• 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:
    An inertial measurement system for an aircraft, comprising at least one processing unit connected to at least two inertial units, namely at least one first type inertial unit comprising at least three angular sensors and three linear sensors aligned on three distinct sensitive axes, and an inertial unit of the second type, comprising at least four angular sensors and four linear sensors aligned on four distinct sensitive axes. The system is arranged to calculate an inertial data from the six sensors of the first type of inertial unit and four inertial data from four different combinations of three angular sensors and three linear sensors selected from the sensors of the second inertial unit. type and the processing unit is arranged to compare with each other the four inertial data from the measurements of the second type of inertial unit so as to determine if the second type of inertial unit fails.
    公开号:FR3038048A1
    申请号:FR1555758
    申请日:2015-06-23
    公开日:2016-12-30
    发明作者:David Roberfroid；Jean-Francois Kieffer
    申请人:Sagem Defense Securite SA；
    IPC主号:
    专利说明:

    [0001] The present invention relates to the inertial detection of flight parameters of an aircraft for the purpose in particular of controlling the mobile flying surfaces of the aircraft. These flight parameters include the attitude and the vector of movement of the aircraft. STATE OF THE ART There are aircraft equipped with an airborne inertial reference system (ADIRS) comprising three inertial units of the same type (ADIRU 1, ADIRU 2 and ADIRU 3, of the English "Air Data Inertial Reference Unit") each comprising a nominal set of three angular inertial sensors and three linear inertial sensors aligned on three sensitive axes of a measurement mark. . The inertial units are connected to a processing unit to each provide the processing unit with inertial measurements that can be used by the processing unit for controlling the mobile flying surfaces of the aircraft. Generally, the inertial measuring system is arranged so that the measurements of the ADIRU 1 are presented to the pilot, the measurements of the ADIRU 2 are presented to the co-pilot and the measurements of the ADIRU 3, which is used as a redundant unit, can be presented either to the pilot or to the co-pilot in case of failure of ADIRU 1 or ADIRU 2 respectively. To detect such a failure, the processing unit continuously compares the measurements of the three inertial units to each other to ensure consistency between them.
    [0002] When the measurements provided by one of the inertial units deviate from the measurements provided by the two other inertial units, said inertial unit is declared to be defective by the processing unit. A problem arises when there are only two inertial units left and their measurements start to diverge because it is not possible to determine which inertial unit fails without using other measuring instruments. OBJECT OF THE INVENTION An object of the invention is to provide a means for facilitating the determination of a failure in an inertial measurement system. BRIEF SUMMARY OF THE INVENTION For this purpose, an aircraft inertial measurement system is provided according to the invention, comprising at least one processing unit connected to at least two inertial units, namely at least one inertial unit of a first type comprising at least three angular sensors and three linear sensors aligned on three distinct sensitive axes, and an inertial unit of a second type comprising at least four angular sensors and four linear sensors aligned on four distinct sensitive axes. The processing unit is arranged to receive an inertial data from the measurements provided by the six inertial sensor sensors of the first type and four inertial data from four distinct combinations of three angular sensors and three linear sensors selected from the sensors of the inertial unit of the second type. The processing unit is further arranged to compare the four inertial data from the measurements of the inertial unit of the second type with one another so as to determine if the inertial unit of the second type fails.
    [0003] Thus, the presence of four angular sensors and four linear sensors in the second type of inertial unit makes it possible to obtain four inertial data (for example the attitude of the aircraft) for the calculation of which have been in turn excluded. four linear sensors and one of four angular sensors. If all the sensors work correctly, the inertial data calculated from the measurements of the inertial unit of the second type are substantially identical. In this case, a discrepancy between the inertial data from measurements of the first type of inertial unit and the inertial data from the second type of inertial unit would reveal a failure of the first type of inertial unit. If at least one of the sensors of the second type of inertial unit fails, the four inertial data will not be affected in the same way by this failure so that the inertial data resulting from the measurements of the inertial unit of second type will diverge. Detecting such a divergence thus reveals the failure. Other features and advantages of the invention will emerge on reading the following description of particular non-limiting embodiments of the invention.
