Patent Publication Number: US-2019186352-A1

Title: Auxiliary power unit with combined cooling of generator

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application is a divisional of U.S. application Ser. No. 15/227,483 filed on Aug. 3, 2016, which claims priority from U.S. application No. 62/202,275 filed Aug. 7, 2015, the entire contents of which are incorporated by reference herein. 
    
    
     TECHNICAL FIELD 
     The application relates generally to compound engine assemblies, more particularly to such assemblies used as auxiliary power units (APU). 
     BACKGROUND OF THE ART 
     Traditional gas turbine engine auxiliary power units including an engine core with a combustor which are used to drive a generator typically require a cooling system for the generator. Such a cooling system may include fans and/or ejectors can represent significant power losses and/or create drag penalties in flight. 
     Moreover, such traditional gas turbine engine auxiliary power units usually have an exhaust with relatively high temperature, requiring the use of high temperature materials in the exhaust duct walls, which may represent a significant cost. 
     SUMMARY 
     In one aspect, there is provided an auxiliary power unit for an aircraft, the auxiliary power unit comprising: an internal combustion engine having a liquid coolant system distinct from any fuel and lubricating system of the auxiliary power unit; a generator drivingly engaged to the internal combustion engine, the generator having a liquid coolant system distinct from the liquid coolant system of the internal combustion engine; a first heat exchanger having first coolant passages in fluid communication with the liquid coolant system of the internal combustion engine and first air passages in heat exchange relationship with the first coolant passages; a second heat exchanger having second coolant passages in fluid communication with the liquid coolant system of the generator and second air passages in heat exchange relationship with the second coolant passages; an exhaust duct in fluid communication with the first and second air passages; and a fan received in the exhaust duct and rotatable by the internal combustion engine for driving a cooling air flow through the first and second air passages. 
     In another aspect, there is provided an auxiliary power unit for an aircraft, the auxiliary power unit comprising: an internal combustion engine having a liquid coolant system; a compressor having an outlet in fluid communication with an inlet of the internal combustion engine; a turbine section having an inlet in fluid communication with an outlet of the internal combustion engine, the turbine section including at least one turbine compounded with the internal combustion engine; a generator drivable by the internal combustion engine and having a liquid coolant system distinct from the liquid coolant system of the internal combustion engine; a first heat exchanger in fluid communication with the liquid coolant system of the internal combustion engine; a second heat exchanger in fluid communication with the liquid coolant system of the generator; and a fan rotatable by the internal combustion engine for driving a cooling air flow through the first and second heat exchangers. 
     In a further aspect, there is provided a method of cooling a generator and an internal combustion engine of an auxiliary power unit for an aircraft, the method comprising: circulating a first liquid coolant through the internal combustion engine; circulating a second liquid coolant through the generator; and driving a cooling air flow in heat exchange relationship with the first and second liquid coolants using a fan driven by the internal combustion engine. 
    
    
     
       DESCRIPTION OF THE DRAWINGS 
       Reference is now made to the accompanying figures in which: 
         FIG. 1  is a schematic cross-sectional side view of an auxiliary power unit in accordance with a particular embodiment; 
         FIG. 2  is a schematic cross-sectional plan view of the auxiliary power unit of  FIG. 1 ; 
         FIG. 3  is a schematic tridimensional view of the auxiliary power unit of  FIG. 1 ; 
         FIG. 4  is a schematic cross-sectional view of a rotary engine which may be used in the auxiliary power unit of  FIGS. 1-3 ; 
         FIG. 5  is a schematic tridimensional view of an auxiliary power unit in accordance with another particular embodiment; 
         FIG. 6  is another schematic tridimensional view of the auxiliary power unit of  FIG. 5 , taken from an opposite side; 
         FIG. 7  is a schematic cross-sectional view of part of the auxiliary power unit of  FIG. 5 ; 
         FIG. 8  is a schematic tridimensional view, partly in transparency, of an end of the auxiliary power unit of  FIG. 5  received in a tail cone of an aircraft; 
         FIG. 9  is a schematic bottom view of an auxiliary power unit and tail cone in accordance with a particular embodiment, with part of the tail cone removed for clarity; 
         FIG. 10  is a schematic side view of the auxiliary power unit and tail cone of  FIG. 9 , with part of the tail cone removed for clarity; 
         FIG. 11  is a schematic cross-sectional view of compressor and turbine sections of the auxiliary power units of  FIG. 5  and of  FIG. 9 ; 
         FIG. 12  is a schematic cross-sectional view of part of an auxiliary power unit showing a cooling inlet and heat exchanger configuration in accordance with another particular embodiment which may be alternately used in any of the above auxiliary power units; 
         FIG. 13  is a schematic cross-sectional view of part of an auxiliary power unit showing a cooling inlet and heat exchanger configuration in accordance with another particular embodiment which may be alternately used in any of the above auxiliary power units; 
         FIG. 14  is a schematic cross-sectional view of a compressor section in accordance with another particular embodiment which may be alternately used in any of the above auxiliary power units; 
         FIG. 15  is a diagram of compressor and turbine configuration in accordance with another particular embodiment which may be alternately used in any of the above auxiliary power units; and 
         FIG. 16  is a schematic cross-sectional view of the compressor and turbine configuration of  FIG. 15 . 
