Patent Publication Number: US-9896998-B2

Title: Compound engine assembly with modulated flow

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application claims priority from U.S. provisional application No. 62/118,891 filed Feb. 20, 2015, the entire contents of which are incorporated by reference herein. 
    
    
     TECHNICAL FIELD 
     The application relates generally to compound engine assemblies and, more particularly, to supercharged or turbocharged compound engine assemblies used in aircraft. 
     BACKGROUND OF THE ART 
     Compound engine assemblies including a compressor used as a supercharger or turbocharger may define a relatively bulky assembly which may be difficult to fit into existing aircraft nacelles, thus creating some difficulty in adapting them for aircraft applications. 
     SUMMARY 
     In one aspect, there is provided a compound engine assembly comprising: a compressor; an engine core including at least one internal combustion engine each having a rotor sealingly and rotationally received within a respective internal cavity to provide rotating chambers of variable volume in the respective internal cavity, the engine core having an inlet in fluid communication with an outlet of the compressor; a turbine section having an inlet in fluid communication with an outlet of the engine core, the turbine section configured to compound power with the engine core; and an air conduit having at least one heat exchanger extending thereacross such that an airflow through the air conduit circulates through the at least one heat exchanger, each of the at least one heat exchanger configured to circulate a fluid to be cooled in heat exchange relationship with the airflow circulating therethrough, the air conduit having opposed inlet and outlets in fluid communication with ambient air around the compound engine assembly; wherein an outer wall of the air conduit has a plurality of openings defined therethrough downstream of the at least one heat exchanger, each of the plurality of openings being selectively closable by a respective pivotable flap movable between a retracted position where the opening is obstructed and an extended position away from the opening, each of the plurality of openings defining a fluid communication between the air conduit and the ambient air around the compound engine assembly when the respective pivotable flap is in the extended position. 
     In another aspect, there is provided a compound engine assembly comprising: a compressor; an engine core including at least one rotary internal combustion engine in driving engagement with an engine shaft, the engine core having an inlet in fluid communication with an outlet of the compressor; a turbine section having an inlet in fluid communication with an outlet of the engine core, the turbine section including at least one turbine rotor engaged on a rotatable turbine shaft, the turbine shaft and the engine shaft being drivingly engaged to a same rotatable load; an air conduit having an inlet in fluid communication with ambient air around the compound engine assembly and an opposed outlet, the air conduit having at least one heat exchanger extending thereacross such that an airflow through the air conduit circulates through the at least one heat exchanger, each of the at least one heat exchanger configured to circulate a fluid to be cooled in heat exchange relationship with the airflow circulating therethrough, the air conduit having an outer wall including openings defined therethrough downstream of the at least one heat exchanger; and a plurality of pivotable flaps each attached to the outer wall adjacent a respective one of the openings and each movable between an extended position away from the opening and a retracted position closing the respective opening. 
     In a further aspect, there is provided a method of directing flow through a compound engine assembly, the method comprising: directing air into an inlet of a compressor of the compound engine assembly; directing compressed air from the compressor to an inlet of at least one rotary internal combustion engine of the compound engine assembly, the at least one rotary internal combustion engine driving rotation of an engine shaft; directing exhaust from the at least one rotary internal combustion engine on at least one turbine rotor of a turbine section of the compound engine assembly to drive rotation of a turbine shaft, the turbine shaft compounding power with the engine shaft to drive a rotatable load; and directing air through an air conduit containing at least one heat exchanger and through the at least one heat exchanger to cool a fluid of the compound engine assembly, including modulating a flow of the air through the at least one heat exchanger by selectively opening and closing a plurality of openings defined in an outer wall of the air conduit downstream of the at least one heat exchanger. 
