HK1115566A - Electric-based secondary power system architectures for aircraft - Google Patents

Description

Electric auxiliary power system structure for airplane

The invention relates to an invention named as an electric auxiliary power system structure for an airplane, which is a divisional application of an invention patent application with the application date of No. 10/22/2003 and the application number of 200380101725.X and applied to Boeing company.

Technical Field

The following disclosure relates generally to an auxiliary power system for an aircraft and, more particularly, to an electric auxiliary power system for an aircraft.

This application claims priority to co-pending U.S. provisional patent application 60/420,637, filed 10/22/2002. And is incorporated by reference into this application in its entirety. This application refers to U.S. Pat. No. 6,526,775, the entirety of which is incorporated by reference herein.

Background

Conventional transport aircraft typically support various aircraft systems during flight using the pneumatic, hydraulic, and electric power of the main engines. In addition, conventional transporters often utilize pneumatic and electric power from onboard Auxiliary Power Units (APUs) to support aircraft systems during ground operations. In commercial transport aircraft, the air conditioning system of the aircraft is typically the largest source of auxiliary power usage. In conventional transporters, these systems use high temperature/high pressure air ("bleed air") extracted from the compressor stages of the engines. Air passes through the air conditioning unit before entering the fuselage to satisfy the needs of temperature, ventilation and pressure boost. The conditioned air is then exhausted from the fuselage through an outflow valve or through a normal cabin leak. During ground operation, the APU can provide bleed air from either a separate shaft driven load compressor or from the power section compressor. Similar to bleeding air from the main engines, the high temperature and high pressure air in the APU passes through the air conditioning unit before entering the fuselage.

Fig. 1 is a view schematically showing a conventional pneumatic auxiliary power system structure 100 constructed in accordance with the prior art. The system architecture 100 may include a jet engine 110 (shown as a first engine 110a and a second engine 110 b) for providing thrust to the aircraft. In addition to thrust, the engines 110 are also capable of providing high temperature/pressure air to the bleed manifold 120 through bleed ports 112 (labeled as first bleed port 112a and second bleed port 112b, respectively). The bleed ports 112 receive air from the compressor stages of the engines 110 and pass the air through a heat exchanger 114 (e.g., a precooler) that cools the air before it flows to the bleed manifold 120.

The high pressure air from the bleed manifold 120 supports most of the auxiliary power requirements of the aircraft. For example, a portion of this air flows to air conditioning packs 140 (shown as first air conditioning pack 140a and second air conditioning pack 140 b) that provide conditioned air to passenger cabin 102 in fuselage 104. The air conditioning packs 140 include a set of heat exchangers, conditioning valves, and air cycle machines that condition air to meet the temperature, ventilation, and pressurization requirements of the passenger compartment 102. Another portion of the air flows from the bleed manifold 120 to the turbine 160 for driving the high capacity hydraulic pump 168. The hydraulic pump 168 provides hydraulic power to the landing gear and other hydraulic systems of the aircraft. Still another portion of the high pressure air is directed to hood ice protection system 152 and wing ice protection system 150.

The wing ice protection system 150 includes a valve (not shown) that controls the flow of bleed air to the leading edge of the wing and "piccolo" tubes (not shown) that distribute the heated air evenly along the protected area of the wing leading edge. If ice protection is required for the leading edge slats, a telescoping tube may be employed to provide hot bleed air to the slats in the extended position. The anti-icing bleed air is discharged through holes in the lower surface of the wing or slat.

In addition to the engine 110, the system architecture 100 may also include an APU 130 for use as an alternative pressure source. The APU 130 is typically powered by a starter motor 134 using a battery 136. The APU 130 drives a compressor 138 that provides high pressure air to the bleed manifold 120 for engine starting and other ground operations. For engine starting, high pressure air is caused to flow from the bleed manifold 120 to a starter turbine 154 operatively connected to each engine 110. As an alternative to the APU, bleed air from one engine 110 operating therein may be used to restart another engine 110. Alternatively, an external air tanker (not shown in FIG. 1) may provide high pressure air for starting the engine on the ground.

The system architecture 100 may also include an engine driven generator 116 operatively connected to the engine 110 and an APU driven generator 132 operatively connected to the APU 130. In flight, the engine-driven generator 116 may support conventional electrical system loads, such as the fuel pump 108, the motor-driven hydraulic pump 178, and various fans, galley systems, in-flight entertainment systems, lighting systems, avionics systems, and the like. The APU-driven generator 132 may support these functions as needed during ground operation and during flight. The engine-driven generator 116 and the APU-driven generator 132 are typically rated at 90-120kVA and produce a voltage of 115 Vac. They can power a transformer-rectifier device that can convert 115Vac to 28Vdc for many of the electrical loads described above. Power distribution is performed by electrical systems that rely heavily on thermal circuit breakers and relays.

