• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    The present invention relates to a control sensor device (1) for a rotary wing aircraft (2) comprising at least two IMU inertial modules (51, 52), at least two GNSS receiving means (11, 12) respectively provided with a first fault detection and exclusion module FDE1 (21,22) and covering separate GNSS satellite navigation systems (100,200), at least two second FDE2 detection and fault exclusion modules (31) at least two hybridization platforms (61,62) and at least one third fault detection and exclusion module (41). The FDE1, FED2, FDE3 (21,22,31,41) Fault Detection and Exclusion modules allow the detection of unambiguous and / or erroneous signals to exclude each faulty GNSS (100,200) system. In addition, each hybridization platform (61, 62) makes it possible to determine a hybridized ground speed in order to finally deliver a precise ground speed and integrates said aircraft (2).
    公开号:FR3030058A1
    申请号:FR1402824
    申请日:2014-12-11
    公开日:2016-06-17
    发明作者:Jean Paul Petillon
    申请人:Airbus Helicopters SAS；
    IPC主号:
    专利说明:

    [0001] BACKGROUND OF THE INVENTION A redundant device for piloting sensors for a rotary wing aircraft The general field of the present invention is that of aircraft flight aids, and rotary wing aircraft in particular. The present invention relates to a redundant device for piloting sensors based on at least one inertial unit hybridized by satellite navigation systems for an aircraft determining the speed relative to the ground of this aircraft, as well as a method for determining this speed. relative to the ground of the aircraft. Such a satellite navigation system comprises an onboard receiver, 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 is a satellite navigation system, generally designated by the acronym GNSS meaning in English "Global Navigation Satellite System". Several GNSS systems are currently operational, such as the Global Positioning System (GPS) of the United States of America 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 general limitation on the use of GNSS systems in aircraft control systems is the possibility of multiple failures affecting multiple satellites simultaneously, or even a complete constellation.
    [0002] The present invention takes advantage of the availability of several independent GNSS systems to overcome this limitation. US7436354 discloses a satellite navigation device simultaneously using a plurality of GNSS constellations. Such a device operates by processing position measurements from separate GNSS systems, to detect single faults and multiple faults and to surpass the integrity level of receivers based on a single constellation and using stand-alone control methods. the integrity of these measurements, such methods being known by the acronym RAIM (Receiver Autonomous Integrity Monitoring). Such a device, however, does not implement individual RAIM methods on each of the GNSS systems and does not achieve the level of integrity required for critical applications. In addition, this device does not guarantee a continuity of navigation information, especially in case of loss of signals from satellites due for example to the environment. Finally, this navigation information includes the position of the aircraft but not its speed.
    [0003] It is further known that GNSS systems, although designed to allow the position of various vehicles to be determined, also make it possible to determine their speed. Other technologies also make it possible to estimate the position and speed of an aircraft, without using satellites.
    [0004] Such devices and in particular IRS inertial units can be combined with the use of a GNSS receiver in order to limit the effects of disturbances to GNSS systems. In particular, the ground speed obtained by means of an inertial unit makes it possible to validate the ground speed provided by a GNSS receiver when these two ground speeds are close. In addition, in the event of a GNSS system failure subsequent to the failure of a satellite or masking, for example, the inertial unit is able to provide the GNSS receiver with a limited time, and to estimate the speed ground of the aircraft.
    [0005] For example, document FR2906893 describes a hybrid system comprising at least two GNSS receivers, at least one IMU inertial module for hybridizing the information provided by at least one GNSS receiver and at least one extended Kalman filter. This hybrid system makes it possible to detect a fault on at least one satellite of the GNSS system. This hybrid system also makes it possible to determine a protection radius corresponding to an error on the position provided by this hybrid system. The information provided by such a hybrid device is accurate. On the other hand, this hybrid device is dependent on a single GNSS system, such as the GPS system for example. As a result, the information provided is not sufficiently honest. In addition, such a hybrid device must include a large number of estimators in order to manage the double failures of 20 satellites. Indeed, this device is based on hypothesis tests. The number of estimators thus corresponds to the number of fault configurations, according to these hypotheses, that this device can detect. For example, a device sized to detect all the failure combinations of two satellites out of twenty four satellites of two GNSS constellations must have a number of estimators equal to cl = 276. The introduction of a third GNSS constellation carries the number of estimators required for cfs = 630. This increase in the number of estimators then results in a significant cost and an impossibility of such a system to evolve with the advent of new constellations.
    [0006] Also known is FR2964468 which discloses a device for detecting and excluding multiple failures of satellites for a multi-GNSS system using several constellations simultaneously. This device comprises a Kalman filter bank 5 provided with at least one satellite filter used to exclude the information provided by a satellite deemed to be defective. In addition, this device can be hybridized by an IMU inertial module. As before, the computing power of this Kalman filterbank increases very rapidly with the number of satellites treated and the number of satellite failure combinations envisaged. The document FR2996647 describes an inertial unit connected to a receiver using measurements from several satellites distributed in at least two distinct subsets of satellites in order to determine at least two hybrid navigation solutions.
    [0007] This inertial unit comprises a satellite failure detector, equipped with a main Kalman filter and several secondary Kalman filters, comparing these hybrid navigation solutions downstream of the Kalman filters in order to detect a failure of a satellite and a satellite. exclude the failed satellite.
