Electric-based secondary power system architecture for aircraft

Cross-reference with related applications The present application claims the priority of simultaneously pending US Provisional Patent Application No. 60 / 420,637 filed October 22, 2002, which is hereby incorporated by reference in its entirety. This application is hereby incorporated by reference in its entirety to US Pat. No. 6,526,775.

Technical field The following disclosure generally relates to secondary power systems for aircraft, and more specifically to electrical-based secondary power systems for aircraft.

background Traditional transport aircraft typically utilize pneumatic, hydraulic and electrical power from the main engine to support a variety of aircraft systems during flight. In addition, conventional transport aircraft typically utilize pneumatic and electrical power from on-board auxiliary power units (APUs) to support aircraft systems during ground operations. Aircraft air conditioning systems typically use the most secondary power in commercial transport aircraft. In conventional transport aircraft, these systems use hot / high pressure air (bleed air) extracted from the engine compressor stage. This air passes through the air conditioning pack and into the airframe to meet temperature, ventilation and pressurization requirements. The conditioned air is then expelled from the aircraft through an outflow valve or a normal cabin leak. During ground operation, the APU can provide bleed air from a separate shaft-driven load compressor or from a power section compressor. Similar to the bleeding from the main engine, the hot and high pressure air from the APU passes through the air conditioning pack to the aircraft.

Figure 1 schematically illustrates a traditional pneumatically based secondary power system architecture 100 configured according to the prior art. System architecture 100 may include a jet engine 110 (shown as first engine 110a and second engine 110b) to provide propulsion to an aircraft (not shown). In addition to thrust, the engine 110 may also provide hot / high pressure air to the bleed manifold 120 via the bleed outlet 112, individually identified as the first bleed port 112a and the second bleed port 112b. it can. The bleed air port 112 receives air from the compressor stage of the engine 110 and passes the air through a heat exchanger 114 (such as a precooler), and the heat exchanger 114 cools the air and then passes it through the bleed air manifold 120.

The high pressure air from the bleed manifold 120 supports most of the aircraft's secondary power requirements. For example, a portion of this air flows into air conditioning pack 140 (shown as first air conditioning pack 140a and second air conditioning pack 140b) that supplies regulated air to cabin 102 of Airframe 104. Air conditioning pack 140 includes a set of heat exchangers, regulating valves and air circulation equipment that regulate the air to meet the temperature, ventilation and pressurization requirements of cabin 102. Another portion of the air from the bleed manifold 120 flows to the turbine 160, which drives the high capacity hydraulic pump 168. The hydraulic pump 168 provides hydraulic power to the landing gear and other hydraulic systems of the aircraft. Yet another portion of this high pressure air is directed to the engine cowl anti-icing system 152 and the wing anti-icing system 150.

The wing protection system 150 includes a valve (not shown) that controls the flow of bleed air to the leading edge of the wing and a "piccolo" duct (similarly) that evenly distributes hot air along the protected area of the leading edge of the wing. (Not shown in) and. Supply hot bleed air to the slat when anti-icing of the leading edge slat is required Therefore, the telescope type duct can be used in the extended position. De-icing bleed air is expelled through holes in the underside of the wings or slats.

System architecture 100 may include APU130 as an alternative power source in addition to engine 110. The APU 130 is typically started by a DC start motor 134 with a battery 136. The APU 130 drives a compressor 138 that provides high pressure air to the bleed manifold 120 for engine start and other ground operations. To start the engine, high pressure air flows from the bleed manifold 120 to the starting turbine 154 operably coupled to each of the engines 110. As an alternative to the APU130, bleeding from one of the engines 110 in operation can be used to restart the other engine 110. As yet another alternative, an external air cart (not shown in Figure 1) can provide high pressure air for starting the engine on the ground.

The system architecture 100 may further include an engine-driven generator 116 operably coupled to the engine 110 and an APU-driven generator 132 operably coupled to the APU 130. During flight, the engine-driven generator 116 can be used with conventional electrical system loads such as fuel pump 108, motor-driven hydraulic pump 178, and various fans, waitroom systems, in-flight entertainment systems, lighting systems, aeroelectronic engineering systems, etc. Can be supported. The APU-powered generator 132 can support these functions on demand during ground operations and flight. Engine-driven generators 116 and APU-driven generators 132 are typically rated at 90-120 kVA and generate a voltage of 115 Vac. They can power transformer-rectifier units that convert 115Vac to 28Vdc for many of the electrical loads mentioned above. This power is largely distributed in electrical systems based on thermal circuit breakers and relays.

