Abstract:
A method and system limiting specific consumption of an aircraft by matching sizing of a power supply to actual power needs of a cabin pressure control system. The method optimizes overall efficiency of energy supplied onboard an aircraft including, in an environment near the cabin, at least one main power-generating engine, sized to serve as a single pneumatic energy-generating source for the cabin and as an at most partial propulsive, hydraulic, and/or electric energy-generating source for the rest of the aircraft. The method minimizes power differential between a nominal point of the power sources when the sources are operating, and a sizing point of non-propulsive energy contributions of the sources when the main engine has failed, by equally dividing power contributions of the main engines and the main power generator under nominal operating conditions and in an event of failure of a main engine.

Description:
TECHNICAL FIELD 
     The invention relates to a method for optimizing the overall efficiency of the energy supplied aboard an aircraft, this energy being propulsive or non-propulsive, as well as to a main power unit for implementing such a method. 
     The invention applies to the engine set of aircrafts, i.e. essentially to the engine set of airplanes (jet engines, turbojet engines, turboprops) as well as to the engine set of helicopters (turboshaft engine). 
     Typically, in an aircraft, the cabin which accommodates the passengers is air-conditioned and/or pressurized. An air inlet of the cabin is connected to an environmental control system ECS (initials for “Environmental Control System”) which adjusts the air-flow rate, temperature and/or pressure in collaboration with a possible recirculating system between the ECS system and the cabin. 
     STATE OF THE ART 
     It is known how to recover energy between the air at the outlet of the cabin, which has high pressure and temperature—typically 0.8 bar and 24° C.—, and the air outside the aircraft, the pressure and temperature of which are substantially lower—typically 0.2 bar and −50° C.—. For instance, the U.S. Pat. No. 5,482,229 suggests increasing the temperature of the air coming from an outlet channel of the cabin by means of a heat exchanger fluxed by air circulating in a duct coming from an engine compressor of the aircraft and coupled with the ECS system of the cabin. The air coming from the cabin, which has been warmed up through the heat exchanger, then drives a turbine of a power conversion unit which supplies mechanical or electric energy to auxiliary equipments (pumps, supercharger, alternators, etc.), before being discharged outside the aircraft. 
     However such a conformation does not make it possible to use the exhaust air from the cabin in a reliable way. Indeed, the pressure of this air is regulated in the cabin at a certain level, for example at 0.8 bar, and the variations of pressure between the inside and the outside of the aircraft—for example 0.8 bar internally and 0.2 bar externally when the aircraft ascends or is at high altitude—lead to pressure drops and to intrusive phenomena: the regulation cannot be correctly made any more because the pressure in the cabin is higher than the initial regulation value and the pressure transients are unacceptable for the passenger ear. The air cannot correctly flow out any more because the turbine causes all the time back-pressures blocking the air at the outlet of the cabin. In these conditions, the turbine of the conversion unit cannot be operational any more, in particular during the transient phases of ascent to altitude and high altitude. 
     Furthermore, the heat exchanger is not operational any more on the ground when the cabin door is open. This architecture requires then a heat installation with an additional heat exchanger coupled with an outside air circuit. 
     Besides, in the event of a failure of an equipment driven by the conversion unit, the latter runs into overspeed. 
     Furthermore, the use of air coming from an engine compressor of the aircraft is disadvantageous in terms of energy balance, due to the loss in the pipes because of the distance between the heat exchanger and the engine outlet. Furthermore, the power supplied by the engines to the ECS system during takeoff is overestimated with regard to its energy requirements. The sizing point of the supply of power to the ECS system is actually determined at minimal speed of the HP (high-pressure) body of the main engine, so that it is always capable of supplying the sufficient power to the ECS system—even at idle speed—. 
     Generally, main engines are sized so that they are able to supply, from time to time, an important propulsive power, for example at takeoff of the aircraft, i.e. when the HP body is at high speed, while in the other phases they supply a medium propulsive power, indeed minimal, for example in descent, i.e. when the HP body runs at a low speed. Propulsive power relates essentially to the thrust supplied by the jet engines and to the mechanical power supplied by airplane turboprops and helicopter turboshaft engines. This oversizing of power supply is generally accompanied by a specific overconsumption, in all flight phases apart from the idle. 