    [0008] The use of a main kalman filter and secondary Kalman filters as well as its application to a navigation device using an inertial unit and a GPS receiver are described in particular in the document "A New Failure Detection Approach and Its Application". Autonomous GPS lntegrity 25 Monitoring "- IEEE Transactions on Aerospace and Electronic Systems - Vol 31, No.1 - January 1995 - pages 499-506. The object of the present invention is to propose a device for piloting sensors for an aircraft making it possible to overcome the limitations mentioned above, this piloting sensor device being capable of providing a ground speed of the aircraft which reaches the levels of integrity, availability and accuracy 3030058 5 required by a flight control system, allowing for safe flights close to the ground and obstacles. This piloting sensor device is more particularly intended for a rotary wing aircraft.
    [0009] According to one aspect of the invention, a steering sensor device for a rotary wing aircraft comprises GNSS receiving means of constellations of at least two independent and distinct GNSS systems as well as at least three detection and detection modules. FDE fault exclusion. Each GNSS receiving means is connected to at least one antenna and receives initial navigation signals from several satellites. Each fault detection and exclusion module FDE receives at least two input signals and delivers an output signal, each output signal including a measurement and a state of integrity. According to one embodiment, the control sensor device according to the invention comprises at least two GNSS reception means and can thus use at least two separate GNSS systems. Preferably, the pilot sensor device according to the invention comprises separate and dissimilar GNSS reception means for separately processing the initial navigation signals from satellites belonging to each GNSS system. Each GNSS receiving means is thus dedicated to a specific GNSS system such as, for example, the GPS system, the GLONASS system, the GALILEO system, the QZSS system and the BEIDOU systems. In fact, the control sensor device according to the invention is redundant both at the level of the GNSS systems and at the level of the GNSS reception means, and can thus overcome failures of one of these GNSS systems or of the one of the reception means.
    [0010] According to another embodiment of the invention, each GNSS receiving means is a sub-function of a single multi-GNSS receiver, that is to say able to use navigation signals from satellites. belonging to different GNSS systems but providing distinct solutions for each constellation. According to yet another embodiment, the piloting sensor device according to the invention can use, in lieu of a GNSS system, a satellite telecommunication system such as the IRIDIUM system which uses its own constellation of satellites. . The ephemeris of these satellites being known, it is possible to exploit the Doppler effect on the signals they emit to determine the ground speed of a vehicle. This use of the IRIDIUM system is possible because the piloting device 15 according to the invention aims to determine the ground speed of an aircraft rather than its position. This piloting sensor device according to the invention is remarkable in that each GNSS receiving means comprises a first fault detection and exclusion module FDE1 and at least one second fault detection and exclusion module FDE2. Each first fault detection and exclusion module FDE1 receives and analyzes the initial signals and detects intact initial signals and / or erroneous initial signals. An exemplary embodiment of this treatment is known by the name RAIM, described for example in the document FR2971857. This RAIM method initially intended for the consolidation of position measurements is named V-RAIM in the remainder of this document when it is applied to the determination of an aircraft ground speed.
    [0011] Each GNSS receiving means then delivers a measurement and a state of integrity of a first ground speed signal of the aircraft in a geographical reference point from the initial signals of integrity, excluding, if necessary, said initial erroneous signals.
    [0012] The availability of a first ground speed signal is dependent on the number of GNSS satellites visible to the GNSS receiving means or in proper operation. For example, four satellites are generally required to determine a three-dimensional position as well as a time offset of the receiver clock, or a three-dimensional speed of an aircraft as well as a frequency offset of the receiver clock. . However, at least one fifth satellite is required to have redundancy and thus detect the presence of a single failure of a satellite. In fact, each first fault detection and exclusion module FDE1 receives at least four initial signals to determine a first ground speed signal and at least five initial signals to ensure the integrity of this first ground speed signal. Advantageously, each GNSS receiving means may comprise a high precision clock, such as an atomic clock, used as a frequency reference. Each first fault detection and exclusion module FDE1 then requires an initial signal of less to determine a first ground speed signal of the aircraft. In this way, such a GNSS receiver means comprising an atomic clock makes it possible to determine a first ground speed signal as soon as three satellites are visible and to detect a single satellite failure from four visible satellites. In fact, each GNSS receiving means can provide a guaranteed ground speed measurement for a first level of integrity, with a first level of autonomous monitoring. This first level of integrity only covers the simple failures of a satellite in the constellation of this GNSS receiving means. Each second fault detection and exclusion module 5 FDE2 is connected and is in communication with at least two GNSS receiving means. Each second fault detection and exclusion module FDE2 receives, analyzes and compares the first ground speed signals delivered by at least two GNSS receiving means, and then detects first ground speed signals and / or first grounding signals. ground speed erroneous. Each second fault detection and exclusion module FDE2 can then detect and possibly exclude each defective GNSS system by locating a first wrong ground speed signal, and then determine and output a measurement and integrity state of a second signal. ground speed of the aircraft from at least two first ground speed signals integrity, excluding, where appropriate, the first erroneous ground speed signals. The second fault detection and exclusion module 20 FDE2 receiving the first ground speed signals makes it possible to detect multiple faults, affecting several satellites simultaneously, as well as those affecting the ground segment of a GNSS system. Indeed, by comparing the first ground speed signals transmitted by at least two GNSS receiving means 25 covering at least two distinct and independent GNSS systems, each second FDE2 detection and fault exclusion module can detect inconsistencies between these first Ground speed signals and, at a minimum, pass the failure. "Passive failure" means to make this failure 30 passive, that is to say without catastrophic or dangerous consequences on the device.
