Method of testing aircraft pedal system and device to this end |
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IPC classes for russian patent Method of testing aircraft pedal system and device to this end (RU 2450310):
 Method of searching for faults in dynamic unit in continuous system / 2450309
Reaction of a properly operating system to an input action is first recorded on an interval at control points at discrete moments in time; output signals of a model for each of the control points obtained as a result of trial deviations of parameters of all units are determined, for which trial deviation is successively introduced into each transfer function parameter for all units of the dynamic system and output signals of the system are found for the same input action; the resultant output signals for each of the control points and each of the trial deviations at discrete moments in time are picked up; deviations of signals of the model, which are obtained as a result of trial deviations of corresponding parameters of all structural units from the reaction of the properly operating system are determined; the system with nominal characteristics is replaced with the controlled system; a similar test signal is transmitted to the input of the system; signals of the controlled system for control points at discrete moments in time are determined; deviations of signals of the controlled system for control points at discrete moments in time from nominal values are determined; diagnostic features for each of the parameters are determined from the relationship; a faulty parameter is determined from the minimum value of the diagnostic feature.
 Method of searching for faulty module in discrete dynamic system / 2444774
Unlike the existing method of searching for a faulty module in a continuous dynamic system, the reaction of a time-discrete system know to be properly functioning is recorded for discrete diagnosis cycles with pitch in control points; integral estimates of output signals are determined, for which at the moment of transmitting a test signal, discrete integration of signals is simultaneously initiated with pitch in each of the control points by transmitting signals to inputs of multiplier units, and a discrete exponential signal to second inputs of the multiplier units; output signals of the multiplier units are transmitted to inputs of the discrete integration units; integration is completed at the control instant; the obtained estimates are recorded; integral estimates of model signals are determined for each control point, for which a sample deviation of the parameter of the discrete transfer function is successively entered into each unit of the system and integral estimates of output signals of the system obtained as a result of integrating output signal estimates for each control point are found and each of the sample deviations is recorded; a defective module of the discrete system is determined by the minimum of the diagnostic feature.
 Determination method for dynamic parameters of marine movement mathematical model / 2442718
FIELD: ship navigation. SUBSTANCE: invention refers to ship navigation and can be used for forecasting the ship movements in the course of maneuvering. The fore and backward points are conditionally used. The fore and backwards points are located on the centerline plane of the ship. On a real time basis the coordinates of the fore and backward points are measured. Measurement of the coordinates is fulfilled with the help of the static shear stress receivers and with differential corrections. On the basis of the coordinate measurement results the current values of kinematic movement parameters are determined: linear speeds of the fore F (υf) and backward A (υa) points and their longitudinal (υfx, υax) and lateral (υfy, υay) components in the moving coordinates ZX0Y connected with the ship; longitudinal centre of the rotation (x0) in the moving coordinates ZX0Y connected with the ship; projection of the linear speed vector in the centre of gravity on the y axis 0Y (υy); linear speed of the ship centre of gravity (υ); curvature of the gravity path (R); angular rate of the ship (ω). The obtained results are used for calculation of the current values of the dynamic parameters of the marine movement mathematical model. On the basis of the mathematical model computer modeling is performed in order to forecast the ship movements in the course of maneuvering. EFFECT: improvement of the accuracy of forecasting of the ship movements in the course of maneuvering on the basis of an adequate mathematical model of its travel. 3 cl, 1 dwg  Method of operating industrial plant, and industrial plant control system / 2440596
Proposed method controlling certain number of plant operating parameters and process component parameters and stored in memory unit. Note here that fatigue index inherent in current state of component fatigue is defined. Note also that forecast fatigue is defined. Besides, component with maximum forecast fatigue is identified as drive component, while for multiple preset changes of states, defined is drive component forecast fatigue. Mind that proceeding from certain forecast fatigue values, one of state changes is selected and initiated.
 Method to search for faulty block in dynamic system / 2439648
Response of an admittedly faultless system is registered at a control interval in control points, several integral estimates are determined for output signals of the system for various integration parameters, the produced integral estimates of output signals are registered; several deviations are determined in integral estimates of model signals for each of control points received as a result of trial deviations of block parameters, for this purpose a trial deviation is introduced alternately in each unit of a dynamic model; deviations of model signal integral estimates are determined, produced as a result of trial deviations of structural block parameters; rated values are determined for deviation of integral estimates of model signals, produced as a result of trial deviations of appropriate block parameters; a system with rated characteristics is substituted with a controlled one, integral estimates of controlled system signals are determined for control points and several integration parameters, deviations are determined for integral estimates of controlled system signals for control points from rated values, rated values are determined for deviations of integral estimates of controlled system signals.