     DISCLOSURE OF THE INVENTION 
     The invention aims precisely at limiting the specific consumption by matching the sizing of the power supply and the actual power requirements of the cabin ECS system and more generally of the aircraft, so as to remove the useless power supplies. 
     The invention also aims at supplying energy in a sufficiently reliable way to face the cases of aircraft failure which might induce overspeeds. Another purpose of the invention is to favour the association of a high number of non-propulsive energy-consuming means, in particular the electric, mechanical and/or hydraulic consumers, in order to keep in all flight phases a positive overall energy balance between energy supply and consumption with regard to the known conformations, in particular in transient phases. Furthermore, the present invention is going to make it possible to recover thermal energy on the outlet side of the cabin without any risk of back-pressure that is harmful to the regulation, with an optimized heat exchange. 
     To do this, the invention consists in supplying energy near the cabin outlet, in particular pneumatic energy to the cabin, by means of an engine-type power-generating means. A power-generating means is said to be of engine type when the architecture of this power-generating means is fit for the certification as engine usable during all flight phases, in the same way as a power-generating means serving as a main engine. 
     More precisely, the object of the present invention is a method for optimizing the overall efficiency of the energy supplied aboard an aircraft, this energy being propulsive or non-propulsive, the aircraft being equipped with a passenger cabin with regulated airflow, and with power sources including the main engines. Such an optimization consists in providing, in an environment located near the cabin, at least one engine-type main power-generating means sized so as to serve as single other pneumatic-energy generating source for the cabin and at most partly as other propulsive, hydraulic and/or electric energy-generating source for the rest of the aircraft, and in minimizing the power difference between the nominal point of the power sources when said sources are functioning and the sizing point of the non-propulsive energy contributions of said sources in a situation of failure of a main engine, namely by equally dividing the power contributions of the main engines and main power-generating means under nominal operating conditions and in the event of a failure of a main engine. 
     The main power-generating means makes it possible to adjust the supply of pneumatic energy according to the strict requirement of the cabin, whereas main engines needlessly supplied a power which was substantially higher than the bare minimum, typically twice higher: they have been oversized as far as the pneumatic-energy balance is concerned because their sizing is based on the minimal speed of the main engine HP body. The supply of pneumatic energy being not a matter for the main engines anymore according to the invention, they have a substantially improved efficiency and the overall efficiency also is then substantially improved. 
     Besides, the overall thermal efficiency of a main power-generating means that has been so sized is substantially equal to that of the main engines for the non-propulsive power supply, in descent phase or in nominal flight phase, typically of the order of 20%. An equally dividing of the amounts of electric power is then applied without any significant detriment to consumption. A contrario, in ascent phase, supply of electric energy by the main engines will be preferred because the efficiency of the main engines is higher due to the fact that the speed of the high-pressure body (HP) is higher than that of a main power-generating means. 
     Furthermore, the contribution of an additional main power-generating means offers a redundancy of engines means and thus strengthens the fault tolerance and the availability of the aircraft. 
     The invention also relates to a main power unit, hereinafter: MPU, capable of optimizing the overall energy efficiency according to the above method. Such a main power unit is based on a power unit of the auxiliary power unit type, in an abbreviated form: APU, which has been made more reliable, in order to belong to the engine category, and combined with an energy-recovery structure. 
     APUs usually fit aircrafts in order to feed the various energy-consuming equipments (electric, pneumatic and hydraulic power, air conditioning) on the ground, and start the main engines. When an engine is out of order, some APUs have been sufficiently secured so that they are able to start up again for trying to restart the failing engine during the flight and/or to supply part of the electric energy to the equipments in flight. 
     APUs typically consist of a gas generator—including at least an intake compressor, a combustion chamber and at least one power turbine—as well as means for driving the equipments (supercharger, fuel and hydraulic pumps, electric generator and/or electric starter/generator, etc.) directly or via a power-transfer box with rotational-speed adaptation. An air bleed at the outlet side of the supercharger or intake compressor is used for pneumatically starting the main engines. 
