| 序号 | 专利名 | 申请号 | 申请日 | 公开（公告）号 | 公开（公告）日 | 发明人 |
|---|---|---|---|---|---|---|
| 41 | Aircraft electrical system tester | US10109 | 1979-02-07 | US4251765A | 1981-02-17 | Shawn P. Mears |
| A portable tester for trouble shooting aircraft electrical systems, particularly condition warning systems including a housing having a plug, adapted to connect a wiring harness from the master controlling display unit, and circuitry which simulates various conditions causing the display panel to react in a known manner. Failure of the display panel to react correctly indicates electrical malfunction. | ||||||
| 42 | ELECTRIC SYSTEM STABILIZING SYSTEM FOR AIRCRAFT | EP13825807.4 | 2013-07-29 | EP2880734A1 | 2015-06-10 | IWASHIMA, Atsushi; SUGIMOTO, Kazushige; MATSUO, Kazuya; BREIT, Joseph, S.; NOZARI, Farhad |
| A power stabilizing device in a system includes as a portion of a power converter section, a second PWM converter provided between a second primary AC bus and a second DC bus in an electric system and configured to perform mutual conversion between DC power and AC power. A power stabilizing control section in the power stabilizing device controls charging and discharging of a secondary battery based on a voltage and a frequency in the second primary AC bus, thereby stabilizing the electric system. | ||||||
| 43 | System and method for detecting faults in an aircraft electrical power system | EP04016951.8 | 2004-07-19 | EP1501167A3 | 2005-08-03 | Kohlmeier-Beckman, Carsten, Dipl.-Ing.; Kessler, Jens |
An aircraft electrical power system includes a monitored load (e.g. group of lightbulbs) connected through a sub_system power unit and a power distribution center to a power bus. A current monitor of the power distribution center is able to detect a total failure of the sub_system power unit and its connected monitored load, but is not able directly to detect the failure of a single load component (e.g. single lightbulb). The sub_system power unit includes a fault detector able to detect a fault or failure of a single load component (e.g. single lightbulb) in the monitored load. When the fault detector detects such a fault, the sub_system power unit generates a modulated current signal in the current drawn by the sub_system power unit, with a magnitude greater than the detection threshold of the power distribution center, which thus recognizes the detected fault based on the modulated current signal. |
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| 44 | ELECTRIC SYSTEM STABILIZING SYSTEM FOR AIRCRAFT | EP13825807 | 2013-07-29 | EP2880734A4 | 2016-04-27 | IWASHIMA ATSUSHI; SUGIMOTO KAZUSHIGE; MATSUO KAZUYA; BREIT JOSEPH S; NOZARI FARHAD |
| A power stabilizing device in a system includes as a portion of a power converter section, a second PWM converter provided between a second primary AC bus and a second DC bus in an electric system and configured to perform mutual conversion between DC power and AC power. A power stabilizing control section in the power stabilizing device controls charging and discharging of a secondary battery based on a voltage and a frequency in the second primary AC bus, thereby stabilizing the electric system. | ||||||
| 45 | Method for high fidelity modeling of an aircraft electrical power system | EP13187576.7 | 2013-10-07 | EP2738630B1 | 2016-08-10 | Huang, Hao; Jia, Xiaochuan |
| 46 | SYSTEME ELECTRIQUE DE DEMARRAGE DES MOTEURS D'UN AERONEF. | EP10706013.9 | 2010-01-08 | EP2377235B8 | 2015-12-02 | DE WERGIFOSSE, Eric |
| 47 | SYSTEME ELECTRIQUE DE DEMARRAGE DES MOTEURS D'UN AERONEF. | EP10706013.9 | 2010-01-08 | EP2377235B1 | 2015-09-02 | DE WERGIFOSSE, Eric |
| 48 | ELECTRIC SYSTEM STABILIZING SYSTEM FOR AIRCRAFT | EP13846974.7 | 2013-07-29 | EP2880738A2 | 2015-06-10 | IWASHIMA, Atsushi; SUGIMOTO, Kazushige; MATSUO, Kazuya; BREIT, Joseph, S.; NOZARI, Farhad |
| An electric system of an aircraft includes a power stabilizing device connected to a primary AC bus and a secondary battery. The secondary battery has a rated voltage which allows the secondary battery to absorb regenerative power from a control surface actuator. Based on a voltage and a frequency in the primary AC bus, charging/discharging of the secondary battery is controlled to stabilize the electric system. | ||||||
| 49 | SYSTEME ELECTRIQUE DE DEMARRAGE DES MOTEURS D'UN AERONEF. | EP10706013.9 | 2010-01-08 | EP2377235A2 | 2011-10-19 | DE WERGIFOSSE, Eric |
| The invention relates to an electrical system for starting up an engine (18, 20, 22), including: an AC/DC rectifier (12), supplied with power by an AC power network (14) in order to generate a first DC voltage Vdc; a DC/AC conversion module (12) for generating, on the basis of said first DC voltage Vdc, an AC voltage for starting up the engine, comprising n phase k inverters arrange in parallel (k>1) and generating power at least two times less than a maximum power Pmax that is required to start up the engine. The two power supply lines of each inverter are connected to an electronic protection device that receives the first DC voltage Vdc, and the n outputs of each inverter generate the AC voltage for starting up the engine by means of n inductors in series. | ||||||
