Abstract: In an embodiment, a method of managing force equalization (FEQ) on a vehicle utilizing redundant actuation systems for one or more control surfaces includes determining, via a force sensor, a measured force applied by a first actuation system in relation to a control surface, where the control surface is redundantly serviced by a plurality of actuation systems. The method also includes updating a measured-force time series for the first actuation system with the measured force. The method also includes analyzing movement over at least a portion of the measured-force time series. The method also includes identifying a force-fight cycle in the measured-force time series. The method also includes indicating the force-fight cycle in cumulative force-fight cycle data for the first actuation system.

Abstract: An adaptive flight control system for controlling the pitch of blades of a propulsive propeller of a hybrid helicopter as a function of the return value of the pitch. The adaptive flight control system comprises control means supplying the pitch control order, a piloting member controlling variation of the pitch, and piloting means applying a control gain in order to transform the control order into a setpoint, and transmitting the setpoint to the piloting member. The piloting means include information return means applying a return gain that is variable to the return value of the variation of the pitch to the piloting means, also modifying the control gain as a function of the return value.

Abstract: A system of an aircraft includes a system pump and a booster pump. The system pump is configured to provide hydraulic fluid to the hydraulic system at a first working pressure. The booster pump is configured to supply hydraulic fluid to at least one boostable actuator at a second working pressure higher than the first working pressure. The boostable actuator is operatively coupled to and configured to actuate at least one flight control surface of an aircraft. The booster pump is configured as a high-pressure accumulator and an accumulator energizer, or as a hydraulic actuator pump and a variable speed motor.

Abstract: A hydraulic system of an aircraft may include a system pump configured to provide hydraulic fluid to the hydraulic system at a first working pressure. The hydraulic system may further include a booster pump configured to supply hydraulic fluid to at least one boostable actuator at a second working pressure higher than the first working pressure. The boostable actuator may be operatively coupled to and configured to actuate at least one flight control surface of an aircraft.

Abstract: A flap skew detection system is configured to detect flap skew of one or more flaps moveably secured to one or more wings of an aircraft. The flap skew detection system includes a flap support assembly that couples a flap to a wing. The flap support assembly includes a fixed portion that is configured to secure to the wing, a moveable portion that is moveably coupled to the fixed portion and configured to securely support the flap, and a link moveably coupled to the fixed portion and the moveable portion. The link includes a cylinder defining an internal chamber including a hydraulic fluid chamber, a piston having a piston head within the internal chamber, and hydraulic fluid retained within the hydraulic fluid chamber. A pressure detector is fluidly coupled to the hydraulic fluid chamber. The pressure detector is configured to detect a fluid pressure of the hydraulic fluid pressure within the hydraulic fluid chamber.

Abstract: Methods and apparatus for operating flight control systems of aircrafts are disclosed. An example apparatus includes a flight control system including a processor to: based on data from first and second sensors, determine first and second values; based on the first and second values, determine a location of a jam in the flight control system, the location of the jam being determined based on a summation of the first and second values.

Abstract: An actuating device to displace a part of an empennage of a helicopter, including a first part including a first casing fixed to a frame of the helicopter, a first element mounted in rotation with respect to the first casing and a first motor the first element. A second part includes a second casing connected to the part of the empennage, a second element mounted in rotation with respect to the second casing and a second motor driving the second element. A connecting element connects the first part with the second part, being mobile in translation with respect to the first element. Rotation of the first element causes a translation of the connecting element, and is mobile in translation with respect to the second element. Rotation of the second element causes a translation of the connecting element with respect to the second element.

Abstract: According to one embodiment, a stability augmentation system includes a master linkage, a stability augmentation motor, and three linkages. A first linkage is coupled to the master linkage and operable to receive movements representative of pilot commands from a pilot command system. A second linkage is coupled between the stability augmentation motor and the master linkage and operable to receive movements representative of augmentation commands from the stability augmentation motor. A third linkage is coupled to the master linkage and operable to transmit movements representative of blade position commands to a blade control system in response to the movements representative of pilot commands and the movements representative of augmentation commands.

