Abstract:
A system for managing heat fluxes of an aircraft is provided, including a fuel tank supplying a turbomachine with fuel via a fuel supply circuit; a cell cooling circuit connected to a thermal effluents source, which integrates a first heat exchanger ensuring heat transfer between a coolant circulating in the cell cooling circuit and air flow channeled into a cooling air channel extending from an air intake to an exhaust; and a second heat exchanger ensuring a heat transfer between the coolant circulating in the cell cooling circuit and the fuel. The air intake includes a blocker to open/close it. The system includes a fuel transfer circuit, which connects the fuel tank to the second heat exchanger and provides fuel to the second heat exchanger, including a valve selectively changing a fuel supply direction from the second heat exchanger between the directions of the tank and the fuel supply circuit.

Description:
BACKGROUND OF THE INVENTION 
     1. Field of the Invention 
     This invention relates to a system for managing heat fluxes of an aircraft. 
     2. Description of the Related Art 
     An aircraft comprises a cell and at least one propulsion system. In  FIGS. 1 and 2 , a cell was shown diagrammatically at  10  and a propulsion system was shown diagrammatically at  12 . 
     According to a widely used embodiment, a propulsion system is suspended under a wing by means of a mast. More generally, the propulsion system is connected to the cell by means of an interface  14  that is embodied by dotted lines in  FIGS. 1 and 2 . 
     A propulsion system  12  comprises a turbomachine  16  that is equipped with a first engine cooling circuit  18 , in which a coolant, in particular oil, circulates. 
     The turbomachine is supplied with fuel by means of a fuel circuit  20  that extends from a tank  22  that is arranged at the cell. As illustrated in  FIG. 1 , it is possible to use several cooling sources to cool the oil of the turbomachine, for example by using at least one oil/fuel exchanger  24  at the first engine cooling circuit  18 , and to use the fuel as coolant so as to cool the oil of the turbomachine. In this case, a recirculation circuit  26  is provided so as to reintroduce the fuel that is heated in the tank  22 . In addition, the propulsion system  12  can comprise another source  28  of thermal effluents, for example one or more electric generators installed close to the turbomachine. 
     So as to optimize the operation of these elements  28 , it is necessary to regulate their temperatures by means of a second engine cooling circuit  30 , in which a coolant that passes through a third engine exchanger  32 , in particular an oil/fuel exchanger according to  FIG. 1 , circulates. 
     The characteristics of each engine cooling circuit, namely the characteristics of the fluid to be cooled, for example its flow rate, the characteristics of the exchanger, for example its dimensions, the characteristics of the fluid that is used for cooling, for example its flow rate, are adjusted based on the requirements for regulation of the temperature at the source, in particular so as to keep the temperature of the source below a certain threshold. 
     In the case of the first engine cooling circuit, these requirements vary according to the operation of the aircraft and are more significant when the aircraft is on the ground to the extent where the outside air can be at a high temperature and there is no air flow linked to the movement of the aircraft. 
     Thus, the characteristics of the engine cooling circuit relative to the turbomachine are generally determined based on the most significant constraints when the aircraft is on the ground. 
     The cell  10  also comprises at least one source of thermal effluents  34  and in general several sources  34 ,  34 ′, for example electrical accessories, electronic power systems, an air-conditioning system, avionics, and client equipment. All of these elements are to be temperature-regulated to ensure their operation and to guarantee the highest availability rate. For this purpose, at least one cell cooling circuit  36  is provided. According to the illustrated example, the cell comprises two cell cooling circuits  36 ,  36 ′, each comprising an exchanger  38 ,  38 ′ that makes it possible to cool the coolant that circulates in each of the circuits. 
     According to an embodiment that is illustrated in  FIG. 1 , the exchangers  38 ,  38 ′ are arranged in at least one cooling air channel  40  that is arranged in the lower portion of the fuselage and that comprises—upstream—one or more air intakes  42 , preferably of the dynamic type, and—downstream—one or more exhausts  44 . 
     The cooling air channels  40  operate according to two primary methods:
         The first mode of operation takes place on the ground when the aircraft is immobile or moves at reduced speeds. In these cases, the natural air flow within said cooling air channels  40  is generally very low and is to be forced using in particular an electric fan  46  to allow the evacuation of the heat fluxes.   The second mode of operation takes place in flight when the aircraft moves at high speeds in a cold atmosphere. In this case, the dynamic pressure at the air intakes is significant, and the ambient temperature is relatively low; the effectiveness of the exchanges is very significant, so that it is necessary to limit the flow rate of air circulating in the cooling air channels to not over-cool the thermal effluent sources.       