The system architecture 100 may also include an engine-driven hydraulic pump 118 operatively connected to the engine 110. The hydraulic pump 118 provides hydraulic power to operate surface actuators (aircraft) and other aircraft systems in flight. The motor driven pump 178 may provide backup hydraulic power to maintain activity on the ground.

FIG. 2 is a top view schematically illustrating a prior art aircraft 202 including the auxiliary power system architecture 100 of FIG. 1. The aircraft 202 includes a forward electronics compartment 210 that distributes electrical power to a plurality of electrical loads 220 connected to the system architecture 100. In flight, the electronics bay 210 may receive electrical power from the engine generator 116, as well as the APU 130. At the ground, the electronics bay 210 may receive power from the APU 130 or from an external power source 212 via a receptacle 213.

One drawback of the auxiliary power system configuration 100 described above is that it is sized with the worst conditions (typically cruise speed, high flight loads, hot weather, and failure of one of the engine bleed air systems) to ensure that there is sufficient air flow to meet the demand at any time. Thus, under normal operating conditions, the engines 110 provide bleed air at a pressure and temperature much higher than that required by the air conditioning packs 140 and other aircraft systems. To compensate, the precooler 114 and the air conditioning packs 140 adjust the pressure and temperature to lower values as needed to meet the needs of fuselage pressurization, ventilation and temperature control. Thus, a significant amount of energy is wasted by the precooler and the regulator valve during the conditioning process. Even under optimal conditions, a significant amount of energy extracted from the engines 110 is wasted in the form of heat and pressure drops that occur in the tubes, valves, and other components associated with the bleed manifold 120 and the air conditioning packs 140.

Disclosure of Invention

The present invention generally relates to an auxiliary power system for an aircraft and a method for providing auxiliary power to an aircraft system. In one embodiment, an aircraft constructed according to one aspect of the invention includes a fuselage and a jet engine for providing propulsive thrust to the aircraft. The aircraft also includes an electrical generator operatively connected to the jet engine, and an environmental control system. The environmental control system may include at least one compressor motor for receiving electrical power from the generator to provide outside air to the fuselage in the absence of bleed air from the jet engine.

In another aspect of this embodiment, an aircraft may include wings extending outwardly from a fuselage and an electrically heated wing ice protection system. The electrically heated wing ice protection system may be configured to receive electrical energy from the generator to at least reduce ice formation on a portion of the wing in the absence of bleed air from the jet engine. In yet another aspect of this embodiment, the generator may be a first generator, and the aircraft may further include an auxiliary power unit and a second generator. The second generator is operatively connected to the auxiliary power unit and is operable to receive shaft work from the auxiliary power unit. In this regard, the compressor motor of the at least one environmental control system is configured to receive electrical energy generated by the second generator to provide outside air to the passenger compartment in the absence of compressed air from the auxiliary power unit.

In another embodiment, a method for providing conditioned air to a fuselage of an aircraft includes providing a compressor fan in fluid communication with the fuselage and operatively connecting a motor to the compressor fan to drive the compressor fan. The method also includes operatively connecting an electrical generator to a jet engine of the aircraft and providing electrical power from the electrical generator to the electrical machine to drive the compressor fan. In one aspect of this embodiment, the compressor fan can be driven to flow air from outside the fuselage into the fuselage in the absence of bleed air from the jet engine.

Drawings

FIG. 1 is a schematic diagram showing the construction of a conventional pneumatic auxiliary power system constructed in accordance with the prior art;

FIG. 2 is a top view schematically illustrating a prior art aircraft having the system architecture of FIG. 1;

FIG. 3 is a schematic diagram illustrating an electric auxiliary power system constructed in accordance with an embodiment of the present invention;

FIG. 4 is a schematic top view of an aircraft having the system architecture of FIG. 3 constructed in accordance with one embodiment of the invention;

FIG. 5 is a schematic illustration of an aircraft electrical power distribution system constructed in accordance with an embodiment of the invention;

FIG. 6 is a schematic illustration of an aircraft electrical power distribution system constructed in accordance with another embodiment of the invention;

FIG. 7 is a schematic illustration of an aircraft power distribution system having only AC generators in accordance with yet another embodiment of the present invention;

fig. 8A-8C are schematic diagrams illustrating an electrical energy distribution system having an engine start circuit constructed in accordance with an embodiment of the invention.

Detailed Description

The following disclosure describes systems and methods for providing power to aircraft systems. Certain details are set forth in the following description and in figures 3-7 to provide a thorough understanding of various embodiments of the invention. Additional details describing known structures and systems associated with aircraft and/or aircraft auxiliary power systems will not be set forth in the following description in order to avoid unnecessarily obscuring the description of the various embodiments of the invention.

