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    • 专利摘要:
    The present invention relates to a method of assisting the piloting of an aircraft (1) comprising at least one actuator (5) acting on a control member (2). The aircraft (1) comprises a main measurement system (30) and a secondary measurement system (40) which respectively determine at least one predicted value and at least one measured value of a parameter used in at least one piloting law . An estimated value of at least one parameter is estimated with a filter unit (50) using the predicted and measured values, and at least one margin of measurement accuracy is estimated. Each margin of measurement accuracy is compared with a corresponding threshold, and the control law to be applied is chosen according to this comparison.
    公开号:FR3044634A1
    申请号:FR1502552
    申请日:2015-12-08
    公开日:2017-06-09
    发明作者:Jose Torralba
    申请人:Airbus Helicopters SAS；
    IPC主号:
    专利说明:

    Method and device for piloting an aircraft
    The present invention relates to a method and a device for controlling an aircraft. The invention is therefore in the technical field of aircraft with a flight aid system. In particular, the invention lies in the technical field of a flight control system using sensors delivering with a given accuracy information relating to the displacement of the aircraft, such as an angular velocity, an attitude, a acceleration, ground speed ...
    The movement of an aircraft is usually controlled by maneuvering the control organs of this aircraft. Each control organ has the function of participating in the control of the position in the space of the aircraft. To maneuver the control organs, the aircraft can have actuators. As an illustration, the blades of a main rotor and the blades of a tail rotor of a helicopter represent such control organs. Actuators of the servocontrol type then make it possible to control the pitch of the blades. In addition, actuators of the fuel metering type can act on control organs in the form of motors rotating the blades.
    A system for assisting the piloting of an aircraft can comprise various modes of piloting which implement control laws for controlling the actuators of the aircraft. A control law makes it possible to give an order to at least one actuator as a function of the current value of various parameters and of at least one objective to be maintained. Among these parameters, a control law may involve at least one component of the ground speed of the aircraft in a terrestrial frame.
    Multiple driving laws are known. For example, an act of piloting by objective aims to hold a particular objective. Document FR 2 814 433 refers to such laws. By way of illustration, a ground speed piloting law of a rotary wing aircraft aims to hold a ground speed objective which is given by means of a command operated by a pilot. When the pilot positions the control means in a central position, this pilot may for example require a zero ground speed. The ground speed control law then allows the aircraft to remain stationary while hovering.
    To maintain a ground speed objective, a flight aid system may use a variety of speed measuring systems to evaluate at least one component of the current ground speed in a ground reference system.
    A first speed measuring system takes the form of an inertial unit. An inertial unit is an instrument used on an aircraft to estimate its attitude, speed or position relative to a starting point.
    An inertial unit is usually provided with an inertial measurement unit known by the acronym "IMU" corresponding to the English expression "Inertial Measurement Unit". The inertial measurement unit includes multiple inertial sensors. An inertial measurement unit can include three gyrometers used to measure the components of a three-axis angular velocity vector. In addition, an inertial measurement unit may comprise three accelerometers for measuring the components of a specific force vector (magnitude equivalent to the load factor) along the three axes of the aircraft, with respect to the terrestrial reference system.
    In addition, the inertial unit comprises a computer connected to the inertial unit of measurements. The computer performs, if necessary, the real-time integration of the measurements made by the inertial measurement unit to determine the components of the ground speed vector of the aircraft, or even a pitch angle and a roll angle and an angle heading of the aircraft as well as the position of the aircraft. More precisely, by integrating the measurements of the gyrometers, the computer determines the attitude of the aircraft and therefore its orientation at a given moment. Moreover, by integrating measurements of the accelerometers, which can be related to a terrestrial reference system outside the aircraft thanks to the knowledge of the orientation of the aircraft, the computer determines the components of the ground speed of the aircraft by example in the terrestrial reference. Finally, by integrating the speed, the computer determines the geographical position of the aircraft.
    In fact, the inertial sensors of the inertial measurement unit have measurement biases, which can also vary during navigation. In addition, these inertial sensors are subject to measurement noise. The electrical signals emitted by the inertial sensors are further processed by electronic circuits which themselves introduce noise.
    Biases and measurement noise then distort the measurements made. The inertial units are interesting because these inertial units have a very good availability but are therefore prone to errors that generate over time drift measurements made, especially the integrated ground speed.
    To optimize the operation of an inertial unit, this inertial unit can be provided with powerful sensors, such as gyrometers with errors not exceeding a few hundredths of an hour and accelerometers with errors not exceeding a few tens of millionths. Earth's gravity. Such an inertial unit is however very expensive.
    A second system takes the form of a satellite positioning system.
    Such a satellite navigation system comprises a receiver embedded in a vehicle, which receives signals from several satellites belonging to a constellation of satellites, this constellation being controlled by a fixed terrestrial infrastructure called "ground segment". The set consisting of the receiver, the constellation and the ground segment constitutes a satellite navigation system, generally designated by the acronym "GNSS" corresponding to the English expression "Global Navigation Satellite System". Several GNSS navigation systems are currently operational, such as the system known by the acronym "GPS" corresponding to the English expression "Global Positioning System", and the GLONASS system of Russia. The BEIDOU Chinese systems, the Japanese QZSS system and the European GALILEO system are currently under development or deployment.
    A satellite navigation system notably makes it possible to determine the position of an aircraft as well as the components of the ground speed of this aircraft in the terrestrial reference system.
    A general limitation on the use of satellite navigation systems in aircraft control systems is the possibility of multiple failures affecting multiple satellites simultaneously, or even a complete constellation.
    The document filed on December 11, 2014 with the National Institute of Industrial Property in France under the reference 14 02824 proposes the use of several different constellations of satellites.
    In addition, a satellite navigation system is sensitive to external disturbances, such as atmospheric disturbances, for example.
    Therefore, a third system consists in coupling an inertial unit with a satellite navigation system to obtain the estimated components of a ground speed tending to limit measurement inaccuracies. Coupling can be achieved using a Kalman filter. The Kalman filter provides the estimated components of a ground speed from ground speed components obtained with the satellite navigation system and ground speed components obtained with the inertial unit.
    The ground speed components obtained with the satellite navigation system are referred to as "measured" because the satellite navigation system does not exhibit measurement errors that change over time. On the other hand, the velocity components obtained with the inertial unit are termed "predicted" because of these errors.
    The availability and accuracy of data from satellite navigation systems and inertial units are not perfect, the control of the aircraft with a control law using these data can be difficult to achieve.
    In addition, the increased level of aircraft assistance may cause some loss of attention by the pilot when an autopilot mode is engaged.
    As a result, the operation of an autopilot system is monitored by a monitoring system.
    The monitoring system of an automatic flight control system of an aircraft fulfills the function of securing the data sent to the autopilot module which generates the control laws applied to control the actuators. Its function is also to send the information necessary for the crew to become aware of a possibly deteriorated situation affecting the sensors used, for example. The consequence of sensor degradations on the pilot's workload is made transparent to the pilot most of the time thanks to the redundancy of the sensors used and the management of the surveillance system, for example, via algorithms known in the art. the acronym "FDIR" corresponding to the English expression "Fault Detection Isolation and Recovery". After a multiple failure, that is to say affecting several of the redundant sensors, the consequence on the workload of the driver can be rather heavy.
    In particular, the value of a parameter used by a control law may be unavailable.
    In the case where a parameter useful to the autopilot is declared unavailable by the monitoring system, the autopilot system is automatically disengaged to leave room for manual control at the end of a delay time. The transition between the autopilot mode and the manual steering mode is accompanied by a large and relatively sudden variation in the workload for the pilot.