    
    
     
       DESCRIPTION OF THE DRAWINGS 
       Reference is now made to the accompanying figures in which: 
         FIG. 1  is a schematic view of a compound engine assembly in accordance with a particular embodiment; 
         FIG. 2  is a cross-sectional view of a Wankel engine which can be used in a compound engine assembly such as shown in  FIG. 1 , in accordance with a particular embodiment; 
         FIG. 3  is a schematic tridimensional view of the compound engine assembly of  FIG. 1  in accordance with a particular embodiment; 
         FIG. 4  is a schematic cross-sectional view of a nacelle installation of the compound engine assembly of  FIG. 3 , in accordance with a particular embodiment; 
         FIG. 5  is a schematic tridimensional view of an intake assembly of the compound engine assembly of  FIG. 4 , in accordance with a particular embodiment; 
         FIG. 6A  is a schematic side cross-sectional view of part of the intake assembly of  FIG. 5 ; 
         FIG. 6B  is an enlarged schematic side cross-sectional view of part of the intake assembly of  FIG. 6A ; and 
         FIG. 7  is a schematic tridimensional view of an inlet lip of the intake assembly of  FIG. 5 , in accordance with a particular embodiment. 
     
    
    
     DETAILED DESCRIPTION 
     Described herein are a compound engine assembly  10  and its installation for a propeller airplane. In the embodiment shown, the compound engine assembly  10  includes a liquid cooled heavy fueled multi-rotor rotary engine core  12  and a turbine section  18  used as an exhaust energy recovery system. As will be detailed below, other configurations for the engine core  12  are also possible. 
     Referring to  FIG. 1 , the engine core  12  has an engine shaft  16  driven by the rotary engine(s) and driving a rotatable load, which is shown here as a propeller  8 . It is understood that the compound engine assembly  10  may alternately be configured to drive any other appropriate type of load, including, but not limited to, one or more generator(s), drive shaft(s), accessory(ies), rotor mast(s), compressor(s), or any other appropriate type of load or combination thereof. The compound engine assembly  10  further includes a compressor  14 , and a turbine section  18  compounding power with the engine core  12 . 
     The engine core  12  may include 2, 3, 4 or more rotary engines drivingly engaged to the shaft  16 . In another embodiment, the engine core  12  includes a single rotary engine. Each rotary engine has a rotor sealingly engaged in a respective housing, with each rotary engine having a near constant volume combustion phase for high cycle efficiency. The rotary engine(s) may be Wankel engine(s). Referring to  FIG. 2 , an exemplary embodiment of a Wankel engine is shown. Each Wankel engine comprises a housing  32  defining an internal cavity with a profile defining two lobes, which is preferably an epitrochoid. A rotor  34  is received within the internal cavity. The rotor defines three circumferentially-spaced apex portions  36 , and a generally triangular profile with outwardly arched sides. The apex portions  36  are in sealing engagement with the inner surface of a peripheral wall  38  of the housing  32  to form three working chambers  40  between the rotor  34  and the housing  32 . 
     The rotor  34  is engaged to an eccentric portion  42  of the shaft  16  to perform orbital revolutions within the internal cavity. The shaft  16  performs three rotations for each orbital revolution of the rotor  34 . The geometrical axis  44  of the rotor  34  is offset from and parallel to the axis  46  of the housing  32 . During each orbital revolution, each chamber  40  varies in volume and moves around the internal cavity to undergo the four phases of intake, compression, expansion and exhaust. 
     An intake port  48  is provided through the peripheral wall  38  for successively admitting compressed air into each working chamber  40 . An exhaust port  50  is also provided through the peripheral wall  38  for successively discharging the exhaust gases from each working chamber  40 . Passages  52  for a glow plug, spark plug or other ignition element, as well as for one or more fuel injectors (not shown) are also provided through the peripheral wall  38 . Alternately, the intake port  48 , the exhaust port  50  and/or the passages  52  may be provided through an end or side wall  54  of the housing; and/or, the ignition element and a pilot fuel injector may communicate with a pilot subchamber (not shown) defined in the housing  32  and communicating with the internal cavity for providing a pilot injection. The pilot subchamber may be for example defined in an insert (not shown) received in the peripheral wall  38 . 
     In a particular embodiment the fuel injectors are common rail fuel injectors, and communicate with a source of Heavy fuel (e.g. diesel, kerosene (jet fuel), equivalent biofuel), and deliver the heavy fuel into the engine(s) such that the combustion chamber is stratified with a rich fuel-air mixture near the ignition source and a leaner mixture elsewhere. 