The system architecture 100 may additionally include an engine driven hydraulic pump 118 operably coupled to the engine 110. Hydraulic pump 118 provides hydraulic power to control surface actuators and other aircraft systems during flight. The pump 178, driven by an electric motor, can provide spare hydraulic power for ground maintenance work.

FIG. 2 is a schematic top view of the prior art aircraft 202, including the secondary power system architecture 100 of FIG. Aircraft 202 includes a forward electronics bay 210 that distributes power to a plurality of electrical loads 220 associated with the system architecture 100 described above. During flight, electronics bay 210 can receive power from engine generator 116 and APU. On the ground, electronics bay 210 can receive power from APU 130 or from external power supply 212 via outlet 213.

<p> One drawback of the secondary power system architecture 100 described above is the worst operating conditions (typically cruising) to ensure that sufficient airflow is available to meet system requirements at all times. It is sized for speed, high aircraft load, hot days, and failure of the bleed system of one engine). As a result, under typical operating conditions, the engine 110 provides bleed air at pressures and temperatures significantly higher than the demands of air conditioning pack 140 and other aircraft systems. To compensate, the precooler 114 and air conditioning pack 140 adjust the pressure and temperature to meet the requirements for airframe pressurization, ventilation and temperature control, lowering the values as needed. Therefore, a significant amount of energy is wasted by the precooler and regulating valve during this adjustment. Even under optimal conditions, a significant amount of energy extracted from the engine 110 is in the form of reduced heat and pressure that occurs in the duct system, valves, and other components associated with the bleed manifold 120 and the air conditioning pack 140. Wasted.</p>

<p> Overview The present invention is generally directed to secondary power systems for aircraft and methods for providing secondary power to aircraft systems. In one embodiment, an aircraft according to one aspect of the invention includes an airframe and a jet engine configured to provide propulsion to the aircraft. The aircraft also includes a generator operably coupled to a jet engine and an environmental control system. The environmental control system may include at least one compressor motor configured to receive power from a generator to provide outside air to the cabin in the absence of bleed air from the jet engine.</p><p> In another aspect of this embodiment, the aircraft may further include wings extending outward from the airframe and an electric wing anti-icing system. The electric wing ice protection system can be configured to receive power from the generator to at least reduce the formation of ice 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 can be a first generator, and the aircraft can further include an auxiliary power unit and a second generator. The second generator is operably coupled to the auxiliary power unit and can be configured to receive axial power from the auxiliary power unit. In this aspect, at least one compressor motor in the environmental control system can be configured to receive power from a second generator to provide outside air to the cabin in the absence of compressed air from the auxiliary power unit. ..</p><p> In another embodiment, the method for providing regulated air to the airframe of an aircraft is to provide a compressor fan with fluid communication with the airframe and a compressor fan operably coupled to the compressor fan with an electric motor. It may include driving steps. The method may further include the step of operably coupling the generator to the jet engine of the aircraft and the step of providing power from the generator to the electric motor to drive the compressor fan. In one aspect of this embodiment, the compressor fan can be driven to allow air to flow from outside the fuselage into the fuselage in the absence of bleed air from the jet engine.</p>

Detailed explanation The following disclosure describes a system and method for powering an aircraft system. To provide a complete understanding of the various embodiments of the invention, some details are given in the description below and in FIGS. 3-7. Other details describing well-known structures and systems often associated with aircraft and / or aircraft secondary power systems are described in the disclosure below so as not to unnecessarily obscure the description of the various embodiments of the invention. Not done.

Many of the details, dimensions, angles and other features shown in the drawings are merely exemplary embodiments of the present invention. Thus, other embodiments may have other details, dimensions and features without departing from the spirit or scope of the invention. In addition, yet another embodiment of the invention may be practiced without some of the details described below.

In the drawings, the same reference numbers identify the same, or at least generally similar, elements. To facilitate the description of any particular element, the leading digit of any reference number refers to the drawing in which that element was first introduced. For example, an element<u style="single">3</u>10 is first introduced and explained with reference to FIG.