Many of the details, dimensions, angles and other features shown in the figures are merely illustrative of particular embodiments of the invention. Accordingly, other embodiments may have other details, dimensions, and features without departing from the spirit and scope of the invention. In addition, other embodiments of the invention are possible without several of the details described below.

In the drawings, like reference numerals designate identical or at least substantially similar components. To facilitate the discussion of any particular element, the first digit of any reference number indicates the figure number in which the element is first introduced. For example, component 310 is first introduced and discussed in FIG. 3.

FIG. 3 is a schematic diagram illustrating an electric auxiliary power system configuration 300 constructed in accordance with an embodiment of the present invention. In one aspect of this embodiment, the system architecture 300 includes a first engine 310a and a second engine 310b that provide thrust to an aircraft (not shown). As described in more detail below, a first starter/generator 316a and a second starter/generator 316b are operatively connected to each engine 310 to provide electrical power to a plurality of aircraft systems as needed. The starter/generators 316 support most of the conventional functions performed by the bleed air system shown in figure 1. Among these functions may include fuselage air conditioning and pressurization, engine starting, and wing ice protection.

In another aspect of this embodiment, the system architecture 300 further includes an APU 330 for powering aircraft systems as needed, either for ground operation or in flight. The power to start the APU 330 may be provided by the aircraft battery 336, an external ground power source (not shown), or one or more engine-driven starter/generators 316. The power for the APU 330 is provided by a first APU starter/generator 332a and a second APU starter/generator 332b, each of which is operatively connected to the APU 330.

In contrast to the conventional APU 130 described above with reference to FIG. 1, the APU 330 provides electrical power only to the various aircraft systems. Therefore, it is simpler than APU 130 since all components associated with pneumatic transfer are eliminated. This feature may greatly improve the reliability of the APU and reduce the required maintenance.

In another aspect of this embodiment, the system architecture 300 includes an environmental control system having a first air conditioning unit 340a and a second air conditioning unit 340 b. The air conditioning packs 340 are used to provide conditioned air to the passenger cabin 302 in the fuselage 304 to meet temperature, pressure, and air conditioning needs. In one embodiment, the air conditioning packs 340 may be at least substantially similar to one or more air conditioning systems described in U.S. Pat. No. 6,526,775, which is incorporated by reference herein in its entirety. In another embodiment, the air conditioning pack 340 may include an adjustable speed electric compressor motor 380 configured to receive power from the engine 310 during flight and to receive power from the APU 330 during ground operation. The compressor motor 380 drives a compressor (not shown) that receives fresh outside air through the ram air inlet 342. Fresh air is compressed and flows from the air conditioning packs 340 into the fuselage 304 to meet the pressurization and temperature control needs of the passenger cabin 302. In one embodiment, the system architecture 300 may include one or more variable speed fans (not shown) to distribute air to various portions of the airframe 304 at different flow rates so that the specific needs of the airframe 304 may be met at any given time. Varying the power extracted from the engine 310 in this manner may further improve fuel efficiency.

The adjustable speed compressor motor 380 varies the air pressure and air flow to the cabin based on the cabin volume, the number of occupants, and/or the desired cabin air pressure altitude. For example, if a lower cabin altitude (higher air pressure) is desired, the electrical ECS system of the present invention can accommodate this by increasing the incoming flow and/or decreasing the outgoing flow from the fuselage 304 using the adjustable speed compressor motor 380. Generally, conventional pneumatic systems do not have the ability to reduce cabin altitude to much lower than the design point (e.g., 8000 feet) because these systems are generally sized at the design point. Another advantage of the electrical method of the air conditioner over the conventional pneumatic method is that: the energy extracted by the engine for the electrical method is not wasted by the pre-cooler and the regulator valve in the air conditioning pack 340. Instead, the compressor motor 380 draws only enough electrical energy from the engine 310 as needed by the variable speed compressor to meet the immediate pressurization needs of the passenger cabin 302. This real-time energy optimization can be extended throughout the aircraft platform to other electrical energy users, thereby improving fuel efficiency. As described below, these users may include, for example, recirculation fans, lavatory and galley ventilation fans, cargo heating, wing ice protection, and hydraulic actuation. Fuel economy is improved since only the required energy is extracted.