 Method to search for faulty block in continuous dynamic system / 2439647
As opposed to the available method of searching for a faulty block in a continuous dynamic system, elements of topological links are determined for each block included into the composition of the system for each control point Pji, j=1, …, k; j=1, …, m, elements Pji are determined from multiple values {-1,0,1}, the value -1 is determined, if the sign of signal transfer from the output of the i block to the j control point is negative, the value 0 is determined, if transfer of a signal from the output of the i block to the j control point is not available, the value 1 is determined, if the signal of signal transfer from the output of the i block to the j control point is positive, rated values are determined for elements of the vector of topological links for each block, diagnostic criteria are calculated, and using minimum value of a diagnostic criterion, the defect is determined.
 Parameter control method of guided missile rotating about angle of roll, and automated control system for its implementation / 2438098
Parameter control method of guided missile rotating about angle of roll involves assignment of signals simulating the commands and rotation of missile about the roll angle, their supply to missile guidance control, comparison of current values of control commands at the outlet of control equipment with pre-set simulating values and evaluation as per comparison results of the compliance of controlled parameters with the specified ones, at which the simulating signal of missile rotation about roll angle is shaped in the form of two pulse signals. Pulse signals are offset relative to each other through 90°. At the required period of the beginning of control process there generated is the signal simulating the beginning of the guided missile flight, which is synchronised with the first front of one of two pulse signals, which corresponds to the beginning of shaping of the pitch command. Synchronised signals are allowed to shape pulse signals at the output of signal simulator of missile rotation about roll angle from the beginning of pitch command shaping; at that, from the beginning of signal shaping or its synchronisation there performed is time count during which the parameter control of guided missile is performed. Also, system for method's implementation is described.
 Method of searching for faulty unit in dynamic system / 2435189
Reaction of a good system fjnom(t) j=1,2,…,k is recorded on the interval t∈[0, TK] in k control points; integral estimations of output signals Fjnom(α), j=1, …,k of the system are determined, estimates of output signals Fjnom(α), j=1, …,k obtained from integration are recorded, integral transforms of dynamic characteristics of the model are determined for each of the k control points obtained from sample deviation of parameters of each of m units, deformations of integral transforms of model dynamic characteristics are determined, the system is replaced with nominal characteristics of the controlled system, an analogue test signal x(t) is transmitted to the input of the system, integral transforms of dynamic characteristics of the controlled system for k control points Fj(α), j=1,…, k for parameter α are determined, deviation of integral transforms of dynamic characteristics of the controlled system for k control points from nominal values ΔFj(α)=Fj(α)-Fjnom(α), j=1,…,k, is determined, normalised deviation values of integral transforms of dynamic characteristics of the controlled system are determined, diagnostic features are determined, and a faulty unit is determined by the minimum diagnostic feature.
 System for determining signal cycle breakdown configuration in flowmetre (versions), method of determining signal cycle breakdown configuration in flowmetre and machine-readable data medium / 2432594
Disclosed are inventions where cycle breakdown configuration during transmission and reception of signals in an acoustic flowmetre is determined, wherein transmission is carried out between corresponding converters of a group of pairs of converters. The propagation time of acoustic signals between corresponding converters of the group of pairs of converters is measured. A set error function values is calculated (each value of the error function is characteristic for the specific cycle breakdown configuration when measuring propagation time of acoustic signals) and the cycle breakdown configuration is determined using, at least partly, the set of error values.
 Device for measuring and monitoring relay and electric interlocking unit parameters / 2432593
Proposed engineering solution which employs feedback enables to correct measurements during each measurement without recourse to calibrating all channels. A stable feedback channel based on elements with highly stable parameters is sufficient. The technical result is achieved owing to that the device has a feedback channel which is included in the circuit between computers through an interface which is connected to the digital-to-analogue converter of the feedback, at the output of which there is an amplifier which is connected to a multichannel switch.