    [0013] If at least three GNSS and three reception means were initially available, the device according to the invention also makes it possible to locate the defective GNSS system, to exclude it, and to continue to operate with the valid GNSS systems. The second fault detection and exclusion module FDE2 can operate according to the known method of the median. Such a method is for example described in US4264955. According to this document, the median value of the input signals is calculated. This median value is integrally guaranteed as long as the number of erroneous first ground speed signals is less than half the total number of first ground speed signals available at the input of the second FDE2 detection and fault exclusion module. In addition, said median value may be used as a reference for comparing each of the values of the other input signals. The input signals deviating from the reference, in absolute value, by more than a predetermined threshold are then considered to be defective. In cases where the second fault detection and exclusion module FDE2 no longer has at its inputs only two valid signals, an excessive deviation between these two signals reveals a failure affecting one of the two. Since it is not possible to locate the fault, the second fault detection and exclusion module FDE2 passes the fault by invalidating its output.
    [0014] Finally, when the second fault detection and exclusion module FDE2 receives a single first valid ground speed signal, no second integrated ground speed signal is provided, since the integrity of this first ground speed signal can not be satisfied. to be determined.
    [0015] For example, the predetermined coherence threshold is equal to 0.2 meters per second (0.2m / s). The sensor device according to the invention thus makes it possible to determine a ground speed signal of the aircraft sufficiently integral for use in a control system. Indeed, the integrity of this second ground speed signal of the aircraft results from the cascading of first modules and a second fault detection and exclusion module FDE1 and FDE2, the second stage FDE2 detecting the failures that were not by the first 10 floors FDE1. In addition, this second ground speed signal is very available when it is determined from more than two separate and independent GNSS systems. It is very unlikely that multiple failures will simultaneously affect several independent GNSS systems. In addition, each GNSS receiving means may advantageously be connected to at least two receiving antennas. Thus, each GNSS receiving means can determine, as described for example in FR2964199, the directions of arrival of the initial signals from the satellites, compare them to the expected directions and reject those of said initial signals for which a discrepancy is noted. In the preferred embodiment of the invention, the availability of an integrated ground speed solution for the control laws is improved by the implementation of inertial measurements. In this embodiment, the control sensor device according to the invention further comprises at least one IMU inertial module and at least one hybridization platform. Each IMU inertial module 30 provides inertial measurement signals 3030058 11 characterizing acceleration and angular velocities of the aircraft. It is then known to estimate by integration of these signals of inertial measurements of accelerations and angular velocities an inertial ground speed of this aircraft. Each hybridization platform is connected to and communicates with an IMU inertial module and a second FDE2 fault detection and exclusion module. The hybridization platform and the IMU inertial module form an inertial chain. Each hybridization platform receives and processes these inertial measurement signals as well as possibly a second ground speed signal, constituting a helper speed, and determines a measurement constituting a third ground speed signal of the aircraft. In known manner, a speed-assisted hybridization platform performs the following operations: integration of the inertial measurements to obtain a hybrid speed estimation, calculation of the difference between said hybrid speed estimation and the speed of assistance, estimation corrections of the inertial measurements from said gap. In the phases of flight where the speed of assistance is not available, the speed obtained is purely inertial, but benefits from the last estimated corrections. The third ground speed signal is thus continuously available even when the second ground speed signal is not available. Preferably, the control sensor device according to the invention comprises at least two hybridization platforms and at least two IMU inertial modules.
    [0016] In addition, the control sensor device according to the invention may comprise at least one third FDE3 detection and fault exclusion module. Each third fault detection and exclusion module FDE3 is connected and is in communication with at least two hybridization platforms. Each third FDE3 Fault Detection and Exclusion Module receives, analyzes, and compares the third ground speed signals delivered by the hybridization platforms and detects third solid ground speed signals and / or third errored ground speed signals. . Each third fault detection and exclusion module FDE3 can then detect a fault on an inertial chain and at least passivate this fault. In embodiments where the pilot sensor device comprises more than two inertial chains, the third FDE3 Fault Detection and Exclusion Module may further locate the faulty inertial chain and exclude it. Each third fault detection and exclusion module FDE3 can thus determine and deliver a measurement and a state of integrity of a fourth ground speed signal of the aircraft from at least two third ground speed signals. correct, excluding, where appropriate, the third erroneous ground speed signals. Each third fault detection and exclusion module FDE3 can implement the median method for detecting and possibly locating and excluding the third defective ground speed signals. The third FDE3 Fault Detection and Exclusion Module may further be connected and in communication with at least one second Fault Detection and Exclusion Module 3030058 13 to receive, analyze and compare to the FDE3. minus a second ground speed signal and the third ground speed signals. Each third fault detection and exclusion module FDE3 can then detect and locate second and / or third intact ground speed signals as well as second and / or third errored ground speed signals. Each third fault detection and exclusion module FDE3 can then determine and deliver a measurement and an integrity status of the fourth ground speed signal of the aircraft from at least one second ground speed signal that is integral and correct. / or at least two third solid ground speed signals excluding, if appropriate, the second and / or third errored ground speed signals. Each third fault detection and exclusion module FDE3 can thus locate the detected fault which can be either a failure of an inertial chain or a failure common to the GNSS systems. Each third fault detection and exclusion module FDE3 can then exclude the defective inertial chain or the GNSS systems from the location of the erroneous second and third ground speed signals. The third fault detection and exclusion module FDE3 delivers the fourth ground speed signal by applying, for example, the method of determination according to the median value.