     In all of the cases, the air intakes are never blocked because the cooling air requirements always exist regardless of the mode of operation. To the extent that the air intakes always interfere with the aerodynamic flow, the cooling air channels prove to be detrimental in terms of aerodynamic drag for the aircraft and therefore in terms of energy consumption of the propulsion systems. 
     According to a first variant, the cooling air channels have set dimensions and are consequently simple, light and reliable. However, whereby their dimensions are calculated based on the most significant requirements, their impact on the aerodynamic drag, and therefore on the consumption of the aircraft, is significant during the high-speed flight phases, whereas the requirements are normally low for these flight phases. 
     According to another variant, the cooling air channels have a variable geometry to adapt their dimensions based on requirements, but in this case, the channels prove complex, heavy, and not very reliable. 
     As illustrated in  FIG. 1 , for the cooling systems that are implanted in propulsion systems, the fuel tanks  22  can constitute heat sinks. Actually, it is known to one skilled in the art that even when the fuel level is at its lowest, the tanks have intrinsic capacities for absorbing the thermal effluents. 
     So as to eliminate the drawbacks linked to the cooling air channels, according to another variant illustrated in  FIG. 2 , the exchangers  38 ,  38 ′ are not placed in a cooling air channel but ensure a heat transfer between the coolant of the cell cooling circuit(s)  36 ,  36 ′ and the fuel for conveying the heat fluxes in the direction of the tanks. For this purpose, at least one circuit  48  is provided between the tank(s)  22  and the exchanger(s)  38 ,  38 ′. 
     This relatively simple solution makes it possible to eliminate the cooling air channels and consequently is not detrimental in terms of aerodynamic drag and therefore energy consumption. 
     However, this solution is not completely satisfactory because its operating period is limited to the extent where it is no longer possible to dissipate the heat in the tanks when the fuel temperature reaches a certain threshold linked to the fuel temperature that is accepted by the turbomachines or to the risks of inflammability of the tank. 
     Consequently, when this threshold is reached, the capacities for heat dissipation are low, so that it is necessary to operate certain sources of thermal effluents in degraded mode; this is, for example, the case of the air-conditioning systems of the cell at the end of the flight. 
     According to other constraints, in terms of aircraft design, the components of the cell and the components of the propulsion systems are segregated for safety reasons. Actually, it is necessary to ensure that a breakdown that appears at the cell and disrupts the operation of the propulsion systems is extremely improbable. 
     BRIEF SUMMARY OF THE INVENTION 
     Also, the purpose of this invention is to eliminate the drawbacks of the prior art by proposing a system for managing the heat fluxes of an aircraft that limits the impact on the aerodynamic drag and therefore on the energy consumption of the aircraft, there being no limit on the operating period. 
     According to another objective, the system for managing the heat fluxes is to preserve the cell/propulsion system segregation. 
     For this purpose, the invention has as its object a system for managing the heat fluxes of an aircraft comprising at least one propulsion system that integrates a turbomachine and a cell that comprises, on the one hand, at least one fuel tank so as to supply the turbomachine via a fuel supply circuit, and, on the other hand, at least one source of thermal effluents connected to a cell cooling circuit that integrates first means for ensuring a heat transfer between a coolant that circulates in said cell cooling circuit and an air flow that is channeled into at least one cooling air channel that extends from at least one air intake up to at least one air exhaust, whereby said system comprises second means for ensuring a heat transfer between a coolant that circulates in said cell cooling circuit and the fuel, a fuel circuit that connects said at least one tank to said second heat transfer means, whereby the air intake comprises blocking means that can occupy a first open state in which they allow air to pass inside the channel and a second closed state in which they block the air intake, characterized in that the fuel circuit comprises at least one control valve that orients the fuel either in the direction of the tank or in the direction of the circuit designed for the propulsion system. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Other characteristics and advantages will emerge from the following description of the invention, a description that is provided only by way of example, with regard to the accompanying drawings, in which: 
         FIG. 1  is a diagram that illustrates a first variant of a system for managing the heat fluxes of an aircraft according to the prior art, 
         FIG. 2  is a diagram that illustrates a second variant of a system for managing the heat fluxes of an aircraft according to the prior art, 
         FIG. 3  is a diagram that illustrates a first variant of a system for managing the heat fluxes of an aircraft according to the invention, 
         FIG. 4  is a diagram that illustrates another variant of a system for managing the heat fluxes of an aircraft according to the invention, 
         FIG. 5  is a cutaway illustrating a cooling air channel that integrates two double-flux exchangers according to a first embodiment of the invention, 
         FIG. 6  is a cutaway that illustrates a cooling air channel that integrates two triple-flux exchangers according to another embodiment of the invention, 
         FIG. 7  is a cutaway of a double-flux exchanger, and 
         FIG. 8  is a cutaway of a triple-flux exchanger. 
