| 序号 | 专利名 | 申请号 | 申请日 | 公开（公告）号 | 公开（公告）日 | 发明人 |
|---|---|---|---|---|---|---|
| 141 | HORIZONTAL TAIL LOAD OPTIMIZATION SYSTEM AND METHOD | US15797576 | 2017-10-30 | US20180043992A1 | 2018-02-15 | Vedad Mahmulyin |
| A method of controlling an elevator of an aircraft includes selecting a factor for increasing or decreasing a predetermined horizontal tail load alleviation (HTLA) authority limit for an elevator based on at least one aircraft parameter. The HTLA authority limit decreases with an increase in Mach number and/or airspeed. The method also includes computing an elevator position limit as a product of the HTLA authority limit and the factor, and moving the elevator to a commanded elevator position that is no greater than the elevator position limit. | ||||||
| 142 | System and method for optimizing horizontal tail loads | US14278868 | 2014-05-15 | US09878776B2 | 2018-01-30 | Vedad Mahmulyin |
| A method of controlling an elevator of an aircraft may include identifying a current stabilizer angle of incidence of a stabilizer of the aircraft. The stabilizer may include an elevator pivotably coupled to the stabilizer. The method may further include comparing the current stabilizer angle of incidence with a threshold stabilizer angle of incidence, and selecting an elevator position limit that is more restrictive if the current stabilizer angle of incidence is greater than or equal to the threshold stabilizer angle of incidence. The method may additionally include moving the elevator to a commanded elevator position that is no greater than the elevator position limit. | ||||||
| 143 | METHOD OF CONTROLLING AN ELECTRICAL TAXIING SYSTEM | US15607724 | 2017-05-30 | US20170351242A1 | 2017-12-07 | Clement GORCE |
| A method of controlling an aircraft electrical taxiing system, the method comprising the steps of: defining a target value (Ld_nmax) for an electrical parameter; generating a nominal force command (Cmd_nom) for the electrical taxiing system; in parallel with generating the nominal force command (Cmd_nom), using a processing system (2) to produce a maximum command force (Force_max) for the electrical taxiing system so that a real value of the electrical parameter reaches the target value (Ld_nmax), the processing system (2) comprising a regulator loop (4); and generating an optimized force command (Cmd_opt) for the electrical taxiing system equal to the smaller of the nominal force command and the maximum command force. | ||||||
| 144 | HAPTIC FEEDBACK FOR REALTIME TRAJECTORY CONSTRAINTS | US15507113 | 2015-09-15 | US20170285662A1 | 2017-10-05 | Igor Cherepinsky; Harshad S. Sane |
| A system for receiving feedback in a flight plan of a vehicle includes a haptic-enabled device comprising a crew seat with an inceptor mounted thereto; and a processor with memory having instructions stored thereon that, when executed by the processor, cause the system to: receive signals indicative of the flight plan for the vehicle; receive deviation signals indicative of a proposed deviation in a trajectory for the flight plan; and transmit signals to the haptic-enabled device representing trajectory constraints in the proposed deviation in response to the receiving of the deviation signals. | ||||||
| 145 | Motionless flight control surface skew detection system | US14675845 | 2015-04-01 | US09764853B2 | 2017-09-19 | Peter A. Padilla |
| A motionless skew detection system for an aircraft is disclosed, and includes a flight control surface of an aircraft wing, two drive mechanisms for operating the flight control surface, a first load sensor and a second load sensor for each of the two drive mechanisms, and a control module. Each of the two drive mechanisms are located on opposing sides of the flight control surface and each of the two drive mechanisms include at least a first linkage including a first outer surface and a second linkage including a second outer surface. The first load sensor is disposed along the first outer surface of the first linkage and the second load sensor is disposed along the second outer surface of the second linkage. The control module is in signal communication with the first load sensor and the second load sensor of each drive mechanism. | ||||||
| 146 | Active impact force/torque control for an electromechanical actuator | US14780344 | 2014-05-29 | US09758236B2 | 2017-09-12 | John David Neely; John Mendenhall White |
| A system that improves on known systems for reducing output torque by a motor in the event of a jam may include an electromechanical actuator (EMA), a motor configured to drive the EMA and a controller. The controller may be coupled to the motor and configured to receive a speed of the EMA and a position of the EMA. The controller may be further configured to determine whether a jam of the EMA is imminent or is occurring according to the EMA speed, EMA position, and a known range of motion of the EMA, and to provide an input signal to the motor to reduce a torque of the motor if a jam of the EMA is imminent or is occurring. | ||||||
| 147 | PILOT ACTIVATED TRIM FOR FLY-BY-WIRE AIRCRAFT | US15404804 | 2017-01-12 | US20170197705A1 | 2017-07-13 | Leonard M. Wengler, III; Kevin L. Bredenbeck; Matthew T. Luszcz; Matthew A. White; William Fell |