    [0017] Furthermore, the use of at least one inertial chain thus makes it possible to ensure the continuity of the supply of the third speed signal and, consequently, of the fourth ground speed signal of the aircraft in the event of unavailability of a second ground speed signal integrates.
    [0018] Advantageously, the use of several inertial channels in the control sensor device makes it possible to pass the simultaneous and coherent faults of all the GNSS systems. Such a situation is almost impossible if the defects considered are unintentional failures. On the other hand, a malicious attempt by an individual or an organization to falsify all the GNSS signals received by the aircraft can not be excluded. Thus, in such situations, a device according to the invention rejects all of the second ground speed signals and continues to deliver a fourth integral, purely inertial ground speed signal. According to a particular embodiment of the invention, each hybridization platform comprises a purely inertial virtual platform and a hybridization error filter 15 communicating with each other. Each purely inertial virtual platform is connected and in communication with an IMU inertial module thus forming an inertial unit. The pilot sensor device comprises two inertial units, a two-way computing calculator, and two inertial unit hybridisation error filters, one in each calculation channel. Each calculation channel calculates two hybridization error filters with firstly a second fault detection and exclusion module FDE2 and secondly a third fault detection and exclusion module FDE3. Each second fault detection and exclusion module FDE2 is connected and in communication with two hybridization error filters and each third fault detection and exclusion module FDE3 is in communication with two hybridization error filters. for each calculation channel. The use of these two calculation channels in parallel makes it possible to detect and passivate a possible malfunction of one of these calculation channels.
    [0019] 3030058 15 Each purely inertial virtual platform receives inertial measurement signals from an IMU inertial module that this purely inertial virtual platform transforms into a pure inertial ground speed. In fact, each inertial unit delivers a pure inertial ground speed of the aircraft. Each hybridization error filter is then connected to an inertial unit as well as to a second fault detection and exclusion module FDE2 to receive a pure inertial ground speed of the aircraft and a second ground speed signal of the aircraft. 'aircraft.
    [0020] In addition, each hybridization error filter is preferably a Kalman filter which estimates errors affecting state variables rather than directly state variables. The present invention also relates to a method for determining a ground speed of an aircraft comprising several steps. In a first step, initial navigation signals from several satellites belonging to at least two separate and independent GNSS systems are received. In a second step, these initial navigation signals are analyzed for each GNSS system. In a third step, honest initial signals and / or erroneous initial signals are detected. In a fourth step, a measurement and a state of integrity of at least two first ground speed signals of the aircraft are provided in a geographical reference for at least two separate and independent GNSS systems from the initial signals. integrity by excluding, as the case may be, erroneous initial signals.
    [0021] In a fifth step, the first ground speed signals are analyzed and compared. In a sixth step, first solid ground speed signals and / or first ground speed signals are detected. During a seventh step, each defective GNSS system is detected and excluded providing a first errored ground speed signal. During an eighth step, a measurement and a state of integrity of at least one second ground speed signal of the aircraft are determined and delivered from at least two first ground speed signals of integrity. excluding, where appropriate, said first errored ground speed signals. During the eighth step, at least a second ground speed signal of the aircraft can be determined if at least two first ground speed signals are available. Each second ground speed signal of the aircraft is determined by a method of the median. The method for determining a ground speed of an aircraft 20 thus makes it possible to detect a simple failure of a satellite of a GNSS system and / or a multiple failure within one or more GNSS systems. The second ground speed signal of the aircraft therefore remains available and integrates despite multiple failures. According to a first variant of this embodiment of the invention, this method may comprise additional steps. During a ninth step, inertial measurement signals are acquired, the inertial measurement signals characterizing acceleration and angular velocities of the aircraft. In a tenth step, each second ground speed signal and the inertial measurement signals are processed.
    [0022] During an eleventh step, at least one measurement constituting at least one third ground speed signal of the aircraft is determined and delivered from the inertial measurement signals and, if appropriate, from a second ground speed signal. integrates, the third ground speed signal then being continuously available. During this eleventh step, each third ground speed signal can be determined according to a known hybridization method and conventionally used in the aeronautical field. Such a hybridization method makes it possible to deliver a third hybrid ground speed signal. During a twelfth step, the third ground speed signals are analyzed and compared. Then, during a thirteenth step, third solid ground speed signals and / or third ground speed signals are detected. During a fourteenth step, a measurement and a state of integrity of a fourth ground speed signal of the aircraft are determined and delivered from at least two third solid ground speed signals excluding the case. The third ground speed signal can be determined by the median method. According to a second variant of this embodiment of the invention, these additional steps take place as follows, the ninth, tenth and eleventh steps being identical to the first embodiment. During the twelfth step, at least one second ground speed signal and the third ground speed signals are analyzed and compared. During the thirteenth step, second and / or third ground speed signals, as well as second and / or third errored ground speed signals, are detected and located.