    For example, during a hover performed by engaging an autopilot mode implementing a pilot law having a ground speed target to hold, a pilot can focus his attention on other aspects, such as an operation winching for example. A malfunction of the system measuring the ground speed of the aircraft may induce an unexpected movement of the aircraft to be countered by the pilot, or even following the failure, a sudden disengagement from the automatic piloting mode. The pilot must then take partial or total control of the aircraft. The effect of impairments affecting sensors used by an automatic flight control system may therefore translate for the pilot by a change of control strategy, possibly passing in a few seconds from an automatic pilot mode to a high assistance to a pilot. manual steering mode.
    The present invention therefore aims to provide a method for avoiding such a sudden change under certain conditions. The invention therefore relates to a method of assisting the piloting of an aircraft comprising at least one actuator acting on a control member making it possible to control the position of the aircraft in space, the aircraft memorizing at least two laws of control that can each be applied to generate an order transmitted to at least one actuator, the aircraft comprising a main measurement system and a secondary measurement system which respectively determine at least one value called "predicted value" and at least one value called " measured value "of a parameter used in one of the control laws called" default control law ", each control law separate from the default control law being called" degraded driving law ".
    This method comprises the following steps: calculation with a filtering unit applying a Kalman algorithm of an estimated value from at least one predicted value and at least one measured value and estimation of at least one precision margin measurement, - comparison of the value of a measurement accuracy margin with a corresponding threshold, - control of said actuator by automatically applying said default control law as long as each margin of measurement accuracy is lower than the corresponding threshold, - control said actuator automatically applying a control law called "degraded driving law" when at least one margin of measurement accuracy is greater than or equal to the corresponding threshold.
    The term "parameter" designates a state parameter of the aircraft to be chosen from a list comprising in particular the components of the ground speed of the aircraft in a given reference frame and for example in the three-dimensional terrestrial reference system, the components of the aircraft. acceleration of the aircraft in this given reference frame, the components of the angular velocity of the aircraft in this given reference frame, an angle of pitch of the aircraft, an angle of roll of the aircraft.
    In particular, the main measurement system and the secondary measurement system can determine at least the predicted values and the measured values of the same parameters, and in particular the components of the ground speed. The filtering unit then estimates so-called "estimated values" of the parameters quantized by the main measurement system, using all the predicted and measured values. The term "corresponding threshold" refers to the threshold corresponding to a particular measurement accuracy margin. Each margin of precision of measurement is indeed to be compared to a particular threshold called "corresponding threshold". Several different measurement accuracy margins can be compared to the same threshold.
    The main measurement system makes it possible to determine predicted values of parameters at each moment. Nevertheless, these predicted values may be tainted with errors.
    For example, the main measurement system comprises a plurality of gyrometers and a plurality of accelerometers. The main measurement system can therefore comprise at least one inertial measurement unit likely to present measurement errors that cause an increase in the error on certain parameters such as the ground speed as a function of the evolution of time. Conversely, the secondary measurement system makes it possible to determine measured values of these parameters. However, these measured values may be unavailable at certain times.
    For example, the secondary measurement system is a satellite positioning means.
    Alternatively, the secondary measurement system comprises an object detection means emitting a detection signal capable of returning to the object detection means following an impact with an object.
    Another variant of the invention is therefore to use for example a system known by the acronym LIDAR corresponding to the English expression "Llght Detection And Ranging", an imager or equipment known by the acronym RADAR.
    These devices are effective but may temporarily be ineffective. Therefore, the filtering unit applies a usual Kalman algorithm for continuously identifying the errors of the main measurement system by means of modeling, conventionally called "error model". Such a Kalman algorithm is usually referred to as a "Kalman filter".
    As a result, the filtering unit makes it possible to determine the estimated components of flight control parameters of the aircraft, such as, for example, the components of the ground speed of the aircraft in a given reference frame and for example the three-dimensional terrestrial reference frame, the components of the acceleration of the aircraft in this given reference frame, the components of the angular velocity of the aircraft in this given reference frame, an angle of pitch of the aircraft, an angle of roll of the aircraft.
    Indeed, the filter unit can quantify at each computation time the errors of the main measurement system. Each piece of data estimated by the filtering unit is then obtained by subtracting from a data item from the main measurement system an estimated error of this main measurement system. As an illustration, the main measurement system estimates the predicted values of the ground speed components in the terrestrial reference system. The secondary measurement system provides the filter unit with the measured values of the ground speed components in the terrestrial frame. The filter unit then determines a ground speed error for each ground speed component based on the predicted and measured values. Each estimated value of a component of the ground speed is then equal to the difference of the predicted value of this component of the ground speed minus the speed error for this component.
    In addition, this filtering unit makes it possible to determine the accuracy of the data exploited. Indeed, the filtering unit establishes a covariance matrix, each margin of measurement precision then being equal to the square root of a corresponding term present on the diagonal of this covariance matrix. For example, the covariance matrix has on its diagonal terms respectively relating to the parameters mentioned above.
    Each margin of precision of measurement represents the margin of precision of a parameter. For example, the filter unit estimates a value of the horizontal component of the ground speed of the aircraft at 100 knots, and estimates the precision margin of this value at 0.5 knots. The horizontal component is therefore equal to 100 nodes plus or minus 0.5 nodes.
    For the record, one knot corresponds to one nautical mile per hour, or 1.852 kilometers per hour or 0.514 meters per second.
    We will refer to the literature to obtain information on the filtering units implementing Kalman algorithms. For example, the Wikipedia website publishes the various equations of a Kalman filter. We can also refer for more details to the document of the ISAE (Higher Institute of Aeronautics and Space) by following the link: https://www.qooqle.fr/url sa=t&rct = i & g = esrc = s & source = web & cd = 1 & cd = 0ahUKEwi0rJmQ36PJAhXlwBQKHZcKDT8QFqqqMAA & url = htt p% 3A% 2F% 2Fpersonal.isae.fr% 2Fsites% 2Fpersonal% 2FIMG% 2F pdf% 2FintroKalman vf 2008.pdf & usq = AFQiCNFQbCt6m4lt4-lnksCO53sPfCavlg & sig2 = 0OnxCiUuQzlviB-mblF8vq & bvm = bv.108194040.d.ZWU & cad = ria.
    As long as the secondary measurement system is operational, each margin of measurement accuracy is below a predetermined threshold. Therefore, the autopilot system controls the aircraft by applying the default piloting law, that is to say, the pilot law that offers the best piloting assistance.
    On the other hand, when the secondary measurement system is temporarily inoperative, the errors of the main measuring system increase. The aircraft is then piloted by automatically applying a degraded piloting law.
    Therefore, this method does not propose to disconnect the autopilot system but proposes to apply a degraded driving law more adapted to the situation.
    In order to maintain an adequate workload for the pilot during periods of non-operation of the secondary measurement system, this method allows automatic transition to degraded driving laws depending on the state of the measurement systems. This transition is automatic and therefore does not require a manual reconfiguration of the piloting system by the pilot. The workload of this pilot is therefore limited in potentially stressful situations for this pilot. The term "degraded driving law" refers to a law increasing the workload of a pilot, but in return more robust to certain sensor errors. Consequently, different degraded piloting laws can be implemented, each degraded control law being associated with usage configurations that depend on measurement accuracy margins.
    These degraded control laws are adapted to the types of errors encountered with the main measurement system, and to the amplitude of these errors as a function of the amplitudes typically acceptable for driving. The greater the error of the main measuring system, the higher the measurement accuracy margins will be. Depending on the errors encountered, the invention then proposes various control laws.