Abstract: A plurality of hydraulically operated actuators that drive a control surface and a plurality of control systems that respectively control the operation of the actuators are included. Each of the control systems includes a control valve, a state switching valve and a test orifice circuit. The test orifice circuit includes a test orifice used for testing the performance of an orifice. The state switching valve is provided such that its position can be switched between an actuator connection position, a damping position and a test orifice position.

Abstract: According to one embodiment, a stability augmentation system includes a master linkage, a stability augmentation motor, and three linkages. A first linkage is coupled to the master linkage and operable to receive movements representative of pilot commands from a pilot command system. A second linkage is coupled between the stability augmentation motor and the master linkage and operable to receive movements representative of augmentation commands from the stability augmentation motor. A third linkage is coupled to the master linkage and operable to transmit movements representative of blade position commands to a blade control system in response to the movements representative of pilot commands and the movements representative of augmentation commands.

Abstract: Some embodiments relate to a method of controlling a flight of a flight vehicle according to a first mode of operation and changing the mode of operation to a second mode of operation having a different bandwidth than the first mode of operation. Other embodiments relate to a flight-control system for a flight vehicle configured to control a flight of a flight vehicle according to a first mode of operation and to control the flight of the flight vehicle according to a second mode of operation to use less energy than the first mode of operation. Other embodiments relate to a control actuation system configured to control positions of aerodynamic elements in a flight vehicle in response to commands from a guidance system according to a first mode of operation and to change the mode of operation to a second mode of operation having a different bandwidth than the first mode of operation.

Abstract: According to one embodiment, a bracket configured to hold a pressure transducer comprises a body, a face plate, and one or more pressure fitting connectors. The body has a rear portion and a front portion that meet to form an angle less than 180 degrees. The face plate is coupled to the front portion of the body, and has a surface defining a plate aperture configured for a cable to pass through. The pressure fitting connectors are coupled to the rear portion of the body. Each pressure fitting connector is configured to receive a pressure hose.

Abstract: A hydraulic system includes at least one hydraulic circuit with at least one pump, one or more consumer equipment units, at least one HP discharge line comprising at least one bypass, the bypass being located between the pump and the units, at least one return line including at least one bypass, the bypass being located between a tank and the units. The hydraulic system further includes at least one closure element on the at least one HP line between the bypass and the units, and at least one nonreturn valve on the at least one return line between the bypass and the units. The closure element is placed automatically in the closed position when a break of the HP line is detected downstream of the closure element of the HP line.

Abstract: A control surface drive system having a plurality of actuator assemblies are coupled to first and second supply lines and to a return line. The first and second supply lines are connected to a source of hydraulic fluid. At least one of the actuator assemblies has a hydraulic actuator movably connectable to an aircraft control surface. A flow control assembly is connected to the return line and to at least one of the first and second supply lines. A bypass line is in fluid communication with the first and second supply lines and positioned to recycle the hydraulic fluid from one of the first and second supply lines back into the other one of the first and second supply lines when the hydraulic actuator moves toward the first position. A computer controller operatively interconnects the plurality of actuator assemblies and the flow control assembly. It is emphasized that this abstract is provided to comply with the rules requiring an abstract.

Abstract: An aircraft having at least one in-flight attitude control system, in turn having at least one actuator; and a hydraulic circuit connected to the actuator and having at least one pump designed to deliver a first flow when the pressure of the hydraulic circuit is above a presettable threshold value. The pump is designed to deliver a second flow greater than the first flow, and the aircraft has a sensor for detecting a quantity associated with the pressure of the hydraulic circuit; and a programmable central control unit, which controls the pump to deliver the second flow when the quantity detected by the sensor corresponds to a pressure of the hydraulic circuit below the threshold value.

Abstract: An aircraft reaction link includes a pair of linear portions, a coupling portion that couples ends of the pair of linear portions on the same side to each other, a fulcrum shaft attachment portion, and a cylinder attachment portion. A metal frame that is made of a metallic material and that is provided so as to extend across the coupling portion and the pair of linear portions and a composite frame that is made of fiber-reinforced plastic and that is provided so as to extend at least in the pair of linear portions are bonded together, and thereby, the pair of linear portions and the coupling portion are formed. The metal frame is provided so as to extend from the fulcrum shaft attachment portion to the cylinder attachment portion.