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FIELD: transport. SUBSTANCE: proposed device comprises aircraft pedal actuator, transducer to measure force applied to aircraft pedal to actuate it and to generate signal corresponding to force applied, device to measure deviation of component in response to actuation of aircraft pedal, and control unit to process signals from force transducer and deviation meters to generate output signals indicating aforesaid deviation depending on force applied to aircraft pedal. EFFECT: reliable and comprehensive test results. 15 cl, 1 dwg
The technical field to which the invention relates. The invention relates to a method of testing the pedal system of the aircraft and to the device for implementing this method. The level of technology Brakes and steering of an aircraft, in particular a passenger aircraft typically operate with pedals located in the cabin of the aircraft. As a rule, establish two identical pairs of pedals to control the brakes and the wheel of the aircraft, one for the commander and one for the first pilot of the aircraft. The right pedal each pair of pedals is to put into action right brake the aircraft and cause the deviation to the right of the steering wheel of an aircraft. The left pedal of each pair of pedals is used to operate the left brake the aircraft and cause the deviation to the left of the steering wheel of an aircraft. To actuate the brakes of the aircraft pedal can be rotated around the axis of rotation, while the actuation of the rudder of the aircraft can be achieved by pressing the pedal forward, i.e. towards the front end of the aircraft. In some types of aircraft, particularly in narrow-body aircraft brake system of an aircraft is a brake system with electric control is. Such a brake system with electric control can contain an electronic control unit, which receives signals from sensors pedals, showing, for example, the force applied to the pedal, and/or the speed of impact on the foot. Then, on the basis of signals received from sensors pedals, electronic control unit generates the control commands that are transmitted to the device enable the brakes to be actuated brake in accordance with the actuation of the pedal. The steering system of an aircraft can be a system with electric control, however, the alternative can also be a mechanical system containing cables that after exposure to the steering pedal mechanically transmit the force applied to the pedals, the steering wheel, to ensure its turn. During final Assembly of the aircraft it is necessary to verify the correct operation of the brake pedal and steering system of the aircraft. In particular, it is necessary to test the brake pedal and steering system of the aircraft in order to detect defects of the individual components or Assembly errors, which could interfere with the correct functioning of the brake pedal and steering system of the aircraft. Disclosure of izopet the deposits The present invention is the provision of a method of testing the pedal system of the aircraft and device for carrying out the method, which will allow you to securely and quickly verify the correct operation of a pedal system of the aircraft. For solving the aforementioned problem, the device for testing the pedal system of an aircraft according to the invention contains a device to actuate the pedal. Device actuation pedal preferably is a mechanical device that serves to mechanical impact on the foot of the aircraft. In addition, the device according to the invention contains a force sensor designed for measuring the force of impact applied to the pedals of the aircraft when subjected to a pedal of an aircraft with a device to actuate the pedal. The force sensor is also configured to supply signals corresponding to the force of the impact, measured by the specified sensor. The unit of measurement deviations present in the device for testing the pedal system of an aircraft according to the invention, is intended for measuring angular deviations of the component deflectable in response to the impact on the foot of the aircraft. Finally, the device according to izopet the tion contains the control unit, which is made by processing signals of the force sensor and device for measuring the deflection with the purpose of generating output signals indicating the angular deviation component deviating from exposure to the pedal of the aircraft depending on the efforts of the impact applied to the pedals of the aircraft. The control unit may represent, for example, a stationary personal computer or laptop computer. As explained above, the brake system and the wheel of the aircraft can be controlled by a single pair of pedals mounted in the cabin of the aircraft. To actuate the brakes of the aircraft pedal rotates around the axis of rotation, while the actuation of the rudder of the aircraft are provided by pressing the pedal forward, i.e. towards the front end of the aircraft. Thus, by appropriate selection of the point of application of force to the pedal device according to the invention can be used to test the operation of the brakes of an aircraft or the operation of the steering wheel. Depending on use whether this device for testing the operation of the brakes of an aircraft or the operation of the steering wheel, the unit of measurement variance measures the deviation component of the braking system of the aircraft or declined the surveillance component of the steering system of the aircraft. The output of the control unit, which shows the angular deviation component deviating from exposure to the pedal of the aircraft depending on the efforts of the impact applied to the pedals of the aircraft, can be obtained in the form of data tables and/or diagrams, showing a graph of the angular deviation component deviating from exposure to the pedal of the aircraft depending on the efforts of the impact applied to the pedals of the aircraft. This table and/or chart may be displayed