| A fly-by-wire aircraft and method of flying a fly-by-wire aircraft is disclosed. The aircraft includes a control system for flying the aircraft in one of a proportional ground control mode and a model following controls mode. A control device at a control interface of the aircraft selectively activates a trim follow up function in the control system. When flying the aircraft in a proportional ground control mode, trim follow up function can be activated. The control system can then transition into the model following controls mode with the trim follow up function activated to reduce transient behavior. | ||||||
| 148 | Method and system for controlling the flight of an aircraft | US14807014 | 2015-07-23 | US09663219B2 | 2017-05-30 | Mathieu Carton; Pierre Debusschere; David Chabe |
| The system comprises at least one lift generator element that is able to modify directly the lift of the aircraft and means for defining a deflection instruction upon actuation by a pilot of the aircraft of a control column of the aircraft generating a vertical load factor control value and applying it to an elevator and simultaneously defining a command instruction and applying it to the lift generator element to generate a direct lift. | ||||||
| 149 | PUSHING DEVICE, MOVING MECHANISM AND AIRCRAFT | US15134711 | 2016-04-21 | US20160311523A1 | 2016-10-27 | Ming LI; Yumin SUN |
| The present application relates to a pushing device, a moving mechanism and an aircraft. According to an aspect of the present application, a pushing device for a moving mechanism of an aircraft is provided, the moving mechanism including a primary moving device and an auxiliary moving device assisting the primary moving device, the pushing device including a support member and a pushing assembly supported by the support member, and the pushing assembly including a pushing element and an energy storage element. The pushing element is adapted to push a broken part of the auxiliary moving device to an offset position from a normal working position by means of energy from the energy storage element when the auxiliary moving device breaks. According to the present application, it is possible to provide an effective fault protection to the moving mechanism of the aircraft. | ||||||
| 150 | METHOD AND SYSTEM FOR CONTROLLING THE FLIGHT OF AN AIRCRAFT | US14807014 | 2015-07-23 | US20160023749A1 | 2016-01-28 | Mathieu CARTON; Pierre DEBUSSCHERE; David CHABE |
| The system comprises at least one lift generator element that is able to modify directly the lift of the aircraft and means for defining a deflection instruction upon actuation by a pilot of the aircraft of a control column of the aircraft generating a vertical load factor control value and applying it to an elevator and simultaneously defining a command instruction and applying it to the lift generator element to generate a direct lift. | ||||||
| 151 | Leading edge variable camber system and method | US14034987 | 2013-09-24 | US09180962B2 | 2015-11-10 | Matthew A. Moser; Mark J. Gardner; Michael R. Finn; Mark S. Good; Adam P. Malachowski; Monica E. Thommen; Stephen R. Amorosi; Dan Onu |
| A system for varying a wing camber of an aircraft wing may include a leading edge device coupled to the wing. The leading edge device may be configured to be actuated in an upward direction and a downward direction relative to a retracted position of the leading edge device. | ||||||
| 152 | AIRPLANE WING, AN AIRPLANE AND A FLAP SYSTEM | US14440272 | 2013-10-29 | US20150291275A1 | 2015-10-15 | Adrianus Marinus Franciscus Bastiaansen; Michel Schoonhoven |
| An airplane wing comprises a main wing and a flap system that has a flap at the trailing edge of the main wing. An elongate flap track member is connected to the main wing in such a manner that it can be moved substantially in its longitudinal direction and is guided by supporting bearing elements relative to the main wing between a forward retracted position and a rearward extended position. The flap is rotatably connected to the rear end of the flap track member in such a manner that it can rotate about a rotation axis that extends substantially parallel to the trailing edge of the main wing, so that the flap moves together with the flap track member when the flap track member is moved and so that the flap can be rotated about the rotation axis mechanically independently of the movement of the flap track member. The flap system comprises an actuator system having two actuators. The first actuator is connected to the main wing and has an engagement member that engages the flap or the flap track member for moving the flap together with the flap track member so that the flap track member is move able between its retracted position and its extended position. The second actuator is connected to the flap track member so that the second actuator moves together with the flap track member when the flap track member is moved by means of the first actuator. The second actuator has an engagement member that engages the flap for rotating the flap about the rotation axis. | ||||||
| 153 | Automatic throttle roll angle compensation | US12274756 | 2008-11-20 | US09085371B2 | 2015-07-21 | Frederick Charles Henry Blechen |
