Abstract: A process and machine configured to predict and preempt an undesired load and/or bending moment on a part of a vehicle resulting from an exogenous or a control input. The machine may include a predictor with an algorithm for converting parameters from a state sensed upwind from the part into an estimated normal load on the part and a prediction, for a future time, of a normal load scaled for a weight of the aerospace vehicle. The machine may: produce, using a state upwind from the part on the aerospace vehicle and/or a maneuver input, a predicted state, load and bending moment on the part at a time in the future; derive a command preempting the part from experiencing the predicted load and bending moment; and actuate the command just prior to the part experiencing the predicted state, thereby alleviating the part from experiencing the predicted load and bending moment.

Abstract: A fluid pressure valve according to an embodiment of the present invention is applicable to a fluid pressure servo mechanism. The fluid pressure valve includes a housing having a housing member, the housing member being formed integrally so as to have a first port, a second port, and a flow path connecting between the first port and the second port.

Abstract: A method of supplying hydraulic power via a zonal hydraulic system to an aircraft having a plurality of operating phases includes determining a current operating phase of the aircraft and setting a power limit (e.g., a pressure set point and/or flow limitation) for at least one hydraulic power unit based on the current operating phase. The hydraulic system includes a plurality of hydraulic zones. Each of the hydraulic zones includes a local controller, a hydraulic power unit controlled by the local controller, and at least one actuator powered by the hydraulic power unit. Energy savings may be realized by altering the pressure set point for and/or limiting the flow to hydraulic systems in inactive and/or less active zones. The zonal hydraulic system may be reconfigured for safety based on sensed failures and/or during emergency or alternate power conditions with limited available power. Duty cycles for multiple electric motor driven pumps of the system may be balanced.

Abstract: To provide a hydraulic driving device configured to smoothly extend a hydraulic cylinder. The present invention includes the open circuit E, the closed circuit A, a control device 57, and a control lever 56a. The open circuit E is provided with a flow passage 202 for connecting an outlet port of the open circuit pump 13 and a head chamber 1a of the boom cylinder 1, a selector valve 44a provided in the flow passage 202, a discharge flow passage 404 for connecting a hydraulic fluid tank 25 and a flow passage 200, and a bleed-off valve 64 provided in the discharge flow passage 404. The control device 57 is configured in such a manner that when a manipulate signal for extending the boom cylinder 1 is input from the control lever 56a, the bleed-off valve 64 is closed, the selector valve 44a is subsequently opened, and the open circuit pump 13 is controlled.

Abstract: A steering actuator is for steering an outboard marine engine about a steering axis. The steering actuator comprises a housing; a piston device that is disposed in the housing, wherein hydraulic actuation of the piston device causes the outboard marine engine to pivot about the steering axis; and a valve device that is disposed in the housing. The valve device controls a flow of a hydraulic fluid to move the piston device in a first piston direction and to move the piston device in an opposite, second piston direction. Movement of the piston device in the first piston direction causes the outboard marine engine to pivot in a first pivot direction and movement of the piston device in the second piston direction causes the outboard marine engine to pivot in an opposite, second pivot direction.

Abstract: An actuator including: a housing; a piston movably disposed in the housing, the piston being movable between an extended and retracted position; a plurality of gas generation charges generating a gas in fluid communication with the housing; and an exhaust port for exhausting gas from the cylinder generated by the plurality of gas generation charges; wherein activation of each of the plurality of gas generation charges results in an increase in pressure in the housing causing the piston to move in the housing from the refracted to the extended position. The actuator can further include a return spring for biasing the piston in the retracted position and the plurality of gas generation charges can be disposed in the housing.

Abstract: An air bleed valve having a housing with hydraulic fluid therein and a first acoustic wave sensor mounted within the housing so as to be covered in the hydraulic fluid and a second acoustic wave sensor mounted within the housing so that it can detect when air is present in the housing sufficient to be vented outside upon the opening of an air vent valve and where both the first acoustic wave sensor and the second acoustic wave sensor and the air vent valve are electrically connected to a air bleed controller. In a preferred embodiment, the first and second acoustic wave sensors are of the type known as a shear horizontal surface acoustic sensor.

Abstract: Example fluid circuits (e.g., within aircrafts) include first and second pump assemblies. The first pump assembly has an electric motor and a first fluid pump. The first fluid pump is coupled to the electric motor and has a case drain port that is in fluid communication with a case drain region of the first fluid pump. The second pump assembly is powered by hydraulic pressure from the first fluid outlet of the first fluid pump and functions to augment flow through the case drain region of the first fluid pump.

