Patent Description:
Turboprop gas turbine engines are most commonly mounted to aircraft in a "tractor" (or "puller") configuration, whereby the engine is mounted with the propeller located forward of the engine (relative to a direction of travel of the aircraft) such that the aircraft is "pulled" through the air by the propeller. Turboprop engines may however also be mounted to an aircraft in a "pusher" configuration, whereby the engine is mounted with the propeller located behind the engine (relative to the direction of travel of the aircraft). Much as engines mounted in a tractor configuration can be mounted to the wings or in the nose of the fuselage, pusher engines can be either wing mounted, mounted to the fuselage tail and/or to pylons on the aircraft. <CIT> discloses a gas turbine engine having a heat exchanger axially aligned with the engine core.

<CIT> and <CIT> disclose engine installations for the prior art.

Challenges exist with some existing pusher turboprop installations, however, which make the use of pusher configurations less common than tractor powerplant installations. In a pusher powerplant installation, for example, the center of gravity of the powerplant may be located further rearward in comparison with those in puller configurations, and the overall envelope of the nacelle required may also be longer than for a comparable tractor powerplant configuration, more complex structural mounting configurations may be required, air inlet ducting may be more difficult, and/or foreign object damage considerations may also be more difficult to manage.

Improvements in pusher gas turbine engine powerplants and their installations are therefore sought.

In a first aspect, there is provided a turboprop gas turbine engine adapted to be mounted to an aircraft according to claim <NUM>.

Optional embodiments are defined by the dependent claims.

Further details of these and other aspects of the subject matter of this application will be apparent from the detailed description and drawings included below.

<FIG> illustrates an exemplary aircraft <NUM> to which one or more gas turbine engines <NUM> may be mounted. The gas turbine engines <NUM> are turboprops, and are mounted to the aircraft in pusher configurations, whereby the engine is mounted with the propeller <NUM> located behind the engine (relative to the direction of travel <NUM> of the aircraft <NUM> - also referred to herein and depicted in the drawings as the "pilot view" direction). In the embodiment of <FIG>, the pusher turboprop gas turbine engines <NUM> (or simply "pusher engines") are mounted to the wings <NUM> of the aircraft.

The terms "upstream" and "downstream" as used herein, unless indicated otherwise, are understood to be relative to the direction of travel <NUM> of the aircraft (i.e. the "pilot view" direction). Similarly, the terms "forward" and "rearward" as used herein are also understood to be relative to the direction of travel <NUM> of the aircraft.

Each pusher engine <NUM> as described herein will generally be referred to in the singular, however it is to be understood that two or more or each of such engines may be provided on the aircraft <NUM>.

Each pusher engine <NUM> defines a longitudinal axis LA (e.g., central axis). In various embodiments, longitudinal axis LA may correspond to an axis of rotation of propeller <NUM> and/or longitudinal axis LA may correspond to an axis of rotation of a low-pressure spool and/or a high-pressure spool of a core <NUM> of the gas turbine engine <NUM>. Each gas turbine engine <NUM> may be housed in a nacelle <NUM>, serving as an aerodynamically-shaped covering for gas turbine engine <NUM>.

The pusher engine <NUM> accordingly includes an engine core <NUM> (including compressor(s) <NUM>, combustor <NUM> and turbine(s) <NUM>) and a reduction gearbox <NUM> which drives the propeller <NUM>, which is located rearward of the gas turbine engine <NUM>, relative to a direction of travel of the aircraft, as is the case for pusher-style engines.

The pusher engines as described in further detail below are generally intended to be wing-mounted, in that the nacelle <NUM> and the engine <NUM> are mounted to, and overtop of, a wing <NUM> of the aircraft <NUM>. However, as mentioned above, it is to be understood that the pusher turboprop engines and installations described herein may also be adapted for being mounted to the fuselage of the aircraft and/or to pylons mounted to the aircraft.

Referring to <FIG>, the gas turbine engine <NUM> comprises an air intake <NUM> for channeling a flow of ambient air into gas turbine engine <NUM>. Air intake <NUM> comprises intake inlet <NUM> that is generally forward-facing and, in at least one particular embodiment, may be substantially aligned with the longitudinal axis LA of the engine <NUM>. In typical pusher installations, the air intake is often offset radially outwardly (e.g. downwardly) relative to the longitudinal axis LA of engine. In at least some of the embodiments described herein, however, the intake inlet <NUM> of the air intake <NUM> is substantially axially aligned with the gas turbine engine <NUM>, thereby allowing for a slimmer (e.g. smaller diameter) nacelle <NUM>. Intake inlet <NUM> may be generally forward-facing, so that ambient air ingested into the air intake <NUM> is directed rearward within the envelope of the nacelle and to the engine air inlet <NUM> of the engine <NUM>.