    This method therefore automatically proposes to the pilot the most appropriate driving law as a function of the estimated state of the main measurement system, this state being estimated through the margins of the measurement accuracies. The term "most suitable" refers to the pilot law that offers the best compromise between the pilot's workload to control the aircraft, and the robustness of the piloting law with regard to measurement errors.
    This automation of switching between various types of control laws is not obvious. Indeed, the definition of thresholds allowing switching between different laws is not obvious because these thresholds must take into account the workload considered acceptable to the pilot, and must make it possible to minimize the increase in workload of this pilot. following a change in pilotage law.
    In addition, the estimated measurement accuracy margins in fact restore the greater or lesser variation in the accuracy of the parameter or parameters of interest as a function of the more or less dynamic trajectory of the aircraft. A dynamic trajectory induces a significant variation in the measurement accuracy margins.
    This rich information is transmitted in real time thanks to the information provided by the main measurement system (eg load factors), in order to optimize the availability of the default control law when the measurement system information is lost. secondary.
    For example, on a weak flight after the loss of the secondary measurement system, measurement accuracy margins may change slowly. The default control law can then possibly be used during the rest of the flight, even if the secondary measurement information does not return.
    On the contrary, if the trajectory is more dynamic, measurement precision margins can evolve rapidly and will inform the system of the highest growth (compared to the dynamic case) of the errors. The default control law will thus be deactivated relatively early, but potentially later than if one used a conventional system of deactivation after a delay starting from the moment of loss of the secondary measurement.
    The method may further include one or more of the following features.
    Thus, the method may comprise an alert step during which an alert is generated when a degraded piloting law is applied.
    A visual and / or audible and / or haptic alert is generated so that the pilot is aware of the applied piloting law. For example, a screen displays a message indicating which driving law is applied.
    Optionally, the default piloting law is a ground speed master control law that generates an order given to at least one actuator to hold a ground speed objective.
    For example, the pilot of a helicopter maneuvers a cyclic pitch control stick to generate a ground speed objective. The autopilot system then applies the main control law in ground speed to achieve the given ground speed objective.
    This main ground speed control law imposes in particular during each turn a lateral speed to achieve relative to the ground of the aircraft in a lateral direction present in a horizontal plane perpendicular to gravity, the lateral direction being perpendicular to the course followed . The expression "lateral speed to be achieved" designates a speed to be reached during the maneuver.
    This lateral speed to achieve during turns can of course be changed by the driver if he wishes. In particular, the lateral speed to be achieved may be a zero speed in the absence of specific order given by the pilot for this parameter, or a particular lateral speed selected by the pilot.
    The main ground speed control law offers the pilot a high level of assistance. In particular, the main ground speed control law makes it possible to assist the pilot to coordinate the turns, ie to maintain a particular lateral speed during the turn. Decoupling terms in the main ground speed control law are then provided. These terms depend on the yaw rate of the aircraft and the ground speed components of the aircraft.
    Such a master ground speed control law is known to those skilled in the art.
    In addition, a degraded piloting law may be a secondary ground speed piloting law which generates an order given to at least one actuator to hold a ground speed objective, this secondary ground speed piloting law not proposing coordination of ground speed. automatic turning and therefore not controlling at each turn a lateral speed to realize the aircraft relative to the ground of the aircraft in a lateral direction present in a horizontal plane perpendicular to gravity, said lateral direction being perpendicular to the course followed . Therefore, the method may comprise the following steps: from each predicted value and from each measured value, computation with a filtering unit applying a Kalman algorithm of an estimated value of a parameter associated with a predicted value and estimating at least one measurement precision margin of a parameter associated with a predicted value, comparing each precision margin of speed with a so-called "speed threshold" threshold, said secondary speed control law not being able to be applied only if at least one speed accuracy margin is greater than or equal to the speed threshold.
    Such a speed threshold is favorably equal to 2 knots. The expression "can not be applied" means that the corresponding condition is a necessary condition, but not necessarily sufficient.
    In the nominal situation, ground speed components are provided by the main measurement system and the secondary measurement system. Under these conditions, a Kalman filter applied by the filtering unit makes it possible to precisely estimate the errors of the main measurement system. A main ground speed piloting law for holding a ground speed objective is proposed to the pilot.
    The Kalman filter produces an estimate of the precision margins of the evaluated parameters via the application of a Riccati equation embodying an error model of the main measurement system, including following the loss of information from the measurement system secondary. In case of loss of information from the secondary measurement system, the measurements made by the main measuring system are no longer fully compensated for their errors. These errors progressively increase, more or less slowly depending on the nature of the main measurement system and the trajectory followed. However, the hybridization carried out with the Kalman filter prior to the loss of the information from the secondary measurement system makes it possible to slow down the growth of the errors for a certain time.
    In addition, the filtering unit also makes it possible, by estimating the accuracy margin of measurement of the monitored parameters, to know when to switch from the control law from predefined thresholds. The more dynamic the trajectory, the faster the switching will occur because these measurement accuracy margins themselves depend on the trajectory followed.
    In this case, as long as the data transmitted by the secondary measurement system is available, the measurement accuracy margins are lower than the corresponding thresholds. The main driving law in ground speed is then applied.
    The autopilot system then correctly controls the aircraft, for example by maintaining a substantially zero lateral speed during turns.
    In case of loss of the data transmitted by the secondary measurement system, the measurement of the components of the ground speed is in fact tainted with a certain error.
    In the absence of the invention, the autopilot system no longer correctly controls the aircraft since a parameter of the applied control law is not correctly measured. For example, the aircraft no longer evolves at a zero lateral speed during turns. The pilot must then intervene manually to maintain the desired trajectory, especially when cornering. The action of the pilot must be all the more dynamic as the rate of turn is strong, resulting in an increase in the pilot workload under these conditions.
    According to the invention, a speed precision margin of at least one component of the estimated ground speed will on the contrary exceed the predefined speed threshold in the event of loss of the data transmitted by the secondary measurement system. Therefore, a switch to the secondary control law in ground speed automatically takes place. As a result, the term of coordination of turn of the main law of piloting in ground speed is inhibited. The consequence of this switching is an increase in the workload of a pilot to make a turn, but in return the pilot completely clears unpredictable trajectory errors resulting from the use of a poorly estimated speed.
    Furthermore, a degraded piloting law may be a stance control law which generates an order given to at least one actuator to maintain a target attitude of the aircraft.
    An attitude objective may take the form of an objective assigned to an angle of roll of the aircraft and / or an angle of pitch of the aircraft.
    Thus, the method may have the following steps: calculation with the component filtering unit of an estimated acceleration of the aircraft and at least one margin of measurement accuracy, said at least one margin of precision of measurement comprising an acceleration margin of accuracy for each component of this estimated acceleration of the aircraft, - comparison of each acceleration accuracy margin with a so-called "acceleration threshold" threshold, said attitude-holding piloting law can be applied only if at least one acceleration precision margin is greater than or equal to the acceleration threshold.
    The acceleration threshold is for example equal to one thousandth of the Earth's gravity.
    An error relating to a component of the ground speed is associated with a certain drift rate, which corresponds to an error of a component of an acceleration of the aircraft.
    This error of a component of an acceleration may, according to the applicant, pose a problem for the control of the aircraft in a stationary situation. During the implementation of a ground speed control law, a variation of an acceleration of the order of 0.5 thousandth of the gravity induces a significant pilot workload to compensate for the parasitic effect of this variation on gravity. the position of the aircraft in space.