Abstract: A control surface drive system having a plurality of actuator assemblies are coupled to first and second supply lines and to a return line. The first and second supply lines are connected to a source of hydraulic fluid. At least one of the actuator assemblies has a hydraulic actuator movably connectable to an aircraft control surface. A flow control assembly is connected to the return line and to at least one of the first and second supply lines. A bypass line is in fluid communication with the first and second supply lines and positioned to recycle the hydraulic fluid from one of the first and second supply lines back into the other one of the first and second supply lines when the hydraulic actuator moves toward the first position. A computer controller operatively interconnects the plurality of actuator assemblies and the flow control assembly. It is emphasized that this abstract is provided to comply with the rules requiring an abstract.

Abstract: A user interface linkage system includes first and second user interfaces that are linked by a single hydraulic circuit. The system is configured to provide compensation for relatively slow changes in hydraulic fluid pressure due, for example, to temperature variations. The system is also configured to allow the user interfaces to be selectively unlinked from each other.

Abstract: A hydraulic system includes at least one hydraulic circuit with at least one pump, one or more consumer equipment units, at least one HP discharge line comprising at least one bypass, the bypass being located between the pump and the units, at least one return line including at least one bypass, the bypass being located between a tank and the units. The hydraulic system further includes at least one closure element on the at least one HP line between the bypass and the units, and at least one nonreturn valve on the at least one return line between the bypass and the units. The closure element is placed automatically in the closed position when a break of the HP line is detected downstream of the closure element of the HP line.

Abstract: A user interface fluid linkage system includes first and second user interfaces that are linked by a single, constant volume fluid line. The system includes identically configured, but oppositely mounted hydraulic fluid chambers to link motion between the first and second user interfaces.

Abstract: A control surface drive system having a plurality of actuator assemblies are coupled to first and second supply lines and to a return line. The first and second supply lines are connected to a source of hydraulic fluid. At least one of the actuator assemblies has a hydraulic actuator movably connectable to an aircraft control surface. A flow control assembly is connected to the return line and to at least one of the first and second supply lines. A bypass line is in fluid communication with the first and second supply lines and positioned to recycle the hydraulic fluid from one of the first and second supply lines back into the other one of the first and second supply lines when the hydraulic actuator moves toward the first position. A computer controller operatively interconnects the plurality of actuator assemblies and the flow control assembly. It is emphasized that this abstract is provided to comply with the rules requiring an abstract.

Abstract: A user interface linkage system includes first and second user interfaces that are linked by a single hydraulic circuit. The system is configured to provide compensation for relatively slow changes in hydraulic fluid pressure due, for example, to temperature variations. The system is also configured to allow the user interfaces to be selectively unlinked from each other.

Abstract: The present invention relates to a trim actuator actuating system having two separate hydraulic power supply systems, through which it is possible for the first time to use a hydraulic linear actuator as a trimmable horizontal stabilizer actuator in an aircraft. The at least one hydraulic linear actuator is coupled to the first power supply system in such a way that it applies hydraulic power to the linear actuator during normal operation. In order to ensure the application of pressure to the hydraulic linear actuator in case of breakdown, in such a case, the pressure application to the hydraulic linear actuator is maintained with the aid of the second hydraulic power supply system.

Abstract: Aircraft control surface operating systems and apparatuses, and associated methods of use. In one embodiment, an aircraft system includes an actuator operably coupled to a control surface in one-to-one correspondence. The actuator can include at least a first fluid port and a second fluid port. A first fluid circuit can be configured to provide pressurized working fluid from a fluid source to the first fluid port on the actuator to retract the actuator. A second fluid circuit can be configured to similarly provide pressurized working fluid from the fluid source to the second fluid port on the actuator to extend the actuator. The aircraft system can further include an accumulator configured to store pressurized working fluid for actuator operation in the event of a pressure drop in the fluid source.