     The use of an APU, even secured, during all the flight phases to supply non-propulsive energy is considered as unrealistic because of an unfavourable energy efficiency in comparison with the main engines: operating an APU during the whole flight duration means additional fuel consumption. 
     Now, if the APU is converted into an engine-type power unit for permanently supplying pneumatic energy according to the strict requirement of the cabin, then an aircraft having such a unit offers a favourable balance. 
     As such, in an aircraft including energy-consuming equipments, in particular a cabin the air of which is renewed and the temperature and/or pressure of which are regulated by means of a regulation system ECS, main power-generating engines and a flight control unit, a main power unit according to the invention built into a compartment which is insulated from the other zones of the aircraft with a fireproof bulkhead and fitted with an outside-air intake and an exhaust nozzle, includes an engine-type power unit of the above described type fitted with a gas generator and with a power turbine for driving equipments including a supercharger. The supercharger is coupled, via a regulation control which communicates with the control unit, with the ECS system in order to supply the necessary pneumatic energy to the cabin. 
     According to particular embodiments:
         the main power unit is coupled with a recovery structure including at least one energy-recovery turbine for driving the equipments with the power turbine and coupled, on the air-inlet side, with the outlet of the cabin to cool, on the air-outlet side, the equipments, the supercharger being built into this recovery structure as a supplier of pneumatic energy to the cabin;   the supercharger includes a variable-pitch air diffuser having blades, the adjustment of which is servo-controlled by the regulation control, capable of strictly adjusting the air flow to the supply of pressure and flow rate required by the ECS in every flight phase;   a variation in the setting of the diffuser of the supercharger results in a variation in the air-flow rate with a substantially constant pressure ratio: so, the balance between need and supply is met without significant wasting;   the supercharger is directly coupled with the power turbine to avoid any loss of energy due to a transfer of power other than a mechanical transfer;   the gas generator includes an intake compressor which can serve as a supercharger;   the recovery turbine is a turbine, preferably centripetal, with variable-pitch guide vane assembly having blades the orientation of which is servo-controlled by the regulation control;   at least one pressure sensor regulates the opening and closing of the blades of the diffuser and guide vane assembly in connection with the servo-control;   the recovery turbine ejects, on the outlet side, an air flow into the compartment of the main power unit which, after it has cooled the equipments and auxiliary equipments contained in the aft compartment, is evacuated into the exhaust nozzle by a jet pump action resulting from the efflux velocity of the hot air flow coming out of the power turbine;   the recovery turbine is coupled with a soundproofing device in order to avoid the propagation of the wind noises into the cabin;   the most open possible setting positions can go beyond full opening into radial position, i.e. the so called zero position;   a regulation of the variable setting, between full opening on the ground and progressive closing of the air flow while gaining altitude, can be automated by means of the regulation control according to the pressurization in the cabin.       

     Generally, the fact was taken into account that the loss of energy supply capacity of the main unit, which increases with height, should be at least partially compensated in flight by an optimization of the positions of the variable settings of the recovery turbine in the most closed position compatible with back-pressures at the outlet side of the cabin and of the supercharger in the most open possible position. 
     The level of thermodynamic power compatible with the in-flight stresses for the main unit is minimized: even if, on the ground, the appropriate positions of the variable settings penalize the efficiency of the recovery turbine and supercharger, the main power unit the thermodynamic power of which has been sized in that way is then capable of supplying enough energy on the ground. So, optimizing efficiency in flight was preferred. In the whole flight envelope, the overall efficiency of the compressor and recovery turbine is thus optimal thanks to the presence of a diffuser and/or a guide vane assembly with variable settings. 
     According to other advantageous embodiments:
         means for transmitting power from the power and recovery turbines to the mechanical, pneumatic, hydraulic and/or electric equipments of the aircraft are provided, in particular in the form of a power-transfer box;   the recovery structure comprises a heat exchanger having two heat-transfer circuits: a primary circuit connected, on the inlet side, with the hot-air-flow outlet of the power turbine and, on the outlet side, with the exhaust nozzle; and a secondary circuit connected, on the inlet side, with an air-flow outlet of the cabin and, on the outlet side, with the recovery turbine;   the variable-pitch guide vane assembly of the recovery turbine, coupled with means of regulation, is capable of orienting the air flow coming from the heat exchanger, in particular during transient phases of the aircraft as well as at altitude—transient phases relating to the phases of takeoff, ascent, descent and landing—.       