FIG. 3 is a schematic diagram showing an electrical-based secondary power system architecture 300 configured according to an embodiment of the present invention. In one aspect of the invention, system architecture 300 includes a first engine 310a and a second engine 310b to provide propulsion to an aircraft (not shown). As described in more detail below, the first starter / generator 316a and the second starter / generator 316b are operably coupled to each of the engines 310 to power on multiple aircraft systems. It can be provided on a demand basis. The starter / generator 316 supports most of the aircraft functions traditionally performed by the bleed system described above in Figure 1. These functions may include, among other things, airframe air conditioning and pressurization, engine starting, and wing deicing.

In another aspect of this embodiment, the system architecture 300 further includes an APU 330 for powering the aircraft system during ground operation and flight, as needed. Power to start the PAU 330 can be provided by one or more of the aircraft battery 336, an external ground power source (not shown), or an engine-powered starter / generator 316. Power from the APU330 is provided by a first APU starter / generator 332a and a second APU starter / generator 332b, each operably coupled to the APU330.

In contrast to the traditional APU130 described above with reference to Figure 1, the APU330 provides only power to various aircraft systems. Therefore, it can be much simpler than the APU130, as it eliminates the need for all the components involved in pneumatic power transmission. This feature can result in a significant improvement in APU reliability and a reduction in required maintenance.

In yet another aspect of this embodiment, the system architecture 300 includes an environmental control system (ECS) having a first air conditioning pack 340a and a second air conditioning pack 340b. Air conditioning pack 340 is configured to provide regulated air to cabin 302 of Airframe 304 to meet temperature, pressure and air conditioning requirements. In one embodiment, the air conditioning pack 340 may be at least generally similar to one or more of the air conditioning systems disclosed in US Pat. No. 6,526,775, which is hereby incorporated by reference in its entirety. In another embodiment, the air conditioning pack 340 may include a variable speed electric compressor motor 380 configured to receive power from the engine 310 during flight and 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 intake 342. The fresh air is compressed and flows from the air conditioning pack 340 to the aircraft 304 to meet the pressurization and temperature control requirements of cabin 302. In one embodiment, the system architecture 300 is one or more variable speed fans that distribute air to different parts of the airframe 304 at different flow rates to meet the specific requirements of the airframe 304 at any given time point. (Not shown) can be included. Adjusting the power draw from the engine 310 in this way can further improve fuel efficiency.

With the variable speed compressor motor 380, cabin air pressure and airflow can be varied based on cabin volume, number of passengers, and / or desired cabin pressure altitude. For example, if lower cabin altitude (higher pressure) is desired, the electric ECS system of the present invention uses a variable speed compressor motor 380 to increase inflow and / or reduce outflow from airframe 304. , This can be dealt with. In general, conventional pneumatic systems cannot lower cabin altitude well below their design point (eg 8000ft). This is because these systems are usually sized for their design points. Another advantage of the electrical approach to air conditioning over the traditional pneumatic approach is that the energy extracted from the engine for the electrical approach is not wasted by the precooler and control valve in the air conditioning pack 340. Instead, the compressor motor 380 only draws sufficient power from the engine 310 as needed by the variable speed compressor to meet the immediate pressurization requirements of cabin 302. This real-time energy optimization can be extended to other power-using devices throughout the aircraft platform to increase fuel efficiency. As described below, for example, such equipment may include recirculation fans, restroom and lavatory ventilation fans, cargo heating, wing deicing, and hydraulic operation. Fuel savings can be increased simply by drawing in the required energy.

In another aspect of the invention, the system architecture 300 further includes a wing ice protection system 350 that utilizes power from the engine 310. The wing ice protection system 350 may be configured according to at least two embodiments of the invention to prevent, or at least reduce, the formation of ice on a portion of the wing 352. In the example of electric heating anti-icing, a heating element such as a blanket (not shown) can be coupled close to the inner portion of the leading edge of the wing or positioned in other ways. To prevent ice on the wing, the heating bracket can be sequentially energized to heat the leading edge of the wing so that any ice buildup will melt and / or disengage from the leading edge of the wing. This method can be significantly more efficient than conventional bleeding systems because the desired parts of the leading edge of the wing are heated sequentially rather than simultaneously. Therefore, the power draw for anti-icing is significantly reduced. In addition, in contrast to the bleed system, the wings do not have bleed outlet holes. As a result, aircraft drag and local noise are reduced compared to conventional systems.

