Aircraft installation for supplying pressurized air

Abstract

An aircraft installation of a pneumatic system of an aircraft with different compressed air sources for supplying pressurized air to air consumer equipment. Either an air bleed system, electrical compressors, or a combination thereof may perform such supplying of compressed air depending on the aircraft operation condition, for instance the flight altitude or specific flight phases. Also, a turbofan engine, and a method for supplying compressed air to air consumer equipment are disclosed.

Claims

1. An aircraft installation for supplying pressurized air to an air consumer equipment of an aircraft, the aircraft installation comprising: a gas turbine engine having a single port located at a low-intermediate compressor stage of such gas turbine engine; a bleed duct, at one end in fluid communication with the single port of the gas turbine engine, and at an opposite end in fluid communication with the air consumer equipment; a cooling duct comprising an inlet by which cooling air enters into the cooling duct, and an outlet; a branch duct in fluid communication with the bleed duct and configured to divert a portion of bleed air from the bleed duct into the outlet of the cooling duct; and an electrical compressor with an intake and an outlet, so that the intake is in fluid communication with the outlet of the cooling duct after the joining with the branch duct, and the outlet is in fluid communication with the bleed duct between the branch duct and the opposite end of the bleed duct which is communicated with the air consumer equipment.

2. The aircraft installation according to claim 1, wherein the inlet of the cooling duct is an air scoop arranged so as to substantially face the incoming cooling air entering into the cooling duct.

3. The aircraft installation according to claim 1, wherein the inlet of the cooling duct is an air scoop arranged so as to be angled relative to the incoming cooling air.

4. The aircraft installation according to claim 1, wherein the intake of the electrical compressor is connected to the outlet of the cooling duct after the joining of the branch duct by a substantially 90° elbow.

5. The aircraft installation according to claim 1, wherein the intake of the electrical compressor is connected to the outlet of the cooling duct after the joining of the branch duct by a substantially 45° elbow.

6. A turbofan engine comprising: a nacelle, and a fairing arranged internally to such nacelle for covering a gas turbine engine, a fan positioned in a foremost part of the nacelle and connected to the gas turbine engine by a shaft, wherein the fan extends up to a diameter of the nacelle; so that a secondary zone is formed between the nacelle and the fairing for passing cooling air from the fan; a pylon for hanging the entire turbofan engine from a wing of an aircraft; and an aircraft installation for supplying pressurized air to an air consumer equipment of the aircraft, wherein the aircraft installation is installed within the pylon, wherein the gas turbine engine has a single port located at a low-intermediate compressor stage of such gas turbine engine, and the aircraft installation comprises: a bleed duct, at one end in fluid communication with the single port of the gas turbine engine, and at an opposite end in fluid communication with the air consumer equipment; a cooling duct comprising an inlet by which cooling air enters into the cooling duct, and an outlet; a branch duct in fluid communication with the bleed duct and configured to divert a portion of bleed air from the bleed duct into the outlet of the cooling duct; and an electrical compressor with an intake and an outlet, so that the intake is in fluid communication with the outlet of the cooling duct after the joining with the branch duct, and the outlet is in fluid communication with the bleed duct between the branch duct and the opposite end of the bleed duct which is communicated with the air consumer equipment.

7. The turbofan according to claim 6, wherein the inlet of the cooling duct is an air scoop arranged so as to substantially face the incoming cooling air entering into the cooling duct; wherein the inlet of the cooling duct is arranged at the foremost part of a pylon bifurcation within the secondary zone formed between the nacelle and the fairing so that the cooling air entering the cooling duct comes from the fan.

8. The turbofan according to claim 7, wherein the pylon is covered by an upper fairing between the external nacelle and the wing, and wherein the inlet of the cooling duct is arranged on such upper fairing by a ram air intake.

9. The turbofan according to claim 6, wherein the inlet of the cooling duct is an air scoop arranged so as to be angled relative to the incoming cooling air; wherein the inlet of the cooling duct is arranged angled relative to the incoming cooling air on a pylon bifurcation within the secondary zone formed between the nacelle and the fairing so that the cooling air entering the cooling duct comes from the fan.