As best seen in <FIG>, the pusher engine <NUM> may be of a type suitable for use in aircraft applications for subsonic flight generally comprising, in serial flow communication, air intake <NUM> through which ambient air is received, a compressor section <NUM> (which may be a multistage compressor) for pressurizing the air, a combustor <NUM> in which the compressed air is mixed with fuel and ignited for generating an annular stream of hot combustion gases, and turbine section <NUM> (which may be a multistage turbine) for extracting energy from the combustion gases. In various embodiments, gas turbine engine <NUM> may have a dual-spool configuration but it is understood that gas turbine engine <NUM> may not be limited to such configuration. For example, gas turbine engine <NUM> may comprise high-pressure spool <NUM> including one or more stages of multistage compressor <NUM> and one or more high-pressure turbines <NUM> of turbine section <NUM>. Gas turbine engine <NUM> may also comprise low-pressure spool <NUM> including one or more stages of multistage compressor <NUM> and one or more low-pressure (i.e., power) turbines <NUM> of turbine section <NUM>. Low-pressure spool <NUM> may be mechanically coupled to output shaft <NUM> via a reduction gearbox <NUM>, to which the propeller <NUM> may be coupled.

Referring still to <FIG>, in one particular embodiment the air intake <NUM> of the pusher engine <NUM> includes a forward-facing intake inlet <NUM> receiving the ambient air, and an inlet duct <NUM> in fluid communication with the intake inlet <NUM> and extending downstream therefrom. More particularly, the air inlet duct <NUM> extends from the intake inlet <NUM> to the engine air inlet <NUM>, such as to provide the main airflow into the core <NUM> of the engine <NUM>. The air inlet duct <NUM> includes a first, or upstream, section <NUM> that extends immediately downstream of the intake inlet <NUM> and may be, in one particular embodiment, centrally located such that an intake axis IA defined by the intake inlet <NUM> is substantially concentric with the longitudinal axis LA of the engine <NUM>. A second, or downstream, section <NUM> of the air inlet duct <NUM> terminates at the engine air inlet <NUM>. As best seen in <FIG>, in one particular embodiment, this second downstream section <NUM> of the air inlet duct <NUM> may be bifurcated into two separate ducts <NUM>' and <NUM>". In the embodiment of <FIG>, the two separate ducts <NUM>' and <NUM>" are bifurcated relative to a central vertical plane extending through the longitudinal engine axis LA. Accordingly, the outlets of each of these ducts <NUM>' and <NUM>" may thereby direct the inlet airflow into opposite sides of an annual quasi-scroll intake <NUM> that that feeds the engine air inlet <NUM>. In the embodiment of <FIG>, therefore, the downstream duct section <NUM> is bifurcated, thereby comprising the bifurcated duct portions <NUM>' and <NUM>" which are split at the top-dead-center (TDC) and may thus be located at <NUM> and <NUM> o'clock, for example. The bifurcated duct portions <NUM>' and <NUM>" of the downstream duct <NUM> feed air to the engine air inlet <NUM> at two circumferential locations, thereby partially wrapping around the annular engine inlet and thus forming a "quasi-scroll" type air inlet duct configuration. This quasi-scroll inlet configuration may contribute to ensuring relatively low losses and enable some ram recovery of the air entering the engine core <NUM>. By integrating the front-facing air inlet <NUM> and inlet duct <NUM>, and such as quasi-scroll type inlet ducting to the engine air inlet <NUM>, overall engine performance (e.g. specific fuel consumption (SPC)) may accordingly be improved.

As seen in both <FIG>, an additional air duct <NUM> (which may also be referred to herein as a "second air outlet duct" or simply as an "oil cooler duct"), directs a portion of the incoming cool air into an air-cooled-oil-cooler (ACOC) <NUM>. In the depicted embodiment, the ACOC air duct <NUM> is fluidly connected to the air inlet duct <NUM> at a point thereon between the first, upstream, section <NUM>, and the second, downstream, section <NUM>. This connection point may also be located over the wing <NUM> and downstream (and axially aft) of an inertial particle separator duct <NUM>, as will be described in further detail below. More particularly, between the intake inlet <NUM> and the engine air inlet <NUM>, a portion of the airflow within the main air inlet duct <NUM> is diverted away into the ACOC inlet duct <NUM>. The air intake <NUM> therefore includes two separate outlet ducts, namely the main air inlet flow within the downstream duct section <NUM> of the air inlet duct <NUM>, which directs a portion of the inlet airflow into the core <NUM> of the engine <NUM>, and the ACOC air duct <NUM>, which directs another portion of the incoming air into an air-cooled-oil-cooler (ACOC) <NUM>. The portion of the inlet airflow directed to the engine core may be a major or majority portion of the incoming airflow, and the portion of the airflow directed into the ACOC <NUM> may be a small portion of the total incoming airflow. As such, the air duct <NUM> directs a portion of the air flowing through the inlet duct <NUM> to the ACOC <NUM>, and a remaining portion of the incoming air is directed through the downstream duct portion <NUM> to the engine core air inlet <NUM>.