     
    
    
     DETAILED DESCRIPTION 
     The present description includes compound engine assembly auxiliary power units for providing supplementary ground and flight pneumatic and/or electric power for airborne auxiliary power unit applications. In a particular embodiment, the auxiliary power units are configured to directly replace a traditional gas turbine engine auxiliary power unit and perform in a more efficient manner, with power/weight and power/volume properties meeting the requirements for airborne application. Application to fixed or mobile ground power units is also possible. 
     Referring to  FIGS. 1-3 , an auxiliary power unit  10  in accordance with a particular embodiment is generally shown. The auxiliary power unit  10  includes an engine core  12 ′ including one or more intermittent internal combustion engines  12  engaged to a common shaft  16  (see  FIG. 2 ). In a particular embodiment, the intermittent internal combustion engine(s)  12  is/are rotary internal combustion engine(s), for example Wankel engine(s); it is however understood that other types of intermittent internal combustion engines may alternately be used. 
     Referring to  FIG. 4 , an example of a Wankel engine which may be used in the engine core  12 ′ is shown. It is understood that the configuration of the engine(s)  12 , e.g. placement of ports, number and placement of seals, etc., may vary from that of the embodiment shown. The engine  12  comprises a housing  32  defining a rotor cavity having a profile defining two lobes, which is preferably an epitrochoid. A rotor  34  is received within the rotor 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 and separate three working chambers  40  of variable volume between the rotor  34  and the housing  32 . The peripheral wall  38  extends between two axially spaced apart end walls  54  to enclose the rotor cavity. 
     The rotor  34  is engaged to an eccentric portion  42  of an output shaft  16  to perform orbital revolutions within the rotor cavity. The output 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 rotor cavity to undergo the four phases of intake, compression, expansion and exhaust. 
     An intake port  48  is provided through the peripheral wall  38  for admitting compressed air into one of the working chambers  40 . An exhaust port  50  is also provided through the peripheral wall  38  for discharge of the exhaust gases from the working chambers  40 . Passages  52  for a spark plug, glow plug or other ignition mechanism, as well as for one or more fuel injectors of a fuel injection system (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 the end or side wall  54  of the housing. A subchamber (not shown) may be provided in communication with the chambers  40 , for pilot or pre injection of fuel for combustion. 
     For efficient operation the working chambers  40  are sealed by spring-loaded peripheral or apex seals  56  extending from the rotor  34  to engage the inner surface of 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 inner surface of the end walls  54 . The rotor  34  also includes at least one spring-loaded oil seal ring  62  biased against the inner surface of the end wall  54  around the bearing for the rotor  34  on the shaft eccentric portion  42 . 
     The fuel injector(s) of the engine  12 , which in a particular embodiment are common rail fuel injectors, communicate with a source of Heavy fuel (e.g. diesel, kerosene (jet fuel), equivalent biofuel), and deliver the heavy fuel into the engine  12  such that the combustion chamber is stratified with a rich fuel-air mixture near the ignition source and a leaner mixture elsewhere. 
     Referring back to  FIGS. 1-3 , the auxiliary power unit  10  includes a supercharger compressor  20  having an outlet in fluid communication with the inlet of the engine core  12 ′ (e.g. intake port  48  of each engine  12 ). Air enters an inlet plenum  19  from the aircraft inlet  14 , and the air is compressed by the compressor  20  which optionally includes variable inlet guide vanes  23  and optionally includes a variable diffuser  25  ( FIG. 2 ), which in a particular embodiment allow for management of a wide range of flow and pressure ratio conditions. The air from the compressor  20  circulates through an intercooler heat exchanger  18  to drop its temperature, for example from about 450° F. to 250° F., prior to entering the engine core. In the embodiment shown, the compressor  20  also provides bleed air for the aircraft; after leaving the compressor  20  and before reaching the intercooler  18 , a portion of the compressed air is directed to a bleed duct  27  to be delivered to the aircraft. 
     At certain operating conditions it may be necessary to bleed excess air from the compressor  20  to avoid surge. In the embodiment shown, the conduit between the compressor  20  and the intercooler  18  is in fluid communication with an excess air duct  29  to bleed this excess air; a diverter valve  31  is incorporated in the excess air duct  29  to manage the flow of air being bled from the compressor  20 . The diverter valve  31  may be scheduled to open based on sensed compressor exit conditions indicating operation close to surge. 