10. The turbofan according to claim 6, wherein the inlet of the cooling duct is arranged on a portion of a peripheral ring of the fairing so that the cooling air entering the cooling duct comes from the fan.

11. The turbofan according to claim 6, wherein the inlet of the cooling duct is arranged on the nacelle by a ram air intake.

12. An aircraft comprising at least one turbofan engine according to claim 6, wherein the aircraft further comprises a vapor cycle machine as the air consumer equipment.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

(1) These and other characteristics and advantages of the invention will become clearly understood in view of the detailed description of the invention which becomes apparent from a preferred embodiment of the invention, given just as an example and not being limited thereto, with reference to the drawings.

(2) FIG. 1 shows a schematic graph of a conventional IP-HP air bleed system power delivering in comparison with power required by air consumers.

(3) FIGS. 2a-b show a schematic representation of an aircraft comprising (a) a conventional pneumatic system, and (b) a pneumatic system according to the present invention.

(4) FIGS. 3a-b show an embodiment of an aircraft installation according to the present invention, in lateral and top view, respectively.

(5) FIGS. 4a-b show an embodiment of an aircraft installation according to the present invention, in lateral and top view, respectively.

(6) FIGS. 5a-b show an embodiment of an aircraft installation according to the present invention, in lateral and top view, respectively.

(7) FIGS. 6a-b show an embodiment of an aircraft installation according to the present invention.

(8) FIG. 7 shows a schematic architecture of a pneumatic system with an embodiment of an aircraft installation according to the present invention.

(9) FIG. 8 shows the internal fairing of a turbofan engine, and the internal upper fairing that surrounds the pylon structure on the top part, the so-called upper bifurcation.

(10) FIG. 9 shows the external nacelle of a turbofan engine, and the external upper fairing that surrounds the pylon structure on the top part between the external nacelle and the wing.

(11) FIG. 10 shows a schematic representation of an aircraft mission profile using a hybrid solution for supplying compressed air according to the present invention throughout the flight phases.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

(12) As it will be appreciated by one skilled in the art, aspects of the present invention may be embodied as an aircraft installation, a pneumatic system, an aircraft, or a method for supplying pressurized air to an air consumer equipment.

(13) FIG. 1 depicts a schematic graph (8) of a conventional IP-HP air bleed system power delivering (8.1) in comparison with power required (8.2) by air consumers throughout a complete flight.

(14) As it can be seen, the power required (8.2) by air consumers in kW is compared vs. the power delivered (8.1) by conventional IP-HP air bleed system (in kW). Superimposed to the former, there is an overview of the flight phases (8.3) through which aircraft passes in a complete flight, in particular taking altitude as a reference to place the aircraft in each of such flight phases.

(15) In this exemplary mission profile, there is a mismatch between power supplied by the air bleed system and required by the air consumers both at the beginning and end of the flight, that is, in principle when aircraft is on-ground or close to it below a certain flight altitude.

(16) Left-ordinate axis of the graph indicates power (in kW), while right-ordinate axis indicates flight altitude (in ft.). Finally, abscissa axis refers to flight time (in minutes).

(17) Typical IP-HP air bleed system is conventionally designed as follows:

(18) IP port extracts air during take-off, climbing, cruise, and holding; and

(19) HP port extracts air on-ground, during descent and even holding if IP port is not capable of providing enough air pressure to meet air consumer requirements.

(20) Therefore, in those phases where HP port is extracting air to supply air consumers, there is a significant energy loss as it can be seen by peaks (8.4) in the graph (selected by dashed circles). Those peaks (8.4) represent a power mismatch which entails an energy loss.

(21) This energy loss is because:

(22) The energy delivered (i.e., in terms of compressed or bleed air) by the HP port during ‘holding’ phase is significantly higher than the energy required by air consumers at that time. In other words, HP port is used under these conditions because the energy delivered by the IP port is not sufficient. On the other hand, the energy surplus not used is transmitted to the external air at the pre-cooler, thus increasing its temperature. This heated air is released directly into the atmosphere becoming a loss of energy for the engine while reducing the efficiency.