The ACOC <NUM> may be located in-line within the ACOC air duct <NUM>. The ACOC <NUM> is used to cool engine oil, by transferring the heat from the engine oil to the cooler air flowing through the second outlet duct <NUM>. After flowing through the ACOC <NUM>, the air (by then heated by the ACOC) can flow out of the second outlet duct <NUM> for ejection - either externally to atmosphere or within the nacelle <NUM> - or re-use for other purposes (e.g. ant-icing over nacelle or engine components.

Accordingly, as described above and shown in <FIG>, air inters the air intake <NUM> through the common inlet duct <NUM>, and at a point located downstream of the intake inlet <NUM> splits into the main engine downstream duct portion <NUM> and the ACOC air inlet duct <NUM>, in order to feed incoming air into the engine core <NUM> and to the ACOC <NUM>, respectively. As shown, in this embodiment, the second outlet duct <NUM> and the ACOC <NUM> that is fed by it are both located axially forward of the engine core <NUM> of the engine <NUM>, but may be substantially aligned therewith (relative to the axis LA) such as to produce an overall relative "long and thin" envelope for the nacelle <NUM>, which therefore has a relatively long forward "snout". As seen in <FIG>, the ACOC <NUM> and the air duct <NUM> feeding it may be axially located in line (or overtop of) the wings <NUM> of the aircraft, with the majority of the engine <NUM> and/or the center of the mass of the engine <NUM> being located axially after of the wings <NUM>.

As shown in <FIG>, the first, most upstream, section <NUM> of the inlet duct <NUM> of the air intake <NUM> is centrally located within the nacelle <NUM> relative to the engine centerline LA. As such, intake axis IA of the first section <NUM> of the inlet duct <NUM> may be substantially coaxial with the main engine axis LA, as seen in <FIG>. The downstream duct section <NUM>, and thus the bifurcated duct portions <NUM>' and <NUM>", may however be radially offset from the inlet duct <NUM> and thus the intake axis IA, extending up and radially outward (over the accessory gear box <NUM> which is also axially aligned with the engine axis LA in order to maintain a slim radial envelope) before redirecting radially inward to the engine air inlet <NUM>.

Given that the turboprop engine <NUM> is a pusher engine with the propeller <NUM> located aft of the engine core <NUM> and having a front-facing air inlet <NUM>, the possible effect of foreign object damage (FOD) caused by particles (e.g. ice, debris, etc.) in the incoming air should be considered. Accordingly, as seen in <FIG>, the air intake <NUM> also optionally includes an inertial particle separator (or "IPS") <NUM>. The IPS <NUM> is located upstream, and axially forward, of the engine air inlet <NUM>, so as to prevent ingestion of ice, other particles and/or other FOD-causing objects into the engine. The terms "particles" or simply "FOD" may be used herein to describe generally all of such FOD particles/objects, even though it will be understood that in some cases the FOD-causing objects in question may be significantly larger than an ice or other particle per se - such as a bird or other airborne object for example that could be unintentionally ingested into the engine's air intake. In the embodiment of <FIG>, the IPS <NUM> is also located upstream, and axially forward, of the ACOC <NUM>. More particularly, the IPS <NUM> includes an IPS duct <NUM> that is fluidly connected with the upstream section <NUM> of the air inlet duct <NUM>, and a FOD particle separator that is capable of separating FOD particles from the incoming airflow and re-directing them into the IPS duct <NUM>. Accordingly, such undesirable FOD particles prevented from flowing unhindered downstream, through the air inlet duct <NUM> and thus into the air inlet of the engine <NUM>.