Abstract: Apparatus for providing actuation power to an aircraft flight control surface. The apparatus includes a first actuator and a second actuator, where the second actuator has a bypass mode and an assist mode. The mode of the second actuator is determined by an internal differential pressure across the first actuator. The apparatus further includes a mode select valve for selecting between the bypass mode and the assist mode for the second actuator, and a solenoid valve that when energized causes the mode select valve to select the other mode for the second actuator. When the second actuator is in the bypass mode, the first actuator provides the actuation power for the aircraft flight control surface. But, when the second actuator is in the assist mode, both the first and second actuators provide the actuation power for the aircraft flight control surface. Accordingly, the apparatus allows the flight control surface to be operated in a more efficient manner with less hydraulic flow being required.

Abstract: Control system for a control surface (100) of an aircraft comprising at least one actuator (102, 104) of the control surface, connected to a hydraulic circuit and equipped with at least two power jacks (110) capable of operating in at least one activated mode and one deactivated mode. According to the invention, the system also comprises a control unit (114) capable of making a number of jacks operate in activated mode depending on the aircraft flight conditions. Application to aircraft.

Abstract: An electrically controlled hydraulic actuating system includes a servocontrol including a cylinder, a piston dividing the cylinder into two chambers each provided with a hydraulic connection port, a control valve connected to a hydraulic circuit. A two-way hydraulic pump is driven in rotation by an electric motor driven by power electronics. A switching device allows the connection ports of the chambers to be connected either to the control valve or the pump. The switching device connects the hydraulic connection ports both to the control valve and the pump, so as to allow the servocontrol to be supplied the sum of the hydraulic power delivered by the hydraulic circuit and the hydraulic power generated by the pump actuated by the electric motor.

Abstract: A functional fluid that is particularly useful as an aircraft hydraulic fluid is provided. The fluid contains a major amount of an organo/phosphate ester base stock; based on the total weight of the fluid of from about 0.5 to 10 wt % of certain alkoxylated polyether amines and from about 4 to about 20 wt % of auxiliary additives selected from the group consisting essentially of antioxidants, VI improvers, rust inhibitors, erosion inhibitors and defoamers.

Abstract: An aircraft flap drive system utilized a variable displacement motor (15) having a fluid pressure actuated device (23) to vary the displacement of the motor. The system includes a displacement control valve (41), and an electrohydraulic control means (55,67,85) operable in response to an electrical input signal (91) to communicate pressurized fluid from a source (11) to the motor. The displacement control valve (41) includes a load sensing arrangement (13,53,49) responsive to an increasing load on the motor to bias the displacement control valve toward a position whereby the motor moves toward its maximum displacement (minimum flow). One result is that the conventional fixed displacement motor, of a particular size and weight, may be replaced by a variable motor which is smaller and lighter.

Abstract: The system includes a hydraulic surface actuator for operating a piston coupled to a structural member of an aircraft for moving the structural member. An electrical feedback sensor is coupled to the actuator piston rod. A summing circuit for receiving a command signal and the feedback signal is provided having an output which is coupled to a loop gain compensator for operating a servo motor. The shaft of the servo motor drives the shaft of a variable displacement pump having ports leading to the ports of the actuator. The variable displacement pump has a swash plate the angle of which is reduced by a hinge moment sensor and a de-stroking actuator when the load on the surface actuator exceeds a given percentage of maximum. A pump gain sensor measures the reduction in pump gain and produces an electrical output which is applied to the loop gain compensator to compensate for reduction in pump gain.

Abstract: A device for the control of a control surface of an aircraft has at least two actuators (110, 112, 114), each having at least one electric control input (111a, 113a, 115a). At least one of the actuators (115), called the mixed actuator, also has a mechanical control input (115b), and an electric control system for the actuators is able to occupy a state corresponding to a fault of an engine in which at least two of the actuators simultaneously operate the control surface and a state corresponding to an electric fault, in which the mixed actuator (115) operates the control surface from the mechanical control input.

Abstract: A flight control surface actuation system is provided in which a plurality of actuators (6-13) are connected to one another in order to synchronize their operation. Each actuator has its own gearbox such that the torque transmitted in the synchronizing element is kept low.