| A method, apparatus, and computer program product for adjusting a thrust for an aircraft. A level of thrust is dynamically identified using a drag on the aircraft needed to substantially maintain a speed of the aircraft in a turn to form an identified level of thrust during turning of the aircraft. The turn of the aircraft is performed using the identified level of thrust. | ||||||
| 154 | AIRCRAFT WEIGHT IDENTIFICATION USING FILTERED TRIM ESTIMATION | US14022623 | 2013-09-10 | US20150073626A1 | 2015-03-12 | Anthony Litwinowicz; Stephen Kubik; Steven W. Hong |
| Embodiments are directed to receiving, by a computing device comprising a processor, at least one control input associated with an aircraft, obtaining, by the computing device, a predicted response to the at least one control input by filtering on a trim position, wherein the predicted response is based on a model of the aircraft, obtaining, by the computing device, an actual response of the aircraft to the at least one control input, comparing, by the computing device, the predicted response and the actual response, and determining, by the computing device, at least one attribute based on the comparison | ||||||
| 155 | Method for predicting a horizontal stabilizer fault | US14012223 | 2013-08-28 | US08972101B2 | 2015-03-03 | Christopher Joseph Catt; Mark John Robbins |
| A method of predicting a horizontal stabilizer system fault in an aircraft, where the method includes receiving data relevant to a characteristic of the pitch of the aircraft during flight, comparing the received data to a reference pitch characteristic, predicting a fault in the horizontal stabilizer system based on the comparison, and providing an indication of the predicted fault. | ||||||
| 156 | Aerodynamic coefficient estimation device and control surface failure/damage detection device | US13575365 | 2011-03-16 | US08954208B2 | 2015-02-10 | Koichi Yamasaki |
| A highly reliable aerodynamic coefficient estimate can be computed, and computation of this aerodynamic coefficient estimate enables accurate detection of control surface failure/damage while reducing a discomfort for passengers. A deflection angle command signal generation means (5) generates a deflection angle command signal for estimating an aerodynamic coefficient indicating the aerodynamic characteristics of an airframe. A kinetic state quantity acquisition means (6) acquires a kinetic state quantity of the airframe that is obtained as a result of a control surface provided on the airframe being moved based on the deflection angle command signal. A candidate value calculation means (7) calculates candidate values for estimating the aerodynamic coefficient from the kinetic state quantity using two or more different estimations. An aerodynamic coefficient estimate determination means (8) determines an aerodynamic coefficient estimate based on the candidate values. | ||||||
| 157 | METHOD FOR PREDICTING A HORIZONTAL STABILIZER FAULT | US14012223 | 2013-08-28 | US20140229056A1 | 2014-08-14 | Christopher Joseph Catt; Mark John Robbins |
| A method of predicting a horizontal stabilizer system fault in an aircraft, where the method includes receiving data relevant to a characteristic of the pitch of the aircraft during flight, comparing the received data to a reference pitch characteristic, predicting a fault in the horizontal stabilizer system based on the comparison, and providing an indication of the predicted fault. | ||||||
| 158 | ACTUATOR-LINK ASSEMBLY MANUFACTURING METHOD, ACTUATOR-LINK ASSEMBLY DESIGNING METHOD, AND ACTUATOR-LINK ASSEMBLY | US14174661 | 2014-02-06 | US20140150605A1 | 2014-06-05 | Toshiaki OGAWA; Koji ITOH; Makoto NAGASHIMA |
| In a material determining step, the material constituting an actuator and the material constituting a link are determined such that at least one of the materials contains fiber reinforced plastic. In a computing step, a computation model that defines the relationship between a control surface, the actuator, and the link is used to compute the change in gain margin with the change in a rigidity ratio, which is the ratio of the rigidity of the link to the rigidity of the actuator. The rigidities of the actuator and the link are determined in a rigidity determining step based on a result of the above-described computation, the shapes of the actuator and the link are determined in a shape determining step, and the actuator and the link are formed in a formation step, and are assembled in an assembly step. | ||||||
| 159 | More electric flight control system onboard an aircraft | US12651603 | 2010-01-04 | US08740155B2 | 2014-06-03 | Marc Fervel; Alexandre Gentilhomme |
| An electric flight control system onboard an aircraft provided with flight control surfaces and controllers for controlling these surfaces, that include at least one local electro-hydraulic generator to supply hydraulic servocontrols connected to flight control surfaces. | ||||||
| 160 | MORE ELECTRIC FLIGHT CONTROL SYSTEM ONBOARD AN AIRCRAFT | US12651603 | 2010-01-04 | US20100170999A1 | 2010-07-08 | Marc FERVEL; Alexandre Gentilhomme |
| The invention relates to a more electric flight control system onboard an aircraft provided with flight control surfaces and means of controlling these surfaces, that comprises at least one local electro-hydraulic generator (HPP1, HPP2) to supply hydraulic servocontrols (34) connected to flight control surfaces (31, 32, 33). | ||||||
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