In the embodiment of <FIG>, the IPS <NUM> includes a one or more FOD-deflectors, or FOD-diverters, that extend at least partially into the air inlet duct and act to redirect any FOD into the IPS duct <NUM>. In one particular embodiment, the FOD-deflectors comprise one or more FOD-diverting plates <NUM> (or simply "plates" hereinbelow) which project into the upstream section <NUM> of the air inlet duct <NUM>, at a location upstream of the IPS duct <NUM>. The FOD diverting plates <NUM>, which in the depicted embodiment include three plates however fewer or more plates may alternately be used, may be either stationary or movable such as to be adjusted into a desired position relative to one another and relative to the incoming airflow through the inlet duct <NUM>. In one particularly embodiment, wherein at least one plate <NUM> is movable, the plate is displaceable between a deployed position and a retracted position. In the deployed position, the plate will protrude into the air inlet duct to redirect FOD. In the retracted position, the plate is substantially withdrawn from the airflow through the air inlet duct and will therefor generate less flow restriction. It is to be understood that in some embodiments, the moveable plate may be positioned not just at these two end points of its travel, i.e. at the fully retracted and fully deployed position, but may also be positioned a number of possible intermediate positions between these two extremes. In the depicted embodiment, each of the three plates <NUM> may be independent movable, and each may comprise either a two-state active control (e.g. being either deployed or retracted, as needed) or a fully variable active control whereby any desired angular position of each of the plates can be selected as required.

Regardless of whether they are fixed or movable, the plates <NUM> are positioned such as to divert any potential FOD-causing particles that might enter the inlet duct <NUM> into the IPS duct <NUM>, for subsequent discharge overboard. Unwanted FOD-causing particles are therefore diverted into the IPS duct <NUM>, and prevented from flowing further downstream and thus from being ingested into the engine air inlet <NUM> or the ACOC <NUM>. In one particular embodiment, a plurality of plates <NUM> are provided and extend into the air inlet duct <NUM>, with at least a first one of the plurality of plates extending away from an inner wall of the air inlet duct in a downstream direction, and at least a second one of the plurality of plates extending from an inner wall of the air inlet duct in an upstream direction. In some embodiments, the plate which is more downstream within the air inlet duct will be the one which extends in the upstream direction away from the wall of the air inlet duct. These first and second ones of the plates may be mounted to the same inner wall of the air inlet duct, or may be located on opposite sides from each other with the duct.

The IPS duct <NUM> may therefore be integrated with the inlet duct <NUM>, and it projects away therefrom downward and radially away from the inlet axis IA. As can be appreciated from <FIG>, the IPS duct <NUM> in this embodiment is separate and distinct from, and located upstream of, the second outlet duct <NUM> feeding air to the ACOC <NUM>. The IPS duct <NUM> may thus be axially located forward of the wing <NUM> of the aircraft, such as to expel any unwanted FOD particles outboard and away from the aerodynamic surfaces of the wings <NUM>. The second outlet duct <NUM> and the ACOC <NUM>, in this embodiment, are axially positioned over the wing <NUM>, with the engine core <NUM> of the gas turbine engine <NUM> being located substantially rearward of the wing <NUM>.

Several other embodiments of the present air intakes will now be described. Unless otherwise indicated, the features of each of the following engines and their respective air inlets will be similar to those of the pusher engine <NUM> and the air inlet <NUM> as described above. Only the differences will be described in more detail below, for the avoidance of repetition.

Referring now to the embodiment of <FIG>, the pusher engine <NUM> includes an air intake <NUM> in which the IPS duct and the ACOC duct are integrated into a single air duct <NUM>, which is connected to the main air inlet duct <NUM> such as to divert a portion of the flow therethrough. More particularly, the IPS duct <NUM> of the present air intake configuration includes the ACOC <NUM> therein, thereby avoiding the need for an additional air duct exclusively for the ACOC. The combined IPS and ACOC duct <NUM> is located forward of the wing <NUM> of the aircraft.

Additionally, unlike the air inlet duct <NUM> described above, the air inlet duct <NUM> is not bifurcated at TDC, as can be best seen in <FIG>. The downstream section <NUM> of the air inlet duct <NUM> therefore connects to the engine air inlet quasi-scroll <NUM> at a single location, for example at TDC. The engine air inlet is therefore a single inlet, quasi-scroll configuration.

Referring now to the embodiment of <FIG>, the pusher engine <NUM> includes an air intake <NUM> which is formed such that the overall axial length of the nacelle <NUM> required is shorter than that of the engines <NUM> and <NUM>, thereby providing a so-called "medium snout" nacelle design, as opposed to the longer nacelle snouts of engines <NUM>, <NUM>. Much as per the air intake <NUM>, the air intake <NUM> includes an integrated IPS and ACOC duct <NUM> that includes the ACOC <NUM>. In this embodiment, however, the IPS <NUM> employs one or more fixed elements projecting into the upstream portion <NUM> of the main air inlet duct <NUM>. For example, the one or more fixed elements may include a fixed icing screen, which will capture ice particles within the incoming airflow. The fixed icing screen <NUM> may be part of a two state, passive system, rather than the active systems of the movable plates <NUM> described above.