     In the engine core  12 ′ air is mixed with fuel and combusted to provide power and a residual quantity of intermediate pressure exhaust gas. The outlet of the engine core  12 ′ (e.g. exhaust port  50  of each engine  12 ) is in fluid communication with an inlet of a turbine section, so that the exhaust gases from the engine core  12 ′ are expanded in the turbine section. The turbine section has one or more turbines  26 ,  22  compounded with the engine core  12 ′. In a particular embodiment, the turbine section includes a first stage turbine  26  having an outlet in fluid communication with an inlet of a second stage turbine  22 , with the turbines  26 ,  22  having different reaction ratios from one another. 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, and which characterize the turbine with respect to “pure impulse” or “pure reaction” turbines: 
     
       
         
           
             
               
                 
                   
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     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 core engine(s)  12  while stabilizing the flow and the second stage turbine  22  is configured to extract energy from the remaining pressure in the flow. Accordingly, in a particular embodiment the first stage turbine  26  has a lower reaction ratio (i.e. lower value) than that of the second stage turbine  22 . In a particular embodiment, the first stage turbine  26  has a reaction ratio of 0.25 or lower (temperature or pressure based) or of 0.2 or lower (temperature or pressure based), and the second stage turbine  22  a reaction ratio higher than 0.25 (temperature or pressure based) and/or is a medium reaction pressure turbine. Other values are also possible. 
     The compressor  20  may be driven by one or more of the turbines  26 ,  22  and/or the engine core  12 ; in the embodiment shown and as can be best seen in  FIG. 2 , the first and second stage turbines  26 ,  22  and compressor  20  are coupled to the same shaft  24 . In a particular embodiment, the turbines  26 ,  22  and compressor  20  coupled on the same shaft  24  allow for a reasonably efficient non dimensional specific speed match between the compressor and turbine section. In a particular embodiment the turbine shaft  24  rotates at approximately 40,000 to 50,000 rpm; other values for the rotational speeds are also possible. 
     In the embodiment shown, the first and second stage turbines  26 ,  22  are both compounded with the engine core  12 ′ by having the turbine and engine shafts  24 ,  16  coupled through a gearbox  28 . In a particular embodiment, the transmission of the gearbox  28  includes a compound gear train such that torque and power may be communicated between the turbine and engine shafts  24 ,  16  in either direction. 
     In a particular embodiment, part of the compressor airflow which is delivered to the aircraft forms the output “load”. A large part of this load is supported by the turbines  26 ,  22  on the same shaft  24  and therefore the load on the engine core  12 ′ transmitted via the gearbox  28  is minimized. Thus losses and additional heat from the gearbox  28  may be minimized. Alternatively if the turbines  26 ,  22  provide more power than the compressor  20  requires the excess torque transmitted to the engine core  12 ′ may be relatively small. 
     In a particular embodiment, the engine core  12 ′ including rotary internal combustion engine(s)  12  runs at approximately 8000 rpm; other values are also possible. In a particular embodiment, the combined step up gear ratio defined by the gearbox  28  between the engine core shaft  16  and the turbine shaft  24  is between about 4:1 and 7:1, for example about 5:1. In a particular embodiment, a two stage compound idler system is used to provide the appropriate ratio and provide offset centres between the engine core shaft  16  and the turbine shaft  24 . The offset between the engine core shaft  16  and the turbine shaft  24  may allow for the hot exhaust output from the ports  50  of the core engines  12  to be ducted directly into the turbine section while minimizing the length of the ducts. 
     A generator  64  is drivable by the engine core  12 ′ to provide aircraft electrical power for accessories and/or control purposes, for example by being driven through mechanical engagement with the engine core  12 ′ directly or through the gearbox  28 , or by mechanical engagement with the turbine shaft  24 . In the embodiment shown, the generator  64  is mounted directly (i.e. without intermediate gearing) to the end of the engine core shaft  16 . In a particular embodiment the generator  64  is a 400 Hz, 6 pole alternator/generator with a design synchronous speed of 8000 rpm; other configurations are also possible. The alternator/generator  64  may serve as a startor. In a particular embodiment, elimination of any intermediate gearing between the engine core shaft  16  and the alternator/generator  64  eliminates heat generation and loss associated with that gearing (which may generally corresponds to approximately 2% of the rated generator load). 