on the display of the control unit and/or to print using a printer that is connected to the control unit and controlled by this unit. Data received by the control unit during the test can be stored in a storage device, which may be an integrated component of the control unit or a separate memory block. Device for testing the pedal system of an aircraft according to the invention can safely and quickly verify the correct operation of a pedal system of the aircraft and to present to the user the results of the test using the control unit. Thus, the control unit forms a human-machine interface device for testing according to the invention. The power control device is TBA for testing the pedal system of an aircraft according to the invention can be performed by the control device to actuate the pedal. The control unit may control the device to actuate the pedals, in particular, using the corresponding computer program that is executed in the control unit and supplying the control unit of the corresponding control commands to the device to actuate the pedal. Control commands, the control unit turns on the device actuating pedal, can be selected in accordance with pre-established testing requirements, which define the parameters to be monitored, for example, the nominal force on the pedal, the speed of impact on the foot pedal, etc. Device for testing the pedal system of an aircraft according to the invention can also be integrated in the test environment at a higher level. Environment testing of higher level can be designed to simulate specific situations, for example, flight situations, while the aircraft is on the ground. In this case, the device according to the invention can be used, for example, to verify the operation of the brake and/or steering system of the aircraft after impact on the brake pedal and rudder pedals, respectively, in a particular situation, which simulates the environment of the test. Device actuation pedal in the device is according to the invention preferably includes a motor, for example, a linear direct current motor, which is designed to move the lever force transmission attached can be rotated to the basic design of the device to actuate the pedal. The motor may be connected to the lever is transmitted through the worm gear. The basic design of the device to actuate the pedals are preferably made with the possibility of joining structural component of an aircraft. The lever is transmitted can connect with the pull of the force intended for the application of force feedback to the pedals of the aircraft. To test the brake pedal of an aircraft or rudder pedals of an aircraft main structure of the device to actuate the pedals can be attached to the structural component of an aircraft, preferably to the floor panel located in the cockpit of the aircraft. Next, the thrust force can be attached to the pedals of an aircraft using a suitable clamping device. The point of application of force to the pedal, i.e. the point of connection of thrust force with a pedal you can choose depending on whether the test of the brake system of the aircraft or the steering system of the aircraft. In SL the tea test the brake system of an aircraft is the point of application of force to the pedal should be chosen in such a way to the actuation device to actuate the pedal caused the rotation of the pedal. Conversely, in the case of testing the steering system of the aircraft point of application of force to the pedals should be chosen in such a way that the actuation device to actuate the pedal was causing translational movement of the pedal in the direction of the front end of the aircraft. To influence the pedal aircraft motor device actuation pedal can function in such a way as to reduce the length of the worm worm gear and hence to cause the lever is transmitted in the forward direction, i.e. in the direction of the pedals of the aircraft. Due to the rotation of the lever is transmitted driving force of the motor is transmitted to the thrust force and, consequently, on the foot of the aircraft. The motor can be controlled by the control unit in accordance with a pre-defined reference parameters, which form the basis of the computer program running in the control unit. Force sensor designed for measuring the force of impact applied to the pedals of the aircraft when subjected to the pedal of the aircraft, preferably connected to the thrust force. Because craving an effort represents the device which directly applies to the pedals of the aircraft, the force of the impact created by the motor of the device to actuate the pedals of the aircraft, the location of the specified sensor on the deadlift effort makes it particularly reliably and accurately measure the force applied to the pedals of the aircraft. As indicated above, the device according to the invention can be used to test the operation of the brakes of the aircraft or for testing the operation of steering the aircraft. If the device is used for testing the operation of the brakes of the aircraft, the unit of measurement deviations preferably is a device for measuring the angular deviation, i.e. turning the pedals of the aircraft in response to the impact on the foot of the aircraft. Since the brake system of an aircraft is usually a braking system, electrically operated, the rotation of the pedals of the aircraft in response to the impact of a pedal of an aircraft is an important parameter for the correct functioning of the braking system. Device for measuring the deflection of the pedal can be a potentiometer, which can be directly attached to the pedals of the aircraft. However, preferably the device for measuring the deflection of the pedal mounted on the lever in front of the Chi force device to actuate the pedals of the aircraft. In other words, the device for measuring the deflection of the pedal preferably measures the deviation or the angle of rotation of the lever is transmitted, which directly corresponds to