Inlet screens as described herein may comprise a metallic screen acts in operation to substantially prevent foreign objects (e.g., pieces of ice) larger than a certain size from exiting intake outlet. The screen(s) may also serve as a surface on which ice is permitted to accrete, thereby preventing or reducing the likelihood of ice accreting further downstream into gas turbine engine.

Referring now to the embodiment of <FIG>, the pusher engine <NUM> includes an air intake <NUM> which is even further simplified, in that it includes no fixed icing screen or other projecting plates as part of an IPS. Much as per the air intakes <NUM> and <NUM>, the air intake <NUM> similarly includes an integrated IPS and ACOC duct <NUM> that includes the ACOC <NUM>.

Referring now to the embodiment of <FIG>, the pusher engine <NUM> includes an air intake <NUM> which is similar to that of air intake <NUM> described above, in that it includes a separate IPS duct <NUM> and ACOC duct <NUM>. In this embodiment, however, no IPS plates or other icing screens (fixed or otherwise) are provided as part of an IPS system.

Referring now to the unclaimed example of <FIG>, the pusher engine <NUM> includes an air intake <NUM> that includes both an IPS duct <NUM> and an ACOC duct <NUM>. However, in this unclaimed example, the ACOC duct <NUM> is not integrated with or connected directly to the main air inlet duct <NUM> feeding the air to the engine core. Rather, the ACOC duct <NUM> containing the ACOC <NUM> is mounted separately within the nacelle <NUM>, over the wing <NUM>, and receives incoming airflow from air flowing through the nacelle <NUM>. In other words, the air directed through the ACOC duct <NUM> is not air which is diverted off from the main air inlet flow to the engine core. All of the air entering the front-facing air inlet <NUM> is therefore directed through the air inlet duct <NUM> to the engine air inlet <NUM>.

<FIG> depicts another unclaimed example, similar to that of <FIG>, except wherein the ACOC duct <NUM> is mounted below the wing <NUM>, instead of above it. As such, the air directed through the ACOC duct <NUM> can be drawn from outside the nacelle <NUM> through inlet <NUM>. However, in an alternate unclaimed example, the inlet <NUM> of the ACOC duct <NUM> may still be located such as to dawn air from outside the nacelle <NUM>, however with the ACOC duct <NUM> located elsewhere within the nacelle.

<FIG> depict alternate unclaimed examples, wherein the ACOC duct is disposed in other locations within the nacelle, generally aft of the wing of the aircraft. The ACOC duct may similarly be separate from (<FIG>) or integrated with (<FIG>) the IPS duct.

<FIG> depicts an alternate unclaimed example, similar to that of <FIG>, except having an air inlet plenum (or "dump box") at the engine air inlet, rather than the quasi-scroll air inlet configuration.

Referring now to the unclaimed example of <FIG>, the pusher engine <NUM> includes an air intake <NUM> that includes an IPS duct <NUM> that and an ACOC duct <NUM>. The IPS duct <NUM> is fluidly connected with the upstream section <NUM> of the air inlet duct <NUM>, and a FOD particle separator that is capable of separating FOD particles from the incoming airflow and re-directing them into the IPS duct <NUM>, for ejection overboard at <NUM>. Accordingly, such undesirable FOD particles are prevented from flowing unhindered downstream, through the air inlet duct <NUM> and thus into the air inlet <NUM> of the engine <NUM>.

In this unclaimed example, the air intake <NUM> and the upstream section <NUM> of the air intake <NUM> is located near the upper portion of the nacelle, and is not axially aligned with the engine. More particularly, the air intake <NUM> of the upstream section <NUM> of the air inlet duct <NUM> defines an intake axis IA that is radially offset (parallel to and spaced apart) from the from the main engine axis LA, as can be seen in <FIG>.

As can also be seen in <FIG>, in this unclaimed example, the ACOC duct <NUM> is split apart from the IPS duct <NUM> and the rest of the air inlet duct <NUM> of the air intake <NUM>. The ACOC duct <NUM> is located, in this unclaimed example, below the engine core, and draws air directly from outside the engine and/or outside the nacelle, for feeding through the ACOC in the ACOC duct <NUM>. Accordingly, the airflow used to cool the ACOC, which flows through the ACOC duct <NUM>, is independent from the airflow through the air intake <NUM>, which feeds the engine.