     In these conditions, energy recovery on the outlet side of the cabin—in the form of pressure and/or temperature—is optimized thanks to the proximity to the main power source, while ensuring an air outflow on the outlet side of the cabin with a controlled back-pressure in the cabin. Besides, connecting the energy recovery means to a main power-generating source, and not to a mere compressor or an alternator, makes it possible to absorb overspeeds that can start in the event of a failure thanks to the inertia resulting from the mass effect due to the components of the power-generating source and all the consumers. 
     Furthermore, recovering energy on the outlet side of the cabin can be undertaken by supplementing the potential energy contained in the air outflow from the cabin by thermal energy used to cool systems dedicated to aircraft equipments before being further enriched by heat exchange between the aforementioned air flows. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Other aspects, characteristics and advantages of the invention will appear in the following non-restrictive description of particular embodiments, in reference to the accompanying drawings which show respectively: 
       in  FIG. 1 , a diagram of an example of a main power unit according to the invention in an aircraft aft compartment, in connection with an aircraft cabin fitted with an environmental control system ECS; 
       in  FIG. 2 , a schematic sectional view of an example of a MPU centripetal recovery turbine provided with a variable-pitch guide vane assembly; 
       in  FIG. 3 , a schematic sectional view of an example of a MPU supercharger provided with a variable-pitch guide vane assembly, and 
       in  FIG. 4 , a graph of the power supplied to an aircraft, according to the thermal efficiency of the power sources, on which the nominal point and the sizing point are shown. 
     
    
    
     DETAILED DESCRIPTION OF EMBODIMENTS 
     In all the Figs., identical or similar elements having the same function are identified with identical or related reference marks. 
     In reference to  FIG. 1  showing a schematic diagram, a main power unit  1  is arranged in an aft compartment  2  situated in the downstream part of the aircraft  3 . The passenger cabin  4  is situated upstream and coupled with the aft compartment  2  via an intermediate compartment  5 . A pressure bulkhead  6  separates the cabin  4  from the intermediate compartment and a fireproof bulkhead  7  insulates the intermediate compartment  5  from the aft compartment  2 , which is fitted with an outside-air intake  21  and an exhaust nozzle  22 . 
     The main power unit  1  includes an engine  10 , of the APU type but of the engine category, combined with an energy-recovery structure. The auxiliary engine consists of: a gas generator or HP body  11 , including an intake compressor  110  for an air flow F 1  coming from the air intake  21 ; a combustion chamber  111 ; and a turbine  112  for driving the compressor  110  by means of a HP shaft  113 . This gas generator is coupled, on the inlet side, with an air-flow duct K 1  mounted on the outside-air intake  21  and, on the outlet side, with a power turbine  12  which produces a hot air flow F 2 , typically of about 500 to 600° C. 
     The energy-recovery structure is centred on a recovery turbine  13  in connection with a soundproofing device  14 , in order to avoid the propagation of the wind noises outside the compartment, in particular into the cabin. 
     This recovery turbine  13  is coupled with the power turbine  12  for driving equipments  100 —mechanical, pneumatic (compressors), electric (alternators) and/or hydraulic (pumps)—especially a supercharger  15  and a starter/generator  16 , via a power-transfer box  17  in the example. This box  17  is fitted with gearboxes and bevel gears (not shown) suitable for power transmission. The power turbine  12  supplies its power to the box  17  via a shaft  121 , i.e. a through-going shaft in the illustrated example. Alternatively this shaft can be a non-through-going shaft or an outside shaft via an appropriate box of reduction (not shown). This box is advantageously fitted with a freewheel intended for its disconnection in the non-recovery phases (for example in the case of an open airplane cabin door). 