     In a particular embodiment, the auxiliary power unit  10  includes a full authority electronic control managing all the operational requirements. The control system manages the compressor inlet guide vanes  23  and/or variable diffuser  25  (if applicable) of the shared supercharger and aircraft bleed compressor  20  to achieve the required bleed pressure and flow to the bleed duct  27  and the required fuel/air ratio in the engine core  12 ′ to maintain the governed speed. In the event of conflict between the aircraft air requirements and the governed speed, the compressor variables are set as required to allow the system to maintain the governed speed and provide priority to the generator power. In the event this action causes excess air flow or excess pressure, these conditions may be managed by opening the diverter valve  31 . A load valve (not shown) can also optionally be provided in the bleed duct  27  and managed by the control system to throttle or cut off the air supply to the aircraft. 
     With a constant volume combustion cycle in the engine core  12 ′ the breakdown of waste heat of the auxiliary power unit  10  is different from a traditional gas turbine engine auxiliary power unit. Less heat is evacuated through the exhaust and more heat is given up to the engine casing. Accordingly, the engine(s)  12  of the engine core  12 ′ have a coolant system which in a particular embodiment is distinct from any fuel and lubricating system of the auxiliary power unit  10 ; in other words, a dedicated coolant is circulated through the engine(s)  12  of the engine core  12 ′, for example through multiple coolant passages defined in the walls of the housing  32 , and this dedicated coolant is circulated separately and independently from the lubricant and the fuel of the auxiliary power unit  10 , including the lubricant of the engine core  12 ′. The dedicated coolant may be a liquid coolant, for example water. A heat exchanger defining an engine core cooler  66  includes coolant passages  66   a  (see  FIG. 1 ) in fluid communication with the coolant system of the engine core  12 ′ and air passages  66   b  (see  FIG. 1 ) in heat exchange relationship with the coolant passages  66   a.    
     The generator  64  also includes a coolant system distinct from the coolant system of the engine(s)  12 ; the coolant system of the generator may be independent from or may be common with a lubrication system of the generator  64 . The generator coolant may be a liquid coolant, for example oil. A second heat exchanger defining a generator cooler  68  includes coolant passages  68   a  (see  FIG. 1 ) in fluid communication with the coolant system of the generator  64  and air passages  68   b  (see  FIG. 1 ) in heat exchange relationship with the coolant passages  68   a . In the embodiment shown, both coolers  66 ,  68  are provided in a common package, with the coolant passages  66   a ,  68   a  of the two coolers  66 ,  68  being distinct from one another. In a particular embodiment where the generator coolant is oil or another suitable lubricant, the generator coolant system is common with (in fluid communication with) the lubrication system of the auxiliary power unit  10 , which distributes lubricant to various components of the auxiliary power unit  10  (e.g. bearings, gears, etc., of the engine core  12 ′, the compressor  20 , the turbines  22 ,  26 , the gearbox  28 ), so the second heat exchanger  68  is also an engine lubricant cooler. Alternately, a separate heat exchanger (not shown) may be provided for the lubrication system of the auxiliary power unit  10 , and the cooler  68  may be configured to cool only the generator lubricant/coolant. 
     The air passages  66   b ,  68   b  of the coolers  66 ,  68  are in fluid communication with an exhaust duct  70  of the auxiliary power unit  10 ; the exhaust duct  70  has an outlet  72  in fluid communication with the environment of the aircraft, so that the cooling air flow can be discharged to atmosphere. The exhaust duct  70  defines a cooling inlet  74  in fluid communication with an aircraft compartment  76  containing the auxiliary power unit  10 . In the embodiment shown, the coolers  66 ,  68  are received in the exhaust duct  70 . The intercooler  18  is also received in the exhaust duct  70 , upstream of the coolers  66 ,  68 . 
     A fan  78  ( FIG. 2 ) is rotatable by the engine core  12 ′ and in fluid communication with the exhaust duct  70  for driving the cooling air flow from the compartment  76 , through the heat exchangers (coolers  66 ,  68  and intercooler  18 ) and out of the exhaust duct  70  to atmosphere. In the embodiment shown, the fan  78  is received in the exhaust duct  70  upstream of the heat exchangers  18 ,  66 ,  68  and is directly driven by the engine core  12 ′, by being mounted on the end of the engine core shaft  16  opposite from the generator  64 . In a particular embodiment, direct drive of the fan  78  by the engine core shaft  16  allows to avoid additional gear loss and heat which would be produced by a gear drive. Alternately, the fan  78  may be driven through a transmission (whether in gearbox  28  or another transmission specific to the fan  78 ), or be electrically or hydraulically driven by a motor obtaining power directly or indirectly from the engine core  12 ′. 
     In a particular embodiment, the blade speed of the fan  78  is sufficiently low such that the fan  78  can be made of a common Al alloy, organic composite or thermoplastic material. In a particular embodiment, the fan  78  rotates at about 8000 rpm; other values are also possible. 