<FIG> depicts an alternate unclaimed example, similar to that of <FIG>, having an ACOC duct <NUM> that is split apart from the IPS duct <NUM> and the rest of the air inlet duct <NUM> of the air intake <NUM>. In this unclaimed example, however, the ACOC duct <NUM> is located above the engine core, rather than below, and is axially located rearward of the IPS duct <NUM> and the air intake <NUM>. Nevertheless, the airflow feeds to the air intake <NUM> and the ACOC duct <NUM> are separate, and each draws ambient air in from different external inlets in the nacelle.

In various embodiments, the air intake may channel the flow of ambient air (represented by the arrow F in the figures) toward engine inlet of gas turbine engine. The engine inlet may have a substantially annular shape and may be disposed upstream of compressor <NUM> of the engine. For the purpose of description and reference with the figures, the air intake may define an intake axis IA, which in certain embodiments may be substantially coaxial with annular engine inlet and/or substantially coaxial with longitudinal axis LA (e.g., center line) of gas turbine engine when air intake is installed on gas turbine engine. The engine inlet may comprise an annular opening into which the flow of air discharged substantially axially rearwardly from intake outlet. In some embodiments, the annular engine inlet may be coaxial with longitudinal axis LA of gas turbine engine. In some embodiments, the longitudinal axis LA of gas turbine engine may correspond to the axis or rotation of high-pressure spool <NUM> and of low-pressure spool <NUM> as shown in <FIG>. Accordingly, in embodiments where an axis of rotation of propeller <NUM> is radially offset from an axis of rotation of high-pressure spool <NUM> and low-pressure spool <NUM>, the longitudinal axis LA may not necessarily correspond to the axis of rotation of propeller <NUM>. In various embodiments, the intake axis IA may be either substantially coaxial the longitudinal axis LA, or may be offset therefrom (e.g. spaced apart therefrom in a radial or transverse direction from the longitudinal axis).

The annual quasi-scroll intake <NUM> that that feeds the engine air inlet <NUM>, as described above with reference to <FIG>, may be in fluid communication with intake duct <NUM> and receive the flow of air F from intake duct <NUM>. A scroll portion of the quasi-scroll intake <NUM> may channel the flow of air F toward engine inlet via the intake outlet. The scroll portion may define one or more converging quasi scroll-shaped passages formed such as to cause acceleration and redirection of the flow of air F toward engine inlet with relatively low energy losses and pressure distortion.

During operation, the present air intakes may be installed on gas turbine engine and used to channel a flow of air to engine inlet with relatively low energy losses and flow distortion (e.g., swirl and pressure distortions). Air intake may define a generally streamlined flow path between intake inlet and the intake outlet. For example, in some embodiments, air intake may not comprise a plenum (i.e., dump box) often found in traditional air intakes and which may cause significant energy losses. In various embodiments, improvements of flow characteristics of the flow or air F may improve engine performance in comparison with some other traditional air intakes.

In various embodiments, the air intakes as described herein may be fabricated according to known or other manufacturing methods using suitable sheet metal or polymeric material. In some embodiments, air intake or part(s) thereof may be cast using a suitable metallic material or molded from a suitable polymeric material. In some embodiments, air intake may comprise a plurality of components (e.g., pieces of sheet metal) pieced (e.g., welded) together to form air intake.

Referring now to <FIG>, in the unclaimed example shown, an accessory gearbox <NUM> is in driving engagement with the low-pressure spool <NUM> (<FIG>) of the engine <NUM>. The accessory gearbox (AGB) <NUM> defines at least one, three in the unclaimed example shown, outputs 11a for accessories (not shown) to be drivingly engaged to the low-pressure spool <NUM> via the accessory gearbox <NUM>. As shown, the outputs 11a face a direction having a radial component relative to the longitudinal axis LA of the engine core <NUM>. The accessories connected to the outputs 11a are herein located radially between the nacelle <NUM> and the outputs 11a. In other words, the disclosed pusher engine <NUM> allows the accessories to be located radially outwardly of the gearbox <NUM> and radially inwardly of the nacelle <NUM> relative to the axis LA. Consequently, an access door 16a of the nacelle <NUM> may allow easy access via an opening 16b defined through the nacelle <NUM> to replace, repair, and perform maintenance of the accessories without having to dismantle the pusher engine <NUM>.