the angular deviation of the pedals of the aircraft. Through the use of devices for measuring the deflection of the pedal, representing the embedded component device actuating pedal, you can prevent installation of devices for measuring the deflection of the pedal on the pedal of the aircraft for testing. This greatly simplifies the installation of the device for testing according to the invention on Board the aircraft. If the device according to the invention is used for testing the operation of the steering wheel of an aircraft, the device for measuring the deflection of the pedal preferably is a device for measuring the angular deflection of the rudder of an aircraft in response to the impact on the foot of the aircraft. Since the steering system of the aircraft is usually a mechanical system of cables, the actual deviation of the rudder in response to the impact of a pedal of an aircraft is an important parameter for the correct functioning of the steering system. Device for measuring the deflection of the steering wheel can be a potentiometer, which can be directly attached to a control surface of an aircraft. Preferably attic deviations steering, which is a component of the steering system of the aircraft, is used as a device for measuring the deflection by the simple connection of a sensor deviation of the steering wheel to the control unit of the device according to the invention. In a preferred embodiment, the device for testing the pedal system of an aircraft according to the invention includes a device for signal amplification. The amplification device signals may be included between the control unit and the force sensor in order to amplify signals exchanged between the control unit and the force sensor. The amplification device signals preferably amplifies the signals exchanged between the control unit and device for measuring deviations. However, you can modify the amplification device signals in such a way as to provide amplification of signals exchanged between the control unit and device to measure deviations. The device according to the invention can also contain a control unit power supply, which is designed to control the power supply of the various device components, in particular components of the device to actuate the pedal. For example, the control unit power supply may be configured to control supply of the electric motor, Dutch is ka efforts and potentiometer for measuring the angular deviation of the pedals of the aircraft after application of force feedback to the pedals of the aircraft. The device according to the invention preferably contains a single front-end component, which serves to amplify the signals exchanged between the control unit and device actuation pedal, and for regulating the supply of various device components, in particular components of the device to actuate the pedal. The signal levels of the device for testing the pedal system of an aircraft according to the invention are relatively low. Therefore, the device according to the invention has a low sensitivity to electromagnetic interference. The device according to the invention may also contain analog-to-digital Converter, designed for converting signals exchanged between the control unit and the interface component of the signal amplification and regulation of power. The control unit of the device for testing the pedal system of an aircraft according to the invention can be connected to the database server. The connection control unit, and the database server can be secured with a local data network (Local Area Network, LAN). The data obtained during the test, you can easily save in the storage device a database server, which reduces the risk of data loss in the event of a malfunction of zapomina the feeder control unit. The control unit may be configured to compare the magnitude of the angular deviation measured with a device measuring the deflection at the preset force of the impact, with the nominal value of the angular deviation at the preset force of the impact. Preferably there is an opportunity to compare two different values of the angular deviation, measured at two different preset values the efforts of the impact, with two different nominal values of angular deviations at the two preset values the efforts of the impact. The nominal value of the angular deviation may be stored, for example, in a storage device control unit or database server. In addition, the control unit may be configured to alarm if the difference between the value of the angular deviation, measured using measuring devices deviations, and the nominal value of the angular deviation when the preset force of the impact exceeds a preset limit. A warning signal may be supplied in the form of messages of certain colors on the screen display control unit. However, it is also possible to supply acoustic warn inogo signal using the control unit or by a separate feeder warning signal. The control unit may be configured to compare, in particular, the measured minimum and/or maximum value of the angular deviation and/or the measured minimum and/or maximum values the efforts of the impact, and/or measured values of the coupling, and/or measured values of symmetry with the corresponding nominal values. A warning signal may be supplied, if the difference between the measured value and the nominal value exceeds a predefined limit. The method of testing the pedal system of an aircraft according to the invention includes the operations affected by the pedal of the aircraft with the help of the device actuating pedal, measure the force applied to the pedals of the aircraft when subjected to the pedal aircraft, and provide a signal corresponding to the force of the impact, using the force sensor, and measure the angular deviation component deviating in response to the impact on the foot of the aircraft, the measurement deviation. Depending on the test operation of the brake or steering device for measuring deviation measures