Abstract: A rotary actuator (16) has an inner body (32) forming a rotor and at least one outer body (34) forming a stator. The inner body (32) and outer body (34) carry blades (40,44) arranged in an alternating manner and defining between them two series of variable volume chambers. The cross-section of the chambers and the height of the triangular blades (40,44) decreases from one end to the other, as a result of a substantially frustum-shape of the inner body (32). Such an actuator (16) can be used for controlling an aircraft rudder or control surface and it is then positioned along the articulation axis of the rudder or control surface.

Abstract: A hydraulic control system for aircraft is disclosed, which has a flight maneuver system mounted on the aircraft and provided with a maneuver signal input and output system for controlling the aircraft depending on flight conditions. The system includes a differentiating circuit responsive to a first output signal for calculating an actuating speed of a servo actuator and for generating a speed signal Vc, a load calculating circuit responsive to the load signal for computing the load on the actuator and for producing a load signal, a pressure calculating circuit responsive to the speed and load signals for computing a pressure Ps to the actuator and for producing a pressure signal, and a setting circuit responsive to the pressure signal for deciding an optimum pressure of the hydraulic pump so as to precisely control the flight surface without energy loss. The pressure Ps is calculated by the following formula:Ps=E.multidot.(Vc/Vm).sup.2 .multidot.

Abstract: An actuator system for a control surface of an aircraft having a hydraulic system and a source of actuation system commands. The actuator system includes hydraulic circuits connected to the aircraft hydraulic system and an electrically-powered hydraulic circuit integrated with the actuator. The electrically-powered hydraulic circuit includes an electric motor connected to a hydraulic fluid pump. The pump can be controlled to adapt to the power requirements to actuate the control surface.

Abstract: An attitude control system for controlling a plurality of attitude control devices, which cooperate to establish the attitude of a vehicle includes redundant control units for outputting attitude control signals to each of the attitude control devices. Additionally, a plurality of discreet power servo units are provided which are located in proximity to and operatively connected to the plurality of attitude control devices. The redundant control units and the plurality of discreet power servo units are connected electrically and/or optically. Each of the discreet power servo units includes an adder for adding an attitude control signal and a positional feed back signal, a rotational drive source for providing a rotational motion in response to the output from the adder, a fluid pressure circuit for hydraulically driving an attitude control device coupled thereto, and a sensor for sensing the position of the attitude control device coupled thereto and for providing to the adder the positive feed back signal.

Abstract: A backdrive system is provided for aircraft flight control surfaces, including a motor for driving a control surface. A fluid pressure responsive engaging/disengaging device is operatively associated between the motor and the control surface. The device may be a clutch normally engaged in the absence of to supply pressure or a brake normally disengaged in response to supply pressure. A supply of fluid pressure is connected to the device. An accumulator is connected in-line between the fluid pressure supply and the engaging/disengaging device for storing hydraulic energy during normal operation of the system and for releasing the stored hydraulic energy for at least a given period of time in the event of loss of supply pressure to allow the control surface to backdrive to a neutral position. A check-valve is provided between the fluid pressure supply and the accumulator to prevent return of hydraulic energy from the accumulator supply of fluid pressure when the supply pressure is lost.

Abstract: This invention relates to a system for controlling the high-lift flaps of an aircraft subjected to the blast of at least one engine. The system includes reversible hydraulic jacks for actuating the flaps arranged so that each flap is actuated by a jack and each of the two chambers of one jack are connected to the corresponding chamber of the other jack, as well as to an orifice in a hydraulic fluid distributing unit, so that the supply and exhaust of the corresponding chambers of the jacks are respectively simultaneous. The high-lift flaps are connected by a mechanical torsional connection which allows a dissymetry of position of the high-lift flaps in extended position. A safety system is also provided to block the high-lift flaps in position as soon as the extent of dissymetry of position between the high-lift flaps in extended position reaches a predetermined threshold.

Abstract: A hydraulic system for an aircraft to provide improved reserve capability for the thrust reverser power system. There is a fluid reservoir having a vent tube with an inlet location located above the bottom wall of the reservoir. In the event that fluid in the reservoir reaches a predetermined low level, supply and pressure isolation valves isolate at least one pump from the rest of the hydraulic system, and also start that pump operating, with a reserve supply of fluid being available for supplying power for the thrust reverser.