Still referring to <FIG>, an unclaimed example, a mounting structure M is used to secure the engine core <NUM> to the wing <NUM>. The mounting structure includes a frame <NUM>. The frame <NUM> is herein secured to both of fore and aft spars 17a, 17b of the wing <NUM>. The fore and aft spars 17a, 17b are structural members that extend in a spanwise direction along a spanwise axis SA of the wing <NUM>, from a root of the wing <NUM> to a tip thereof. The fore spar 17a is located proximate a leading edge of the wing <NUM> whereas the aft spar 17b is located proximate a trailing edge of the wing <NUM>.

In the unclaimed example shown, the frame <NUM> includes a plurality of members 100a that are secured to one another. The members 100a are assembled to define an engine-receiving space S sized to contain the engine core <NUM> and the AGB <NUM>. The members 100a may be tubular members.

However, having a side mounted accessory gearbox <NUM>, with side mounted accessories, might complicate installation of the engine core <NUM> and AGB <NUM> on the wing <NUM> as a radial dimension of the engine core <NUM> with the AGB <NUM> might be too big to allow the pusher engine <NUM> to be inserted in the engine-receiving space S by moving said engine <NUM> along its axis LA. Some of the frame members 100a may prevent penetration of the engine core <NUM> and AGB <NUM> into the engine-receiving space S if the engine and AGB <NUM>, <NUM> are moved axially along the longitudinal axis LA.

Referring now to <FIG>, unclaimed examples, the engine core <NUM> is installed by moving said engine core in a vertical direction VD relative to a ground G. The vertical direction may correspond to a radial direction relative to the longitudinal axis LA of the engine core <NUM>. In the unclaimed example shown, the engine core <NUM> is inserted in the desired position by decreasing an elevation of the engine core <NUM> relative to the ground G. In other words, the engine core <NUM> is lowered into the desired position into the engine-receiving space S by bringing the engine core <NUM> closer to the ground G.

In the depicted unclaimed example, the frame <NUM> has an upper frame portion 100b and a lower frame portion 100c both including some of the tubular members 100a. As shown more clearly on <FIG>, the mounting structure M further includes a mount ring <NUM> for connecting the engine core <NUM> to the frame <NUM>. Each of the upper and lower frame portions 100b, 100c is secured to a respective one of upper and lower U-shaped mounts 102a, 102b of the mount ring <NUM>. The U-shaped mounts are also referred to as horseshoe-shaped mount. Both of the U-shaped mounts 102a, 102b have a U-shape circumferentially extending around the engine core <NUM> relative to the longitudinal axis LA. The lower frame portion 100c and the lower U-shaped mount 102b are secured to the wing <NUM> via the fore and aft spars 17a, 17b whereas the upper frame portion 100b and the upper U-shaped mount 102a are secured to the engine core <NUM> prior to moving the engine core <NUM> in the desired position. The lower U-shaped mount 102b defines an opening O oriented away from the ground G when the aircraft is on the ground G. The upper and lower frame portions 100b, 100c may be securable to one another and the upper and lower U-shaped mounts 102a, 102b may be securable to one another once the engine core <NUM> and AGB <NUM> are in the desired position, within the engine-receiving space S.

The engine core <NUM> includes a plurality of pads 13a that are circumferentially distributed around the longitudinal axis LA. The pads 13a are herein secured to a gas generator case 13b of the engine core <NUM>. The gas generator case 13b is a case that surrounds the combustor <NUM> of the engine core <NUM>. It will be appreciated that the pads 13a may be secured to the compressor case and/or to the turbine-exhaust case. The engine core <NUM> may be secured to the U-shaped mounts 102a, 102b and the frame <NUM> via any suitable case of the engine core <NUM>. Elastomeric material may be located between the upper and lower U-shaped mounts 102a, 102b and the pads 13a for dampening vibrations occurring with operation of the gas turbine engine <NUM>.

Referring to <FIG>, unclaimed examples, for mounting the gas turbine engine <NUM> to the wing <NUM> of the aircraft <NUM>, the mounting structure S is secured to the wing <NUM> of the aircraft <NUM>; the engine core <NUM> is inserted within the engine-receiving space S by changing the elevation of the engine core <NUM> relative to the ground G until the engine core <NUM> is at least partially enclosed by the lower U-shaped mount 102b; and the engine core <NUM> is secured to the lower U-shaped mount 102b.