    As a result, the method proposes a second automatic switching to a behavior control law beyond an acceleration error, and therefore when an acceleration margin of accuracy is greater than or equal to the acceleration threshold. . In fact, an acceleration error of 1 mg (one thousandth of the acceleration of gravity) represents an attitude error equivalent of 1 mrad (milliradian), ie 0.057 deg. This error of attitude has only an almost imperceptible consequence on the attitude held with a law of piloting in attitude. As a result, the application of an Attitude Control Act makes sense.
    Moreover, in the context of a main measuring system of the inertial measurement unit type, each gyrometer belongs to a gyrometer class having an accuracy of at least 0.1 deg / h (degree per hour).
    This realization involves robust inertial sensors of a class better than 10 Nm / h (nautical mile per hour) with a precision gyrometer at least equal to 0.1 deg / h. In addition, the sensitivity of the inertial sensors to the potentially important dynamics of the aircraft is favorably the lowest possible. In terms of the specifications of the inertial sensors, this constraint means that the scale factor of the gyrometers as well as the inter-axis non-orthogonality of the sensors must be small.
    As a result, another switching depending for example on a pitch error threshold may then not be necessary. In fact, the attitude control law presents, with such inertial sensors, a notable robustness to attitude errors and angular velocity. Consequently, the attitude control law induces negligible pilot errors due to the accuracy of the inertial sensors.
    However, a degraded piloting law may be an angular speed-holding control law which generates an order given to at least one actuator to achieve an angular velocity objective of the aircraft by using only components of an angular velocity of the aircraft.
    The verb "achieve" means that the angular velocity behavior control law is intended to allow the aircraft to achieve an angular velocity objective.
    Such a control law can be interesting when less accurate inertial sensors are used. Therefore, the method has the following steps: - calculation with the filtering unit of an attitude angle roll and an attitude angle pitch estimated aircraft and at least one margin of measurement accuracy, said at least one measurement precision margin comprising a roll angle accuracy margin and a pitch angle accuracy margin, - a comparison of the roll angle accuracy margin and the margin pitch angle accuracy with a so-called "angular threshold" threshold, said angular velocity control law being applicable only if at least one of said roll angle and pitch angle precision margin is greater than or equal to the angular threshold.
    This attitude difference can be equal to a value between 3 and 5 degrees.
    As regards the inertial sensors, other inertial classes than those previously mentioned are possible.
    If a better inertial class of gyrometers is adopted, 0.01 deg / h for example, an attitude holding piloting law or even the secondary pilot law in ground speed may be useless.
    On the other hand, if a less good inertia of the gyrometers is chosen, of a class greater than 1 deg / h, the assistance level will be degraded in a less continuous manner than in the solution described above. Indeed, the automatic switching would quickly move the autopilot system from a ground speed control law to a steering attitude control law. The attitude holding piloting law has a limited robustness with respect to significant errors on the roll and pitch angles. Typically, the Applicant notes that the maximum permissible error level for steering is between 3 and 5 degrees of roll or pitch error. Beyond this, the pilot has difficulty in manually compensating for the movements of the aircraft resulting from measurement errors.
    Therefore, if the inertial class of the inertial sensors used is likely to exceed this level of error, then the invention proposes switching to an angular rate control law using only angular velocity information.
    In the absence of angular velocity control law, the method can then have the following steps: control of an actuator by applying the default control law as long as each margin of measurement accuracy is less than the corresponding threshold, controlling the actuator by applying the secondary pilot law in ground speed if at least one speed accuracy margin is greater than or equal to the speed threshold and if no acceleration margin of accuracy is greater than or equal to the threshold acceleration. actuating the actuator by applying said attitude holding piloting law if at least one acceleration precision margin is greater than or equal to the acceleration threshold.
    On the other hand, and in the presence of an angular velocity control law, the method can then have the following steps: control of an actuator by applying the default control law as long as each margin of precision of measurement is below the threshold corresponding, - controlling the actuator by applying the secondary control law in ground speed if at least one speed accuracy margin is greater than or equal to the speed threshold and if no acceleration margin of accuracy is greater than or equal to acceleration threshold and if none of said roll angle and pitch angle accuracy values is greater than or equal to the angular threshold, - controlling the actuator by applying said attitude holding piloting law if at least one Acceleration accuracy margin is greater than or equal to the acceleration threshold and if none of said roll angle angle and tanga angle accuracy values ge is greater than or equal to the angular threshold, - controlling said actuator by applying the angular velocity control law if at least one of said margin of precision of roll angle and pitch angle is greater than or equal to the angular threshold.
    In addition to a method, the invention provides a device for assisting the piloting of an aircraft configured to apply this method. The invention thus aims at a device for assisting the piloting of an aircraft, this aircraft comprising at least one actuator acting on a control member making it possible to control the position of the aircraft in space, the device for assisting the aircraft control comprising an automatic control module storing at least two control laws that can each be applied to generate an order transmitted to at least one actuator, the piloting assistance device comprising a main measurement system and a secondary measurement system which determine respectively at least one value called "predicted value" and at least one value called "measured value" of a parameter used in one of the control laws called "default control law".
    This piloting aid device comprises a filtering unit connected to the automatic control module as well as to the main measurement system and to the secondary measurement system, the filtering unit applying a Kalman algorithm to implement the described method. previously.
    Finally, the invention is an aircraft comprising at least one actuator acting on a control member for controlling the position of the aircraft in space,
    This aircraft comprises a pilot assistance device of the type described above. The invention and its advantages will appear in more detail in the context of the following description with examples given by way of illustration with reference to the appended figures which represent: FIG. 1, a view of an aircraft according to the invention, FIGS. 2 and 3, diagrams showing exemplary embodiments of a secondary measurement system, FIGS. 4 to 6, diagrams showing exemplary embodiments of a main measurement system, FIGS. 7 to 9. , diagrams explaining the method according to the invention, and - figure 10 a diagram showing various angles of an aircraft, - figure 11, a diagram illustrating a lateral direction of an aircraft.
    The elements present in several separate figures are assigned a single reference.
    Figure 1 shows an aircraft 1 according to the invention.
    With reference to FIG. 10, this aircraft may be a rotorcraft equipped with rotors 3. In particular, the rotorcraft may comprise at least one rotor participating at least partially in the lift or even the propulsion of the aircraft, or even at least one rotor participating in the control of the yaw movement of this aircraft. At each instant, the aircraft 1 has a pitch angle TANG about a pitch axis AX1, a roll angle ROL about an axis of roll AX2 and a yaw angle LCT about an axis of yaw AX3. .
    In addition, this aircraft 1 moves relative to the ground 200 with a speed called "ground speed Vsl" and an acceleration ysl. Ground speed and acceleration have three components in the terrestrial reference system, namely a lateral component as well as a longitudinal component and an elevation component, respectively called "EAST", "NORTH", "DOWN" in the English language.
    With reference to FIG. 11, the aircraft may have a so-called "lateral velocity" with respect to the ground 200. The lateral velocity represents a projection of the velocity vector Vs1 of the aircraft 1 in a lateral direction DLAT, this lateral direction being present in a horizontal plane PH perpendicular to gravity PES and being perpendicular to the cap CP followed.
    Referring to Figure 1, the aircraft 1 comprises control members 2 for controlling the position of the aircraft in space. For example, such control members include the blades of a rotor 3 of a rotorcraft.
    To move these control members 2, the aircraft comprises actuators 5. These actuators 5 can take the form of electric or pneumatic or hydraulic cylinders, electric motors.
    The actuators 5 may be controlled by a pilot using flight controls 6. For example, flight controls of a helicopter may comprise a handle called "cyclic pitch control sleeve 7", a lever called "lever collective pitch, a rudder, a keyboard to capture a goal ...