     
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS 
       FIGS. 3 and 4  diagrammatically show an aircraft with a cell  50  and at least one propulsion system  52 . 
     Cell is broadly defined as all of the elements of the aircraft, in particular the fuselage, the wings, and the tail assemblies, with the exception of the propulsion systems. 
     According to a widely used embodiment, a propulsion system is suspended under a wing by means of a mast. More generally, the propulsion system is connected to the cell by means of an interface  54  that is embodied by dotted lines in  FIGS. 3 and 4 . 
     A propulsion system  52  comprises a turbomachine  56  that is equipped with a first engine cooling circuit  58 , in which a coolant, in particular oil, circulates. 
     The turbomachine  56  is supplied with fuel by means of a fuel supply circuit  60  that extends from at least one tank  61  that is arranged at the cell. 
     To ensure the cooling of the coolant of the engine cooling circuit  58 , said circuit  58  comprises at least one first engine exchanger  62 . According to an embodiment that is illustrated in  FIGS. 3 and 4 , the exchanger(s)  62  ensure(s) a heat transfer between the coolant that circulates in the cooling circuit  58  of the engine and the fuel. 
     When the fuel is used as a vector to transfer heat, means  64  are provided upstream from the turbomachine  56  to regulate the fuel that is injected into the turbomachine  56  as well as a return circuit  66  of the fuel in the direction of the tank  61 . This regulation and this return circuit are necessary when, for example, the quantity of fuel to ensure adequate cooling is more than the quantity of fuel injected into the turbomachine. 
     In addition, the propulsion system  52  can comprise another source  68  of thermal effluents, for example one or more electric generators installed close to the turbomachine. 
     So as to optimize the operation of these sources  68 , it is necessary to regulate their temperatures by means of a second engine cooling circuit  70 , in which a coolant that passes through at least one second engine exchanger  72  circulates. 
     According to an embodiment that is illustrated in  FIGS. 3 and 4 , the exchanger(s)  72  ensure(s) a heat transfer between the coolant that circulates in the cooling circuit  70  of the engine and the fuel. 
     In addition, at the propulsion system  52 , the fuel circuit  60 ,  66  can comprise at least one pump  74 , at least one recirculation valve  76  to manage the fuel flux between the tanks, the turbomachines  56 , and the different exchangers  62 ,  72 . 
     The invention is not limited to the embodiment that is shown in  FIGS. 3 and 4  that relates to the portion of the system for managing the heat fluxes that is integrated in the propulsion system. Thus, the fuel circuit  60  can be simplified and can comprise only one connection between a tank and a turbomachine. In the same way, the exchangers  62  and  72  can use air to evacuate the heat instead of fuel, whereby these exchangers comprise at least one surface that is in contact with the aerodynamic fluxes so as to reduce the impact on the drag. 
     The cell  50  also comprises at least one source of thermal effluents  78 , and generally several sources  78 ,  78 ′, for example electrical accessories, electronic power systems, an air-conditioning system, avionics, and client equipment. All of these elements are to be temperature-regulated to ensure their operation and to guarantee the highest availability rate. For this purpose, at least one cell cooling circuit  80  is provided. According to the illustrated example, the cell  50  comprises two cell cooling circuits  80 ,  80 ′, one for each source of thermal effluents. 
     According to the invention, the cell cooling circuit  80  comprises first means  82  for ensuring a heat transfer between the coolant that circulates in said circuit  80  and the air, as well as second means  84  for ensuring a heat transfer between the coolant that circulates in said circuit  80  and the fuel. 
     According to an embodiment that is illustrated in  FIG. 3 , the cell cooling circuit  80  comprises a first fluid/air exchanger  86  that is arranged in a cooling air channel  88  as well as a second fluid/fuel exchanger  90  that is distant and different from the first exchanger  86 , a circuit  92  that ensures the supply of said second exchanger  90  with fuel, whereby the latter is either reinjected in the tank  61  or oriented toward the circuit  60  that is designed for a propulsion system  52 . According to this variant, the exchangers  86  and  90  are of the double-flux type. 