(23) The energy delivered by the IP port during take-off and climb phases is significantly higher than the energy required to meet the requirements of the air consumers at that flight phases. Likewise, the energy surplus not used is transmitted to the external air at the pre-cooler, thus increasing its temperature. This heated air is again released directly into the atmosphere becoming a loss of energy for the engine while reducing its efficiency.

(24) FIG. 2a depicts a schematic representation of an aircraft comprising a conventional pneumatic system exclusively based on air bleed system (2).

(25) In particular, the aircraft (10) comprises two turbofan engines (7), each hanging from a wing by respective pylons. It is schematically represented the ducting or channeling from the two ports, IP (2.1) and HP, coming from different compressor stages of the gas turbine engines of the turbofans. It is to be noted that valves and other hydraulic equipment are not shown through these figures.

(26) It is shown that bleed ports (IP and HP) are in fluid communication (by channels or ducts (12.1.1, 12.2.1)) with WAIS (6.3) and Air Conditioning Packs (6.1) of the ECS in order to convey pressurized air thereto.

(27) Furthermore, the aircraft (10) comprises an Auxiliary Power Unit (‘APU’) (16) at the tailcone of the aircraft (10). This APU (16) is also in fluid communication (by APU bleed ducting (16.1)) with the WAIS (6.3) and Air Conditioning Packs (6.1) of the ECS in order to provide either pneumatic or electrical energy thereto.

(28) Typical APU bleed ducting (16.1) for pneumatic mode is also associated with OverHeat Detection System for safety reasons.

(29) On the other hand, FIG. 2b depicts an example of a schematic representation of a similar aircraft (10) as the one shown in FIG. 2a but comprising a pneumatic system with an aircraft installation (1) according to the present invention.

(30) Instead of IP-HP port for each turbofan engine as shown in FIG. 2a, the air bleed system (12) according to the present invention only extracts air from a single port, which is in fluid connection with the Air Conditioning Packs (6.1) of the ECS. In this particular embodiment, the Air Conditioning Packs are replaced by Vapor Cycle Machine Packs (6.2) which need lower air pressure in comparison with conventional Air Conditioning Packs (6.1).

(31) Further, two electrical compressors (5) are positioned within the belly fairing of the aircraft (10) along with the Air Conditioning Packs or Vapor Cycle Machine Packs of the ECS.

(32) In particular embodiments, the wing anti-ice system (‘WAIS’) may be electrical (6.4) so the ducting for conveying pressured air is no longer needed. Instead, wiring connections (which are lighter than ducts) should be deployed.

(33) Similarly, APU bleed ducting (16.1) is deleted since pneumatic mode is no longer needed. Only electrical mode for supplying power to the electrical compressor (5), for instance, is envisaged. Besides, other power consumers such as batteries, electrical WAIS (6.4), or the like, may be supplied by the APU (16) running in electric mode or any other source.

(34) Deletion of APU bleed ducting (16.1) (that is, APU only works in ‘electrical mode’) bring the following advantages along:

(35) Significant weight reduction, around 170 kg. (in a short-range aircraft (10)).

(36) Removal of the harmful installation of a high pressure and temperature duct running through the pressurized fuselage.

(37) Deletion of OHDS associated to the APU ducting.

(38) The formerly needed surplus of compressed air provided by the APU (16) is, within the present invention, exclusively provided by the air bleed system (12) through the single port (e.g., IP) after optimization and modelling works. This can be easily done by the person skilled in the art knowing temperature and pressure constraints of the aircraft installation of the pneumatic system, with the aim to meet air consumers (6) requirements acknowledged beforehand.

(39) For example, the combination of electrical WAIS (6.4) and Vapor Cycle Machine Packs (6.2) in the ECS permits reducing by 2 or 3 compressor stages the location of the single port due to low pressure requirement of Vapor Cycle Machine packs above 15000 ft. (8 to 12 psig nominal conditions and up to 14 psig in failure cases).