    In a purely mechanical architecture, each flight control is connected to an actuator 5 by a mechanical chain 8. In an at least partially electrical architecture, each flight control can generate an order transmitted by a wired or non-wired link to an actuator.
    Furthermore, the aircraft 1 comprises a piloting aid device 10.
    This piloting aid device 10 includes an automatic piloting module 20. The automatic piloting module 20 can comprise for example a processor, an integrated circuit, a programmable system, a logic circuit, these examples not limiting the scope given to the term "autopilot module".
    For example, the automatic control module 20 may comprise a computer 21 in the form of a processor or the like and a memory unit 22. This memory unit 22 may have a non-volatile memory storing non-modifiable information and a volatile memory able to store variable values that vary over time.
    In particular, control laws L1, L2, L3, L4 can be applied by the automatic control module 20. Therefore, these control laws can be stored in the non-volatile memory.
    Each control law allows the automatic control module to generate an order given to at least one actuator 5. In a purely mechanical architecture, the automatic control module can for this purpose control a jack 9 connected to a mechanical chain 8. In a At least partially electrical architecture, the automatic control module 20 can generate an order transmitted by a wired or non-wired link to an actuator 5.
    Pilotage laws include a default piloting law.
    This default control law can be a law called "main driving law in ground speed L1". The application of this principal law in ground speed L1 allows to transmit an order to at least one actuator 5 to hold a ground speed objective. This law of piloting in ground speed imposes the maintenance of a particular lateral speed in turns called "lateral speed to achieve".
    By default, the side speed to achieve is for example zero but can be modified by a driver.
    In addition, the autopilot laws include at least one degraded piloting law.
    A degraded piloting law is called "secondary driving law in ground speed L2". The secondary ground speed control law L2 makes it possible to generate an order transmitted to at least one actuator 5 to hold a ground speed objective. However, the ground speed secondary control law is not intended to maintain a lateral speed to be performed in turns unlike the main ground speed control law L1.
    The cyclic pitch control stick is for example maneuvered by a pilot to transmit a ground speed objective, the main ground speed control law L1 or the secondary ground speed pilot law L2 generating an order to tend to reach this objective. .
    Another degraded piloting law is called the L3 Attitude Control Law. The attitude control law L3 generates an order transmitted to at least one actuator 5 to maintain a target attitude of the aircraft 1. Therefore, the attitude control law L3 makes it possible to maintain an objective of ROL roll angle and / or pitch angle lens TANG.
    The cyclic pitch control stick is for example maneuvered by a pilot to transmit a roll angle objective ROL and / or pitch angle TANG, the L3 attitude hold piloting law generating an order to tend to reach this goal.
    Another degraded control law is called "L4 angular speed behavior control law". The angular speed-holding control law L4 generates an order transmitted to at least one actuator 5 for holding an angular speed objective of the aircraft 1 using only components of an angular speed of the aircraft 1. The joystick cyclic pitch control is for example maneuvered by a pilot to transmit an angular velocity objective, the angular velocity behavior control law generating an order to tend to achieve this objective.
    For example, the L4 angular speed-holding control law aims at maintaining a zero angular velocity when the pilot does not give a piloting command, and generates an angular velocity objective when the pilot maneuvers a flight control. Therefore, when the piloting aid device 10 is activated by conventional means not shown, the automatic control module 20 determines the appropriate control law, then controls each actuator using this appropriate control law for each actuator.
    The automatic piloting module 20 may be connected to an alerting means 70 to indicate to a pilot the applied piloting law. Such alert means 70 may have a screen or equivalent signaling in writing the applied control law, an enclosure indicating audibly the driving law applied, or a body indicating to a pilot tactilically applied control law. For example, a flight control may vibrate differently depending on the law applied.
    To implement a control law, the automatic control module 20 is connected to multiple organs.
    Thus, the automatic control module 20 can be connected to flight controls 6 that allow a pilot to set a goal to achieve. The control law then makes it possible to establish the orders to be given to the actuators as a function of the objective to be achieved and the current state of the aircraft.
    In addition, the automatic control module 20 can be connected to various sensors 60 to estimate this current state. These sensors 60 may include sensors 61 for establishing the current state of an actuator, such as a sensor measuring the position of an actuator member. These sensors 60 may also include means 62 which measure current parameter values of the aircraft such as its angular velocities, plates or load factors.
    Moreover, in order to determine the appropriate driving law, the automatic control module compares measurement precision margins of certain parameters with stored thresholds.
    The piloting aid device 10 has a processing subassembly for estimating the value of these parameters and the corresponding measurement accuracy margins.
    This processing subset includes a main measurement system 30 and a secondary measurement system 40 which respectively determine at least one value called "predicted value" and at least one value called "measured value" of at least one parameter used in at least one of the pilotage laws.
    The main measurement system 30 is described as "main" insofar as this main measurement system 30 continuously provides the required measurements, although sometimes tainted with errors. Because of these errors, the measurements made by the main measurement system 30 are referred to as "predicted values".
    With reference to FIG. 4, the main measurement system 30 may in particular comprise a plurality of inertial sensors including gyrometers 32 and accelerometers 31.
    Thus, the main measurement system 30 may have at least one inertial measurement unit 37 provided with several gyrometers and several accelerometers. In particular, an inertial measurement unit 37 may comprise, for example, three gyrometers 32 arranged respectively along axes orthogonal to each other, and three of accelerometers 31 arranged respectively along axes orthogonal to each other.
    Each gyrometer 32 belongs optionally to a class of gyrometers having an accuracy of at least 0.1 deg / h.
    Such an inertial measurement unit may be part of an inertial unit.
    For example, the gyrometers 32 and the accelerometers 31 are connected to an organ commonly referred to as "virtual platform 33". This virtual platform 33 may comprise a processor, an integrated circuit, a programmable system, a logic circuit, these examples not limiting the range. given to the expression "virtual platform". For example, the virtual platform comprises a processor 34 executing instructions stored in a memory 35.
    The virtual platform determines, by integrating the measurements of the gyrometers, the attitude of the aircraft at a time of calculation. Moreover, by integrating measurements of the accelerometers, which can be related to a terrestrial reference system outside the aircraft thanks to the knowledge of the attitude of the aircraft, the virtual platform determines the predicted values Vp of the three components of the speed. ground of the aircraft in the terrestrial reference system.
    In the example of Figure 4, a single inertial measurement unit is used.
    According to the example of FIG. 5, the piloting aid device can comprise several inertial measurement units 371, 372. Each inertial measurement unit 371, 372 communicates with a computing platform 360 that makes it possible in particular to carry out a filtering. For example, a first inertial measurement unit 371, and a second inertial measurement unit 372 transmit to the computing platform a first predicted value Vp1 and a second predicted value Vp2 for each ground speed component, as well as a first value predicts ωρ1 and a second predicted value ωρ2 for each component of the angular velocity, then a first predicted value yp1 and a second predicted value yp2 for each component of the acceleration of the aircraft.
    Independently of the embodiment, the primary measurement system makes it possible to determine predicted values Vp of each component of the ground speed of the aircraft, predicted values yp of each component of the acceleration of the aircraft, and predicted values ωρ of each component of the angular velocity of the aircraft. Similarly, the primary measurement system makes it possible to determine predicted values of a pitch angle TANGp as well as a roll angle ROLp and a heading angle of the aircraft.
    Moreover and with reference to FIG. 2, the secondary measurement system 40 is qualified as "secondary" insofar as this secondary measurement system 40 does not provide the required measurements at all times. Nevertheless, when the secondary measurement system 40 is operating, this secondary measurement system 40 provides precise values which are therefore referred to as "measured values".