     According to another embodiment that is illustrated in  FIG. 4 , the cell cooling circuit  80  comprises a single fuel/fluid/air exchanger  94  that is arranged in a cooling air channel  96 , a circuit  98  ensuring the supply of said exchanger  90  with fuel, whereby the latter is either reinjected into the tank  61  or oriented toward the circuit  60  that is designed for a propulsion system  52 . According to this variant, the exchanger  94  is of the triple-flux type. 
     Preferably, to preserve the principle of segregation of the elements, each source of thermal effluents  78 ,  78 ′ comprises a cooling circuit  80 ,  80 ′ that is specific thereto, whereby each circuit  80 ,  80 ′ comprises either two double-flux exchangers  86  and  90 , respectively  86 ′ and  90 ′, or is a triple-flux exchanger  94 , or  94 ′. 
     Preferably, the exchangers  86 ,  86 ′ (or  94 ,  94 ′) are arranged in a single cooling air channel  88  (or  96 ). As a variant, the exchangers each use a cooling air channel that is specific thereto. 
     The cooling air channel  88  or  96  extends from at least one air intake  100  and at least one air exhaust  102 . 
     According to an important characteristic of the invention, the air intake  100  comprises blocking means  104  that can occupy a first open state (in dotted lines in  FIGS. 3 and 4 ), in which they allow the air to pass inside the channel and a second closed state (in heavy lines in  FIGS. 3 and 4 ), in which they block the air intake so as to minimize the impact of said intake on the aerodynamic drag. 
     Preferably, means  106  are provided to force the air flow into the cooling air channel, for example a fan. 
     According to an embodiment that is illustrated in  FIGS. 5 and 6 , the air exhaust  102  is of the leveling type and is located at a surface that is in contact with the aerodynamic fluxes flowing outside of the aircraft. Advantageously, the air exhaust  102  comprises a grid that makes it possible to limit the impact of said exhaust on the aerodynamic drag of the aircraft. 
     According to one embodiment, the air intake  100  is of the leveling type and is located at a surface that is in contact with the aerodynamic fluxes that flow outside of the aircraft. The blocking means  104  come in the form of a door  108  that is connected to the aircraft by means of a hinge  110 , whose opening and closing are controlled by an actuator  112 , whereby said door is in closed position at a surface that is in contact with the aerodynamic fluxes. Of course, the shapes of the door in closed position as well as those of the hinge  110  are defined so as to reduce the impact on the aerodynamic drag. 
     As illustrated in  FIGS. 5 and 6 , the door can open toward the outside so as to project relative to the surface of the aircraft that is in contact with the aerodynamic fluxes. However, as will be explained below, this selection has only a slight influence on energy consumption to the extent where the door is open when the aircraft is revving up or moves at a low speed. 
     This arrangement with an opening toward the outside makes it possible to limit the space occupied by the cooling air channel, which makes it possible to have a large passage section. 
     The cooling air channel  88  or  96  can have different shapes. It generally has the shape of a more or less tapered U or V. 
     The shape of the channel and the arrangement of the exchanger(s) are adapted so as to promote the heat exchanges at the exchanger(s) and to reduce the differential heads. The channel is optimized for the ground phases. 
     Preferably, the exchangers are arranged perpendicularly to the flow of air circulating in the cooling air channel. 
     According to a first embodiment, as illustrated in  FIG. 5 , the cooling air channel  88  has a V shape and comprises two double-flux exchangers  86 ,  86 ′, coupled, arranged at the point of the V, whereby the fan  106  is inserted between the exchangers  86 ,  86 ′ and the exhaust  102 . 
     According to another embodiment, as illustrated in  FIG. 6 , the cooling air channel has a U shape and comprises two coupled triple-flux exchangers  94 ,  94 ′, arranged at the branch of the U that is connected to the air exhaust  102 , whereby the fan  106  is inserted between the exchangers  94 ,  94 ′ and the exhaust  102 . 
       FIG. 7  shows in section a double-flux plate exchanger  86 , which comprises a first series of pipes  114  designed for coolant and connected to the cooling circuit  80 ,  80 ′, whereby said pipes  114  are kept spaced from one another using separators  116  that allow air to pass. 