the deviation component of the braking system of an aircraft or a component of the steering system of the aircraft. The control unit processed the t signals of the force sensor and device for measuring deviations. Then, the control unit generates an output signal indicating the angular deviation component deviating in response to the impact on the foot of the aircraft depending on the efforts of the impact applied to the pedals of the aircraft. The method according to the invention preferably includes an operation at which the trigger lever is transmitted directly connected to the pull of the force intended for the application of force feedback to the pedals of the aircraft, by means of an electric motor, for example, linear direct current motor that can be attached to the lever is transmitted through the worm gear. The force applied to the pedals of the aircraft, when exposed to the pedal of the aircraft can be measured by force sensor connected to the thrust force. In the case of test operation of the brakes of the aircraft measuring device preferably deviation measures the angular deviation, i.e. the rotation of the pedals of the aircraft in response to the impact on the foot of the aircraft. If brake system of an aircraft is a brake system with electric control, the deviation of the pedals of the aircraft in response to the impact of a pedal of an aircraft is an important parameter for the correct options is onirovaniya brake system. However, if you want to test the operation of the steering wheel of an aircraft, the device for measuring the deflection preferably measures the angular deviation of the rudder of an aircraft in response to the impact on the foot of the aircraft. If the steering system of the aircraft represents the mechanical system of cables, the actual deviation of the rudder in response to the impact of a pedal of an aircraft is an important parameter for the correct functioning of the steering system. The method according to the invention may also contain an operation, which amplify the signals exchanged between the control unit and the force sensor. In addition, can be regulated supply device components bringing into action of the pedal. The signals exchanged between the control unit and device for signal amplification and regulation of power supply, can be converted using an analog-to-digital Converter. The data obtained during the test the pedal system of the aircraft, preferably retain in memory the database server, which is connected to the control unit. You can minimize the risk of data loss in case of failure of the storage device control unit. Finally, the method according to the invention may contain operations which, which compares the magnitude of the angular deviation measured with a device measuring the deflection at the preset force of the impact, with the nominal value of the angular deviation at the preset force of the impact. Additionally, you may be served a warning signal if the difference between the value of the angular deviation, measured using measuring devices deviations, and the nominal value of the angular deviation when the preset force of the impact exceeds a preset limit. In particular, it is possible to compare the measured minimum and/or maximum value of the angular deviation and/or the measured minimum and/or maximum value of the effort of the impact, and/or the measured value of the coupling, and/or the measured value of symmetry with the corresponding nominal values. A warning signal may be supplied, if the difference between the measured value and the nominal value exceeds a predefined limit. During final Assembly of the aircraft test, described above, can be produced for both pedals of a pedal of a brake device of a pair of pedal commander for both pedals of a pedal of a brake device of a pair of pedals of the first pilot. Analogion is that you can test for both pedals pedal steering device, a pair of pedal commander for both pedals pedal steering device, a pair of pedals of the first pilot. Further the present invention are explained with reference to a schematic drawing, which shows a preferred implementation of the device for testing the pedal system of an aircraft according to the invention. The implementation of the invention The device 10 for testing a pedal system 12 of an aircraft contains a device 14 of the actuating pedal, designed to influence the pedal 12', 12” of the aircraft. The pedal system 12 of the aircraft shown in the drawing, represents one of two identical pairs of pedals 12, installed in the cabin of the aircraft, with the first pair of pedals 12 is intended for the commander, and the second pair of pedals 12 for the first pilot of the aircraft. The right pedal 12” each pair of pedals 12 serves to actuate the right (not shown in the drawing) of the brakes of the aircraft and to reject the right of the steering wheel 13 of the aircraft. The left pedal 12' of each pair of pedals 12 serves to actuate the left (not shown in the drawing) of the brakes of the aircraft and for deviations to the left of the steering wheel 13 of the aircraft. To actuate the brakes of the aircraft pedal 12', 12” rotate around the axis T of the turn while the wheel 13 of the aircraft can be actuated by pressing on the pedal 12', 12” forward, i.e. in which upravlenii the front end of the aircraft. Brake system of an aircraft is a brake system with electric control, i.e. the signals from sensors pedal (not shown in the drawing)indicating, for example, the force on the pedal and/or the speed of impact on the foot, are processed by the control unit, which controls a device to actuate the electromagnetic brake in dependence on signals received from sensors pedals. In contrast, the wheel 13 is driven mechanical system cables (not shown in the drawing), which transmits the steering wheel 13, the force on the pedal to cause the turning of the steering wheel 13. The device 14 to actuate the pedal contains a linear motor 16 DC, which