In another embodiment of the present invention, the system architecture 300 further includes a wing ice protection system 350 that utilizes electrical power from the engine 310. Wing ice protection system 350 may be constructed in accordance with at least two embodiments of the present invention to prevent or at least reduce ice formation on a portion of wing 352. In an electrothermal ice protection embodiment, a heating element, such as an electrothermal film (not shown), may be incorporated into or disposed proximate to the interior of the leading edge of the wing. For wing ice protection, the thin heating layer may be energized to heat the wing leading edge so that any ice build up may melt and/or detach from the wing leading edge. This method is significantly more efficient than conventional bleed air systems because only the desired portion of the leading edge of the wing is heated, rather than in real time. Thus, the extraction energy for ice protection is significantly reduced. Furthermore, in contrast to bleed air systems, there are no bleed air discharge holes on the wing. Thus reducing aircraft drag and common noise as compared to conventional systems.

The wing ice protection system 350 also operates as an electromechanical system in accordance with another embodiment of the present invention. In this embodiment, an electromechanical actuator (not shown) inside the wing leading edge may be used to briefly vibrate the wing leading edge to break any ice off and fall. This embodiment requires much less electrical energy than the electrothermal embodiments described above. In one embodiment, the wing ice protection system 350 can be divided into different portions that are applied to different areas of the wing or slat leading edge. In this way, if a portion of the wing leading edge does not require ice protection, that portion of the wing ice protection system 350 can be turned off, thereby further reducing the power requirements on the engine. In addition, the anti-icing system may be periodically operated as needed according to different periods of time to substantially reduce icing and optimize power usage.

In another aspect of this embodiment, starter/generator 316 may be a dual function device that provides electrical power to aircraft systems when operating as a generator and shaft work for engine starting when operating as a starter. This electric starting capability may enhance the results of the internal start of the engines 310 in the event that one or more of the primary engines 310 are shut down during normal flight. For example, typical high bypass flow rate engines may experience starting difficulties in all flight conditions because the windmilling effect in flight may not provide sufficient torque. Rather, the starter/generators 316 of the present invention are configured to receive electrical energy from any number of power sources onboard the aircraft to assist the engines 310 by providing additional starting torque during an in-flight start.

To start the engine 310, the starter/generator 316 may be operated as a synchronous starter motor, with the starting process being controlled by an engine start converter (not shown). The engine start converter may provide regulated power (e.g., adjustable voltage and frequency) to the starter/generator 316 during starting to optimize starting performance. The engine start converter may also function as a motor controller for controlling the cabin booster compressor motor 380 and/or other adjustable speed motors on the aircraft. Similarly, the APU start converter (not shown) may function as a motor controller for controlling other adjustable speed motors on the aircraft, such as the cabin inert gas generation system (OBIGGS) 309. The power that the starter/generator 316 must provide for engine starting may come from the aircraft battery 336, the APU 330, the Ram Air Turbine (RAT)367, the fuel cell (not shown), or other energy source. The dual function aspect of the starter/generator 316 is not provided by the air turbine starter 154 described above with reference to FIG. 1. Unlike starter/generator 316, air turbine engine starter 154 does not function during operation of engine 110.

In another aspect of this embodiment, the starter generators 316 may be directly coupled to the gearbox of the engine 310 such that they operate at a frequency proportional to engine speed (e.g., 360 and 700 Hz). Such a generator may be the simplest and most efficient method, since the engine does not include a complex constant speed drive. Thus, in this embodiment, starter/generator 316 is more reliable and can save more cost than a conventional generator with a complex constant speed drive. However, in other embodiments, other types of generators may be used. For example, in one other embodiment where a constant speed is desired, a constant speed generator may be employed.

In another aspect of this embodiment, the system architecture 300 includes a hydraulic system having left, right and center channels. Hydraulic power for the left and right channels may be provided by engine-driven hydraulic pumps 318 operatively connected to each engine 310. In addition, the smaller motor driven hydraulic pump 319 can also provide hydraulic power to the right and left ground operated channels to make up for the engine driven pump 318. The engine driven pump 318 can provide hydraulic power for flight control actuators, stabilizer trim actuators, and other functions. Hydraulic power for the center channel is provided by two large capacity electric motor driven hydraulic pumps 368. In contrast to the hydraulic pump 168, which is driven by the engine bleed air to meet peak hydraulic demand as described with reference to FIG. 1, the hydraulic pump 368 is driven by the power source of the engine 310. The hydraulic pump 368 may provide hydraulic power for the landing gear system 369 and other systems, including flight control actuation, thrust reversers, braking, nose/tail flap, and nose gear steering systems (not shown). In another aspect of this embodiment, only one hydraulic pump 368 operates throughout the flight, while the other pumps only operate at takeoff and landing.

