Abstract:
A method for optimizing operability of an aircraft propulsive unit, and a self-contained power unit implementing the method. The method removes mechanical bleed constraints in engines during transient flight phases of an aircraft to optimize operability of the engine assembly during the phases. To this end, a supply of power is provided, particularly during the phases, by an additional indirectly propulsive engine power source. The method for optimizing operability of the propulsive unit of an aircraft including main engines as main drive sources includes, using a main engine power unit GPP as a power source, providing all the non-propulsive power and, during the transient engine phases, at most partially providing additional power to the body of the main engines.

Description:
TECHNICAL FIELD 
     The invention relates to a method for optimizing the operability of the engine set of an aircraft, as well as a main power unit capable of implementing such a method. 
     The invention applies to the engine set of aircrafts, that is essentially to the engine set of airplanes (jet engines, turbojet engines, turboprops) as well as to the engine set of helicopters (turboshaft engine). 
     An aircraft engine comprises, typically and in a simplified way, a compressor-chamber of combustion-turbine set forming a gas generator. After their combustion, the hot gases expand in the turbine which mechanically drives the compressor via a high-pressure (HP in an abbreviated form) shaft or HP body, and they supply potential energy. 
     In the case of an airplane, this potential energy generates the propulsion energy in the form of thrust, either directly (in jet engines) or indirectly via the LP low-pressure body (in turbojet engines with fan or turboprops with propeller). In the case of a helicopter, this propulsion energy is transmitted to the rotary wing through a main transmission gearbox (generally known as MGB). Turbine engines produce energy within the time corresponding to a level of instantaneous power transmission. 
     The flows of air in the compressor and the turbine can cause, in certain operating conditions, a phenomenon referred to as compressor surge which makes the hot gases from the gas generator go back towards the air inlet of the compressor and can lead to the most serious consequences (sudden drop in the lift, reversal of thrust, breaking of the blades, destruction of the engine). It is thus necessary to keep some surge margin. Such a phenomenon is to be banned in the aeronautical field. 
     For each flight phase, a surge line can be drawn dependent on the air intake/outlet pressure ratio and the air-flow rate. The working line of the engine must stay below this surge line in order to avoid, in particular, any thrust loss. The difference between the working line and the surge line, called surge margin, decreases when the HP-body speed is low. 
     The surge margin decreases all the more so as mechanical bleeds are taken from the HP bodies to feed the electric and hydraulic equipments (alternators, pump, etc.). The present specifications tend towards a substantial increase in these bleeds. Now a sufficient surge margin must make it possible to accelerate the HP body when the flight circumstances require to re-throttle up. 
     A similar functioning is expected as far as helicopter turboshaft engines are concerned. However, in the case of a failure of an engine (“One Engine Inoperative”, OEI in the abbreviated form), a rapid acceleration is required from the HP body of the sound engine. 
     STATE OF THE ART 
     Generally, the importance of the needs for mechanical bleeds limits the acceleration capacities of the engines during transient phases, i.e. the operability of the power source. To obtain the intended accelerations, it would be useful to increase the surge margin by lowering the working line. It would be also possible to bleed air from the main engine. But then, in either case, the overall efficiency of the turbine engine is substantially modified. 
     The aircraft being an isolated system, the only solution is to temporarily limit the requests for mechanical bleed from the HP body. But the consequences can be harmful to the equipment and to the functions fulfilled by the equipments (cabin air conditioning, landing gear, etc.). 
     DISCLOSURE OF THE INVENTION 
     The invention aims at removing the constraint of mechanical bleed in the engines during the transient phases in order to optimize the operability of the engine group during these phases. To do this, power is supplied, especially during these phases, by an additional, engine-type, power source, supplying indirectly-propulsive power. A power-generating means is said to be of the engine type when the architecture and the performance of this power-generating means are fit for the certification as engine usable during all flight phases, in the same way as a power-generating means serving as an aircraft main engine. 
     More precisely, the object of the present invention is method for optimizing the operability of the engine set of an aircraft including main engines as main drive sources. Such an optimization consists, by means of an engine-type main power source as a drive source, in supplying the totality of the non-propulsive energy and, during the transient phases of the engines, in supplying at most partially an additional power to the high-pressure body HP of the main engines. The transient phases of the engines are especially acceleration phases, failure cases and functioning at idle speed. 