through the worm gear 20 is connected with the lever 18 is transmitted. The lever 18 is transmitted installed with the possibility of rotation of the core structure 22, while the main structure 22 can be attached to a structural component of an aircraft. Thrust 24 effort is connected with the pivoting lever 18 is transmitted and is used for the force of impact generated by the motor 16, the pedal 12” aircraft. The force sensor 26 is installed on the deadlift 24 effort and is intended for measuring the force of impact applied to pedal the 12" aircraft electric motor 16 through the worm gear 20, the lever 18 is transmitted and thrust 24 effort. The first potentiometer, which serves as the device 28 measuring the deflection of the pedal, mounted on the rotary arm 18 of the power transmission device 14 to actuate the pedal. Thus, the device 28 measuring the deflection of the pedal provides a measurement of the deviation or angle of rotation of the lever 18 is transmitted, which directly corresponds to the angular deviation of the pedal 12” aircraft after application of the appropriate efforts impact to the pedal 12” aircraft. The second potentiometer, which serves as the device 30 of measuring the deflection of the steering wheel attached directly to the steering wheel 13 and thus directly measures the deviation of the steering wheel 13 upon application of a corresponding force impact to the pedal 12” aircraft. The device 14 to actuate the pedal connected to the control unit, which is a portable computer 32. Unit 32 control controls the motor 16 of the device 14 to actuate the pedal in accordance with a computer program, running in block 32 of control and forcing unit 32 of the management to apply appropriate control signals to the electric motor 16 of the device 14 to actuate the pedal. Control signals from the unit 32 control choose based on pre-established requirements to test that define the various parameters to be monitored, for example, the nominal force on the pedal, the speed of impact on the foot pedal, etc. Unit 32 performs control also receiving and processing signals from the sensor 26 efforts, device 28 measuring the deflection of the pedal and the device 30 of measuring the deflection of the steering wheel. The signals exchanged between the block 32 control and sensor 26 efforts are amplified by the device 34 signal amplification and regulation of power. The device 34 signal amplification and regulation of the power supply is also used to regulate the power supply to the motor 16, the sensor 26 efforts and devices 28 measuring the deflection of the pedal; an Analog-to-digital Converter 36 is used to convert the signals exchanged between the block 32 and control unit 34 of the signal amplification and regulation of power. Block 32 contains a control display 38 and is connected to a printer (not shown in the drawing). In addition, the block 32 control connects to the server 40 of the database, the data received by the unit 32 controls during testing the pedal system of the aircraft can be stored in a storage device of the server 40 of the database. The test item is given brake or rudder pedals of an aircraft main structure 22 of the device 14 to actuate the pedals are attached to the floor panel, located in the crew cabin of the aircraft. Then rod 24 application of force added to the pedal 12” of the aircraft by means of a suitable clamping device. The point of application of force to the pedal 12, i.e. the point of connection between the thrust 24 application of force to the pedal 12”, choose depending on that, do test the braking system of the aircraft or the steering system of the aircraft. In the case of testing the brake system of an aircraft is the point of application of force to the pedal 12” is chosen so that the actuation device 14 to actuate the pedal caused the rotation of the pedal 12” around the axis T of the rotation. Conversely, in the case of testing the steering system of the aircraft, the point of application of force to the pedal 12” is chosen so that the actuation device 14 to actuate the pedal caused the translational movement of the pedal 12 in the direction of the front end of the aircraft. For actuation of the pedal 12” aircraft motor 16 of the device 14 to actuate the pedal rotates in such a way as to reduce the length of the worm of the worm gear 20 and thereby cause the lever 18 is transmitted in the forward direction, i.e. in the direction of the pedal 12” aircraft. In the rotary movement of the lever 18 is transmitted driving force ele is tradigital 16 is transmitted to the rod 24 of application of force and, therefore, the pedal 12” aircraft. Motor 16 controls unit 32 of the management in accordance with pre-defined reference parameters, which form the basis of the computer program running in block 32 of the control. After actuation of the pedal 12” aircraft impact force applied to the pedal 12” electric motor 16 through the worm gear 20, the lever 18 is transmitted and thrust 24 effort, measured by the sensor 26 efforts. The sensor 26 effort signals corresponding to the force applied to the pedal 12”, at block 32, the control signals coming from the sensor 26 efforts to block 32 control, enhanced device 34 signal amplification and regulation of power. In the case of testing the brake pedal aircraft unit 32 control receives signals from the device 28 measuring the deflection of the pedal, which correspond to the deviation or angle of rotation of the lever 18 is transmitted and, therefore, the angular deviation of the pedal 12” aircraft after application of force feedback to the pedal 12” aircraft. When tested rudder pedals of an aircraft unit 32 control receives signals from the device 30 of measuring the deflection of the steering wheel, which correspond to angular deviation ru is I 13 aircraft after application of force feedback to the pedal 12” aircraft. Unit 32 of the management processes the signals from the sensor 26 efforts, as well as