The wing ice protection system 350 can also operate as an electromechanical system according to another embodiment of the present invention. In this embodiment, an electromechanical actuator (not shown) located inside the leading edge of the wing can be configured to vibrate the leading edge of the wing for a short period of time to allow any ice buildup to fall off. .. This embodiment may require significantly less power than the electric heating embodiment described above. In both embodiments, the wing protection system 350 can be divided into separate segments, which correspond to separate areas of the front edge of the wing or slats. That way, if part of the leading edge of the wing does not require anti-icing, that section of the wing anti-icing system 350 can be turned off, resulting in a further reduction in power demand from the engine. In addition, separate sections of the protection system can be circulated according to separate schedules, if desired, in order to adequately reduce ice while optimizing power use.

In another aspect of this embodiment, the starter / generator 316 provides power for the aircraft system when operating as a generator and axially powered for engine start when operating as a starter. Can be a dual function device. This electric start function can enhance the in-flight starting procedure of the engine 310 if one or more of the main engines 310 are stopped during normal flight operation. For example, a typical, high bypass ratio engine may be difficult to restart throughout the flight regime, as the wind turbine effect in flight may not provide sufficient torque. In contrast, the starter / generator 316 of the present invention provides additional starting torque from any number of power sources on the aircraft to assist the engine 310 during an in-flight restart. It is configured to receive power.

To start the engine 310, the starter / generator 316 can be operated as a synchronous start motor whose starting process is controlled by an engine start converter (not shown). The engine starting converter can provide controlled power (eg, adjustable voltage and frequency) to the starter / generator 316 during the starting process for optimum starting performance. The engine start converter can also function as a motor controller for the cabin pressurized compressor motor 380 and / or other variable speed motors on the aircraft. Similarly, the APU start converter (not shown) can function as a motor controller for other variable speed motors on the aircraft, such as the onboard Inert Gas Generating System (OBIGGS) 309. The power required to energize the starter / generator 316 to start the engine comes from an aircraft battery 336, APU330, ram air turbine (RAT) 367, fuel tank (not shown), or other source. obtain. The dual function aspect of the starter / generator 316 is not provided by the air turbine engine-starter 154 described above with reference to FIG. Unlike the starter / generator 316, the air turbine engine-starter 154 serves no purpose while the engine 110 is in operation.

In yet another aspect of this embodiment, the starter / generator 316 can be coupled directly to the gearbox of the engine 310 to operate at a frequency proportional to the engine speed (eg 360-700 Hz). This type of generator does not involve complex constant speed drive and may be the simplest and most efficient approach. As a result, in this embodiment, the starter / generator 316 can be more reliable and the cost of spare parts can be lower than conventional generators with complex constant speed drives. However, in other embodiments, other types of generators can be used. For example, in other embodiments where constant speed is desirable, a constant speed generator can be used.

In yet another aspect of the invention, the system architecture 300 includes a hydraulic system having a left channel, a right channel and a central channel. Hydraulic power for the left and right channels can be provided by an engine-driven hydraulic pump 318 operably coupled to each of the engines 310. In addition, a smaller, electrically motor driven hydraulic pump 319 can also provide hydraulic power to the left and right channels for ground operation and to supplement the engine driven pump 318. The engine-driven pump 318 can provide hydraulic power for flight control activators, stabilizer trim activators, and other functions. The hydraulic power for the central channel is provided by two large capacity hydraulic pumps 368 driven by an electric motor. The hydraulic pump 368 is driven by power from the engine 310, in contrast to the hydraulic pump 168 described above with reference to FIG. 1, which is driven by engine bleed air to meet peak hydraulic demand. Hydraulic pump 368 provides hydraulic power for the landing gear system 369 and other systems including flight control actuation, thrust reverse devices, brakes, front / trailing edge flaps, and front wheel maneuvering system (not shown). It is possible. In yet another aspect of the invention, throughout flight, only one of the hydraulic pumps 368 operates, while the other pump operates only during takeoff and landing.

In another aspect of the invention, the system architecture 300 provides multiple variable speed fuel pumps 308 for transferring fuel from the fuel tank 390 to one or more of the engines 310, or to another fuel tank (not shown). Can include. A typical commercial aircraft uses a fuel pump to transfer fuel from one area of the wing to another. This allows the aircraft to maintain a center of gravity that maximizes the performance of the aircraft. The traditional pneumatic-based system architecture 100 described above with reference to Figure 1 typically includes a constant speed fuel pump to transfer fuel from one tank to the next or to the engine 110. .. These constant speed fuel pumps are typically configured to operate at maximum pressure at all times, and the flow rate corresponding to this maximum pressure may be required to properly transfer fuel between tanks or to the engine 110. Even if it is rare, it is. For this reason, such fuel systems typically include a pressure regulator that only removes excess fuel pressure. This excess fuel pressure corresponds to the waste of engine power. In contrast, in the electrical-based architecture 300 of Figure 3, the speed of the fuel pump transfers fuel from one tank to the next, based on the amount of fuel required to be transferred and the speed required to perform the transfer. Can be modified to transfer to, optimizing the overall center of gravity of the aircraft under normal flight conditions. This ability to maintain an optimal center of gravity throughout the flight segment further improves aircraft range and fuel efficiency by simply drawing in the amount of power actually required by the fuel pump at any given time point. Can be done.