    The secondary measurement system 40 may include object detection means transmitting a detection signal capable of returning to the object detection means following an impact with an object, such as a LIDAR or RADAR system.
    However, the secondary measurement system 40 may comprise a satellite positioning means 45.
    As a result, the aircraft carries at least one satellite reception means 41 able to communicate with a satellite constellation 42.
    According to FIG. 2, a single satellite reception means 41 makes it possible to determine the measured values of the components of the ground speed of the aircraft.
    According to FIG. 3, the aircraft carries a plurality of satellite reception means 41. Each satellite reception means 41 communicates with its own satellite constellation 42.
    As a result, a consolidation unit 43 is connected to each satellite reception means 41. Thus, each satellite reception means 41 determines the measured values Vm1, Vm2 of the ground speed components of the aircraft. The consolidation unit 43 then derives a consolidated measured value Vm from the components of the ground speed of the aircraft, for example by applying an algorithm of the median.
    It should be noted that the teaching of the document filed on December 11, 2014 with the National Institute of Industrial Property in France under reference 14 02824 can be used to consolidate data.
    Referring to Figure 1, the predicted values Vp, γρ, ωρ, ROLp, TANGp and measured Vm are transmitted to a filter unit 50 of the processing subset.
    This filtering unit comprises at least one filtering module 500 which applies a Kalman algorithm. Such a filtering module 500 may comprise for example a processor, an integrated circuit, a programmable system, a logic circuit, these examples not limiting the scope given to the expression "filtering module".
    For example, the filtering module 500 may comprise a computer 51 in the form of a processor or the like and a memory unit 52. This memory unit 52 may have a non-volatile memory 53 storing non-modifiable information and a volatile memory. 54 able to store variable values that vary over time.
    To apply a Kalman algorithm, the usual equations of a Kalman filter are stored, such as the Riccati equation, for example, which makes it possible to calculate the covariance of the estimation errors.
    In addition, an error model is stored. This error model includes a plurality of states that represent the defects of gyrometers and accelerometers, as well as their consequences on the horizontality and verticality of the inertial platform. Typically, these states include the gyrometer biases along the bearing axes of these gyrometers, the accelerometer biases along the carrier axes, an attitude error of the inertial platform with respect to the real geographical axes, an error of each component of the speed. ground resulting from the errors of the gyrometers via a double integration and errors of the accelerometers via an integration. The Kalman algorithm makes it possible to hybridize the data coming from the main measurement system 30 and the secondary measurement system 40. Therefore, this Kalman algorithm makes it possible to determine estimated values of the components of the ground speed V, the components of the acceleration y, components of the angular velocity ώ, the roll angle ROI and the pitch angle TANG.
    In addition, this Kalman algorithm makes it possible to determine measurement accuracy margins for each of these parameters.
    Thus, this Kalman algorithm notably makes it possible to determine a speed accuracy margin σεν for each estimated ground speed component, an acceleration accuracy margin a pour for each estimated component of the acceleration of the aircraft, a margin oeROL roll angle accuracy for the estimated roll angle of the aircraft and a pitch angle accuracy margin σεΤΑΝα for the estimated pitch angle of the aircraft.
    The data supplying the filtering module therefore comes from the main measurement system 30 and the secondary measurement system 40.
    As long as the data of the secondary measurement system 40 is present, the real-time estimated states are used to compensate for the errors of the inertial sensors of the secondary measurement system 40. The estimation of the states is accompanied by an estimation of their covariance. and their precision margin, via the Riccati equation. This Riccati equation is fed by the data delivered by the gyrometers and the accelerometers and information of loss or presence of the ground speed components provided by the secondary measurement system 40.
    Each margin of precision can be equal to the square root of a term present on the diagonal of the covariance matrix. Even if the values of the ground speed components provided by the secondary measurement system 40 are lost, the precision margins and the estimated values continue to be estimated in open loop, ie in the absence measured values of the ground speed. Nevertheless, the accuracy margins will then increase, and all the faster as the trajectory of the carrier is dynamic. This dependence of precision margins on the trajectory is taken into account in the Riccati equation via the data delivered by the gyrometers and the accelerometers.
    These precision margins are then used according to the method of the invention to determine the control law to be applied at each moment.
    As a result, the data estimated by the filtering module are transmitted to the automatic control module 20.
    With reference to FIG. 6, the filter unit may comprise a plurality of filtering modules 501, 502, 503. For example, the filtering unit comprises a filtering module per inertial measurement unit 371, 372, 373.
    The data estimated by each filtering module is then transmitted to a consolidation means 505, such as a consolidation means of the type described above.
    For example, the consolidation means 505 applies a median algorithm to carry out a monitoring of the precision margins of each variable of interest (for example speed and acceleration) calculated with the data of several inertial measurement units 371. 372, 373, before the transmission of consolidated precision margins, thus declared integrity, to the autopilot module. If a margin of precision falls out of a predefined threshold with respect to the median value between all the values of all the filter modules501, 502, 503, then if the redundancy is sufficient, the system is automatically reconfigured to automatically discard the defective value. . If the redundancy of the inertial measurement units 371, 372, 373 does not allow it, then the most degraded piloting law (that is to say that requiring no parameter whose margin of accuracy is declared unintegrated) is automatically selected by the device.
    Figures 7 to 9 illustrate the method of the invention.
    Referring to FIG. 7, the invention associates a main measurement system 30 providing data to a secondary measurement system 40 for continuously identifying faults in the main measurement system 30 through a unit. filtering 50.
    The errors of the main measurement system 30 induce measurement accuracy margins which are estimated by the filter unit 50.
    The automatic piloting module 20 then determines the piloting law to be applied as a function of these precision margins. When a consolidation means 505 is used downstream of the filtering unit, the consolidation means 505 can determine whether the precision margins of a variable determined by each filtering module are substantially equal.
    Indeed, these different margins of precision of the same variable are supposed to be identical in the absence of sensor failure because the accuracy margins depend only on the trajectory and the inertial class of the sensors. The consolidation means 505 thus makes it possible to automatically discredit the erroneous accuracy margins if the redundancy of the sensors is sufficient. In the negative, an automatic reversion to the most degraded piloting law is caused if there is an indeterminacy on the location of the error.
    Figure 8 illustrates the interest of the invention.
    This FIG. 8 presents a diagram showing the workload WK of a pilot on the ordinate, and the EROR errors of the main measurement system 30 on the abscissa.
    In addition, the main driving laws in ground speed L1, secondary speed ground L2, holding L3 attitudes and L4 angular speed holding are schematized.
    It is understood that switching COM1, COM2, COM3 between two control laws made at the appropriate time optimizes the workload of a pilot.
    In a first step, the main driving law in ground speed L1 is for example applied. From a given instant, the automatic control module automatically performs a first switch COM1 to apply the secondary control law in ground speed L2. The errors of the main measurement system 30 then continue to increase, which induces a second commutation COM 2 towards the L3 attitude-holding control law. Optionally, a third commutation COM3 may induce the automatic implementation of the angular speed-holding control law L4.
    FIG. 9 makes it possible to illustrate more precisely the method applied by the piloting aid device 10.
    With reference to FIG. 9, the main measurement system 30 determines the predicted values of parameters, and in particular of each component of the ground speed of the aircraft, of each component of the acceleration of the aircraft, of each component of the angular velocity of the aircraft, a pitch angle TANGp as well as a roll angle ROLp of the aircraft.
    The secondary measurement system 30 determines so-called "measured values" of the ground speed components. Therefore, from each predicted value and from each measured value, the filter unit determines an estimated value of at least one parameter measured by the primary measurement system, and estimates at least one margin of measurement accuracy.