    [0023] During the fourteenth step, a measurement and a state of integrity of the fourth ground speed signal of the aircraft are determined and delivered from at least one second ground speed signal and / or from minus two third smooth ground speed signals excluding, if appropriate, the second and / or third errored ground speed signals. All the steps can be sequentially linked. However, the ninth step may proceed simultaneously to at least one of the first eight steps. The invention and its advantages will appear in more detail with reference to the following description with exemplary embodiments given by way of illustration with reference to the appended figures which represent: FIG. 1, a rotary wing aircraft equipped with a control sensor device according to the invention, and FIGS. 2 to 4, several embodiments of a control sensor device according to the invention. The elements present in several separate figures are assigned a single reference.
    [0024] FIG. 1 represents a rotary wing aircraft 2 equipped with a steering sensor device 1 according to the invention. This pilot sensor device 1 is provided with four antennas 111, 112, 121, 122. Two antennas 111, 121 are positioned above the fuselage 3 of the aircraft 2 and two antennas 112, 122 are positioned on the tail beam 4 of the aircraft 2. These antennas 111, 112, 121, 122 make it possible to pick up initial navigation signals emitted by different satellites. 101,102,201,202,301,302,401,402.
    [0025] The 101,102,201,202,301,302,401,402 satellites belong respectively to a 100,200, 300,400 GNSS system such as the GPS system, the GLONASS system, the GALILEO system, the QZSS system and the BEIDOU systems. According to a first embodiment of the piloting sensor device 1 shown in FIG. 2, the piloting sensor device 1 comprises four GNSS receiving means 11, 12, 13, 14 dedicated respectively to a 100,200,300,400 GNSS system and a second fault detection and exclusion module FDE2 31 connected to each GNSS receiving means 11,12,13,14. The pilot sensor device 1 is thus redundant at 100,200,300,400 GNSS systems and capable of covering four 100,200,300,400 GNSS systems and thus compensate for any failure of at least one of these 100,200,300,400 GNSS systems. Each GNSS receiving means 11, 12, 13, 14 is connected to two antennas 111, 112, 121, 122, 131, 132, 141, 142 and comprises a first detection and fault exclusion module FDE1 21, 22, 23, 24. Each first fault detection and exclusion module FDE1 21,22,23,24 receives and analyzes the initial navigation signals of a GNSS receiving means 11,12,13,14 in order to detect initial signals. navigational rules and initial erroneous navigation signals. Then, each first fault detection and exclusion module FDE1 21,22,23,24 can determine from these initial signals integrity a first ground speed signal of the aircraft 2. This first ground speed signal of the aircraft 2 can be determined for example by a V-RAIM method of independent control of the integrity. In fact, each GNSS receiving means 11, 12, 13, 14 can deliver a first ground speed signal of the aircraft 2 by guaranteeing a first level of autonomous surveillance and the integrity of this first ground speed signal in case of simple failures of a satellite of a GNSS system 100,200,300,400. The second fault detection and exclusion module FDE2 31 receives and compares these first ground speed signals from the four GNSS receiving means 11, 12, 13, 14. The second FDE2 detection and exclusion module 31 can then detect multiple failures of at least one 100,200,300,400 GNSS system, exclude each 100,200,300,400 GNSS system undergoing this multiple failure and determine a second ground speed signal from the aircraft 2. The second FDE2 Fault Detection and Exclusion Module 31 may apply the known method of the median to determine, from the first two ground speed signals, the second ground speed signal. According to a second embodiment of the pilot sensor device 1 shown in FIG. 3, the pilot sensor device 1 comprises two GNSS receiving means 11, 12, each GNSS receiving means 11, 12 being connected to two GNSS receivers 11, 12, antennas 111, 112, 121, 122, a second FDE2 detection and exclusion module 31, two IMU 51.52 inertial modules, two hybridization platforms 61, 62 and a third FDE3 detection and exclusion module 41.
    [0026] In addition, each GNSS receiving means 11, 12 comprises a first FDE1 21,22 detection and fault exclusion module as well as a 115,125 atomic clock. This atomic clock 115, 125 is used as a frequency reference making it possible to reduce by one unit the number of satellites 10 required for each GNSS reception means 11, 12 to determine, on the one hand, a simple satellite failure and, on the other hand, a first ground speed signal. The second fault detection and exclusion module FDE2 31 is connected to the two GNSS receiving means 11, 12, and 15 to the two hybridization platforms 61, 62 and delivers a second ground speed signal from the aircraft 2. Each IMU 51.52 inertial module provides inertial measurement signals of accelerations and angular velocities and is connected to a hybridization platform 61, 62. An IMU inertial module 51,52 and the hybridization platform 61,62 to which it is connected thus form an inertial chain 71,72. Each hybridization platform 61, 62 receives the inertial measurements of accelerations and angular velocities and can then determine a pure inertial ground speed of the aircraft 2.