     According to preferred embodiments, the power supplied to the HP body of the main engines is produced by an electric generator fitting the main power source in collaboration with an electric starter of the main engines converted into a drive—like during the ground start of the main engines—or by ways of a bleed of compressed air from the main power source in collaboration with an air starter of the main engines. 
     In particular, the main power source can provide power to the HP body of the main engines in order to obtain higher acceleration rates with, possibly, an idle adjusted at a lower level than the nominal idle. 
     More particularly, in airplane use, in the phase of flight descent as a transient phase, the main power source supplies power to the HP body of the main engines. An optimized operability is then achieved, in terms of capacity of maximized acceleration, by an increase of the surge margin of the main engines in a flight phase where this margin would be minimal without supply of additional power, with a working line at the nearest to surge. 
     More particularly, in helicopter use, in the event of a failure of a main engine, the main power source supplies the electric power to the HP body of the sound main engine so that the latter can have an acceleration capacity such that its surge margin is sufficient. Consequently, the surge margin is thus preserved while the operability has become optimal. 
     Thus, in aircraft use, in steady-state phase as well as in transient phase, the main power source supplies power to the HP body of the main engines. 
     The invention also relates to a main power unit, hereinafter: MPU, capable of optimizing the operability of the engine set of an aircraft 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 means for regulating the compressed air in the cabin. 
     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 and 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 the flight phases to supply non-propulsive energy should be considered as unrealistic because of an unfavourable energy efficiency in comparison with the main engines: operating an APU during the whole flight duration is fuel-consuming. 
     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 and makes it possible to optimize the operability of the engine set of the aircraft. 
     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 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 that includes an energy-recovery turbine for driving the equipments with the power turbine and that is coupled, on the air-inlet side, with the outlet of the cabin in order 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 recovery turbine ( 13 ) 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 wind noises into the cabin;   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.       

     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 the 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: 
         FIG. 1 , a graph of the variations of the working line of an aircraft engine; 
         FIGS. 2 a  and 2 b   , two block diagrams of the distribution of the propulsive and non-propulsive energies, notably in transient phase ( FIG. 2 b   ), supplied by the main engines and the main power unit of an aircraft in nominal flight; 
         FIG. 3 , a block diagram of the distribution of the energies in the event of a failure of a helicopter engine; and 
         FIG. 4 , 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. 
     
    
    
     DETAILED DESCRIPTION OF EMBODIMENTS 
     In all the Figs., identical or similar elements having the same function are identified with identical or similar reference marks. 
     In reference to  FIG. 1 , the graph of the variations in the working line LF of an aircraft engine is showed in a system of reference of the air pressure ratio P/P dependent on the adjusted air-flow rate D for a given flight phase. The air-flow rate is said to be adjusted in order to enable a significant graphic representation including the influence of the various bearing parameters. A surge line LP and the working lines LF, LF 2  of the engine are shown in this system of reference. The working line LF stays below this surge line LP in order to avoid any thrust loss. The surge margin MP, i.e. the difference between the working line and the surge line, decreases with the speed (or air-flow rate) of the engine HP body, for example from the maximal speed N M  down to the idling speed N R  which are permissible for this flight phase. 
     A reduced surge margin increases the efficiency of the engine but can entail risks of surge if the working line gets too close to the surge line. For example, during an acceleration starting from the idling speed N R , the working points Pf 1  in transient describe, on the graph, the working line LF 1  from N R  to N M . The decrease of the surge margin MP along this line LF 1  is due to the fuel injection into the combustion chamber necessary for the acceleration of the HP body. The position of point Pf 1  showed on the graph corresponds to the minimal surge margin. Mechanical bleeds from the HP body to feed equipments (arrow PM) also reduce the surge margin. The present specifications tend towards a substantial increase of said bleeds which substantially increase the risks of surge. 
     Overshoots during accelerations is thus delicate to handle. Besides, air bleeds (arrow PA), for instance at the level of the engine compressor, in order to supply energy to other equipments (cabin air conditioning, etc.) increase the surge margin. The working line migrates then from line LF to line LF 2 , this migration resulting in a loss of efficiency at constant flow rate as shown by arrow P R . 
