at the first point <NUM>. The plateau surface axial length <NUM> may have a second (e.g., minimum) value at the second point <NUM>, which second value is different (e.g., less) than the first value.

The center body <NUM> of <FIG> and <FIG> is arranged within the nacelle inlet structure <NUM> and projects axially through the inlet orifice <NUM>. The center body <NUM> at least partially (or completely) forms an inner peripheral boundary of the inlet passage <NUM>. The plateau surface <NUM> is axially aligned with the inlet lip <NUM>.

The center body <NUM> may be configured to rotate (e.g., clockwise or counter-clockwise) about the axial centerline <NUM>, <NUM>, <NUM> between the first position (see <FIG>) and the second position (see <FIG>). The center body <NUM>, for example, may be rotationally supported by one or more bearing structures, track assemblies and/or other suitable slidable / movable / pivotable connectors (not shown).

In the first position of <FIG>, an inner peripheral boundary of the metering portion <NUM> is (e.g., completely) formed by the plateau surface <NUM>. In this position, a minimum distance <NUM> between the inlet lip <NUM> and the center body <NUM> and, more particularly, between the inner lip portion <NUM> and the plateau surface <NUM> may be at least substantially or completely equal / uniform about the axial centerline <NUM>, <NUM>, <NUM>. An axial distance <NUM> between the nacelle leading edge <NUM> and the plateau surface trailing edge <NUM> may also or alternatively be at least substantially or completely equal / uniform about the axial centerline <NUM>, <NUM>, <NUM>.

In the second position of <FIG>, the inner peripheral boundary of the metering portion <NUM> is formed by at least (or only) a portion of the plateau surface <NUM> and a portion of the second surface <NUM>. In this position, the minimum distance <NUM> may (e.g., continuously or intermittently) change about the axial centerline <NUM>, <NUM>, <NUM>. For example, the minimum distance <NUM> at a first point <NUM> along the metering portion <NUM> may be different (e.g., less) than the minimum distance at a second point <NUM> along the metering portion <NUM>. The first point <NUM> of <FIG> may be at the gravitational top of the aircraft propulsion system <NUM>. The second point <NUM> of <FIG> may be diametrically opposed to the first point <NUM> and/or at the gravitational bottom of the aircraft propulsion system <NUM>. The minimum distance <NUM> at the first point <NUM> may be equal to the minimum distance <NUM> when the center body <NUM> is in the first position of <FIG>. The axial distance <NUM> between the nacelle leading edge <NUM> and the plateau surface trailing edge <NUM> may also or alternatively (e.g., continuously or intermittently) change about the axial centerline <NUM>, <NUM>, <NUM>. For example, the axial distance <NUM> at the first point <NUM> may be different (e.g., greater) than the axial distance <NUM> at the second point <NUM>. With the foregoing arrangement, the second area 74B of the metering portion <NUM> (see <FIG>) is different (e.g., greater) than the first area 74A of the metering portion <NUM> (see <FIG>).

In some embodiments, the axial centerline / rotational axis <NUM> of the center body <NUM> is coaxial with the axial centerline / rotational axis <NUM> of the aircraft propulsion system <NUM> and its gas turbine engine <NUM> as described above. However, in other embodiments, the axial centerline / rotational axis <NUM> of the center body <NUM> may be eccentric / non-coaxial with the axial centerline / rotational axis <NUM> of the aircraft propulsion system <NUM> and its gas turbine engine <NUM> as shown in <FIG>. More particularly, the axial centerline <NUM> may be displaced from and/or angularly offset from the axial centerline <NUM>.

In some embodiments, referring to <FIG>, the center body <NUM> may rotate one-hundred and eighty degrees (<NUM>°) between the first position of <FIG> and the second position of <FIG>. However, in other embodiments, the center body <NUM> may rotate more than hundred and eighty degrees (e.g., between hundred and eighty degrees (<NUM>°) and two-hundred and seventy degrees (<NUM>°)) between the first and the second positions. In still other embodiments, the center body <NUM> may rotate less than hundred and eighty degrees (e.g., between ninety degrees (<NUM>°) and hundred and eighty degrees (<NUM>°)) between the first and the second positions.

In some embodiments, the center body <NUM> may be actuated by a gear drive system. In other embodiments, the center body <NUM> may be actuated by one or more other types of actuators such as, but not limited to, one or more worm and gear arrangements and/or one or more linear actuators arranged around a periphery of the center body <NUM>.

The center body <NUM> is described above as a movable structure, and the nacelle inlet structure <NUM> is described above as a stationary structure. However, it is contemplated that the functionality / operation of these structures <NUM> and <NUM> may be reversed. For example, in some embodiments, the center body <NUM> may be configured as a stationary structure, and the nacelle inlet structure <NUM> and/or its inlet lip <NUM> may be configured as a movable structure; e.g., a rotatable structure. The nacelle inlet structure <NUM> and/or its inlet lip <NUM> may thereby move (e.g., rotate clockwise or counter-clockwise about the centerline <NUM>, <NUM>, <NUM>) between the first and the second positions to increase or decrease the area of the metering portion <NUM>.

The aircraft propulsion system <NUM> and its variable area inlet assembly <NUM> may be configured with various gas turbine engines other than the one described above. The gas turbine engine <NUM>, for example, may be configured as a geared or a direct drive turbine engine. The gas turbine engine <NUM> may be configured with a single spool, with two spools (e.g., see <FIG>), or with more than two spools. The gas turbine engine <NUM> may be configured as a turbofan engine, a turbojet engine or any other type of turbine engine. The present invention therefore is not limited to any particular types or configurations of gas turbine engines. The present disclosure is also not limited to applications where the aircraft is capable to traveling supersonic speeds.

Claim 1:
An assembly (<NUM>) for an aircraft propulsion system (<NUM>), comprising:
a center body (<NUM>);
a scarfed inlet structure (<NUM>) extending circumferentially about the center body (<NUM>); and
an inlet passage (<NUM>) radially between and formed by at least the center body (<NUM>) and the scarfed inlet structure (<NUM>), the inlet passage (<NUM>) comprising a metering portion (<NUM>);
a first component of the assembly (<NUM>) configured to rotate about an axis (<NUM>) relative to a second component of the assembly (<NUM>) between:
a first position where the metering portion (<NUM>) has a first area (74A); and
a second position where the metering portion (<NUM>) has a second area (74B) that is different than the first area (74A);
wherein the first component of the assembly (<NUM>) comprises one of the center body (<NUM>) or the scarfed inlet structure (<NUM>), and the second component of the assembly (<NUM>) comprises the other one of the center body (<NUM>) or the scarfed inlet structure (<NUM>);
characterised in that:
the center body (<NUM>) includes a first tapered surface (<NUM>), a second tapered surface (<NUM>) and a plateau surface (<NUM>) that extends axially between the first tapered surface (<NUM>) and the second tapered surface (<NUM>);
the metering portion (<NUM>) has an inner peripheral boundary;
the inner peripheral boundary is formed by the plateau surface (<NUM>) when the first component of the assembly (<NUM>) is in the first position;
the inner peripheral boundary is formed by the plateau surface (<NUM>) and the second tapered surface (<NUM>) when the first component of the assembly (<NUM>) is in the second position;
the plateau surface (<NUM>) comprises a trailing edge (<NUM>); and
a first point (<NUM>) on the trailing edge (<NUM>) is axially displaced from a second point (<NUM>) on the trailing edge (<NUM>) along the axis (<NUM>).