     For efficient operation the working chambers  40  are sealed, for example by spring-loaded apex seals  56  extending from the rotor  34  to engage the peripheral wall  38 , and spring-loaded face or gas seals  58  and end or corner seals  60  extending from the rotor  34  to engage the end walls  54 . The rotor  34  also includes at least one spring-loaded oil seal ring  62  biased against the end wall  54  around the bearing for the rotor  34  on the shaft eccentric portion  42 . 
     Each Wankel engine provides an exhaust flow in the form of a relatively long exhaust pulse; for example, in a particular embodiment, each Wankel engine has one explosion per 360° of rotation of the shaft, with the exhaust port remaining open for about 270° of that rotation, thus providing for a pulse duty cycle of about 75%. By contrast, a piston of a reciprocating 4-stroke piston engine typically has one explosion per 720° of rotation of the shaft with the exhaust port remaining open for about 180° of that rotation, thus providing a pulse duty cycle of 25%. 
     In a particular embodiment which may be particularly but not exclusively suitable for low altitude, each Wankel engine has a volumetric expansion ratio of from 5 to 9, and operates following the Miller cycle, with a volumetric compression ratio lower than the volumetric expansion ratio, for example by having the intake port located closer to the top dead center (TDC) than an engine where the volumetric compression and expansion ratios are equal or similar. Alternately, each Wankel engine operates with similar or equal volumetric compression and expansion ratios. 
     It is understood that other configurations are possible for the engine core  12 . The configuration of the engine(s) of the engine core  12 , e.g. placement of ports, number and placement of seals, etc., may vary from that of the embodiment shown. In addition, it is understood that each engine of the engine core  12  may be any other type of internal combustion engine including, but not limited to, any other type of rotary engine, and any other type of non-rotary internal combustion engine such as a reciprocating engine. 
     Referring back to  FIG. 1 , the rotary engine core  12  is supercharged with the compressor  14  mounted in-line with the engine core, i.e. the compressor rotor(s)  14   a  rotate co-axially with the engine shaft  16 . In the embodiment shown, the compressor rotor(s)  14   a  are engaged on a compressor shaft  15 , and the engine shaft  16  is in driving engagement with the compressor shaft  15  through a step-up gearbox  20 . In a particular embodiment, the gearbox  20  is a planetary gear system. In a particular embodiment, the compressor shaft  15  includes a sun gear  20   s  which is drivingly engaged to carrier-mounted planet gears  20   p , which are drivingly engaged to a fixed ring gear  20   r . The rotating carrier assembly is connected to the engine shaft  16 , for example through a splined connection. In a particular embodiment, the planetary gear system elements (sun gear, planet gears and ring gear) within the gearbox  20  are configured to define a speed ratio of about 7:1 between the compressor shaft  15  and engine core shaft  16 . It is understood that any other appropriate configuration and/or speed ratio for the gearbox  20  may alternately be used. 
     In the embodiment shown and referring particularly to  FIG. 1 , the compressor  14  is a centrifugal compressor with a single rotor  14   a . Other configurations are alternately possible. The compressor  14  may be single-stage device or a multiple-stage device and may include one or more rotors having radial, axial or mixed flow blades. 
     The outlet of the compressor  14  is in fluid communication with the inlet of the engine core  12 , which corresponds to or communicates with the inlet of each engine of the engine core  12 . Accordingly, air enters the compressor  14  and is compressed and circulated to the inlet of the engine core  12 . In a particular embodiment, the compressor  14  includes variable inlet guide vanes  22  through which the air circulates before reaching the compressor rotor(s)  14   a.    
     The engine core  12  receives the pressurized air from the compressor  14  and burns fuel at high pressure to provide energy. Mechanical power produced by the engine core  12  drives the propeller  8 . 
     Each engine of the engine core  12  provides an exhaust flow in the form of exhaust pulses of high pressure hot gas exiting at high peak velocity. The outlet of the engine core  12  (i.e. the outlet of each engine of the engine core  12 ) is in fluid communication with the inlet of the turbine section  18 , and accordingly the exhaust flow from the engine core  12  is supplied to the turbine section  18 . 