       FIG. 8  shows in section a triple-flux plate exchanger  94  that comprises a first series of pipes  118  that are designed for coolant and connected to the cooling circuit  80 ,  80 ′, whereby each two are coupled to a pipe  120  of a second series of pipes designed for fuel and connected to the fuel circuit  98 , whereby the pairs of pipes  118  and  120  are kept spaced from one another using separators  122  that allow air to pass. 
     The air/fluid exchangers are optimized for the ground phases whereas the fuel/fluid exchangers are optimized for the flight phases. 
     According to one embodiment, at the output of the tank  61 , the fuel circuit  92  or  98  comprises a pump  124  that is followed by a first control valve  126  that orients the fuel either in the direction of the exchangers  84 ,  84 ′,  94 ,  94 ′ of the cell or in the direction of a second control valve  128  that orients the fuel either in the direction of the tank  61  or in the direction of the circuit  60  that is designed for a propulsion system. 
     Between the first control valve  126  and the exchangers  84 ,  84 ′,  94 ,  94 ′, at least one third control valve  130  makes it possible to control the flow rate of the fuel toward each of said exchangers. 
     At the output of the exchangers  84 ,  84 ′,  94 ,  94 ′, at least one fourth control valve  132  is provided to collect the fuel fluxes that come from the exchangers and to orient them toward the second control valve  128 . 
     Other technical solutions can be considered to regulate the fuel fluxes in the direction of the exchangers, of the tank and propulsion systems. 
     The operation of the system for managing the heat fluxes is now described. 
     During the ground phases, the cooling air channel(s)  88 ,  96  are open, and the means  106  for generating an air flow inside the channel are activated. The thermal effluents of the aircraft are thus dissipated in the atmosphere using exchangers  86 ,  86 ′ or  94 ,  94 ′. 
     The cooling air channel, the exchangers and the means  106  are thus sized for this ground phase and not for the cruising phases. This channel can optionally be activated at low speeds. 
     Consequently, the design of the cooling air channel and more particularly the air intake is simplified, whereby the air intake does not need to be of the dynamic type to recover the kinetic energy of the incoming air. 
     Relative to the geometry of the channel, it is no longer optimized based on two different flight regimes, namely at high speed and when revving up. Consequently, the channel does not have variable geometry but rather a stationary geometry adapted to the flight phases when revving up or at low speeds, which makes it possible to obtain a reliable, light and compact channel. 
     According to another advantage, the positioning of the air intake  100  is no longer dictated by considerations of aerodynamic order but by installation constraints. 
     During the flight phases, in particular at high speeds, the cooling air channel is closed, so that it does not induce any impact on the aerodynamic drag. In this case, the thermal effluents are dissipated via the fuel in the tank(s) or in the propulsion systems or other elements that are linked to the fuel circuit. 
     In the case where the thermal effluents are dissipated in the fuel tanks, certain sizing precautions should be taken. However, to the extent that the thermal effluents are dissipated in the fuel during flight phases at high altitudes and at high speeds, the heating of the fuel and therefore of the tank from which the heat exchanges originate is counterbalanced by the cooling of tanks from which aerodynamic fluxes that flow outside of the wings originate. 
     When the thermal effluents are dissipated in the fuel tanks, there is no limitation on use other than the fuel&#39;s heat storage capacity itself and maximum limitations of fuel temperatures in the piping. 
     During certain flight phases at low speeds for which the heat absorption capacities of the fuel circuit are limited either because the quantity of fuel remaining in the tanks is low or because the turbomachines operate while idling and do not absorb much fuel, the cooling air channel(s) can be open to supplement the cooling capacity. At these speeds, the channels induce only a low impact on the aerodynamic drag. If necessary, the fans can force the flux in the cooling air channels without it being necessary to provide air intakes of the dynamic type. 
     In the case where the effluents are dissipated in the fuel that supplies the turbomachines, the possible increase of the aerodynamic drag that originates from the air/fluid exchangers will be compensated by the additional thrust that is produced because of the higher temperature of the fuel. 
     If, upon the departure of the aircraft, the flight personnel see that the dissipation of the thermal effluents via the cooling air channel cannot be used due to a breakdown of the fan, the actuator controlling the opening of the door, the pilot can decide to use the tanks to dissipate the heat. In this case, as for the prior art illustrated in  FIG. 2 , this configuration has operational limits in terms of operating period. However, it makes it possible for the aircraft to accomplish its mission. 
     As a variant or in a complementary manner, the pilot can decide to use propulsion systems to dissipate heat. In this case, even if this configuration can have an impact on the sizing of the engine exchangers, it makes it possible for the aircraft to accomplish its mission.