     Rotation of the fan  78  induces flow from the compartment  76 , which also provides a compartment ventilation function. In a particular embodiment, side openings from the main aircraft inlet  14  allow cooling air to flow into the compartment  76  under the driving action of the fan  78  to cool the surfaces of the auxiliary power unit  10  exposed within the compartment  76 . In a particular embodiment, the fan inlet is protected by a screen to prevent larger objects from damaging the fan  78 . 
     Although multiple distinct coolers are shown in series on  FIGS. 1-3 , alternately only one integrated cooler unit may be used with areas sub divided dedicated to the engine lubricant/generator coolant, engine core liquid coolant, and inter-cooling functions. The heat exchangers  18 ,  66 ,  68  may also be angled at an angle more than 90° to the flow direction, for example to optimize the area presented to the airflow. Although not shown, the coolers  66 ,  68  may include a thermal bypass system to prevent over-cooling at lower ambient temperatures, for example managed by the electronic control system based on sensed coolant temperatures, or by any other suitable thermostat concept. 
     The cooling system of the engine core  12 ′ is thus integrated with that of the generator  64  and with the cooling system for the lubricant of the auxiliary power unit  10 . In a particular embodiment, this integration allows for a reduction or minimization of the power loss from fans and ejectors traditionally used, and/or to avoid cooling drag penalties in flight. In a particular embodiment, the auxiliary power unit  10  is configured to reduce or avoid the generation of additional heat, for example from gear train losses. 
     Through the integrated cooling system, the same fan  78  drives the cooling air flow through the compartment  76 , engine core cooler  66 , intercooler  18 , and generator/engine lubricant cooler  68 , and then discharges the cooling air out to atmosphere through the exhaust duct  70 ; in a particular embodiment, the entire auxiliary power unit  10  and its cooling system can be installed and removed as a single assembly with interconnects and aircraft inlet and exhaust similar to that of a traditional gas turbine engine auxiliary power unit. In use and in a particular embodiment, the generator  64  and the engine core  12 ′ are thus cooled by circulating a first coolant (e.g. water) through the engine(s)  12  of the engine core  12 ′, circulating a second coolant (e.g. oil) through the generator  64 , and driving the cooling air flow in heat exchange relationship with the first and second coolants using the fan  78  driven by the auxiliary power unit  10 . 
     If applicable any diverted air from the compressor  20  can also be introduced in the exhaust duct  70 . Accordingly, in the embodiment shown, the excess air duct  29  provides a direct fluid communication between the compressor  20  and a portion of the exhaust duct  70  located downstream of the fan  78  and heat exchangers  18 ,  66 ,  68 . 
     In a particular embodiment, the exhaust duct  70  is located in a tail cone of the aircraft. As can be best seen in  FIGS. 1-2 , an intermediate duct  80  extends in fluid communication with the exhaust of the engine core  12 ′, by being connected to an exhaust of the second stage turbine  22 . The intermediate duct  80  has an outlet  82  positioned in the exhaust duct  70 , downstream of the fan  78  and upstream of the outlet  72  of the exhaust duct  70 . The outlet  82  of the intermediate duct  80  is spaced radially inwardly from a peripheral wall  70 ′ of the exhaust duct  70 . The air and exhaust gases are thus discharged in the exhaust duct  70  so that the flow of cooling air surrounds the flow of exhaust gases. The mass flow and/or volume of flow of exhaust gases is/are smaller than the flow of cooling air. In a particular embodiment, the mass flow of exhaust gases is 20% or less of the mass flow of cooling air. An open cross-sectional area of the outlet  82  of the intermediate duct  80  is smaller than an open cross-sectional area of the exhaust duct  70  around the outlet  82  of the intermediate duct  80  (where “open cross-sectional area of the exhaust duct  70 ” refers to the cross-sectional area of the exhaust duct  70  not occupied by the intermediate duct  80 ). In a particular embodiment, the ratio of the diameter of the intermediate duct  80  on the diameter of the exhaust duct  70  is from 0.2 to 0.4, for example around ⅓. Other values are also possible, depending for example on the optimisation of the weight and cost of the auxiliary power unit  10  as a whole. 
     In the embodiment shown, the intermediate duct  80  is concentric with the peripheral wall  70 ′ of the exhaust duct  70 ; the flow of exhaust gases is thus discharged along a central axis C of the exhaust duct  70 . 