    [0027] Each hybridization platform 61, 62 also receives the second ground speed signal from the aircraft 2 and can then process this second ground speed signal and the pure inertial ground speed of the aircraft 2 in order to determine a third signal of ground speed of the aircraft 2.
    [0028] This third ground speed signal is a hybridized ground speed of this second ground speed signal and the pure inertial ground speed when this second ground speed signal is integral. This third ground speed signal is equal to the pure inertial ground speed when this second ground speed signal is erroneous or unavailable. This third ground speed signal is thus continuously available. The third fault detection and exclusion module FDE3 41 is connected to the two hybridization platforms 61, 62. This third fault detection and exclusion module FDE3 41 thus receives, analyzes and compares two third ground speed signals from the aircraft 2 to determine a fourth ground speed signal of the aircraft 2 using, for example, the method of the median.
    [0029] According to this second embodiment, the integrity and the availability of the fourth ground speed signal of the aircraft 2 is improved by using two IMU 51,52 inertial modules and two GNSS receivers 11,12 of two independent 100,200 GNSS systems and distinct.
    [0030] According to a third embodiment of the control sensor device 1 shown in FIG. 4, the control sensor device 1 comprises, like the second embodiment, two GNSS reception means 11, 12, a second detection module and FDE2 31, two IMU 51.52 inertial modules, two hybridization platforms 61, 62, and a third FDE3 fault detection and exclusion module 41. Each GNSS receiving means 11, 12 is dedicated to a single 100,200 GNSS system, thus making it possible to cover two 100,200 GNSS systems, for example the GPS system and the GALILEO 30 system.
    [0031] The pilot sensor device 1 also comprises a computer 200 provided with two calculation channels 201, 202. Each hybridization platform 61, 62 comprises a purely inertial virtual platform 81, 82 and two hybridization error filters 91, 91 ', 92, 92', a hybridization error filter 91, 91 ', 92, 92 'being located in each calculation channel 201,202. According to each calculation channel 201, 202, the second fault detection and exclusion module FDE2 31 is connected to the two GNSS reception means 11, 12 and to two hybridization error filters 91, 91 ', 92, 92'. as well as the third FDE3 detection and exclusion exclusion module 41. The second FDE2 detection and exclusion exclusion module 31 thus delivers a second integrated ground speed signal from the aircraft 2. A purely inertial virtual platform 81,82 is connected to an IMU 51,52 inertial module and thus forms with this IMU 51,52 inertial module an inertial unit 101,102 providing a pure inertial ground speed of the aircraft 2. Each hybridization error filter 91, 91 ', 92, 92' is preferably a Kalman filter.
    [0032] Each hybridization error filter 91, 91 ', 92, 92' receives, analyzes and compares the second ground speed signal of the aircraft 2 and the pure inertial ground speed of the aircraft 2, and then determines the third signal. ground speed of the aircraft 2 which can be a hybridized ground speed or the speed pure inertial soil. This third ground speed signal is thus continuously available. The third fault detection and exclusion module FDE3 41 is connected to the hybridization error filters 91, 91 ', 92, 92' and to the second fault detection and exclusion module 3030058 24 FDE2 31. This third fault detection and exclusion module FDE3 41 then receives, analyzes and compares two third ground speed signals and the second ground speed signal 2, and then determines a fourth ground speed signal of the aircraft 2 according to these conditions. two calculation channels 201,202. The third FDE3 detection and exclusion module 41 uses, for example, the median method. In fact, the control sensor device 1 makes it possible to ensure continuity of supply of a fourth ground speed signal. The operation of this third embodiment is similar to the operation of the second embodiment. The use of two GNSS receiving means 11, 12 and two inertial units 101, 102 makes it possible to guarantee the availability and integrity of the fourth ground speed signal of the aircraft 2, even in the event of the unavailability of a second signal of Ground speed integrates. Advantageously, the comparison of the second and third ground speed signals at the third fault detection and exclusion module FDE3 41 makes it possible to detect anomalies on several GNSS 100,200 systems which could have gone unnoticed according to the second embodiment of the invention. pilot sensor device 1, such as interference or decoys, for example. The integrity of this fourth ground speed signal is thus increased and is then sufficient for a device to assist the piloting of the aircraft 2. Finally, this piloting sensor device 1 uses standard components, such as in particular two GNSS receiving means 11,12 and two inertial units 101,102 for example, in order to reduce its cost.
    [0033] 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 possible modes. It is of course conceivable to replace a means described by equivalent means without departing from the scope of the present invention.