Many other systems can be incorporated into the system architecture 300 to further reduce the power extracted from the engine 310. For example, in one embodiment, an air conditioning pack 340 that adjusts the power extracted from the engine 310 based on the speed of the fan may use a speed adjustable or variable speed fan. In another embodiment, it is possible to use a resistance heater instead of the bleed air used in conventional systems to warm the cargo hold (not shown). These resistor heaters can have the ability to be pulse width modulated, better controlling temperature and further reducing energy consumption. Similarly, freight air conditioning systems can be configured to rely less on external air, but on recirculated air, for cabin cooling. In this way, the energy loss associated with the external air is eliminated and only the power required to cool the recirculated air is consumed.

As explained above with reference to FIG. 1, in the traditional system architecture 100, the engine 110 provides most of the secondary aircraft power in the form of pneumatics from the bleed. In contrast, in the system architecture 300 of the present invention, the engine 310 provides most of the secondary aircraft power in the form of electricity from the starter / generator 316. Eliminating the pneumatic bleed from the compressor portion of the engine 310 results in a more efficient engine design by reducing the compressor capacity requirements and improving the operating cycle. Furthermore, eliminating the labor-intensive bleeding system is expected to reduce the need for aircraft maintenance and increase aircraft reliability. This is because there are fewer engine components and fewer pneumatic ducts, precoolers and valves in the distribution system. In addition, the electrical-based system architecture 300 does not require any means of protection against duct rupture and above-appropriate temperatures.

A further advantage of the electrical-based system architecture 300 is that it can utilize motor control to regulate individual loads and extract only the minimum amount of power required from the engine 310 under any operating conditions. is there. These loads are adjustable rather than simply turned on or off, so less power is drawn from the engine 310. The ability to adjust power consumption for any power load can directly improve the fuel efficiency of an aircraft. Other benefits associated with the electrical-based system architecture 300 are real-time power extraction optimization, and non-recurring engineering potential related to eliminating waste associated with engine extraction, improving air quality, and licensing multiple engine extraction systems. Reduction can be included.

The system architecture 300 described above with reference to FIG. 3 includes two engines and two starters / generators for each engine, but in other embodiments, the system architecture configured according to the present invention It is possible to include more or less engines, with more or less generators, depending on the requirements of the particular application. For example, in another embodiment, a system architecture configured according to the present invention may include four jet engines, each of which may include one starter / generator. In yet another embodiment, the system architecture configured according to the present invention can include a single engine with two or more generators or starters / generators. Thus, the invention is not limited to aircraft with a particular number of engines or starters / generators.

FIG. 4 is a schematic top view of an aircraft 402 including the electrical based system architecture 300 of FIG. 3 configured according to an embodiment of the present invention. Aircraft 402 may include four starters / generators 316 coupled to two engines 310 and two starters / generators 332 coupled to APU 330 mounted on the tail of aircraft 402. In one aspect of this embodiment, the aircraft 402 is further configured to receive 115 Vac or 230 Vac of power from external power sources 412a and 412b, respectively. 2 can include (identified as ground outlet 413b).

In another aspect of the invention, aircraft 402 may include forward electrical equipment bay 410a and rear electrical equipment bay 410b. The four remote power distribution units (RPDUs) 424a-d can distribute power from device bay 410 to multiple system loads 420 associated with system architecture 300. RPDU424 can be largely based on solid-state power controllers instead of traditional thermal circuit breakers and relays.