Abstract: An aircraft hydraulic air bleed valve system having an air vent valve connected to an electronic controller where the opening of the air vent valve is permitted only when the aircraft is in a pre-determined flight mode where the aircraft flight mode is determined by a gyroscope connected to the controller.

Abstract: A method for predicting aerodynamic impact for small appendages on aircraft, wherein the improvement comprises using an adaptable computational fluid dynamic model of airflow adjacent the appendage by isolating a patch surrounding the small appendage, measuring the load on the patch without the appendage in place and with the appendage in a place and subtracting the two for increasing computational accuracy of the load predictions for the small appendage to be able to measure the effect of the small appendage.

Abstract: A double redundancy electro hydrostatic actuator system includes two hydraulic pumps; two fail safe valves connected with the two hydraulic pumps, respectively; one dual tandem hydraulic cylinder connected with the two fail safe valves and having a piston rod, wherein the piston rod is moved by switching supply and discharge of the fluid; two switching valves connected with the two fail safe valves; two accumulators connected with the two switching valves and the two hydraulic pumps, respectively; and two chambers connected with the two switching valves, respectively. Each of the two accumulators accumulates the fluid from a corresponding one of the two hydraulic pumps, and sends the fluid to a corresponding one of the two fail safe valves. The two chambers receive the fluid from the two fail safe valves, respectively.

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: Systems and methods for providing back-up hydraulic power for aircraft are disclosed. A system in accordance with one aspect of the invention includes a hydraulic supply line, a hydraulic return line, and a hydraulic actuator coupled between the supply line and the return line. The system can further include a fluid energy storage device coupled in parallel with the hydraulic actuator between the supply line and the return line, and a first valve (e.g., check valve) coupled in series between the supply line and the fluid energy storage device. A second control valve can control the flow of fluid from the fluid energy storage device to the actuator. In particular embodiments, the fluid energy storage can be coupled to an actuator that deploys and retracts an aerial refueling device.

Abstract: A backup system is provided that has a local electric motor and pump for some or all of the hydraulic actuators on an aircraft. A local backup hydraulic actuator has two power sources, hydraulic as primary and electrical as backup. During normal operation, the hydraulic actuator receives pressurized fluid from a hydraulic system and the fluid flow to the chambers is controlled by a servo valve. If the hydraulic system fails, the electronic controller detects the failure by observing the signal indicative of the pressure from the pressure sensor, and the controller powers the local hydraulic pump to provide high pressure hydraulic fluid to the hydraulic actuator via the servo valve.

Abstract: An actuator control system has a dual concentric servo valve having a spool and at least one motor adapted to selectively displace the spool.

Abstract: The system includes a hydraulic powered actuator. A pump system having first and second hydraulic lines is coupled to the first and second ports, respectively and is capable of providing hydraulic fluid to either the first and second lines. First and second pressure sensors are coupled to the first and second lines, respectively. A third line is coupled between the first and second lines and includes first and second valves mounted in series therein. Preferably, the first and second valves are capable of latching in the open position. A reservoir is coupled to the third line between the first and second valves. A shut off valve is included for cutting off the flow from the reservoir.

Abstract: A control device includes a main servocontrol loop having a servocontrolled system, and a compensation device on the input side of the control device. The control device includes an internal loop that receives a first signal derived from the main servocontrol loop as input, and generates, as a function firstly of the first signal and secondly of a second signal characteristic of a parameter of the servocontrolled system, an error signal characteristic of the difference between said second signal and a theoretical signal determined using a model. The error signal being injected into the main servocontrol loop to correct the control of the servocontrolled system. The internal loop and/or the main loop include a variable gain amplifier, and the control device controls the gain of the variable gain amplifier.

Abstract: A flight control actuation system comprises a controller, electromechanical actuator and a pneumatic actuator. During normal operation, only the electromechanical actuator is needed to operate a flight control surface. When the electromechanical actuator load level exceeds 40 amps positive, the controller activates the pneumatic actuator to offset electromechanical actuator loads to assist the manipulation of flight control surfaces. The assistance from the pneumatic load assist actuator enables the use of an electromechanical actuator that is smaller in size and mass, requires less power, needs less cooling processes, achieves high output forces and adapts to electrical current variations. The flight control actuation system is adapted for aircraft, spacecraft, missiles, and other flight vehicles, especially flight vehicles that are large in size and travel at high velocities.