In another aspect of this embodiment, the system architecture 300 can include a plurality of adjustable speed fuel pumps 308 that deliver fuel from a fuel tank 390 to one or more engines 310 or to another fuel tank (not shown). Typical commercial aircraft use fuel pumps to transport fuel from one wing region to another. This enables the aircraft to maintain its centre of gravity to optimise aircraft performance. In the conventional pneumatic system configuration 100 described above with reference to FIG. 1, a constant speed fuel pump is typically included for delivering fuel from one tank to the next or to the engine 110. These constant speed fuel pumps typically operate at a maximum pressure at all times, even when there is little need to deliver a flow rate corresponding to the maximum pressure between fuel tanks or between fuel delivered by the fuel tanks and engine 110. To this end, these fuel systems typically include a pressure regulator for simply relieving excess fuel pressure. Excessive fuel pressure equates to wasted engine energy. In contrast, in the electrically powered configuration 300 of FIG. 3, the speed of the fuel pump may be varied to deliver fuel from one fuel tank to another according to the amount of fuel to be delivered and the flow rate required for delivery to optimize the overall center of gravity of the aircraft during normal flight conditions. The ability to maintain the optimum of center of gravity throughout the flight phase further improves the range and fuel efficiency of the aircraft by pumping only the energy actually required at any given time by the fuel pump.

Many other systems may be incorporated into the system architecture 300 to further reduce the energy extracted from the engine 310. For example, in one embodiment, a variable speed or variable speed fan may be used in the air conditioning pack 340 to vary the energy extracted from the engine 310 based on the fan speed. In another embodiment, the cargo compartment (not shown) may be heated with resistance heating wires instead of bleed air as used in conventional systems. These heating resistance wires may have pulse width modulation capability in order to better control the temperature and further reduce the energy consumption. Similarly, the cargo air conditioning system may be configured to rely less on outside air, and more on circulating air to cool the compartment. In this way, the energy losses associated with the outside air are eliminated and only the energy required to cool the circulating air is expended.

As discussed above with reference to FIG. 1, in the conventional system architecture 100, the engine 110 provides the primary secondary flight power in the form of pneumatic bleed air. In contrast, in the system architecture 300 of the present invention, the engine 310 provides auxiliary flight power in an electric form primarily comprised of a starter/generator 316. By eliminating the pneumatic bleed ports of the compressor section of the engine 310, the engine design is made more efficient by reducing the required compressor capacity and improving the duty cycle. Furthermore, eliminating the maintenance intensive bleed air system would be desirable to reduce the maintenance requirements and increase the reliability of the aircraft due to fewer components of the engine and fewer pneumatic conduits, precoolers and valves in the distribution system. Furthermore, it is not necessary for the electrical system architecture 300 to take steps to prevent temperature conditions of duct cracking and overheating.

Another advantage of electric-based system architecture 300 is that it can utilize the motor controller to regulate individual loads so that only the minimum amount of energy necessary is drawn from engine 310 under any operating condition. Since these loads are adjustable, rather than simply on or off, less energy is drawn from the engine 310. The ability to adjust the power consumption of any electrical load can directly improve the fuel efficiency of the aircraft. Other advantages associated with the electrical system architecture 300 are as follows: real-time energy extraction is optimized and waste associated with engine bleed air is eliminated; the air quality is improved; potentially reducing non-repetitive engineering associated with certification of multi-stage engine bleed air systems.

While the system architecture 300 described above with reference to FIG. 3 includes two engines each having two starter/generators, in another embodiment, a system architecture constructed in accordance with the present invention may include more or less engines having more or less individual starter/generators, as desired for a particular application. For example, in an alternative embodiment, a system configuration constructed in accordance with the present invention may include four jet engines, each having a single starter/generator. In another embodiment, a system architecture constructed in accordance with the present invention may include only one engine having two or more generators or starter/generators. Thus, the present invention is not limited to aircraft having a particular number of engines or starter/generators.

Fig. 4 is a schematic top view of an aircraft having the system architecture 300 of fig. 3 constructed in accordance with one embodiment of the invention. The aircraft 402 may include four starter/generators 316 coupled to two engines 310 and two starter/generators 332 coupled to an APU 330 mounted aft of the aircraft. In one aspect of this embodiment, the aircraft 402 may also include two ground sockets 413 (labeled as a first ground socket 413a and a second ground socket 413b) for receiving 115Vac or 230Vac power from external power sources 412a and 412b, respectively.

In another aspect of this embodiment, the aircraft 402 can include a forward electrical equipment bay 410a and an aft electrical equipment bay 410 b. Four Remote Power Distribution Units (RPDUs)424a-d may distribute power from equipment bays 410 to a plurality of system loads 420 connected to system architecture 300. The RPDUs 424a may rely primarily on solid state power controllers, rather than traditional thermoelectric breakers and relays.