     The turbine section  18  includes at least one rotor engaged on a turbine shaft  19 . Mechanical energy recovered by the turbine section  18  is compounded with that of the engine shaft  16  to drive the propeller  8 . The turbine shaft  19  is mechanically linked to, and in driving engagement with, the engine shaft  16  through a reduction gearbox  24 , for example through an offset gear train with idler gear. In a particular embodiment, the elements of the reduction gearbox  24  (e.g. offset gear train) are configured to define a reduction ratio of approximately 5:1 between the turbine shaft  19  and the engine shaft  16 . The engine shaft  16  is also mechanically linked to, and in driving engagement with, the propeller  8  through the same reduction gearbox  24 . In a particular embodiment, the reduction gearbox  24  includes two gear train branches: a compounding branch  24   c  mechanically linking the turbine shaft  19  and the engine shaft  16  and a downstream planetary branch  24   p  mechanically linking the engine shaft  16  and propeller  8 . In another embodiment, the turbine shaft  19  and engine shaft  16  may be engaged to the propeller  8  through different gearboxes, or the turbine shaft  19  may be engaged to the engine shaft  16  separately from the engagement between the engine shaft  16  and the propeller  8 . In particular embodiment, the turbine shaft  19  is engaged to the compressor gearbox  20 . 
     As can be seen in  FIGS. 1 and 3 , the turbine shaft  19  is parallel to and radially offset from (i.e., non-coaxial to) the engine shaft  16  and compressor shaft  15 . The compressor rotor(s)  14   a  and engine shaft  16  are thus rotatable about a common axis (central axis of the compressor and engine shafts  15 ,  16 ) which is parallel to and radially offset from the axis of rotation of the turbine rotor(s)  26   a ,  28   a  (central axis of the turbine shaft  19 ). In a particular embodiment, the offset configuration of the turbine section  18  allows for the turbine section  18  to be enclosed in a casing separate from that of the engine core  12  and the compressor  14 , such that the turbine section  18  is modular and removable (e.g. removable on-wing) from the remainder of the compound engine assembly  10 . 
     Referring particularly to  FIG. 1 , the turbine section  18  may include one or more turbine stages. In a particular embodiment, the turbine section  18  includes a first stage turbine  26  receiving the exhaust from the engine core  12 , and a second stage turbine  28  receiving the exhaust from the first stage turbine  26 . The first stage turbine  26  is configured as a velocity turbine, also known as an impulse turbine, and recovers the kinetic energy of the core exhaust gas while creating minimal or no back pressure to the exhaust of the engine core  12 . The second stage turbine  28  is configured as a pressure turbine, also known as a reaction turbine, and completes the recovery of available mechanical energy from the exhaust gas. Each turbine  26 ,  28  may be a centrifugal or axial device with one or more rotors having radial, axial or mixed flow blades. In another embodiment, the turbine section  18  may include a single turbine, configured either as an impulse turbine or as a pressure turbine. 
     A pure impulse turbine works by changing the direction of the flow without accelerating the flow inside the rotor; the fluid is deflected without a significant pressure drop across the rotor blades. The blades of the pure impulse turbine are designed such that in a transverse plane perpendicular to the direction of flow, the area defined between the blades is the same at the leading edges of the blades and at the trailing edges of the blade: the flow area of the turbine is constant, and the blades are usually symmetrical about the plane of the rotating disc. The work of the pure impulse turbine is due only to the change of direction in the flow through the turbine blades. Typical pure impulse turbines include steam and hydraulic turbines. 
     In contrast, a reaction turbine accelerates the flow inside the rotor but needs a static pressure drop across the rotor to enable this flow acceleration. The blades of the reaction turbine are designed such that in a transverse plane perpendicular to the direction of flow, the area defined between the blades is larger at the leading edges of the blades than at the trailing edges of the blade: the flow area of the turbine reduces along the direction of flow, and the blades are usually not symmetrical about the plane of the rotating disc. The work of the pure reaction turbine is due mostly to the acceleration of the flow through the turbine blades. 