    权利要求:
    Claims (11)
    [0001]
    REVENDICATIONS1. Pilot sensor device (1) for a rotary wing aircraft (2) comprising: - GNSS receiver means (11, 12, 13, 14) constellations of at least two independent GNSS systems (100, 200, 300, 400); said GNSS receiving means (11,12,13,14) receiving initial signals from a plurality of satellites (101,102,201,202,301,302,401,402), -at least three FDE fault detection and exclusion modules (21,22,31,41), each fault detection and exclusion module FDE (21,22,31,41) receiving at least two input signals and delivering an output signal, each output signal including a measurement and a state of integrity, characterized in each GNSS receiving means (11, 12) comprises a first fault detection and exclusion module FDE1 (21,22) per GNSS system (100,200), each first fault detection and exclusion module FDE1 (21,22) receives and analyzes said initial signals and detects embedded initial signals and / or erroneous initial signals, - each GNSS receiving means (11, 12) delivers a measurement and a state of integrity of a first ground speed signal from said aircraft (2) to a geographical reference for at least one system GNSS (100,200) from said integrity-free initial signals, possibly excluding said erroneous initial signals, - said pilot sensor device (1) comprises at least a second FDE2 fault detection and exclusion module (31), each second module for detecting and excluding 27 FDE2 failures (31) being in communication with at least two of said GNSS receiving means (11, 12) and receiving, analyzing and comparing said first ground speed signals delivered by said at least two two GNSS receiving means 5 (11,12), detecting first ground speed signals integrity and / or first ground speed signals erroneous, each second fault detection and exclusion module FDE2 (31) can then de detect and exclude each defective GNSS system (100,200) providing a first erroneous ground speed signal, and then determine and deliver a measurement and a state of integrity of a second ground speed signal from said aircraft (2) from minus two first ground speed signals, excluding, where appropriate, said first wrong ground speed signals. 15
    [0002]
    2. Control sensor device (1) according to claim 1, characterized in that said control sensor device (1) comprises at least one IMU inertial module (51,52) and at least one hybridization platform (61 , 62), each IMU inertial module (51, 52) providing inertial measurement signals characterizing accelerations and angular velocities of said aircraft (2), each hybridization platform (61, 62) being in communication with a second FDE2 detection and fault exclusion module (31) and an IMU inertial module (51,52), - each hybridization platform (61,62) receives and processes said inertial measurement signals and possibly a second signal ground speed, then determines and delivers a measurement constituting a third ground speed signal of said aircraft (2) from said inertial measurement signals and optionally 3030058 28 of a second integrated ground speed signal, said third speed signal e sol being a pure inertial sol speed when said hybridization platform (61,62) receives no second integrated ground speed signal and a hybridized ground speed when said hybridization platform (61,62) receives a second hybridization signal (61,62). ground speed integrates, said third ground speed signal then being continuously available.
    [0003]
    3. Control sensor device (1) according to claim 2, characterized in that said control sensor device (1) comprises at least two hybridization platforms (61,62), at least two IMU inertial modules ( 51, 52) and at least one third FDE3 detection and exclusion module (41), a hybridization platform (61, 62) and an IMU inertial module (51, 52) forming an inertial chain (71). , 72), each third fault detection and exclusion module FDE3 (41) being in communication with at least two hybridization platforms (61,62) for determining and outputting a fourth ground speed signal, each third FDE3 fault detection and exclusion module (41) receiving, analyzing and comparing said third ground speed signals delivered by said hybridization platforms (61,62) and detecting third ground speed signals of integrity and / or third wrong ground speed signals, each t third fault detection and exclusion module FDE3 (41) can then detect a failure on an inertial chain (71,72) and possibly exclude said inertial chain (71,72) from said erroneous third ground speed signals, and then determining and delivering a measurement and a state of integrity of said fourth ground speed signal of said aircraft (2) from at least two third solid ground speed signals 3030058 29 excluding, if appropriate, said third ground speed signals. wrong.
    [0004]
    4. Control sensor device (1) according to claim 3, characterized in that each third fault detection and exclusion module FDE3 (41) is in communication with at least one second detection and exclusion module FDE2 failures (31) to receive, analyze and compare at least a second ground speed signal and said third ground speed signals, to detect and locate second and / or third ground speed signals as well as second and / or third ground speed signals erroneously, then to determine and deliver a measurement and a state of integrity of said fourth ground speed signal of said aircraft (2) from at least one second signal of ground speed and / or at least two third ground speed signals, excluding, where appropriate, said second and / or third errored ground speed signals.
    [0005]
    5. Piloting sensor device (1) according to any one of claims 3 to 4, characterized in that - each hybridization platform (61,62) comprises a purely inertial virtual platform (81,82) and two Hybridization error filters (91, 92) communicating with each other, a purely inertial virtual platform (81, 82) communicating with an IMU Inertial Module (51, 52), thereby forming an inertial unit (101, 102); said pilot sensor device (1) comprises two inertial units (101, 102), a calculating two-way calculator (200) (201, 202) and four hybridization error filters (91, 91, 92, 92). '), each calculation channel (201,202) relating a hybridization error filter (91,91', 92,92 ') to a second fault detection and exclusion module FDE2 ( 31) and on the other hand a third fault detection and exclusion module FDE3 (41), - each second detection module and fault exclusion FDE2 (31) being in communication with two hybridization error filters (91, 92) for each calculation channel (201, 202), each third fault detection and exclusion module FDE3 ( 41) being in communication with two hybridization error filters (91, 92) for each calculation channel (201, 202).
    [0006]
    Pilot sensor device (1) according to one of Claims 3 to 5, characterized in that at least one second fault detection and exclusion module FDE2 (31, 32) and / or minus a third fault detection and exclusion module FDE3 (41,42) uses (s) a method of determination according to the median value. 20
    [0007]
    7. Control sensor device (1) according to any one of claims 1 to 6, characterized in that at least one GNSS receiving means (11,12) comprises an atomic clock (115,125).