     In a particular embodiment, the larger and cooler cooling air flow surrounding the exhaust gas flow allows for the peripheral wall  70 ′ of the exhaust duct  70  to be made of materials requiring a lower resistance to high temperature than materials which would be in direct contact with the exhaust gas flow, where “resistance to high temperature” refers to the ability of a material to keep their strength, rigidity and durability when submitted to high temperatures. This may allow for the use of less expensive materials for the peripheral wall  70 ′ of the exhaust duct  70 . In a particular embodiment, the temperature of the flow against the peripheral wall  70 ′ of the exhaust duct  70  is lower than that against the exhaust duct of a traditional gas turbine engine auxiliary power unit, so that the use of high temperature materials (e.g. nickel or titanium alloy) is not required for the peripheral wall  70 ′. For example, the temperature of the exhaust gases may be 800° F. or more, potentially up to 1200° F.−1400° F., while the cooling air flow temperature may be 250° F. or less; surrounding the exhaust gas flow with the cooling air flow thus significantly reduces the temperature of the flow in contact with the peripheral wall  70 ′. In a particular embodiment, the peripheral wall  70 ′ of the exhaust duct  70  is made of any suitable aluminum alloy, any suitable light metal alloy, any suitable composite material including, but not limited to, carbon fiber composite materials, or any suitable type of polymer. 
     In a particular embodiment, the fan  78  can be designed to deliver enough kinetic energy to act as an ejector pump for the exhaust from the turbines  26 ,  22  and increase the energy delivered by the turbines  26 ,  22 . 
     In particular embodiment, the exhaust of the turbine section is configured so that the flow of exhaust gases expelled from the intermediate duct  80  has a higher velocity than the surrounding cooling air flow circulating in the exhaust duct  70 . In a particular embodiment, the difference in velocity is selected to create an entrainment effect in the cooling air flow, so as to help circulation of the cooling air flow through the heat exchangers  18 ,  66 ,  68  driven by the fan  78 . This may allow for the size of the fan  78  to be reduced, as compared to a configuration without such an entrainment effect. 
     In a particular embodiment, the inlet and exhaust of the auxiliary power unit  10  are located on the aircraft skin such that the inlet ram pressure significantly exceeds the static pressure at the exhaust plane; this pressure may be used with a venturi effect to depress the static pressure at the exhaust plane of the turbines  26 ,  22  in flight, and/or the fan  78  may be reversible such that it can act as a turbine and recover energy in high ram conditions where it is not needed to boost cooling flow. 
     In a particular embodiment, the auxiliary power unit inlet  14  at the aircraft fuselage is provided with a door to prevent unintended wind-milling and drag when the auxiliary power unit is not operating. Where high speed performance is required in flight this door can be shaped to act as a ram air scoop. 
     In a particular embodiment, additional aircraft thrust is gained or the drag penalty is reduced by taking credit for the waste thermal energy transferred to the cooling. In order to maximise this effect (comparable to the Meredith effect in liquid cooled propulsion engines) the sizing of the outlet  82  of the intermediate duct  80  is optimized and the exhaust vector set to provide the maximum propulsive benefit to the aircraft. 
     Referring to  FIG. 1 , in a particular embodiment the auxiliary power unit  10  includes mounts  84  on the gearbox  28  and near the inlet  74  of the exhaust duct  70 ; a single inlet flange and a single exhaust flange are provided for ease of mounting. The integrated cooling system also facilitates installation of the auxiliary power unit  10  in the compartment  76 . 
       FIGS. 5-8 and 11  show an auxiliary power unit  110  in accordance with another embodiment, where elements similar to that of the embodiment of  FIGS. 1-3  are identified with the same reference numerals and will not be further described herein. 
     In this embodiment, the engine core cooler  166  and the generator/engine lubricant cooler  168  are disposed in parallel with respect to one another. As can be best seen in  FIG. 7 , a cooling air duct  186  extends radially outwardly around a circumference of the exhaust duct  70 . The cooling air duct  186  has an outlet in fluid communication with the exhaust duct  70  and an inlet disposed radially outwardly of the outlet and in fluid communication with the compartment  76  through the coolers  166 ,  168 . The engine core cooler  166  and the generator/engine lubricant cooler  168  each extend around a respective portion of a circumference of the cooling air duct  186 . The fan  78  is located in the exhaust duct  70 , thus downstream of the coolers  166 ,  168 . As can be seen from  FIG. 6 , the two coolers  166 ,  168  together extend around only part of the circumference of the exhaust duct  70 , with the intermediate duct  80  and excess air duct  29  extending adjacent the exhaust duct  70  in the circumferential portion free of the coolers  166 ,  168 . The coolers  166 ,  168  can be mounted directly to the auxiliary power unit  110  as shown, or could alternately be installed on the aircraft and linked to the auxiliary power unit  110  with tubing (e.g. flexible tubing). 
     Referring back to  FIG. 7 , it can be seen that the air passages  166   b ,  168   b  of the coolers  166 ,  168  extend along a radial direction R of the auxiliary power unit  110 . Alternately, other orientations for the coolers  166 ,  168  are possible. 
     Still referring to  FIG. 7 , variable pitch blades or variable inlet guide vanes  188  can be provided in the cooling air duct  186  and its junction with the exhaust duct  70 , immediately upstream of the fan  78 , so as to be able to modulate the airflow through the coolers  166 ,  168  and/or control fan power absorption at lower heat load conditions. 