| 50 | A GROUND FAULT INTERRUPT CIRCUIT APPARATUS FOR 400-Hz AIRCRAFT ELECTRICAL SYSTEMS | PCT/US1998/019515 | 1998-09-17 | WO00079667A1 | 2000-12-28 | |
| A resettable ground fault circuit interrupter most advantageously suited for use with 400-Hz aircraft electrical systems and designed to interrupt the AC power between an aircraft electrical system and one or more electrical appliances being operated therein when a ground fault occurs. The circuit relies upon an imbalance sensor (L1) which is implemented in the preferred embodiment disclosed herein as a multi-turn coil of a transformer through which wires connected to an AC input terminal are threaded so that an imbalance resulting from a ground fault can be detected by the coil. The signal generated by the coil as a result of such an imbalance is then amplified (U1) and applied to a thyristor (Q2), the conducting or non-conducting state of which determines whether or not the circuit is in a configuration for interrupting the AC signal or not interrupting the AC signal. Light emitting diodes (OP1) are utilized to indicate the interruption status of the circuit. A reset switch is provided to reset the thyristor after an imbalance ceases because of removal of a ground fault that had previously occurred. | ||||||
| 51 | Aircraft electrical system and associated management method | US13871404 | 2013-04-26 | US09487166B2 | 2016-11-08 | Renaud Loison; Olivier Savin |
| The system according to the invention comprises: an electrical network, and a regulating assembly. The regulating assembly includes a secondary electric power source, a conversion assembly for an additional electric power injected on the electrical network, the conversion assembly being able, in a first configuration, to consume the additional electric power present on the network to create a supply fluid of the secondary source. The assembly includes a reservoir for each supply fluid, to collect the supply fluid produced by the conversion assembly. | ||||||
| 52 | AIRCRAFT ELECTRICAL SYSTEM AND ASSOCIATED MANAGEMENT METHOD | US13871404 | 2013-04-26 | US20130285443A1 | 2013-10-31 | RENAUD LOISON; OLIVIER SAVIN |
| The system according to the invention comprises: an electrical network (20), and a regulating assembly (24). The regulating assembly (24) includes a secondary electric power source (58), a conversion assembly (60) for an additional electric power injected on the electrical network (20), the conversion assembly (60) being able, in a first configuration, to consume the additional electric power present on the network (20) to create a supply fluid of the secondary source (58). The assembly (60) includes a reservoir (62A, 62B) for each supply fluid, to collect the supply fluid produced by the conversion assembly (60). | ||||||
| 53 | Method and apparatus for testing aircraft electrical systems | US12744955 | 2007-11-29 | US08564300B2 | 2013-10-22 | Sven Knoop; Mike Galinski |
| A method, and corresponding apparatus, for testing an aircraft control system is disclosed. The method includes simultaneously coupling a test device to a plurality of separate test points in an aircraft control system, selecting each test point individually such that the test device is enabled for electrical connection with the selected test point, conducting a test on the control system at each selected test point using the test device, detecting a signal or voltage at the selected test point, and indicating a result of the test at the selected test point to an operator. | ||||||
| 54 | ELECTRICAL SYSTEM FOR STARTING UP AIRCRAFT ENGINES | US13143848 | 2010-01-08 | US20110273011A1 | 2011-11-10 | Eric de Wergifosse |
| An electrical system (18, 20, 22) for starting an engine, the system comprising: an AC/DC rectifier (12) powered by an AC power network (14) to deliver a first DC voltage Vdc, and a DC/AC converter module (16) for delivering an AC voltage for starting the engine from said first DC voltage Vdc, the system comprising k n-phase inverters arranged in parallel (k>1) and each delivering power no greater than half a maximum power Pmax required for starting the engine, and the two power supply lines of each of the inverters are connected to an electronic protection device receiving the first DC voltage Vdc, and the n outputs of each of the inverters delivering the AC voltage for starting the engine via n respective series inductors. | ||||||
| 55 | METHOD AND APPARATUS FOR TESTING AIRCRAFT ELECTRICAL SYSTEMS | US12744955 | 2007-11-29 | US20110050244A1 | 2011-03-03 | Sven Knoop; Mike Galinski |