     Most aeronautical turbines are not “pure impulse” or “pure reaction”, but rather operate following a mix of these two opposite but complementary principles—i.e. there is a pressure drop across the blades, there is some reduction of flow area of the turbine blades along the direction of flow, and the speed of rotation of the turbine is due to both the acceleration and the change of direction of the flow. The degree of reaction of a turbine can be determined using the temperature-based reaction ratio (equation 1) or the pressure-based reaction ratio (equation 2), which are typically close to one another in value for a same turbine: 
                     Reaction   ⁡     (   T   )       =       (       t   S3     -     t   S5       )       (       t   S0     -     t   S5       )               (   1   )                 Reaction   ⁡     (   P   )       =       (       P   S3     -     P   S5       )       (       P   S0     -     P   S5       )               (   2   )               
where T is temperature and P is pressure, s refers to a static port, and the numbers refers to the location the temperature or pressure is measured: 0 for the inlet of the turbine vane (stator), 3 for the inlet of the turbine blade (rotor) and 5 for the exit of the turbine blade (rotor); and where a pure impulse turbine would have a ratio of 0 (0%) and a pure reaction turbine would have a ratio of 1 (100%).
 
     In a particular embodiment, the first stage turbine  26  is configured to take benefit of the kinetic energy of the pulsating flow exiting the engine core  12  while stabilizing the flow and the second stage turbine  28  is configured to extract energy from the remaining pressure in the flow while expanding the flow. Accordingly, the first stage turbine  26  has a smaller reaction ratio than that of the second stage turbine  28 . 
     In a particular embodiment, the second stage turbine  28  has a reaction ratio higher than 0.25; in another particular embodiment, the second stage turbine  28  has a reaction ratio higher than 0.3; in another particular embodiment, the second stage turbine  28  has a reaction ratio of about 0.5; in another particular embodiment, the second stage turbine  28  has a reaction ratio higher than 0.5. 
     In a particular embodiment, the first stage turbine  26  has a reaction ratio of at most 0.2; in another particular embodiment, the first stage turbine  26  has a reaction ratio of at most 0.15; in another particular embodiment, the first stage turbine  26  has a reaction ratio of at most 0.1; in another particular embodiment, the first stage turbine  26  has a reaction ratio of at most 0.05. 
     It is understood that any appropriate reaction ratio for the second stage turbine  28  (included, but not limited to, any of the above-mentioned reaction ratios) can be combined with any appropriate reaction ratio for the first stage turbine  26  (included, but not limited to, any of the above-mentioned reaction ratios), and that these values can correspond to pressure-based or temperature-based ratios. Other values are also possible. For example, in a particular embodiment, the two turbines  26 ,  28  may have a same or similar reaction ratio; in another embodiment, the first stage turbine  26  has a higher reaction ratio than that of the second stage turbine  28 . Both turbines  26 ,  28  may be configured as impulse turbines, or both turbines  26 ,  28  may be configured as pressure turbines. 
     In an embodiment where the engine core  12  includes one or more rotary engine(s) each operating with the Miller cycle, the compressor pressure ratio and the turbine section pressure ratio may be higher than a similar engine assembly where the engine core includes one or more rotary engine(s) having similar or equal volumetric compression and expansion ratios. The higher pressure ratio in the turbine section may be accommodated by additional axial turbine stage(s), an additional radial turbine, and/or a combination of axial and radial turbines suitable to accept the higher pressure ratio. 
     Referring to  FIG. 4 , a nacelle installation of the compound engine assembly  10  according to a particular embodiment is shown. The installation includes an intake assembly  66  which features a common inlet  68  and air conduit  70  for the engine assembly (through the compressor  14 ) and the oil and coolant heat exchangers  72 ,  74 . The air conduit  70  extends from the inlet  68  to an opposed outlet  76 . The inlet  68  and outlet  76  of the air conduit  70  communicate with ambient air outside of or around the assembly  10 , for example ambient air outside of a nacelle receiving the assembly. In the embodiment shown, the ambient air penetrates the compound engine assembly  10  through the inlet  68  of the air conduit  70 —the inlet  68  of the air conduit  70  thus defines a nacelle inlet, i.e. an inlet of the assembly  10  as a whole. 