     As can be best seen in  FIGS. 5-6 , the intercooler  118  is not in fluid communication with the exhaust duct  70 , and is instead configured as an air to liquid cooler; the intercooler  118  includes fluid passages receiving the coolant from the engine core  12 ′ through one or more conduits  118 ′ (for example at about 200° F.) and circulating the coolant in heat exchange relationship with the compressed air from the compressor  120  (for example at 450° F.) before the coolant is circulated to the engine core cooler  166  through one or more conduits  118 ″. The intercooler  118  is thus located upstream of the engine core cooler  166  and downstream of the engine core  12 ′ in the coolant circulation path. 
     As can be best seen in  FIGS. 6 and 11 , in this embodiment two compressors are provided: a supercharger compressor  120  to provide compressed air to the engine core  12 ′, and a bleed compressor  121  to provide bleed air for the aircraft. The two compressors  120 ,  121  are connected to the same shaft  124 , which also receives the turbines  26 ,  22  of the turbine section. The compressor inlets can be connected to a common plenum  119  ( FIG. 11 ) or to a respective plenum  119   a ,  119   b  ( FIGS. 5-6 , dotted lines in  FIG. 11 ), with the plenum(s)  119 ,  119   a ,  119   b  being connected to the main inlet  14 . In a particular embodiment, such a configuration allows for accommodating different functional requirements for the supercharging flow (to the engine core  12 ′) and the aircraft flow (to the bleed duct  27 ). 
       FIGS. 9-10  show an auxiliary power unit  210  similar to that of  FIGS. 5-8 , where elements similar to that of the embodiment of  FIGS. 1-3  and/or to that of the embodiment of  FIGS. 5-8  are identified with the same reference numerals and will not be further described herein. The compartment  76  is shown as defined by the tail cone  290  of the aircraft, with the exhaust duct outlet  72  located at the tip of the tail cone  290 . The tail cone  290  defines the main inlet  14  to the compartment  76 , to which the compressor inlets are connected. The auxiliary power unit of  FIGS. 1-3  and/or of  FIGS. 5-8  may be similarly installed. 
     The engine core cooler  266  and the generator/engine lubricant cooler  268  have a rectangular configuration and are circumferentially and axially offset from one another about the exhaust duct  70 ; each is connected to the exhaust duct  70  through a respective cooling air duct  286  ( FIG. 10 ) extending radially outwardly from the exhaust duct  70 . One or both of the coolers  266 ,  268  can have air conduits angled with respect to the radial direction of the auxiliary power unit  210 . 
       FIG. 12  shows an alternate configuration for the cooling inlet and heat exchangers  318 ,  366 ,  368 , which may be used in any of the auxiliary power units  10 ,  110 ,  210  described above. A bifurcated inlet system includes two separate cooling air ducts  386   a ,  386   b , which in a particular embodiment may allow minimizing the length of the cooling air ducts  386   a ,  386   b  and/or of the coolant/lubricant conduits connected to the coolers  366 ,  368  and/or of the compressed air conduits connecting the intercooler  318  to the compressor  320  and to the engine core  12 ′. The cooling air duct  386   a  closest to the engine core inlet manifold  392  is dedicated to the intercooling function and accordingly receives the intercooler  318 , which is this embodiment is air cooled. The other cooling air duct  386   b  receives one or both of the engine core cooler  366  and the generator/engine lubricant cooler  368 . The position of the heat exchangers within the cooling air ducts  386   a ,  386   b  (e.g. how the heat exchangers are grouped in each cooling air duct) may vary, for example depending on the relative demand for cooling air. The pressure losses in each cooling air duct  386   a ,  386   b  of the bifurcated system are balanced to avoid distorting the inlet flow of the fan  78 , which is located in the exhaust conduit  70  downstream of the heat exchangers  318 ,  366 ,  368 . In a particular embodiment, the generator/engine lubricant cooler  368  is positioned in the same cooling air duct  386   a  as the intercooler  318 , with the engine core cooler  366  located in the second cooling air duct  386   b . In another particular embodiment, a whole or a part of the engine core cooler  366  is positioned in the same cooling air duct  386   a  as the intercooler  318 , with the generator/engine lubricant cooler  368  located in the second cooling air duct  386   b.    
       FIG. 13  shows another alternate configuration for the cooling inlet and heat exchangers  418 ,  466 ,  468 , which may be used in any of the auxiliary power units  10 ,  110 ,  210  described above. A bifurcated cooling air duct  486  extends non perpendicularly and at a non-zero angle with respect to the exhaust conduit  70 , with an outlet of the cooling air duct  486  being in fluid communication with the exhaust conduit  70  upstream of the fan  78 . The heat exchangers are received in the cooling air duct, with the engine core cooler  466  and generator/engine lubricant cooler  468  being located upstream of the intercooler  418 . In a particular embodiment, the heat exchangers  418 ,  466 ,  468  are placed as close to the engine core  12 ′ as possible, and weight, volume and losses associated with piping the cycle air as well as the lubricant and liquid coolant is minimized. 