Fig. 5 is a schematic illustration of an aircraft power distribution system 500 constructed in accordance with an embodiment of the invention. In one aspect of this embodiment, the electrical power distribution system 500 includes a first electrical generator 516a and a second electrical generator 516b operatively connected to the aircraft engine 510. In one embodiment, the first generator 516a and the second generator 516b may be high voltage AC generators (e.g., 230Vac generators). In another embodiment, one of the two generators 516 may be a high voltage DC generator (e.g., ± 270Vdc generator). The AC generator 516a may provide electrical power to aircraft equipment that is not affected by the power supply frequency. The DC generator 516b may provide power to components of the aircraft system, including the adjustable speed motor. In other embodiments, the generator 516 may be other types of generators. For example, in one other embodiment, both generators 516 may be AC generators. In this embodiment, the DC power requirements of the system can be met with appropriate ac-DC conversion means. In another embodiment, both generators 516 may be DC generators, and the AC power requirements of the system may be met by appropriate DC-to-AC conversion devices.

In another aspect of this embodiment, the power distribution system 500 further includes a first bus 515a for receiving power from the first generator 516a and a second bus 515b for receiving power from the second generator 516 b. In one embodiment, the first bus 515a may be a high voltage AC bus, such as a 230Vac bus, for providing power directly to a plurality of large rated AC loads 550. These loads may be associated with wing ice protection equipment, hydraulic pumps, fuel pumps, galley systems, or the like. In addition, the first bus 515a may also provide power directly to the third bus 515c through a step-down transformer 522. In one embodiment, the third bus 515c may be a low voltage AC bus, such as a 115Vac bus. The third bus 515c may provide power to a plurality of small-rated AC device loads 544 through a plurality of RPDUs (labeled as a first RPDU 524a and at least a second RPDU 524 b). Such a small rated load 544 may be associated with in-flight entertainment systems, interior and exterior lighting systems, sensor heaters, or the like.

In yet another aspect of this embodiment, the second bus 515b can provide power to a plurality of adjustable speed motors 552 on the aircraft. These motors may include compressors for cabin pressurization, climate control system fans, steam or air cycle ECS units, large hydraulic pumps, flight actuators, or the like. The use of a high voltage DC system avoids potential harmonic distortion problems often associated with motor controllers and provides a means for receiving renewable energy often associated with electrohydraulic actuators. Furthermore, the use of high voltage DC systems can also result in significant weight savings by employing a lightweight DC generator and eliminating harmonic distortion handling and renewable energy absorption devices.

Fig. 6 is a schematic diagram of an aircraft power distribution system 600 constructed in accordance with another embodiment of the invention. In one aspect of this embodiment, the power distribution system 600 includes a first aircraft engine 610a, a second aircraft engine 610b, and an APU 630. The power distribution system 600 may also include three AC generators 616 (identified as a first AC generator 616a, a second AC generator 616b, and a third AC generator 616c, respectively), and three DC generators 618 (identified as a first DC generator 618a, a second DC generator 618b, and a third DC generator 618c, respectively). A first AC generator 616a and a first DC generator 618a are operatively connected to the first engine 610 a. Likewise, a second AC generator 616b and a second DC generator 618b are operatively connected to the second engine 610 b. The third AC generator 616c and the third DC generator 618c are operatively connected to the APU 630. The third AC generator 616c may provide power from the APU630 to two AC buses 615a to service AC loads (not shown) as needed during ground operation and during flight. A third DC generator 618c, operatively connected to APU630, may provide electrical power to both DC buses 615b to service adjustable speed motors (not shown) as needed during ground operation and during flight. In addition, a third AC generator 616c may also provide power to the two DC buses 615b via an AC-to-DC conversion device 624. The two DC buses 615b are each operatively connected to a corresponding motor controller 660 (identified as a first motor controller 660a and a second motor controller 660b, respectively). The motor controller 660 may be used to selectively charge the cabin booster compressor 680 (labeled as a first compressor 680a and a second compressor 680b, respectively) or the engine starter circuit 662 (labeled as a first starter circuit 662a and a second starter circuit 662b, respectively). In another aspect of this embodiment, power distribution system 600 can include a second power receptacle 613a and a second power receptacle 613b for receiving power from an external ground power source. In one embodiment, the first receptacle 613a may be configured to receive 115Vac power from a ground power source and the second receptacle 613b may be configured to receive 230Vac power from an external ground power source.

In one embodiment, high voltage (e.g., 230Vac) ground power received through the second socket 613b may be used to start the engine 610. In this embodiment, the motor controller 660 converts the electrical energy from the DC bus 615b to be directed to the corresponding engine start circuit 662. The electrical energy is directed to a corresponding AC generator 616 and used to run the AC generator 616 as a synchronous machine to turn the corresponding engine 610 for starting. Once the engine 610 is started, the motor controller 660 transitions back to provide electrical power to the cabin pressurization compressor 680.