     The supercharger  15  supplies an environmental control system, called ECS system,  41  of the cabin  4  with air and transfers to it, via a recycling mixing valve  42 , compressed air coming from the outside-air intake  21  through a branch K 11  of duct K 1 . The supercharger  15  is regulated by a regulation control  19  which communicates with the control unit (not shown) so as to supply the necessary pneumatic energy to the cabin. As a variant, the intake compressor  110  can serve as a supercharger  15  by appropriately bleeding air. 
     At least one variable valve  40 , called cabin-pressure-regulation valve, circulates air flow F 3  from the outlet  43  of the cabin  4  to the energy-recovery structure via duct K 2 . Advantageously, duct K 2  goes into the intermediate compartment  5  so that air flow F 3  cools the power electronics  50  inside a cabinet  51 —these auxiliary equipments being dedicated to various systems made for the functioning of the aircraft (landing gear, etc.), which, of course, are non-operational when the cabin door is open—. At the outlet of the compartment  5 , air flow F 3  has a temperature about 40° C. The variable-pitch guide vane assembly can advantageously replace the pressure-regulation valves at the cabin outlet. 
     The recovery structure comprises, in this example, a heat exchanger  18  fitted with a primary circuit C 1 , connected, on the inlet side, with the outlet of hot air flow F 2  and, on the outlet side, with the nozzle  22 —the temperature of flow F 2  being then typically reduced from ca. 550° C. to 300° C.—and with a secondary circuit C 2  connected, on the inlet side, with air flow F 3  coming from the cabin  4  and, on the outlet side, to the recovery turbine  13 . Flow F 3  has then a temperature substantially higher than at the inlet (approximately 40° C.), for example of the order of 150° C. At the outlet of the recovery turbine  13 , air flow F 3  is dispersed in the aft compartment  2  in order to cool the equipments  100  (down to approximately 40° C.) and then collected in the form of flow F 3 ′, by reflection on walls  200  of the compartment, into the nozzle  22 . Collection takes place because of a jet pump action, at the widened intake  221  of this nozzle, resulting from the efflux velocity of hot air flow F 2 , coming from the power turbine  12 , at the outlet of the heat exchanger  18 . 
     The recovery turbine  13  is explained in detail in reference to the schematic sectional view of  FIG. 2 . The recovery turbine is a centripetal turbine fitted with a ring chamber  131  for bringing in air (flow F 3 ). This air is then directed by the variable-pitch guide vane assembly  136 . The turbine  133  has a stator blading  132 . Outlet-side air flow F 3  is acoustically processed and distributed in the aft compartment  2  so that it controls the temperature of the equipments  100  and other non-shown auxiliary equipments (fire, jacks, etc.). Alternatively, other types of turbines can be used: axial or reaction-impulse (inclined). 
     The guide vane assembly  136  is composed of variable-pitch mobile blades  134  which guide and accelerate the air flow coming from the heat exchanger  18 . These blades have a variable pitch and their orientation is adjusted by the regulation control  19  during the transient phases of the aircraft as well as at altitude. In operation, a pressure sensor  135  regulates the opening and closing of the blades  134  of the guide vane assembly  132  in collaboration with control  19 . 
     The supercharger  15  is explained in detail hereinafter in reference to the schematic sectional view of  FIG. 3 . This supercharger has a structure which is similar to that of the recovery turbine but inverted with regard to the circulation of air flow F 1 : ring chamber  151 —variable diffuser  156  with mobile blades  154 —and a centrifugal compressor  153  fitted with fixed blades  152 . The variable-pitch mobile blades  154  are piloted by the regulation control  19 , in particular during the transient phases and at altitude. A pressure sensor  155  regulates the orientation of the blades  154  via the control  19  in order to meet the characteristics defined by the ECS system, namely an air-flow rate  151  adjusted to the required supply of pressure and flow rate (arrow F 1 ). 
     In a concrete example, the pneumatic-power need for the ECS system of a standard airplane is typically 180 kW. A main engine is sized to supply these 180 kW at idle speed whereas in normal operation it produces 360 kW in the quasi-totality of the flight phases. A main power unit according to the invention is thus sized to supply the 180 kW of pneumatic power that are strictly sufficient to meet the needs of the ECS system. 