     It can be seen that the heat exchangers  72 ,  74  extend across the air conduit  70 , such that the airflow through the air conduit  70  circulates through the heat exchangers  72 ,  74 . In the embodiment shown, the heat exchangers  72 ,  74  include an oil heat exchanger  72  which receives the oil from the engine assembly oil system and circulates it in heat exchange relationship with the airflow, such as to cool the oil; and a coolant heat exchanger  74  which receives the coolant from the engine core  12  (e.g. water, oil or other liquid coolant) and circulates it in heat exchange relationship with the airflow, such as to cool the coolant. Although two heat exchangers  72 ,  74  are shown, it is understood that alternately a single heat exchanger or more than two heat exchangers may be provided in the air conduit  70 . The two heat exchangers  72 ,  74  are shown as being placed in parallel, such that a portion of the airflow separately circulates through each heat exchanger. Alternately, the heat exchangers  72 ,  74  may be placed in the air conduit  70  in series such that the same portion of the airflow circulates through one than through the other of the heat exchangers, although such a configuration may necessitate the use of larger heat exchangers. It is also understood that the angle of the heat exchangers  72 ,  74  within the conduit  70  may be different from that shown. In a particular embodiment, the angle of the heat exchangers  72 ,  74  with respect to the airflow within the conduit  70  is selected to obtain a desired balance between pressure losses and effectiveness of the heat exchangers, in consideration of the available space within the conduit  70 . 
     The intake assembly  66  includes an intake plenum  78  configured for connection to and fluid communication with the inlet of the compressor  14 . In the embodiment shown and as can be more clearly seen in  FIG. 5 , the intake plenum  78  is annular. Other configurations are possible. 
     Referring to  FIGS. 4, 5 and 6A , the intake assembly  66  includes first and second intake conduits  80 ,  82  providing fluid communication between the air conduit  70  and the intake plenum  78 . The first intake conduit  80  is connected to the air conduit  70  upstream of the heat exchangers  72 ,  74 , so that the portion of the air conduit  70  upstream of the heat exchangers  72 ,  74  defines a first source of air. The second intake conduit  82  is connected to the air conduit  70  downstream of the heat exchangers  72 ,  74 , so that the portion of the air conduit  70  downstream of the heat exchangers  72 ,  74  defines a second source of air warmer than the first source. In the embodiment shown and as can be more clearly seen in  FIG. 4 , the air conduit  70  is configured to define a diffuser upstream of the heat exchangers  72 ,  74 , such as to decelerate the flow to a low velocity flow at the inlet of the heat exchangers  72 ,  74 . The first intake conduit  80  is connected in the diffuser; in a particular embodiment, the first intake conduit  80  is connected to the air conduit  70  where air velocity is at a minimum. Such a configuration may allow for minimizing of pressure losses. 
     Referring to  FIGS. 6A-6B , in a particular embodiment, the intake conduits  80 ,  82  are in fluid communication with the intake plenum  78  through an engine intake  84  containing an air filter  86 . An air filter bypass valve  88  is provided in the engine intake  84  to allow airflow to the intake plenum  78  around the air filter  86  in case of inadvertent air filter blockage. In a particular embodiment, the air filter bypass valve  86  is a spring loaded pressure differential operated valve. 
     The intake assembly  66  further includes a selector valve  90  positioned upstream of the air filter  86  and allowing for the selection of the intake conduit  80 ,  82  used to circulate the air from the air conduit  70  to the intake plenum  78 . The selector valve  90  is thus configurable between a configuration where the fluid communication between the intake plenum  78  and the air conduit  70  through the first intake conduit  80  is allowed and a configuration where the fluid communication between the intake plenum  78  and the air conduit  70  through the first intake conduit  80  is prevented. 
     In the particular embodiment shown in  FIG. 4 , the selector valve  90  only acts to selectively block or prevent the communication through the first intake conduit  80 , i.e. the intake conduit connected to the air conduit  70  upstream of the heat exchangers  72 ,  74 . The communication through the second intake conduit  82  remains open in both configurations. 