    In particular, the filtering unit determines estimated values of the ground speed components V, the components of the acceleration y, the components of the angular velocity ώ, the roll angle RÔL and the pitch angle TÂNG. from the predicted values and the measured values received.
    In addition, the filtering unit determines a speed accuracy margin σεν for each estimated ground speed component, an acceleration accuracy margin σ pour for each estimated component of the acceleration of the aircraft, a margin of aeR0L roll angle accuracy for the estimated roll angle of the aircraft and a pitch angle accuracy margin σεΤΑΝα for the estimated pitch angle of the aircraft.
    The various estimated values and estimated measurement accuracy margins are transmitted to the autopilot module. Therefore, the autopilot module compares the measurement accuracy margins with a corresponding threshold.
    As a result, the autopilot module controls at least one actuator 5 by automatically applying the default control law L1 as each margin of measurement accuracy is lower than the corresponding threshold.
    By cons, the autopilot module controls at least one actuator 5 by automatically applying a degraded driving law L2, L3, L4 when at least one margin of measurement accuracy is greater than or equal to the corresponding threshold.
    In particular, the filtering unit compares a speed precision margin σεν for each component of the estimated ground speed of the aircraft with a threshold called "speed threshold". Such a speed threshold may be equal to 2 Kts (knots).
    As a result, the secondary ground speed control law L2 can only be applied if at least one speed accuracy margin is greater than or equal to the speed threshold.
    In addition, the filter unit compares each acceleration accuracy margin σεγ with an acceleration threshold, for example equal to one thousandth of the Earth's gravity.
    As a result, the attitude holding control law can only be applied if at least one acceleration accuracy margin is greater than or equal to the acceleration threshold.
    In addition, the filter unit can compare a roll angle accuracy margin aER0L and a pitch angle accuracy margin σεΤΑΝβ to an angular threshold, for example between 3 and 5 degrees.
    As a result, the autopilot module controls at least one actuator by applying the default control law as long as each margin of measurement accuracy is lower than the corresponding threshold,
    On the other hand, the automatic control module controls this actuator by applying the secondary pilot law in ground speed only if at least one speed precision margin is greater than or equal to the speed threshold and if no acceleration accuracy margin n ' is greater than or equal to the acceleration threshold, or if none of the roll angle and pitch angle accuracy values is greater than or equal to the angular threshold.
    Moreover, the automatic control module controls this actuator by applying said attitude-holding piloting law if at least one acceleration accuracy margin is greater than or equal to the acceleration threshold, or even if none of said angle precision values roll and pitch angle is greater than or equal to the angular threshold.
    Finally, the autopilot module controls an actuator by applying the angular velocity control law if at least one of said roll angle and pitch angle accuracy margins is greater than or equal to the angular threshold.
    This method thus makes it possible to smooth the degradation of the piloting assistance level by successive commutations between various control laws when the estimated error of the primary measurement system exceeds certain thresholds. A variant of the invention consists in smoothing the commutations themselves by progressively limiting the authority of terms bringing a high level of assistance, until canceling them to finalize the switching. An example is the progressive limitation of turn coordination assistance (which itself depends on ground speed) as a function of the estimated ground speed error, up to actually switching to a secondary driving law in speed. ground.
    Naturally, the present invention is subject to many variations as to its implementation. Although several embodiments have been described, it is well understood that it is not conceivable to exhaustively identify all the possible modes. It is of course conceivable to replace a means described by equivalent means without departing from the scope of the present invention.
    For example, the aircraft may comprise several actuators. As a result, two different actuators can be associated with separate sets of driving laws, each driving law set having a default driving law and degraded driving laws.
    权利要求:
    Claims (22)
    [1" id="c-fr-0001]
    A method of assisting the piloting of an aircraft (1) comprising at least one actuator (5) acting on a control member (2) which makes it possible to control the position of the aircraft (1) in space, said aircraft (1) storing at least two driving laws that can each be applied to generate an order transmitted to at least one actuator (5), said aircraft (1) comprising a main measurement system (30) and a secondary measurement system (40) which respectively determine at least one so-called "predicted value" value and at least one so-called "measured value" value of a parameter used in one of said control laws called "default control law (L1)", each control law distinct from the default control law being said "degraded control law (L2, L3, L4)", characterized in that said method comprises the following steps: - from each predicted value and from each measured value , calculation with a filtering entity (50) applying a Kalman algorithm of an estimated value of a parameter associated with a predicted value and estimating at least one accuracy margin of measurement of a parameter associated with a predicted value, - comparison of the accuracy margin of measurement with a corresponding threshold, - controlling said actuator (5) by automatically applying said default control law (L1) as long as each margin of measurement accuracy is lower than the corresponding threshold, - controlling said actuator (5). ) automatically applying a degraded control law (L2, L3, L4) when at least one margin of measurement accuracy is greater than or equal to the corresponding threshold.
    [2" id="c-fr-0002]
    2. Method according to claim 1, characterized in that said method comprises an alerting step during which an alert is generated when a degraded driving law (L2, L3, L4) is applied.
    [3" id="c-fr-0003]
    3. Method according to any one of claims 1 to 2, characterized in that said default control law is a main ground speed control law (L1) which generates an order given to at least one actuator (5) for hold a ground speed goal.
    [4" id="c-fr-0004]
    4. Method according to claim 3, characterized in that said master law in ground speed imposes during each turn a lateral speed to achieve the aircraft (1) relative to the ground (200) in a lateral direction (DLAT ) present in a horizontal plane (PH) perpendicular to the gravity (PES), said lateral direction (DLAT) being perpendicular to the course (CP) followed.
    [5" id="c-fr-0005]
    5. Method according to any one of claims 1 to 4, characterized in that a degraded control law is a secondary ground speed control law (L2) which generates an order given to at least one actuator (5) for to maintain a ground speed objective, said ground speed secondary piloting law not controlling, at each turn, a lateral speed to be achieved of the aircraft with respect to the ground in a lateral direction (DL) present in a horizontal plane perpendicular to the gravity, said lateral direction being perpendicular to the course followed.
    [6" id="c-fr-0006]
    6. Method according to claim 5, characterized in that said method comprises the following steps: calculation with the filtering unit (50) of components of an estimated ground speed of the aircraft (1) and at least a precision margin of measurement, said at least one measurement precision margin comprising a speed precision margin for each component of this estimated ground speed of the aircraft (1), - comparison of each speed accuracy margin with a so-called "speed threshold", said ground speed secondary control law can only be applied if at least one speed accuracy margin is greater than or equal to the speed threshold.
    [7" id="c-fr-0007]
    7. Method according to claim 6, characterized in that said speed threshold is equal to 2 knots.
    [8" id="c-fr-0008]
    8. Method according to any one of claims 1 to 7, characterized in that a degraded control law is an attitude control law (L3) which generates an order given to at least one actuator (5) for maintain a target attitude of the aircraft (1).
    [9" id="c-fr-0009]
    9. Method according to claim 8, characterized in that said method has the following steps: calculation with the filter unit (50) of components of an estimated acceleration of the aircraft (1) and at least one measurement precision margin, said at least one measurement precision margin comprising an acceleration precision margin for each component of this estimated acceleration of the aircraft (1), - comparison of each acceleration accuracy margin with a threshold called "acceleration threshold", said attitudinal behavior control law can only be applied if at least one accuracy acceleration margin is greater than or equal to the acceleration threshold.
    [10" id="c-fr-0010]
    10. The method of claim 9, characterized in that said acceleration threshold is equal to one thousandth of the Earth's gravity.