from the device 28 measuring the deflection of the pedal and the device 30 of measuring the deflection of the steering wheel, respectively, and generates output signals indicating the angular deviation of the pedal 12 and 13 steering of the aircraft, respectively, depending on the efforts of the impact applied to the pedal 12” aircraft. As illustrated in the drawing, the relationship between the force applied to the pedal 12” aircraft, and the angular deviation of the pedal 12 or 13 steering of the aircraft, respectively, the output unit 32 of the control in the form of a diagram showing a graph of the angular deflection of the pedal 12 and 13 steering of the aircraft, respectively, from the efforts of the impact applied to the pedal 12” aircraft. This chart is displayed on the display 38 of the frame 32 of the control, and is printed using a printer that is connected to the block 32 of the control and controlled by this block. In addition, the block 32, the control compares the measured minimum and maximum values of the angular deviation of the measured minimum and maximum values the efforts of the impact, measured values of coupling and measured values of symmetry with the corresponding nominal values. Unit 32 of the control will emit a warning signal in the IDA messages of a specific color on the display 38 of the frame 32 management if the difference between the measured value and the nominal value exceeds a predefined limit. In conclusion, the data obtained by the block 32 for the test, and the results of the above comparison values of angular deflection of the pedal 12 and 13 steering of the aircraft, respectively, with two preset values the efforts of the impact with the corresponding nominal values of angular deflection of the two preset values the efforts of the impact is stored in the storage device of the server 40 of the database. 1. The device (10) for testing a pedal system (12) of an aircraft, comprising: 2. The device according to claim 1, characterized in that the device (14) to actuate the pedal includes a motor (16), designed to actuate lever (18) is transmitted, which is connected with a pull (24) application of force to the force of the impact to the pedal (12', 12") of the aircraft. 3. The device according to claim 2, characterized in that the sensor (26) efforts designed for measuring the force of impact applied to the pedal (12', 12") of the aircraft when subjected to the pedal (12', 12") of the aircraft, connected to the thrust (24) effort. 4. The device according to claim 1, characterized in that the first device (28) measurement deviation is a device for measuring angular deviations of the pedal (12', 12") of the aircraft in response to attack by the s on the pedal (12', 12") of the aircraft, while the second device (30) measurement deviation is a device for measuring angular deviations of the steering wheel (13) of the aircraft in response to the impact on the foot pedal (12', 12") of the aircraft. 5. The device according to claim 1, characterized in that it comprises a device (34) the signal amplification and regulation of power supply, made with the possibility of amplification of the signals exchanged between the block (32) of the control and sensor (26) efforts, and with the possibility of regulation of power supply components (16, 26, 28) of the device (14) to actuate the pedals. 6. The device according to claim 5, characterized in that it comprises an analog-to-digital Converter (36)that is designed to convert signals that are exchanged (32) the control unit and the device (34) the signal amplification and regulation of power. 7. The device according to claim 1, characterized in that the block (32) of the control is connected to the server (40) database and/or configured to compare the measured minimum and/or maximum value of the angular deviation of the measured minimum and/or maximum values the efforts of the impact, and/or measured values of the coupling, and/or measured values of symmetry with the corresponding nominal values and with the possibility of filing a warning signal if the differential is between the measured value and the nominal value exceeds a predefined limit. 8. The method of testing a pedal system (12) of an aircraft, comprising the steps are: 9. JV is the sob of claim 8, characterized in that it includes a stage on which by means of the electric motor to operate the lever (18) is transmitted, is connected with a pull (24) application of force to the force of the impact to the pedal (12', 12") of the aircraft. 10. The method according to claim 9, characterized in that it includes a stage on which measure the force applied to the pedal (12', 12") of the aircraft when subjected to the pedal (12', 12") of an aircraft, by means of the sensor (26) efforts connected with craving (24) effort. 11. The method according to claim 8, characterized in that the first device (28) measurement deviation measure angular deviation pedal (12', 12") of the aircraft in response to the impact on the foot pedal (12', 12") of the aircraft, and using a second device (30) measurement deviation measure the angular deviation of the rudder (13) of the aircraft in response to the impact on the foot pedal (12', 12") of the aircraft. 12. The method according to claim 8, characterized in that it includes the operations on which: 13. The method according to item 12, characterized in that it includes a stage on which convert the signals exchanged between the block (32) and control device (34) amplification of the signals and regulated the I power, with the help of analog-to-digital Converter (36). 14. The method according to claim 8, characterized in that it includes a stage on which retain data obtained during testing of a pedal system (12), a storage server device (40) a database connected to the block (32) of the control. 15. The method according to claim 8, characterized in that it comprises the steps that:
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