     In the particular embodiment shown in  FIGS. 6A and 6B , the selector valve  90  is provided at a junction between the two intake conduits  80 ,  82 , and acts to selectively block or prevent the communication through both intake conduits  80 ,  82 . Accordingly, in the configuration shown in  FIG. 6A , the selector valve  90  allows the fluid communication between the intake plenum  78  and the air conduit  70  through the first intake conduit  80  while preventing the fluid communication between the intake plenum  78  and the air conduit  70  through the second intake conduit  82 ; and in the configuration shown in  FIG. 6B , the selector valve  90  prevents the fluid communication between the intake plenum  78  and the air conduit  70  through the first intake conduit  80  while allowing the fluid communication between the intake plenum  78  and the air conduit  70  through the second intake conduit  82 . In the embodiments shown, the selector valve  90  includes a flap pivotable between the two configurations, and blocks the communication through one or the other of the intake conduits  80 ,  82  by blocking the communication between that intake conduit  80 ,  82  and the intake plenum  78 . Other types of valves  90  and/or valve positions are also possible. 
     The selector valve  90  thus allows for the selection of cooler air (first intake conduit  80 , taking air upstream of the heat exchangers  72 ,  74 ) or warmer air (second intake conduit  82 , taking air downstream of the heat exchangers  72 ,  74 ) to feed the compressor  14  and engine assembly  10 , based on the operating conditions of the engine assembly  10 . For example, in icing conditions, the fluid communication through the second conduit  82  may be selected by blocking the fluid communication through the first conduit  80 , so that that the warmer air from downstream of the heat exchangers  72 ,  74  is used to feed the compressor  14 , such as to provide de-icing capability for the engine intake  84 , air filter  86 , intake plenum  78  and compressor inlet with fixed and variable geometries; and in non-icing flight conditions, the fluid communication through the first conduit  80  may be selected so that colder air is used to feed the compressor  14  to provide for better engine performance (as compared to hotter air). 
     Also, selection of the flow through the second intake conduit  82  to extract the engine air downstream of the heat exchangers  72 ,  74  can be used to generate airflow through the heat exchangers  72 ,  74 . For example, for a turboprop engine at ground idle, there is no inlet ram pressure to force air through the air conduit  70  and heat exchangers  72 ,  74 , and the propeller pressure rise may not be sufficient to draw enough air to provide sufficient cooling in the heat exchangers  72 ,  74 ; similar conditions may occur at taxi operations on the ground (engine at low power). Extracting the engine air downstream of the heat exchangers  72 ,  74  produces a “sucking” effect pulling the air through the heat exchangers  72 ,  74 , which in a particular embodiment may allow for sufficient cooling without the need of a fan or blower to provide for the necessary air circulation. A bleed-off Valve  75  can optionally be provided downstream of the compressor  14  and upstream of the engine core  12  (i.e. in the fluid communication between the compressor outlet and the engine core inlet), and opened during idle or taxi operation to increase compressor flow such as to increase the “sucking” effect of extracting the engine air downstream of the heat exchangers  72 ,  74 , and accordingly increase the airflow through the heat exchangers  72 ,  74 . Moreover, an intercooler may optionally be provided just upstream of the engine core  12  to cool the compressor flow prior to routing it to the engine core. 
     In a particular embodiment, the engine intake assembly  66  can be configured as an inertial particle separator when the fluid communication through the first conduit  80  is selected, so that when the air from upstream of the heat exchangers  72 ,  74  is used to feed the engine, the heavy particles are entrained downstream of the heat exchangers  72 ,  74 . In the embodiment shown in  FIG. 4 , the junction between the first conduit  80  and the air conduit  70  is configured as the inertial particle separator: the first conduit  80  defines a sharp turn with respect to the air conduit  70  (e.g. by extending close to or approximately perpendicular thereto), extending at a sufficient angle from the air conduit  70  such that the heavier particles (e.g. ice, sand) continue on a straight path while the air follows the sharp turn, and by the first conduit  80  and air conduit  70  are sized to achieve adequate air velocities to ensure separation of the particles. 
     In the embodiment shown, the air conduit  70  is configured such that all of the air entering the air conduit  70  is circulated through the heat exchangers  72 ,  74  and/or to the intake plenum  78 . Alternately, a bypass conduit could be provided such that a portion of the air entering the conduit  70  is diverted from (i.e. bypasses) the heat exchangers  72 ,  74  and the intake plenum  78  and is instead directly circulated to the outlet  76 . In a particular embodiment, the junction between the bypass conduit and the air conduit  70  is configured as the inertial particle separator, through selection of an appropriate orientation and relative sizing of the bypass conduit with respect to the air conduit  70 . 