    [11" id="c-fr-0011]
    11. Method according to any one of claims 1 to 10, characterized in that a degraded control law is an angular rate of speed control law (L4) which generates an order given to at least one actuator (5). to achieve an angular velocity objective of the aircraft (1) using only components of an angular velocity of the aircraft (1).
    [12" id="c-fr-0012]
    12. Method according to claim 11, characterized in that said method has the following steps: calculation with the filtering unit (50) of a roll attitude angle and an estimated pitch attitude angle of the aircraft and at least one measurement precision margin, said at least one measurement accuracy margin comprising a roll angle precision margin and a pitch angle accuracy margin, - comparison of the roll angle precision margin and the pitch angle accuracy margin with a so-called "angular threshold" threshold, said angular speed control law being applicable only if at least one of said precision margin of angle of roll and pitch angle is greater than or equal to the angular threshold.
    [13" id="c-fr-0013]
    13. The method of claim 12, characterized in that said attitude difference is equal to a value between 3 and 5 degrees.
    [14" id="c-fr-0014]
    14. The method of claim 6 and claim 9, characterized in that said method has the steps of: - controlling said actuator by applying said default control law as each margin of measurement accuracy is less than the corresponding threshold, controlling said actuator by applying said secondary driving law in ground speed only if at least one speed accuracy margin is greater than or equal to the speed threshold and if no acceleration margin of accuracy is greater than or equal to the threshold of acceleration. control of said actuator by applying said attitude-holding steering law if at least one acceleration accuracy margin is greater than or equal to the acceleration threshold.
    [15" id="c-fr-0015]
    15. The method of claim 6, and claim 9 and claim 12, characterized in that said method has the following steps: control of an actuator by applying the default control law as long as each margin of measurement accuracy is lower than the corresponding threshold, - controlling said actuator by applying the secondary control law in ground speed only if at least one speed accuracy margin is greater than or equal to the speed threshold and if no acceleration margin of accuracy is greater or equal to the acceleration threshold and if none of said roll angle and pitch angle accuracy values is greater than or equal to the angular threshold, - controlling said actuator by applying said attitude holding piloting law if at least one Acceleration accuracy margin is greater than or equal to the acceleration threshold and if none of said roll angle accuracy values and pitch angle is not greater than or equal to the angular threshold - controlling said actuator by applying said angular velocity control law if at least one of said margin of roll angle and pitch angle accuracy is greater than or equal to the angular threshold.
    [16" id="c-fr-0016]
    16. The method according to claim 1, wherein said main measurement system comprises a plurality of gyrometers and a plurality of accelerometers.
    [17" id="c-fr-0017]
    17. The method of claim 16, characterized in that each gyrometer (32) belongs to a gyrometer class having an accuracy of at least 0.1 deg / h.
    [18" id="c-fr-0018]
    18. Method according to any one of claims 1 to 17, characterized in that said secondary measurement system (40) comprises a satellite positioning means (45).
    [19" id="c-fr-0019]
    The method according to any one of claims 1 to 18, characterized in that said secondary measurement system (40) comprises object detection means emitting a detection signal able to return to the object detection means. following an impact with an object.
    [20" id="c-fr-0020]
    20. A method according to any one of claims 1 to 19, characterized in that said filtering unit (50) applying a Kalman algorithm establishing a covariance matrix, each margin of measurement accuracy is equal to the square root of a term present on a diagonal of said covariance matrix.
    [21" id="c-fr-0021]
    21. Device for assisting the piloting (10) of an aircraft (1), said aircraft (1) comprising at least one actuator (5) acting on a control member (2) for controlling the position of the aircraft (1) in space, said piloting aid device (10) comprising an automatic piloting module (20) storing at least two driving laws that can each be applied to generate an order transmitted to at least one actuator (5). ), said driver assistance device (10) comprising a main measurement system (30) and a secondary measurement system (40) which respectively determine at least one value called "predicted value" and at least one value called "value measured "of a parameter used in one of said control laws called" default control law ", characterized in that said driver assistance device (10) comprises a filter unit (50) connected to the automatic control module (20) and the sys main measurement system (30) and the secondary measurement system (40), said filtering unit (50) applying a Kalman algorithm for implementing the method according to any one of claims 1 to 20.
    [22" id="c-fr-0022]
    22. Aircraft (1) comprising at least one actuator (5) acting on a control member (2) for controlling the position of the aircraft in space, characterized in that said aircraft (1) comprises a device flight aid according to claim 21.
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    同族专利:
    公开号 | 公开日
    EP3179328A2|2017-06-14|
    FR3044634B1|2017-12-22|
    US10268209B2|2019-04-23|
    US20170160749A1|2017-06-08|
    EP3179328B1|2018-08-08|
    EP3179328A3|2017-06-21|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    US20130261853A1|2004-11-08|2013-10-03|Bell Helicopter Textron Inc.|Vehicle management system using finite state machines|
    US20140316615A1|2008-10-03|2014-10-23|Bell Helicopter Textron Inc.|Method and apparatus for aircraft sensor and actuator failure protection using reconfigurable flight control laws|
    FR2978858A1|2011-08-01|2013-02-08|Airbus Operations Sas|METHOD AND SYSTEM FOR DETERMINING FLIGHT PARAMETERS OF AN AIRCRAFT|
    US20140027565A1|2012-02-10|2014-01-30|Merlin Technology, Inc.|Rotorcraft advanced autopilot control arrangement and methods|
    FR2814433B1|2000-09-28|2002-12-13|Eurocopter France|DEVICE FOR CONTROLLING THE FLIGHT OF AN AIRCRAFT, PARTICULARLY A HELICOPTER|
    FR2866423B1|2004-02-13|2006-05-05|Thales Sa|DEVICE FOR MONITORING THE INTEGRITY OF THE INFORMATION DELIVERED BY AN INS / GNSS HYBRID SYSTEM|
    KR100761011B1|2006-05-30|2007-09-21|학교법인 인하학원|Aiding inertial navigation system using a camera type sun sensor and method there of|
    FR3030058B1|2014-12-11|2016-12-09|Airbus Helicopters|REDUNDANT DEVICE FOR STEERING SENSORS FOR ROTATING CAR AIRCRAFT|KR101684364B1|2015-01-09|2016-12-21|주식회사 대한항공|Method for stabilizing of missoion payload using unmanned air vehicle instruction and position informaion|
    CN105253330B|2015-10-30|2017-04-05|中国空间技术研究院|A kind of information fusion GEO satellite control system menu-type design method based on optimization|
    US10793260B1|2017-10-30|2020-10-06|The Boeing Company|Methods and systems for controlling aircraft flight performance|
    法律状态:
    2016-12-22| PLFP| Fee payment|Year of fee payment: 2 |
    2017-06-09| PLSC| Publication of the preliminary search report|Effective date: 20170609 |
    2017-12-21| PLFP| Fee payment|Year of fee payment: 3 |
    2018-12-19| PLFP| Fee payment|Year of fee payment: 4 |
    2020-10-16| ST| Notification of lapse|Effective date: 20200914 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1502552A|FR3044634B1|2015-12-08|2015-12-08|METHOD AND DEVICE FOR CONTROLLING AN AIRCRAFT|FR1502552A| FR3044634B1|2015-12-08|2015-12-08|METHOD AND DEVICE FOR CONTROLLING AN AIRCRAFT|
    EP16199028.8A| EP3179328B1|2015-12-08|2016-11-16|A method and a device for piloting an aircraft|
    US15/370,111| US10268209B2|2015-12-08|2016-12-06|Method and a device for piloting an aircraft|
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