Fig. 7 is a schematic diagram of an aircraft power distribution system 700 having only AC generators in accordance with yet another embodiment of the invention. The power distribution system 700 includes a first engine 710a, a second engine 710b, and an APU 730. In one aspect of this embodiment, first and second AC generators 716a and 716b are operatively connected to the first engine 710a, third and fourth generators 716c and 716d are operatively connected to the second engine 710b, and fifth and sixth AC generators 716e and 716f are operatively connected to the APU 730. To meet the DC voltage requirements, one or more AC-to-DC conversion devices, such as Automatic Transformer Rectifier Units (ATRUs)724, may be utilized to convert the high voltage AC power from the engines 710 and APUs 730 to high voltage DC power for receipt of the AC power from the AC bus 715. The use of the ATRUs724 enables the power distribution system 700 to provide high voltage AC and DC power to support the conventional 115Vac and 28Vdc bus structures. Further, the unregulated rectifier unit 725 and the regulated rectifier unit 726 may also be utilized to convert AC power from one or more AC generators 716 to DC power for a 28Vdc bus 719. In another aspect of this embodiment, having two AC generators 716 connected to each generator 710 allows both AC generators 716 to be used as synchronous starter motors to increase engine starting energy when needed. As described above with reference to FIG. 3, the embodiments of the present invention described above with reference to FIGS. 5-7 are not limited to the particular number of engines and/or starter/generators shown in the figures, but may be extended to other numbers of engines and starter/generators in different configurations.

Fig. 8A-8C are schematic diagrams illustrating an electrical power distribution system 800 having an engine start circuit constructed in accordance with an embodiment of the invention. Referring first to FIG. 8A, in one aspect of this embodiment, an electrical energy distribution system 800 includes an electrical generator 816 operably coupled to an engine 810 and a compressor motor 880 operably coupled to an environmental control system 840. The generator 816 may provide electrical power to a motor controller 860 via an AC bus 815a, an AC-to-DC converter 824, and a high voltage DC bus 815 b. During normal operation, as shown in fig. 8A, the motor controller 860 may selectively direct electrical power to the compressor motor 880 for operation of the ECS 840.

Fig. 8B shows an engine start configuration of the power distribution system 800. Since the engine 810 is not initially operated in this configuration, power is provided to the motor controller 860 by an alternative AC power source 830 rather than by the generator 816. In one embodiment, the alternate power source 830 may comprise an APU or an external power source. In one aspect of this embodiment, the motor controller 860 selectively directs electrical energy from the alternate power source 830 to the engine start circuit 862. The engine start circuit 862 provides electrical power to the generator 816, and the generator 816 acts as a synchronous motor to start the engine 810.

Fig. 8C shows another engine starting configuration of the power distribution system 800. Here, electric power for starting the engine 810 is supplied from the battery 836. In one aspect of this embodiment, the motor controller 860 selectively completes the circuit to the high voltage DC bus 815b to enable it to receive power from the battery 836. After being connected to the battery 836, the motor controller directs electrical energy to the generator 816 through the engine start circuit 862 as described above.

From the foregoing, it will be appreciated that specific embodiments of the invention have been described above for purposes of illustration only, and that various changes may be made without departing from the spirit and scope of the invention. Accordingly, the invention is not limited except as by the appended claims.

Claims (15)

1. An aircraft, comprising:

a body;

a jet engine configured to provide propulsive thrust to an aircraft;

a generator operatively connected to the jet engine and configured to receive shaft work from the engine;

an environmental control system configured to provide conditioned air to at least a portion of the fuselage in the absence of bleed air from the jet engine, the environmental control system including at least one fan motor configured to receive electrical power from the generator; and

a hydraulically actuated landing gear system configured to movably support at least a portion of the aircraft on the ground, the landing gear receiving hydraulic power from a hydraulic pump driven by an electric motor that receives electrical power from the generator in the absence of bleed air from the jet engine.

2. The aircraft of claim 1, wherein the generator operatively connected to the jet engine is a DC generator, wherein the aircraft further comprises:

an AC generator operatively connected to the jet engine and configured to receive shaft work from the jet engine;

a wing extending outwardly from the fuselage; and

an electrically heated wing ice protection system configured to at least reduce ice formation on a portion of the wing, the electrically heated wing ice protection system configured to receive electrical power from the AC generator in the absence of bleed air from the jet engine.

3. The aircraft of claim 1, wherein the generator operatively connected to the jet engine is a DC generator, wherein the aircraft further comprises an AC generator operatively connected to the jet engine and configured to receive shaft work from the jet engine, wherein the AC generator is operable as a synchronous machine to start the jet engine in the absence of pneumatic power.