     In a particular embodiment and as shown in  FIG. 7 , the lip of the assembly inlet  68  is de-iced by circulating hot coolant through a coil tube  98  disposed in the lip and made of material having appropriate heat conduction properties. The coil tube  98  has an inlet in fluid communication with the coolant system of the engine core  12  and an outlet in fluid communication with the coolant heat exchanger  74 , such that a fraction of the hot coolant flowing out of the engine core  12  is routed to the coil tube  98  of the inlet lip  68  for de-icing, and then rejoins the remainder of the hot coolant flow from the engine core  12  prior to sending the flow to the heat exchanger  74 . 
     Although in the embodiment shown the heat exchangers  72 ,  74  and engine assembly  10  have a common inlet  68  and the first and second intake conduits  80 ,  82  communicate with a same air conduit  70  extending from that inlet, it is understood that alternately the engine assembly  10  and heat exchangers  72 ,  74  may have separate inlets. The first intake conduit  80  may thus communicate with a source of fresh air separate from that feeding the heat exchangers  72 ,  74 . 
     Alternately, the common inlet  68  and air conduit  70  used to feed the heat exchangers  72 ,  74  and the compressor  14  may be used with a single intake conduit providing the fluid communication between the intake plenum  78  and the air conduit  70 , and connected to the air conduit  70  at any appropriate location (downstream or upstream of the heat exchangers). 
     Referring back to  FIG. 4 , in a particular embodiment, variable cowl flaps  92  are pivotally connected to an outer wall  94  of the air conduit  70  downstream of the heat exchangers  72 ,  74 , each adjacent a respective opening  96  defined through the outer wall  94 . The flaps  92  are movable between an extended position (shown) where they extend away from the respective opening  96  and a retracted position where they close the respective opening  96 , such as to modulate the airflow through the air conduit  70  and heat exchangers  72 ,  74 . The openings  96  communicate with ambient air outside of or around the assembly  10  when the flaps are extended, for example ambient air outside of a nacelle receiving the assembly, such that air from the air conduit  70  may exit the conduit through the openings  96 . In a particular embodiment, the cowl flaps  92  are positioned in accordance with the power demand on the engine assembly  10 , such as to regulate the temperature of the oil and coolant being cooled in the heat exchangers  72 ,  74  while reducing or minimizing cooling drag; for example, the cowl flaps  92  are open at take-off and closed at cruise speed. 
     The cowl flaps  92  may have any appropriate configuration. For example, in a particular embodiment, the cowl flaps  92  have a straight airfoil shape; in another embodiment, the cowl flaps  92  have a cambered airfoil shape, configured to flow the exit air horizontally to produce a more effective thrust. In a particular embodiment, the cowl flaps  92  are configured as louvers, each connected to a rod, and an actuator slides the rod to pivot the cowl flaps  92  between the extended and retracted positions to open or close the louvers. Other configurations are also possible. 
     In a particular embodiment, the air conduit outlet  76  downstream of the cowl flaps  92  is shaped to define a nozzle, to form an exit jet opening. In a particular embodiment, the configuration of the nozzle is optimized to minimize the drag induced by the heat exchangers  72 ,  74  at the cruise speed operating conditions. 
     Although any of the above described and shown features and any combination thereof may provide for a suitable configuration to be used as a turboprop engine and/or be received in an aircraft nacelle, in a particular embodiment, the combination of all of the above described and shown features of the compound engine assembly provide for an engine configuration specifically tailored for use as an aircraft turboprop engine. 
     The above description is meant to be exemplary only, and one skilled in the art will recognize that changes may be made to the embodiments described without departing from the scope of the invention disclosed. For example, although the engine assembly has been described as a compound engine assembly, it is understood that elements of the compound engine assembly can be used with non-compounded engine assemblies, and with compound engine assemblies having different configurations, for example engine assemblies where the compressor is in driving engagement with the turbine section without being directly engaged to the engine core; such elements include, but are not limited to, the intake assembly and its components. Still other modifications which fall within the scope of the present invention will be apparent to those skilled in the art, in light of a review of this disclosure, and such modifications are intended to fall within the appended claims.