    [0008]
    8. A control sensor device (1) according to any one of claims 1 to 7, characterized in that said GNSS receiving means (11, 12) are in communication with GNSS systems (100, 200), chosen from a list containing the GPS system, the 3030058 31 GLONASS system, the GALILEO system, the QZSS system and the BEIDOU systems as well as the IRIDIUM system.
    [0009]
    9. A method for determining a ground speed of an aircraft (2), characterized in that, in a first step, initial navigation signals are received from several satellites (101, 102, 201, 201) belonging to constellations of two or more independent GNSS systems (100,200),
    [0010]
    In a second step, said initial signals are analyzed for each GNSS system (100,200); in a third step, intact initial signals and / or initial signals are detected; a fourth step, providing a measurement and a state of integrity of at least two first ground speed signals of said aircraft (2) at a geographical reference point for at least two independent GNSS systems (100,200) from said initial signals with integrity excluding, if appropriate, said erroneous initial signals, during a fifth step, said first ground speed signals are analyzed and compared, during a sixth step, first ground speed signals are detected which are integral. and / or first erroneous ground speed signals, - in a seventh step, each defective GNSS (100,200) system providing a first errored ground speed signal is detected and excluded, and - in the course of one hour In the second step, a measurement and a state of integrity of a second ground speed signal of said aircraft (2) are determined and delivered from at least two first ground speed signals 300058 32 excluding, where appropriate said first ground speed signals erroneous. 10. A method for determining a ground speed of an aircraft (2) according to claim 9, characterized in that, during a ninth step, one acquires inertial measurement signals, said inertial measurement signals. characterizing accelerations and angular velocities of said aircraft (2), during a tenth step, each second ground speed signal and said inertial measurement signals are processed, during an eleventh step, and at least one measurement constituting at least one third ground speed signal from said aircraft (2) is provided from said inertial measurement signals and, if appropriate, from a second integrated ground speed signal, said third ground speed signal being then continuously available, in the course of a twelfth step, the said third ground speed signals are analyzed and compared, during a thirteenth step, third signaling is detected. ground zero and / or third ground speed signals erroneous, and during a fourteenth step, a measurement and a state of integrity of a fourth ground speed signal of said aircraft (2 ) from at least two solid third ground speed signals, excluding, if appropriate, said third errored ground speed signals.
    [0011]
    11. A method for determining a ground speed of an aircraft (2) according to claim 9, characterized in that, during a ninth step, one acquires inertial measurement signals, said measurement signals. inertial data characterizing accelerations and angular velocities of said aircraft (2), during a tenth step, each second ground speed signal and said inertial measurement signals are processed, during an eleventh step, determining and at least one measurement constituting at least one third ground speed signal of said aircraft (2) is provided from said inertial measurement signals and, if appropriate, from a second integrated ground speed signal, said third ground speed signal being then continuously available, - during said twelfth step, at least one second ground speed signal and said third ground speed signals are analyzed and compared, - during said in the thirteenth step, second and / or third ground speed signals are detected as well as second and / or third ground speed signals 20 are erroneously detected, and during said fourteenth stage, it is determined and provides a measurement and a state of integrity of a fourth ground speed signal of said aircraft (2) from at least one second ground speed signal and / or at least two third solid ground speed signals excluding, if appropriate, said second and / or third errored ground speed signals.
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    同族专利:
    公开号 | 公开日
    RU2668077C1|2018-09-26|
    WO2016092160A1|2016-06-16|
    EP3230767A1|2017-10-18|
    EP3230767B1|2018-08-01|
    FR3030058B1|2016-12-09|
    US10935672B2|2021-03-02|
    CN107110975B|2021-01-05|
    CN107110975A|2017-08-29|
    US20170336517A1|2017-11-23|
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    法律状态:
    2015-12-21| PLFP| Fee payment|Year of fee payment: 2 |
    2016-06-17| PLSC| Search report ready|Effective date: 20160617 |
    2016-12-22| PLFP| Fee payment|Year of fee payment: 3 |
    2017-12-21| PLFP| Fee payment|Year of fee payment: 4 |
    2018-12-19| PLFP| Fee payment|Year of fee payment: 5 |
    2020-10-16| ST| Notification of lapse|Effective date: 20200906 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1402824A|FR3030058B1|2014-12-11|2014-12-11|REDUNDANT DEVICE FOR STEERING SENSORS FOR ROTATING CAR AIRCRAFT|FR1402824A| FR3030058B1|2014-12-11|2014-12-11|REDUNDANT DEVICE FOR STEERING SENSORS FOR ROTATING CAR AIRCRAFT|
    RU2017116970A| RU2668077C1|2014-12-11|2015-12-09|Reserved piloting device with sensors for a rotary-wing aircraft|
    PCT/FR2015/000223| WO2016092160A1|2014-12-11|2015-12-09|Redundant device of piloting sensors for a rotary-wing aircraft|
    CN201580061339.5A| CN107110975B|2014-12-11|2015-12-09|Redundant device for a piloting sensor of a rotorcraft|
    EP15823344.5A| EP3230767B1|2014-12-11|2015-12-09|Redundant device of piloting sensors for a rotary-wing aircraft|
    US15/534,053| US10935672B2|2014-12-11|2015-12-09|Redundant device of piloting sensors for a rotary-wing aircraft|
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