     In a particular embodiment, having the heat exchangers  166 ,  168 ,  266 ,  268 ,  318 ,  366 ,  368 ,  418 ,  466 ,  468  located upstream of the fan  78  allows for the heat exchangers to be smaller, since the air circulated therethrough is cooler. However, the fan  78  downstream of the heat exchangers is exposed to warmer air than a fan upstream of the heat exchangers, and accordingly the power requirement for the fan  78  downstream of the heat exchangers may be greater. 
       FIG. 14  shows an alternate configuration for the two compressors, which may be used in replacement of the compressor(s) of any of the auxiliary power units  10 ,  110 ,  210  described above. The supercharger compressor  520  providing the compressed air to the engine core  12 ′ and the bleed compressor  521  providing the compressed air to the aircraft are arranged on both sides of a single rotor  594 , which in a particular embodiment is manufactured by forging. The rotor  594  may be received on a shaft  524  driven by the turbine section. Tip seals  596  (e.g. labyrinth or fin type air seals) with a low pressure “sink” (exhaust)  596  below either of the impeller delivery pressures (e.g. to ambient) are arranged at the impeller tips to prevent interference between the two compressors  520 ,  521  which might result in premature stall or surge, when the two sides are operating at different pressures. 
       FIGS. 15-16  show an alternate configuration for the compressors and turbines, which may be used in replacement of the compressor(s) and turbines of any of the auxiliary power units  10 ,  110 ,  210  described above. The supercharger compressor  620  is mounted on a separate turbocharger shaft  698  with the second stage (e.g. pressure) turbine  622 , and where the first stage turbine  626  drives the bleed compressor  621  through a turbine shaft  624  and is compounded with the engine core  12 ′ through the gearbox  28 . In a particular embodiment, such a configuration allows for the turbocharger  620  to find its own match point and possibly eliminate the need for variables on one of the compressors  620 ,  621 . Variable nozzle geometry (e.g. variable area turbine vanes  699 , see  FIG. 16 ) could be introduced on the second stage turbine  622  to improve controllability of the degree of supercharge. In a particular embodiment, such a configuration allows for the speed of the second stage turbine  622  to be selected independent of the requirements for the first stage turbine  626 . As can be seen in  FIG. 16 , in a particular embodiment the turbocharger shaft  698  is concentric with the shaft  624  of the first stage turbine  622 , and a common inlet plenum  619  is provided for both compressors  620 ,  621 . It is understood that although the second stage turbine  622  is shown as a radial turbine, it could alternately be an axial turbine. 
     Size effects, material capability and cost considerations generally limit the efficiency of typical present gas turbine engine auxiliary power units. In a particular embodiment, the auxiliary power unit  10 ,  110 ,  210  including some measure of constant volume combustion aided by variable supercharging to preserve high altitude performance provides for an increase in efficiency with minimal complexity or need for sophisticated materials requirements and/or improved specific cost as compared to a traditional gas turbine engine auxiliary power unit. 
     Like typical auxiliary power unit installations, the auxiliary power unit  10 ,  110 ,  210  can be used to provide both medium pressure air for aircraft use and constant speed shaft power to drive a generator, for example at synchronous speed for 400 Hz. The auxiliary power unit  10 ,  110 ,  210  may be operated for air alone, electrical power alone or some combination of both types of load at the same time. Normally combined load occurs in ground or low altitude operation. In flight, at altitudes up to the aircraft ceiling, the auxiliary power unit is typically required to be operable for electrical power only, as an additional electrical power source after the main engine(s). In a particular embodiment, the present auxiliary power unit  10 ,  110 ,  210  includes variable supercharging to sustain the required power output in the less dense air at high altitude. 
     In a particular embodiment, the auxiliary power unit  10 ,  110 ,  210  is configured with simple inlet and exhaust connections (including main, load and cooling gas paths) to facilitate quick removal and replacement comparable to the traditional gas turbine engine auxiliary power units. 
     It is understood that the engine assemblies shown as auxiliary power units  10 ,  110 ,  210  may alternately be configured as other types of engine assemblies, including, but not limited to, turboshaft engine assemblies where the engine core  12 ′ is configured as or drivingly engaged to an output shaft, and turboprop engine assemblies where the engine core  12 ′ is drivingly engaged to a propeller. 
     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. Each rotor shown may be a centrifugal or axial device, and may be replaced by two or more rotors having radial, axial or mixed flow blades. 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.