(40) It is to be noted that, although only WAIS (6.3, 6.4) and ECS (6.1, 6.2) are represented as air consumers (6), other minor air consumers may be used such as: fuel tank inerting system, engine starting system, water and waste, and/or hydraulic reservoirs pressurization.

(41) Also, a control unit (controller 9) is electrically connected to some valves of the aircraft installation or the electrical compressor(s) itself to selectively operate them based on an aircraft (10) operational condition. Therefore, it is allowed that the bleed air coming from the single port or cooling air pass through, or be cut-off or the flow rate be reduced.

(42) Particularly, such aircraft (10) operation condition may be a pre-determined flight altitude, for instance 15000 ft., and/or any of the flight phases seen in FIG. 10.

(43) FIGS. 3a and 3b depict a first embodiment of an aircraft installation according to the present invention, in lateral and top view, respectively. It is shown an aircraft installation (1) for supplying pressurized air to an air consumer equipment (6) of an aircraft (10). Particularly, the aircraft installation is arranged on top of the pylon (7.3), that is, the structure from which a turbofan engine (7) hangs from a wing.

(44) The aircraft installation (1) comprises a bleed duct (2) in fluid communication at one end with a compressor stage of a gas turbine engine of the turbofan engine (7) via a single port (2.1), running through the pylon (7.3) and finally connected at opposite end with the air consumer (6) equipment.

(45) As it was explained before, the single port (2.1) is a low-intermediate pressure port preferably located at the first half of the compressor stages of the compressor (not shown).

(46) Taking the turbofan engine (7) coordinate system, where longitudinal axis is the rotation axis of the gas turbine engine (i.e., substantially parallel to the X-axis of the aircraft in body-axis), the bleed duct extends upwards following the foremost part of the pylon (7.3) up to the wing where it is deviated to the belly fairing of the aircraft (10).

(47) In its path, the bleed duct (2) is intersected by a cooling duct (3) adapted to let cooling air enter therein. In this particular example, the cooling duct inlet (3.1) is immersed in the secondary zone of the turbofan engine (7), that is between the external nacelle (7.1) and the internal fairing (7.2) receiving air already compressed by the fan of the turbofan.

(48) As it can be seen, the cooling duct inlet (3.1) is arranged so as to face the incoming cooling air entering into the cooling duct (3). In particular, this inlet is formed by an air scoop that permits regulating the flow rate entering the cooling duct.

(49) Further, close to the inlet (3.1), the cooling duct (3) surrounds the abovementioned bleed duct (2) forming a thermal contact for exchanging heat. Passing such thermal contact point, the cooling duct (3) extends so as to reach the intake (5.1) of an electrical compressor (5).

(50) The bleed air already cooled down may be diverted into a branch duct (4) that communicates such bleed duct (2) with the cooling duct (3) before the electrical compressor intake (5.1). Therefore, the electrical compressor (5) may receive air from the cooling duct which is either cooling air or bleed air (cooled down or not) diverted through the branch duct (4).

(51) Alternatively, this diversion into the branch duct (4) may be done upstream such thermal point.

(52) Finally, the outlet (5.2) of the electrical compressor (5) communicates with the bleed duct (2) to supply compressed air to the air consumer (6) equipment.

(53) In the top view provided by FIG. 3b, it can be seen that the intake (5.1) of the electrical compressor is connected to the outlet (3.2) of the cooling duct after the joining of the branch duct (4) by a substantially 90° elbow. In other words, the cooling duct (3) has a 90° elbow which permits the outlet (5.2) of the electrical compressor to be properly aligned with the bleed duct (2).

(54) At the outlet (5.2) of the electrical compressor (5), there may be a reverse flow protection valve.