FIG. 5 is a schematic diagram of an aircraft power distribution system 500 configured according to an embodiment of the present invention. In one aspect of this embodiment, the power distribution system 500 includes a first generator 516a and a second generator 516b operably coupled to the aircraft engine 510. In one embodiment, the first generator 516a and the second generator 516b may be high voltage AC generators (eg, 230Vac generators). In another embodiment, one of the two generators 516 may be a high voltage DC generator (eg ± 270 Vdc generator). The AC generator 516a can provide power to aircraft equipment that is not affected by the supply frequency. The DC generator 516b can power these components of an aircraft system, including variable speed motors. In other embodiments, the generator 516 may be another type of generator. For example, in another embodiment, both generators 516 may be AC generators. In this embodiment, the DC power requirements of the system can be met with a suitable AC-DC converter. In another embodiment, both generators 516 may be DC generators, and the AC power requirements of this system can be met with a suitable DC-AC converter.

In another aspect of this embodiment, the power distribution system 500 is further configured to be powered by a first bus 515a configured to be powered by a first generator 516a and a second generator 516b. Also includes a second bus 515b. In one embodiment, the first bus 515a can be a high voltage AC bus, such as a 230 Vac bus, configured to directly power a plurality of rated AC loads 550. Such loads may be associated with wing anti-icing equipment, hydraulic pumps, fuel pumps, wait room systems, and the like. In addition, the first bus 515a can also power the third bus 515c directly via the step-down transformer 522. In one embodiment, the third bus 515c can be a lower voltage AC bus, such as a 115Vac bus. The third bus 515c can power a plurality of underrated AC device loads 544 on an aircraft via a plurality of RPDUs (identified as a first RPDU 524a and at least a second RPDU 524b). Such a low rated load 544 may be associated with in-flight entertainment systems, internal and external lighting systems, sensor heaters, and the like.

In yet another aspect of this embodiment, the second bus 515b can power a plurality of variable speed motors 552 on an aircraft. Such motors may include cabin pressurization compressors, environmental control system fans, gas or air circulation ECS packs, large hydraulic pumps, flight actuators and the like. The use of high voltage DC systems can avoid potential harmonic distortion issues often associated with motor controllers and provide a means to address the regenerative energy often associated with electrostatic actuation devices. Can be done. In addition, the use of high voltage DC systems can also provide significant weight savings through the use of lightweight DC generators and the elimination of harmonic distortion processors and regenerated energy absorbers.

FIG. 6 is a schematic diagram of an aircraft power distribution system 600 configured according to another embodiment of the present 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 APU630. The power distribution system 600 further includes three AC generators 616 (identified individually as first AC generator 616a, second AC generator 616b and third AC generator 616c) and three DC generators. It may include a machine 618 (individually identified as a first DC generator 618a, a second DC generator 618b and a third DC generator 618c). The first AC generator 616a and the first DC generator 618a may be operably coupled to the first engine 610a. Similarly, the second AC generator 616b and the second DC generator 618b may be operably coupled to the second engine 610b. A third AC generator 616c and a third DC generator 618c can be operably coupled to the APU630. The third AC generator 616c can provide power from the APU630 to the two AC buses 615a to supply AC loads (not shown) as needed during ground operation and flight. A third DC generator, 618c, operably coupled to the APU630, provides power to two DC buses, 615b, with variable speed motors (also not shown) as needed during ground operation and flight. Can be supplied to. In addition, the third AC generator 616c can also power the two DC buses 615b via the AC-DC converter 624. Each of the two DC buses 615b may be operably connected to the corresponding motor controller 660, individually identified as the first motor controller 660a and the second motor controller 660b. The motor controller 660 is individually identified as the cabin pressurized compressor 680 (individually identified as the first compressor 680a and the second compressor 680b) or the engine starting circuit 662 (as the first starting circuit 662a and the second starting circuit 662b). Identified in) It can be configured to selectively provide power to either. In another aspect of this embodiment, the power distribution system 600 may include a first electrical outlet 613a and a second electrical outlet 613b that are configured to receive power from an external ground power source. In one embodiment, the first outlet 613a can be configured to receive 115 Vac of power from a terrestrial power source and the second outlet 613b can be configured to receive 230 Vac of power from an external terrestrial power source. ..

In one embodiment, the engine 610 can be started using the high voltage (eg 230 Vac) ground power received through the second outlet 613b. In this embodiment, the motor controller 660 is switched so that the power from the DC bus 615b is directed to the corresponding engine starting circuit 662. This power is directed to the corresponding AC generator 616 and is used to operate the AC generator 616 as a synchronous motor to start the engine of the corresponding engine 610 for start-up. Once the engine 610 is started, the motor controller 660 is switched back and forth to provide power to the cabin pressurizing compressor 680.