     An optimized operability is achieved by a supply of power from a main source, in terms of maximized capacity of acceleration: such a supply makes it possible to increase the power provided to the HP body shaft by simultaneously supplying power by fuel injection into the chamber of combustion and power from the MPU. This additional supply increases the acceleration rate of the HP body, while maintaining the surge margin MP of the main engines in a flight phase whereas this margin would at its minimum without the supply of additional power, with a working line LF at the nearest to surge. Furthermore, this additional supply makes it possible to reduce the idling speed N R  while the surge margin stays identical in transient phase. Besides, such a supply makes it possible, in steady-state phase, to maintain an idle level N RO  lower than that determined by the capacity of autonomy of the gas generator. 
     In the absence of a main power unit, each main engine supplies, equally in nominal conditions, propulsive energy Ep and of non-propulsive energy Enp. As it appears in detail hereinafter, a MPU can distribute the totality or part of the non-propulsive energy between each main engine and the MPU. This MPU also provides a few percent of propulsive energy while supplying power to the HP body of the main engines in transient phases, in particular in the event of a failure a main engine. 
     In reference to  FIG. 2 a   , relating to an initial state in steady-state flight of an aircraft, the non-propulsive energy Enp is supplied to the equipments  100  by the main engines  200 , MP 1  and MP 2 , as well as by the MPU  1  according to an equally dividing Enp/3. This equally dividing is also recommended to define the sizing points of the turbine engines. In a phase of descent, during which the main engines are not needed, it is preferable that the non-propulsive energy should be mainly or totally supplied by the MPU. 
     As a variant ( FIG. 2 b   ), the non-propulsive energy Enp is supplied only by the MPU  1  in nominal flight in order to enable the switching means (in particular, the electric switches) to have enough response time in the event of a failure of an engine. Indeed, response times during accelerations to be supplied by a single engine can be insufficient if said engine does not mobilize its whole power (see below the case of engine failure). 
     Besides, the thrust  300  (2xEp) in nominal flight is equally supplied by each of the main engines. The MPU  1  can provide a proportion of propulsive energy kxEp by supplying power to each of the engine HP bodies, where k can amount a few percentage points. 
     In the event of a failure of an engine MP 2  in an aircraft, in particular a helicopter fitted with a MPU  1 , as shown in the example of  FIG. 3 , the remaining sound MP 1  engine  200  is at first configured to supply the whole propulsive energy: its share in the supply of non-propulsive energy Enp to the equipments  100  goes then from Enp/3 (no failure of the other engine, see  FIG. 2 a   ) down to 0xEnp, and its supply in propulsive energy goes from Ep up to 2xEp in order to produce the whole thrust  300 . The MPU  1  is then configured to supply all the non-propulsive energy Enp, while increasing from Enp/3 to Enp. Advantageously, the MPU  1  continues to supply a proportion kxEp of propulsive energy to the HP body of the sound engine with an adjusted coefficient k and an acceleration rate such that the surge margin of the sound engine should be sufficient. In transient, the adjusted coefficient has a value k t  and the MPU supplies then a transient propulsive energy k t Ept to the HP body of the sound engine MP 1  which supplies the whole thrust 2xEpt. 
     In the case of an aircraft in transient flight phases ( FIG. 2 b   ), particularly an airplane, the MPU  1  provides transient propulsive energy k t Ept, with a coefficient k t , to the HP body of each of the main engines  200  supplying 1xEpt, 2xEpt being the total propulsive energy to be supplied in transient phase. The MPU still supplies the totality of the non-propulsive energy, i.e. Enp, beyond the energy 2k t Ept. 
     The power supplied to the HP body of the main engines is produced by the generator, in this case an alternator, fitting the MPU in collaboration with the electric starter of the main engines converted into a drive, like during the ground start of the main engines. 
     In particular, the MPU provides power to the HP body of the main engines so as to carry out accelerations with an idle level adjusted at the lowest. 
     In reference to  FIG. 4  showing a schematic diagram, a MPU  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 MPU  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 thereto, 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 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 .