(55) Furthermore, from this FIG. 3b it can be seen the pylon (7.3) bifurcation, and particularly that the aircraft installation (1) is contained within the pylon (7.3) bifurcation, and from FIG. 3a that the installation is supported on the upper part of the pylon structure

(56) FIGS. 4a and 4b depict a second embodiment of an aircraft installation (1) according to the present invention, in lateral and top view, respectively.

(57) This aircraft installation (1) only differs from the one shown in FIGS. 3a and 3b in that the intake (5.1) of the electrical compressor (5) is connected to the outlet (3.2) of the cooling duct (3) after the joining of the branch duct (4) by a substantially 45° elbow.

(58) Therefore, when cooling air coming from the cooling duct (3) or bleed air diverted by the branch duct (4) passes by this 45° elbow instead by a 90° elbow that entails higher pressure losses, as consequence the pressure drop at that elbow is significantly lower, thus permitting to uphold the air conditions feeding the electrical compressor (5).

(59) FIGS. 5a and 5b depict a third embodiment of an aircraft installation (1) according to the present invention, in lateral and top view, respectively

(60) This aircraft installation (1) only differs from the one shown in FIGS. 3a and 3b in that the inlet (3.1) of the cooling duct (i.e., the air scoop) is arranged so as to be angled relative to the incoming cooling air.

(61) That is, the inlet instead of being frontal relative to the pylon bifurcation (7.3.1) (preferably the upper pylon bifurcation), is positioned in a lateral thereof. Therefore, as the upper pylon bifurcation (7.3.1) is substantially oblong, the inlet (3.1) is arranged on any of its opposite larger lateral areas.

(62) Consequently, since the inlet (3.1) of the cooling duct (3) is arranged at a lateral of the upper pylon bifurcation (7.3.1), the cooling duct circumvents the bleed duct (2) and less thermal contact is provided there between.

(63) The pylon bifurcation (7.3.1, 7.3.2) is the fairing located on the secondary flow area between the internal fairing and external nacelle that surrounds the pylon as well as the systems installed on the top.

(64) It is to be noted that this embodiment shown in FIGS. 5a and 5b may comprise a 45° elbow duct connection alike FIGS. 4a and 4b.

(65) FIGS. 6a and 6b depict a fourth embodiment of an aircraft installation (1) according to the present invention. In particular, FIG. 6a depicts an aircraft installation (1) in a lateral view where neither the internal nacelle fairing (7.2) nor the external nacelle (7.1) of the turbofan (7) is shown.

(66) In particular, the attachment between the front attachment of the pylon and the engine core may influence the position of the aircraft installation. Therefore, two options may be envisaged:

(67) ‘fan case mounted’ (upwards). That is, the aircraft installation may be positioned either in the bifurcation below the pylon, in the internal nacelle fairing, or in a combination thereof; or

(68) ‘core mounted’ (downwards). That is, the aircraft installation may be positioned up or within the internal nacelle fairing since the remainder space is occupied by the pylon itself.

(69) For instance, throughout FIGS. 3 to 5, the pylon frontal part is attached to the engine core (option (ii)) so that the pylon structure is located on the lower part close to the engine core and the compressor. Thus, both the bleed and cooling ducting of the aircraft installation are positioned on top of the pylon structure (7.3).

(70) Alternatively, on other engine attaching systems where the pylon frontal part is attached to the engine fan case (7.2), option (i), the pylon structure (7.3) is located on the upper part so that the compressor (5), bleed (2) and cooling (3) ducting are installed in below the pylon (7.3). In this case, the bleed duct (2) does not go through the pylon structure (7.3) before the compressor (5). Instead, it goes though the pylon structure (7.3) after the compressor (5).

(71) Throughout this description, ‘engine fan case’ and ‘internal nacelle fairing’ (7.2) will be understand as equivalent terms.

(72) Particularly in FIGS. 6a and 6b, regardless the attachment of the pylon, the aircraft installation is either inside the internal nacelle fairing or in the bifurcation (i.e., under the ‘fan case mounted’).