| The present invention provides a method and apparatus for testing an aircraft control system. The method comprises the steps of: coupling a test device (1) to a plurality of separate test points in an aircraft control system simultaneously; selecting each of the plurality of test points individually, such that the test device is enabled for electrical connection with the selected one of the plurality of test points; conducting a test on the control system at each of the plurality of test points selected with the selector means (20,22) using the test device, including detecting a signal or voltage at the test point selected with the selector means; indicating a result of the test at the selected test point to an operator. | ||||||
| 56 | Power supply unit for use with an aircraft electrical system | US12052821 | 2008-03-21 | US07750496B2 | 2010-07-06 | John D. Cottingham, III; Scott A. Davis; James R. Berliner; Robert A. Lockmiller |
| A power supply unit that provides uninterrupted power to an aircraft power bus during switchovers of the aircraft bus from one power source to another. The power supply unit includes circuitry to distinguish between loss of bus power due to switchovers between power sources and loss of bus power due to shutdown of the aircraft. The power supply unit includes a low voltage lockout circuit to interrupt the supply of power when the aircraft bus voltage falls below a minimum value. | ||||||
| 57 | Electrical systems architecture for an aircraft, and related operating methods | US11944199 | 2007-11-21 | US07687927B2 | 2010-03-30 | Mark S. Shander; Richard A. Cote; Michael L. Drake; Howard Carter, III; John T. Peters; Casey Y. K. Ng; Michael S. Hoag; John T. Paterson |
| An electrical architecture for an aircraft is provided. The electrical architecture is particularly suitable for relatively small, compact, and lightweight aircraft. In one embodiment, the electrical architecture includes an electrical generator component coupled to the aircraft engine, and an air compression system coupled to the electrical generator component. The electrical generator component is configured to receive mechanical power from the engine and to generate a constant frequency AC electrical power from the engine mechanical power, and the air compression system is configured to receive the constant frequency AC power as an input and, in response thereto, produce a pressurized air output having variable characteristics (for example, a variable flow rate or a variable air pressure). | ||||||
| 58 | POWER SUPPLY UNIT FOR USE WITH AN AIRCRAFT ELECTRICAL SYSTEM | US12052821 | 2008-03-21 | US20080238191A1 | 2008-10-02 | John D. Cottingham; Scott A. Davis; James R. Berllner; Robert A. Lockmiller |
| A power supply unit that provides uninterrupted power to an aircraft power bus during switchovers of the aircraft bus from one power source to another. The power supply unit includes circuitry to distinguish between loss of bus power due to switchovers between power sources and loss of bus power due to shutdown of the aircraft. The power supply unit includes a low voltage lockout circuit to interrupt the supply of power when the aircraft bus voltage falls below a minimum value. | ||||||
| 59 | Auxiliary electrical generating system for jet aircraft | US3662975D | 1970-09-21 | US3662975A | 1972-05-16 | DRISKILL ROBERT E |
| An auxiliary electrical generating system for jet aircraft wherein compressed air is withdrawn from a jet engine and used to operate a pneumatic turbine which in turn drives a generator. An ancillary gas source may be included for supplying compressed gas to the turbine in the event of engine failure. This abstract is neither intended to define the invention of the application which, of course, is measured by the claims, nor is it intended to be limiting as to scope of the invention in any way.
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| 60 | METHOD FOR INSTALLING AN ELECTRIC SYSTEM INTO AN AIRCRAFT FUSELAGE | US14016829 | 2013-09-03 | US20140167500A1 | 2014-06-19 | Eckart Frankenberger |
| A method for installing an electric system into an aircraft fuselage. The object to provide a method for simply and quickly installing an electric system into an aircraft fuselage, wherein the electric system is as little heavy and space consuming as possible, and wherein a wide range of customized alternatives of the electric system is possible, is achieved by the steps of providing a fuselage element extending along an aircraft longitudinal axis and including an outer skin and an interior surface, providing a foil sheet of an electrically insulating substrate material, attaching said foil sheet on the inner side of said fuselage element, applying particles of electrically conductive material onto the inner surface of said attached foil sheet opposite the outer skin in a predetermined pattern, such that the accumulated particles of electrically conductive material form electric conductor elements along said inner surface of said foil sheet. | ||||||
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