4. The aircraft of claim 1, wherein the generator operatively connected to the jet engine is a DC generator, wherein the aircraft further comprises:

an AC generator operatively connected to the jet engine and configured to receive shaft work from the jet engine; and

an AC-DC converter device is configured to receive AC power from the AC generator and direct DC power to the fan motor of the environmental control system.

5. The aircraft of claim 1, wherein the generator operatively connected to the jet engine is a DC generator, wherein the aircraft further comprises:

an AC generator operatively connected to the jet engine and configured to receive shaft work from the jet engine;

an AC-dc converter configured to receive AC power from the AC generator; and

a motor controller configured to receive the DC electrical power from the AC-DC converter device and direct the DC electrical power to a fan motor of the environmental control system, wherein the motor controller is further configured to selectively direct the DC electrical power from the power source to the AC generator instead of from the AC generator to operate the AC generator as a synchronous motor to start the jet engine in the absence of pneumatic power.

6. A method of providing conditioned air to a fuselage of an aircraft, the aircraft including a jet engine configured to provide propulsive thrust to the aircraft, the method comprising:

operatively connecting an electrical generator to the jet engine, the electrical generator configured to receive shaft work from the jet engine;

operatively connecting an electrical generator to the compressor fan to drive the compressor fan, the compressor fan being positioned in an airflow in communication with the airframe; and

the electrical machine is supplied with electrical power from the generator to drive the compressor fan and to cause air to flow from outside the fuselage into the fuselage in the absence of bleed air from the jet engine.

7. The method of claim 6, further comprising:

placing a heating resistance wire downstream of a compressor fan; and

electric power is transmitted from the generator to the heating resistance wire to heat the outside air before the outside air flows into the body.

8. The method of claim 6, wherein operatively connecting the motor to the compressor fan comprises operatively connecting an adjustable speed motor to the compressor fan, and wherein the method further comprises adjusting the speed of the motor in response to a change in a pressurization demand of the airframe.

9. The method of claim 6, wherein operatively connecting the motor to the compressor fan comprises operatively connecting an adjustable speed motor to the compressor fan, and wherein the method further comprises adjusting the speed of the motor in response to a change in a temperature demand of the airframe.

10. A method for providing supplemental power from a jet engine to a plurality of aircraft systems on a transport aircraft having a fuselage from which wings extend outwardly and the jet engine is configured to provide propulsive thrust, comprising an environmental control system, a wing anti-icing system and a landing gear system, the method for providing supplemental power comprising:

operatively connecting an electrical generator to the jet engine, the electrical generator configured to receive shaft work from the jet engine;

providing electrical power from the generator to a fan motor of an environmental control system configured to provide conditioned air to at least a portion of the fuselage in the absence of bleed air from the jet engine;

providing electrical power from a generator to a heating element of a wing ice protection system configured to at least reduce ice formation on a portion of a wing in the absence of bleed air from a jet engine;

the electric motor of the landing gear system is provided with electric power from an electric generator that is configured to operate the landing gear in the absence of bleed air from the jet engine.

11. The method of claim 10, wherein the generator operatively connected to the jet engine is a first generator, and wherein the method further comprises:

mounting an auxiliary power unit on the aircraft;

operatively connecting a second generator to the auxiliary power unit, wherein the second generator is configured to receive shaft work from the auxiliary power unit; and

the fan motor of the climate control system is supplied with electrical power from the second generator, and the climate control system is configured to supply conditioned air to at least a portion of the fuselage in the absence of compressed air from the auxiliary power unit.

12. The method of claim 10, wherein the generator operatively connected to the jet engine is a first generator, and wherein the method further comprises:

mounting an auxiliary power unit on the aircraft;

operatively connecting a second generator to the auxiliary power unit, wherein the second generator is configured to receive shaft work from the auxiliary power unit; and

the electric motor of the landing gear system is provided with electric power from the generator to a hydraulic pump configured to operate the landing gear without pneumatic power from the auxiliary power unit.

13. A system for providing conditioned air to a fuselage, the system being on an aircraft having a fuselage and a jet engine configured to provide propulsive thrust, the system comprising:

means for providing outside air to a compressor fan in an air stream in communication with the fuselage;

means for extracting electrical energy from the jet engine;

means for supplying at least a portion of the electrical energy to the compressor fan for flowing outside air from the compressor fan into the fuselage of the aircraft in the absence of bleed air from the jet engine.

14. The system of claim 13, further comprising means for providing at least a portion of the electrical power from the generator to a heating element of a wing ice protection system configured to at least reduce ice formation on a portion of the wing in the absence of bleed air from the jet engine.

15. The system according to claim 13, further comprising means for supplying at least a portion of the electrical power of the generator to an electric motor driven hydraulic pump of the landing gear system, the hydraulic pump being configured to operate the landing gear in the absence of bleed air from the jet engine.