(73) In particular, the aircraft installation (1) is positioned within the fairing (7.2) of the turbofan. Accordingly, the inlet (3.1) of the cooling duct (3) is arranged on a portion of a peripheral ring of the fairing (7.2) so that the cooling air entering the cooling duct comes from the fan.

(74) In FIG. 6b, it is pointed out the peripheral ring on the fairing (7.2) where the inlet (3.1) of the cooling duct may be positioned. The size of the inlet may correspond, for instance, to the available size for the air scoop door to be opened within the fairing where such door opens outwards.

(75) Alternatively the air scoop door may be flush-type and therefore opens inwards.

(76) FIG. 7 depicts schematic architecture of a pneumatic system with an embodiment of an aircraft installation (1) according to the present invention. As an example, it may form the hydraulic scheme of the aircraft installation (1) shown in FIGS. 3 to 5. This may be also applicable to FIG. 6, considering that the electrical compressor is in the internal nacelle instead of on top of the pylon.

(77) In particular, the single port (i.e., IP port) (2.1) of the bleed duct (2) is shown connected to a low-intermediate compressor stage of the gas turbine engine. Further, an additional port (3.1) coming from the fan connects with the cooling duct (3).

(78) The intersection of the cooling duct (3) and the bleed duct (2) is schematically represented herein by a box, which may encompass different embodiments such as a thermal contact, etc. However, for illustrative purposes, the branch duct (4) is hidden by such a box.

(79) From such box, a duct (in principle forming part of the cooling duct (3)) extends to the electrical compressor (5). Therefore, electrical compressor (5) may be fed by bleed air or cooling air as explained above.

(80) Once compressed by the electrical compressor (5), air returns to the bleed duct (2) in order to be supplied to the air consumer (6) equipment such as ECS, WAIS, engine start valves, etc.

(81) Furthermore, a dotted line represents the interface between gas turbine engine and pylon.

(82) A pressure sensor located in the bleed duct after the box provides pressure information to a unit that further operates a pressure regulation valve located downstream.

(83) FIG. 8 depicts the internal nacelle fairing (7.2) of a turbofan engine (7) and the upper pylon bifurcation (7.3.1), with an aircraft installation (1) according to the present invention. It is to be noted that external nacelle (7.1) and fan are not shown in these figures for illustrative purposes.

(84) In particular, the inlet (3.1) of the cooling duct (3) is shown arranged on the upper pylon bifurcation (7.3.1) or over internal nacelle fairing (7.2) itself. Suitable locations of a ram air intake (3.1) are pointed out therein (A—frontal-, B—lateral-, C—peripheral-). Final location mainly depends on local flow direction influenced by the aircraft (10) angle of attack.

(85) The internal nacelle fairing (7.2) shown in FIG. 8 further connects with the external nacelle (7.1) via an upper (7.3.1) and lower pylon bifurcation (7.3.2).

(86) Potential locations of the cooling duct inlet (3.1) are shown:

(87) Possible installations on the pylon bifurcation (7.3.1):

(88) at the foremost part of the upper pylon bifurcation (7.3.1) (frontal option);

(89) so as to be angled relative to the incoming cooling air (lateral option);

(90) Possible installation on the internal nacelle fairing (7.2):

(91) on a portion of a peripheral ring of the fairing (7.2).

(92) In a preferred embodiment for the installation on the upper pylon bifurcation (7.3.1), the cooling duct inlet (3.1) is arranged in the front part of the bifurcation (‘frontal option’). This is a higher pressure area which improves inlet efficiency.

(93) Briefly, the cooling duct inlet (3.1) may be located either on the front part of the upper pylon bifurcation (7.3.1) where the pressure is higher or on the lateral side (i.e., so as to be angled relative to the incoming cooling air) where the pressure is lower but enough.

(94) FIG. 9 depicts additional possible location for the cooling duct inlet (3.1) on the external nacelle (7.1) and on upper fairing (7.4) between the external nacelle (7.1) and the wing, where the available pressures are suitable for its installation.

(95) The most suitable location for the cooling air outlet is on the upper part of the upper fairing (7.4).