     The power supply by the main power unit according to the invention is not limited to the supply of pneumatic energy. This unit can indeed supply power to the HP body of the main engines via the starter/generator  16  used as an electric generator coupled with the starter/generator of the main engines used in driving mode. 
     So, with a global need for power of typically 420 kW—i.e. 180 kW of pneumatic power for the ECS system, 60 kW of hydraulic power for the jacks and 180 kW of electric power for the alternators, pump, etc.—the use of a supercharger, a recovery turbine and/or a heat exchanger according to the recovery structure of the invention makes it possible to substantially lower the loss of energy which would be generated by the exclusive use of main engines to carry out these functions. For instance, a supercharger with a variable-pitch diffuser makes it possible to save 180 kW, a variable-pitch recovery turbine typically 90 kW and a heat exchanger from 15 to 20 kW, i.e. 285 to 290 kW altogether. The main engines contribute then only one third to the total of these power supplies (420 kW), pneumatic power excepted (180 kW), i.e. approximately 80 kW, that is to say a substantially lower supply than that of the main power unit which supplies, in this example, 150 kW (70 kW plus one third of the remaining 240 kW, i.e. 80 kW, to supply pneumatic and electric/hydraulic energy respectively). 
     Considering an efficiency of the main power unit (typically 20%) which is similar to that of a main engine in the flight phases other than ascent or failure of one of the engines and lower than that of the main engine (40%) in full use (ascent or the other engine out of order), an equally dividing of the supply of energy between the engines, whether it is a main engine or the main power unit, makes it possible to optimize the overall efficiency covering all the flight phases, under nominal operating conditions or in the event of a failure: for example, the equally dividing of the supply of hydraulic and electric power is ⅓, ⅓, ⅓ for two main engines and a main power unit in operation, and ½, ½ in the event of a failure of a main engine. 
     Furthermore, the equally dividing makes it possible to optimize the efficiency of all the power sources forming a turbine engine as shown, in  FIG. 4 , by the graph G representing the variation in the thermal efficiency dependent on the power Pw supplied by an engine. On this graph, we can see:
         the power sizing point (Pd) 0  of the turbine engine: this sizing point is established in the most severe conditions of need for power (generally in the case of failure of an engine or a particularly difficult takeoff);   the nominal point (Pn) 0  of the turbine engine without the main power unit, and the nominal point (Pn) 1  of the turbine engine with the main power unit with equally dividing;       

     The variation in the thermal efficiency related to the consumption of fuel is optimized when the turbine engine includes the main power unit, namely for the following reasons. Without main power unit, efficiency variation D 0  between points (Pn) 0  and (Pd) 0  is higher than variation D 1  between points (Pn) 1  and (Pd) 0  when the aircraft includes a main power unit, but with substantially lower amounts of power supplied. This situation is the expression of the optimization obtained with the equally dividing by minimizing the difference between the nominal point and the sizing point. Indeed, the first D 0  corresponds to the transition from 50 to 100% (corresponding to 200% to be supplied in the event of a failure) of power supplied by an engine going from nominal conditions to sizing conditions, i.e. a difference of 50%. The second variation D 1  corresponds to the transition from 33% (more exactly ⅓) to 50% in order to go from the first type to the second type of conditions. With a main power unit, the turbine engine shows a decrease of the power to be supplied of ⅓, i.e. 33% for all the main engines, with an overall efficiency (corresponding to the efficiency variation) increased by the difference (D 0 −D 1 ). This example does not take into account the possibility of load shedding which can be applied to the cases of failure. Whether with or without load shed, the efficiency is improved. 
     The above statement refers to the functioning of a main power unit. The case of failure of this unit has not been evoked but, should that arise, it is of course possible to provide for other emergency equipments which can substitute for this unit, for instance in degraded mode, in particular: at least one of the two main engines which will then supply an additional power, or a spare APU or equivalent, or a combination of these sources. 
     Besides, the equally dividing which is evoked in the present statement implies that the power sources have been conceived to enable such an equally dividing in the set out conditions. The statutory constraints and physical stresses, in particular mechanical, to be taken into account generally make it only possible to strive as far as possible towards the ideal conditions for equally dividing.