Abstract: A hydraulic system for an aircraft to provide improved reserve capability for the steering and reserve brake power system. There is a fluid reservoir having a vent tube with an inlet location located above the bottom wall of the reservoir. In the event that fluid in the reservoir reaches a predetermined low level, supply and pressure isolation valves isolate at least one pump from the rest of the hydraulic system, with a reserve supply of fluid being available for supplying power for the steering wheel and reserve brakes.

Abstract: A fault detector for a triplexed mechanical or hydromechanical system which provides for improvement in cost, reliability and weight over previously known balance beam arrangements for achieving redundant two-channel operation to provide a fail-operating first failure mode for operation of aircraft components which require reliable operation for flight safety. The fault detector for a triplexed mechanical or hydromechanical system has a tripodal summing member which receives three mechanical position signals in three different channels with the tripodal summing member retaining a position perpendicular to the force transmitted by the tripodal summing member in response to the position signals and tipping when one of the input signals deviates significantly from the other two.

Abstract: An electromagnetically interference-proof flight control device having an hydraulic positioning drive for a control surface, the position of which is to be influenced. Provided are at least one signal line for feeding in optical control signals, at least one hydraulic pressure line and a return line, an electrohydraulic valve arrangement which precedes the hydraulic positioning drive and is connected to the hydraulic pressure line, an optical-electronic converter which converts the optical control signals into corresponding electronic signals for the electrohydraulic valve arrangement and a hydraulic-electric energy converter. The hydraulic-electric energy converter serves as an electric energy source for the opto-electronic converter and for addressing the electrohydraulic valve arrangement.

Abstract: A system for altering the lift/drag characteristics of powered general aviation aircraft whereby flight path control may be exercised. The control device operates in conjunction with such lift/drag altering devices, spoilers, dive brakes and the like, and is physically unitary with the pilot's operation of the throttle control. Utilizing a single control, the pilot may selectively maintain the craft without lift/drag modification, with automatic lift/drag modification at lower power settings such as under landing approach conditions and lift/drag modifications at higher power settings and without modifying the power setting. A vacuum operated valve having simple selectable operational modes provides flight path control desired. Vacuum distribution means, i.e., vacuum lines are provided from a vacuum source, to the vacuum valve and to the spoilers and dive brakes for bellows actuation of the lift/drag altering devices.

Abstract: A bifurcated feel unit for use in a dual path flight control system includes separate input levers for the pilot and copilot control paths. The input levers pivot about a common axis but are capable of independent rotation upon override of the breakout pogos associated with the control system external to the feel unit. Each input lever is coupled to a source of resistive force, preferably a piston within a hydraulic chamber. Normally the feel force supplied is the sum of the resistive forces applied to each input lever. In the event of a jam in one path of the flight control system, the feel force supplied to the remaining operable path is reduced in proportion to the reduction in available control surface.

Abstract: A redundant steering mechanism for a craft, especially an aircraft, comprs redundant control components which become effective when the normal steering components become disabled as a result of an electrical power supply failure. The redundant control components do not affect the normal control system but are effective even if the electrical power supply of the craft should fail. The redundant control components are operated by a hydraulic or pneumatic medium acting as a signal carrier. The control signal is a multicomponent signal, one component of which is derived from the pilot's control stick and other signal components are derived from such factors as the speed, the angle of attack and so forth. These signal components are superimposed to form the mechanical emergency control signal for operating a hydraulic or pneumatic steering element.

Abstract: The present invention is in the field of aircraft flight spoiler control mechanisms. The invention enables the conventional, primary spoiler control system to retain its operational characteristics while accommodating a secondary input controlled by a conventional computer system to supplement the settings made by the primary input. This is achieved by interposing springs between the primary input and the spoiler control unit. The springs are selected to have a stiffness intermediate the greater force applied by the primary control linkage and the lesser resistance offered by the spoiler control unit. Thus, operation of the primary input causes the control unit to yield before the springs, yet, operation of the secondary input, acting directly on the control unit, causes the springs to yield and absorb adjustments before they are transmitted into the primary control system.