(96) FIG. 10 depicts an exemplary aircraft (10) mission profile using a hybrid solution according to the present invention for supplying compressed air throughout the flight phases.

(97) If the air consumer (6) is fed directly with the bleed air extracted from the single port (2.1), it is represented by a continuous line. On the other hand, it is represented in dashed lines when the air consumer (6) is fed by air extracted via the single port (2.1) and further compressed by the electrical compressor (5). Also, it is represented in dotted lines when air consumer (6) is fed by cooling air of the cooling duct (3) and further compressed by the electrical compressor (5).

(98) It is to be noted that, for illustrative purposes, no overlap between operation of the air bleed system (2) and operation of the at least one electrical compressor (5) (either supplied by cooling or bleed air) is shown, but this situation of overlap is of interest at the interphase when the compressed air source switches.

(99) In particular, the criteria follow by the control unit (controller 9) to operate the air bleed system and/or the electrical compressor (5) upon receiving an aircraft (10) operation condition input (i.e., flight altitude or flight phase) is summarized as follows:

(100) Below a pre-determined altitude, preferably 15000 ft.:

(101) in taxiing and taking-off, the electrical compressor (5) is fed only with cooling air from the cooling duct (3), so that the electrical compressor (5) exclusively supplies compressed air to the air consumer equipment (6);

(102) in climbing, the electrical compressor (5) continues exclusively supplying compressed air to the air consumer (6) equipment up to the pre-determined altitude;

(103) Above the pre-determined altitude:

(104) still in climbing, the air consumer equipment (6) is directly fed by bleed air from the bleed duct (2) (i.e., without passing through the electrical compressor);

(105) in cruise, the air consumer (6) equipment continuous being directly fed by bleed air from the bleed duct (2) (i.e., without passing through the electrical compressor); and

(106) Once cruise phase ends:

(107) in descending and holding still above the pre-determined altitude, the electrical compressor (5) is fed only with bleed air from the bleed duct (2), so that the electrical compressor (5) exclusively supplies compressed air to the air consumer (6) equipment; and

(108) Below the pre-determined altitude:

(109) still in approaching or landing, the electrical compressor (5) is fed only with cooling air from the cooling duct (3), so that the electrical compressor (5) exclusively supplies compressed air to the air consumer (6) equipment.

(110) In other words, as the aircraft (10) passes from one phase to another, a control unit (controller 9) receives the corresponding input and operates the corresponding valves or directly the electrical compressor (5), which affects the particular operation of a compressed air source.

(111) As it was already mentioned, since air bleed system exclusively operates in favorable conditions from energy cost point of view (high altitude and relative high speed), the air bleed duct (2) together with the associated valves or regulators is sized according to cruise phase flight conditions, which encompasses the majority of the flight.

(112) Energy-demanding flight phases such as on-ground operation, take-off, or even the first portion of climbing, as well as other phases like descent (or approaching) and holding relies exclusively in pressurized air supplied by the electrical compressor(s) (5).

(113) Therefore, the electrical compressor(s) (5) adapts the delivered pressure to the required pressure by the air consumer (6) upon indication from the control unit (9).

(114) Throughout the entire description, the person skilled in the art would recognize that specific figures of aircraft (10) operation, or parameters of air bleed systems highly depend on specifics of the aircraft (10) model.

(115) While at least one exemplary embodiment of the present invention(s) is disclosed herein, it should be understood that modifications, substitutions and alternatives may be apparent to one of ordinary skill in the art and can be made without departing from the scope of this disclosure. This disclosure is intended to cover any adaptations or variations of the exemplary embodiment(s). In addition, in this disclosure, the terms “comprise” or “comprising” do not exclude other elements or steps, the terms “a” or “one” do not exclude a plural number, and the term “or” means either or both. Furthermore, characteristics or steps which have been described may also be used in combination with other characteristics or steps and in any order unless the disclosure or context suggests otherwise. This disclosure hereby incorporates by reference the complete disclosure of any patent or application from which it claims benefit or priority.