FIG. 7 is a schematic diagram of a power distribution system 700 having only an AC generator, according to another embodiment of the present 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, the first AC generator 716a and the second AC generator 716b are operably coupled to the first engine 710a, the third AC generator 716c and the fourth AC generator. The 716d is operably coupled to the second engine 710b, and the fifth AC generator 716e and the sixth AC generator 716f are operably coupled to the APU730. To meet the DC voltage requirement, the high voltage AC power from the engine 710 and APU 730 is high by one or more AC-DC converters such as the automatic transformer rectifier unit (ATRU) 724, which receives AC power from the AC bus 715. Can be converted to voltage DC power. With the use of ATRU724, the power distribution system 700 will provide both high voltage AC and DC power to support traditional 115Vac and 28Vdc bus architectures. In addition, AC power from one or more of AC generators 716 can also be converted to DC power for the 28Vdc bus 719 by the untuned transformer rectifier unit 725 and the tuned transformer rectifier unit 726. In another aspect of this embodiment, two AC generators 716 are coupled to each of the engines 710 so that both AC generators 716, if desired, as synchronous starting motors for additional engine starting power. Will be used. As described above with reference to FIG. 3, the embodiments of the invention described above with reference to FIGS. 5-7 are limited to the specified number of engines and / or starters / generators shown. It also applies to other quantities of engines and starters / generators with different configurations.

8A-8C are schematic views showing a power distribution system 800 having an engine starting circuit, configured according to an embodiment of the present invention. First, referring to FIG. 8A, in one aspect of this embodiment, the power distribution system 800 includes a generator 816 operably coupled to the engine 810 and a compressor motor 880 operably coupled to the environmental control system 840. including. The generator 816 can provide power to the motor controller 860 via the AC bus 815a, the AC-DC converter 824, and the high voltage DC bus 815b. During normal operation, as shown in FIG. 8A, the motor controller 860 can selectively direct power to the compressor motor 880 for the operation of the ECS 840.

FIG. 8B shows the engine starting configuration of the power distribution system 800. Since the engine 810 is not initially operating in this configuration, power is provided to the motor controller 860 by an alternative AC power source 830 rather than the generator 816. In one embodiment, the alternative power supply 830 may include an APU or an external power supply. In one aspect of this embodiment, the motor controller 860 selectively directs power from the alternative power supply 830 to the engine starting circuit 862. The engine start circuit 862 provides power to the generator 816, which is configured to act as a synchronous motor to start the engine 810.

Figure 8C shows another engine start configuration for the power distribution system 800. Here, the power to start the engine 810 is provided by the battery 836. In one aspect of this embodiment, the motor controller 860 selectively opens a circuit on the high voltage DC bus 815b so that it can receive power from the battery 836. After connecting to the battery 836, the motor controller directs power to the generator 816 via the engine starting circuit 862 as described above.

From the above, it is understood that while certain embodiments of the invention have been described herein for illustration purposes, various modifications may be made without departing from the spirit and scope of the invention. Will. Therefore, the present invention is not limited except as claimed in the accompanying claims.

<figref num="1">It is a schematic diagram of a conventional pneumatically based secondary power system architecture constructed according to the prior art.</figref><figref num="2">It is a schematic top view of a conventional aircraft having the system architecture of FIG.</figref><figref num="3">FIG. 5 is a schematic diagram showing an electrical-based secondary power system architecture configured according to an embodiment of the present invention.</figref><figref num="4">FIG. 3 is a schematic top view of an aircraft having the system architecture of FIG. 3 configured according to an embodiment of the present invention.</figref><figref num="5">It is the schematic of the aircraft power distribution system configured according to one Embodiment of this invention.</figref><figref num="6">FIG. 5 is a schematic diagram of an aircraft power distribution system configured according to another embodiment of the present invention.</figref><figref num="7">FIG. 5 is a schematic representation of an aircraft power distribution system having only an AC generator, according to yet another embodiment of the present invention.</figref><figref num="8A">It is a schematic diagram which shows the electric power distribution system which has an engine start circuit, which was constructed according to one Embodiment of this invention.</figref><figref num="8B">It is a schematic diagram which shows the electric power distribution system which has an engine start circuit, which was constructed according to one Embodiment of this invention.</figref><figref num="8C">It is a schematic diagram which shows the electric power distribution system which has an engine start circuit, which was constructed according to one Embodiment of this invention.</figref>