In the present case, the lower frame portion 100c is secured to the wing <NUM>; the lower U-shaped mount 102b is secured to the lower frame portion 100c; the upper frame portion 100b and the upper U-shaped mount 102a are secured to the engine core <NUM>; and the engine core <NUM> is inserted into the engine-receiving space S by moving the engine core <NUM> in the vertical direction VD toward the ground G until the engine core <NUM> is received within the lower U-shaped mount 102b via its opening O facing away from the ground G. The upper and lower U-shaped mounts 102a, 102b may be secured to one another and the upper and lower frame portions 100c, 100d, may be secured to one another.

Referring to now to Fig. <NUM>, in an alternate unclaimed example, the mounting structure includes only one U-shaped mount <NUM> defining an opening O' facing toward the ground G. The engine core <NUM> is receivable within the engine-receiving space S in an upward direction away from the ground G.

As shown in Fig. <NUM>, an unclaimed example, for mounting the engine core <NUM> to the wing <NUM>, the engine core <NUM> is elevated away from the ground G until the engine core <NUM> is at least partially enclosed by the U-shaped mount <NUM>. At which point, the engine core <NUM> is secured to the U-shaped mount <NUM> via the pads 13a secured to the gas generator case 13b.

In the unclaimed example shown, the engine core <NUM> may be inserted in the engine-receiving space S without having to mount any frame portion thereto. In other words, a frame may include only a single portion that is secured to the wing <NUM> via the spars 17a, 17b and the horseshoe-shaped mount <NUM> is secured to the frame before installing the engine core <NUM>. After the engine core <NUM> is inserted in the engine-receiving space S, the engine core <NUM> is secured to the U-shaped mount <NUM> via the pads 13a.

Referring to <FIG>, unclaimed examples, after the engine core <NUM> is at the desired position, the air inlet <NUM> and other components such as the nacelle <NUM> may be installed. In the unclaimed example shown, the two separate ducts <NUM>', <NUM>" (<FIG>) are connected to the annular quasi-scroll <NUM> (<FIG>) at a connection point C such that the quasi-scroll <NUM> may be secured to the engine core <NUM> separately from the ducts <NUM>', <NUM>". In the present case, after the engine core <NUM> is installed in the desired position, the quasi-scroll <NUM> is installed and then the ducts <NUM>', <NUM>" are installed and connected to the quasi-scroll <NUM> at the connection point C.

Claim 1:
A turboprop gas turbine engine adapted to be mounted to an aircraft, the turboprop gas turbine engine comprising:
a nacelle (<NUM>);
a propeller (<NUM>);
an air-cooled-oil-cooler (ACOC) (<NUM>);
an engine core (<NUM>) and a gearbox (<NUM>) driving the propeller (<NUM>), the engine core (<NUM>) and the gearbox (<NUM>) being enclosed within the nacelle (<NUM>), the propeller (<NUM>) configured to be located rearward of the gearbox (<NUM>) and the engine core (<NUM>) relative to a direction of travel of the aircraft, the turboprop gas turbine engine being a pusher engine (<NUM>), the engine core (<NUM>) defining a central longitudinal axis (LA); and
an air intake (<NUM>) disposed within the nacelle (<NUM>) and formed to direct ambient air into the engine core (<NUM>) of the turboprop gas turbine engine, the air intake (<NUM>) including an air inlet duct (<NUM>) having a forward-facing intake inlet (<NUM>) receiving the ambient air, the air inlet duct (<NUM>) including an upstream section (<NUM>) and a downstream section (<NUM>), the upstream section (<NUM>) of the air inlet duct (<NUM>) in fluid communication with the intake inlet (<NUM>) and extending downstream from the intake inlet (<NUM>), the downstream section (<NUM>) of the air inlet duct (<NUM>) fluidly connected to and directing air from the upstream section (<NUM>) into an engine air inlet (<NUM>) of the engine core (<NUM>), wherein
the upstream section (<NUM>) of the air inlet duct (<NUM>) is centrally located within the nacelle (<NUM>) relative to the central longitudinal axis (LA), and the intake inlet (<NUM>) of the air intake (<NUM>) extends along an intake axis (IA) that is either substantially coaxial with the central longitudinal axis (LA) or radially spaced apart therefrom, and
a second air outlet duct (<NUM>) is located within the nacelle (<NUM>), the second air outlet duct (<NUM>) is connected to the air inlet duct (<NUM>) at a location between the upstream section (<NUM>) and the downstream section (<NUM>) thereof, the location being between the intake inlet (<NUM>) and the engine air inlet (<NUM>), the second air outlet duct (<NUM>) directing air from the upstream section (<NUM>) of the air inlet duct (<NUM>) into the ACOC (<NUM>),
characterised in that the ACOC (<NUM>) and the second air outlet duct (<NUM>) are located axially forward of the engine core (<NUM>).