    [0004] BRIEF DESCRIPTION OF THE FIGURES Reference will be made to the accompanying single figure schematically showing a system according to the invention. DETAILED DESCRIPTION OF THE INVENTION The inertial measurement system according to the invention is here described in application to an aircraft piloted by a pilot and a co-pilot and is arranged to respectively provide the pilot and the co-pilot with inertial data calculated from the data inertial values of two distinct inertial cen- trals. Other applications are of course possible. This inertial system comprises a processing unit 1 connected to two inertial units, namely an inertial unit 10 of a first type and an inertial unit 20 of a second type. The inertial measurement system 3038048 4 of the invention is mounted on the aircraft and is connected to a first control unit 31 having, in a manner known per se, flight parameters to the pilot of the aircraft and to a second control unit. 32 in a manner known per se flight parameters to the co-pilot of the aircraft. These flight parameters are here displayed on a screen of each control unit 31, 32. The inertial measurement system is also connected to an assistance computer for piloting the aircraft, such a computer being known in itself. . The inertial unit 10 here comprises three angular sensors 11 and three linear sensors 12 aligned on three distinct sensitive axes 13 (the sensors and the axes are here identified individually by means of the indices 1, 2, 3 associated respectively with the references 11, 12 and 13). The three sensitive axes 13 are perpendicular to each other. The inertial unit 10 comprises a control circuit 14 for the sensors 11, 12. The control circuit 14 is arranged to retrieve measurement signals from the sensors 11, 12 and to transmit, to the processing unit 1, these measurement signals or signals representative of the measurement signals. In the case where at least one of the sensors 11, 12 needs to be energized to operate, the control circuit 14 is arranged to send excitation signals to said sensor. Here the angular sensors 11 are vibrating resonator gyrometers and the linear sensors 12 are vibrating resonator accelerometers. The control circuit 14 is thus arranged to exclude the angular sensors 11 and the linear sensors 12. The inertial unit 10 is designed to calculate an inertial data from the three angular sensors 11 and three linear sensors 12. The Inertial unit 20 here comprises four angular sensors 21 and four linear sensors 22 aligned on four distinct sensitive axes 23 (the sensors and the axes are here identified individually by means of indices 1, 2, 3, 4 associated respectively with the references 21 , 22 and 23). The four sensitive axes 23 5 form the diagonals of a cube. The inertial unit 20 further comprises a first control circuit 24.1 of a first set of sensors comprising the angular sensors 21.1, 21.2, 21.3 and the linear sensors 22.1, 22.2, 22.3 and a second control circuit 24.2 of a second one. sensor set comprising the angular sensor 21.4 and the linear sensor 22.4. In the case where at least one of the sensors 21, 22 needs to be excited to operate, the control circuit 24.1, 24.2 is arranged to send to said sensor excitation signals. The angular sensors 21 are here vibrating resonator gyrometers and the linear sensors 22 are here vibrating resonator accelerometers. The control circuit 24 is thus arranged to excite the angular sensors 21 and the linear sensors 22.
    [0005] The inertial unit 20 is arranged to compute four inertial data from four distinct combinations of three angular and three linear sensors selected from the sensors 21, 22, using a different algorithm than the one used. in the inertial unit 10. Each inertial data includes an attitude of the aircraft and a vector of movement of the aircraft. These four inertial data are transmitted to the processing unit 1. The processing unit 1 transmits to the first control unit 31 the inertial data calculated from the measurements provided by the six sensors 11 and 12 (this data inertial is hereinafter referred to as the first main inertial data), and the processing unit 1 transmits to the second control unit 32 the inertial data calculated from the measurements provided by the six sensors 21.1, 21.2, 21.3 and 22.1. , 22.2, 22.3 (this inertial data is hereinafter referred to as the second main inertial data). Prior to the transmission of the main inertial data to the control units 31, 32, the processing unit 1 compares the first main inertial data and the second main inertial data calculated at each instant. If the first main inertial data and the second main inertial data computed at a time are identical or have a difference between them less than a predetermined threshold (taking into account the expected accuracy of the calculated inertial data), the main inertial data are transmitted to the piling units 31, 32. If the first main inertial data and the second main inertial data computed at a time have a difference greater than the predetermined threshold between them, the processing unit 1 compares the data with one another. following - the second main inertial data, - the calculated inertial data, at the same time as the second main inertial data, from the measurements provided by the six sensors 21.1, 21.2, 21.4 and 22.1, 22.2, 22.4 (this data inertial is called the first additional inertial data), - the calculated inertial data, at the same time as the second main inertial data, from the measurements provided by the six sensors 21.1, 21.4, 21.3 and 22.1, 22.4, 22.3 (this inertial data is called the second additional inertial data), the calculated inertial data, at the same time as the second main inertial data, from the measurements provided by the six sensors 21.4, 21.2, 21.3 and 22.4, 22.2, 22.3 (this inertial data is called the third additional inertial data).
    [0006] If the additional inertial data and the second main inertial data converge (they are identical or the difference between them is less than the predetermined threshold), the processing unit 1 transmits an alert signal to the control units 31, 32. informing 10 that the first inertial unit 10 has failed and that only the second main inertial data will now be presented to the pilot and co-pilot. If the additional inertial data and the second main inertial data diverge (the difference between them is greater than the predetermined threshold), the processing unit 1 transmits an alert signal to the control units 31, 32 informing that the second unit Inertial 20 is defective and only the first main inertial data will now be presented to the pilot 20 and the co-pilot. In the first case, it is possible to maintain a comparison of the additional inertial data and the second main inertial data to detect a failure of the second inertial unit 20.
    [0007] The processing unit 1 is thus arranged to calculate a difference between the inertial data resulting from the measurements of the first inertial unit and one of the inertial data from the second inertial unit and to issue an alert when the difference is greater than 30 to a predetermined value. It will be understood that the comparison of the inertial data makes it possible to verify their homogeneity and coherence, and thus to verify the integrity of the inertial measurements of the inertial units and the correct operation of the inertial units and their sensors.
    [0008] Of course, the invention is not limited to the embodiments described but encompasses any variant within the scope of the invention as defined by the claims.
    [0009] In particular, the inertial units may comprise more angular sensors and more linear sensors than in the example described. The system can include multiple processing units.
    [0010] The system may include more inertial units. For example, one can have two inertial units of the first type whose measurements are used to provide two inertial data. The invention is applicable to any type of algorithm, for example of the ARS or AHRS type with stabilization of the attitudes by accelerometric measurements, with or without the air speed, for the second type of inertial unit. The invention is applicable to purely inertial measurements or hybrid measurements with inertial measurements and measurements of another type, for example satellite geolocation measurements or measurements from a magnetometer. The second type of inertial unit may comprise a single control circuit for controlling all the sensors of the control unit or comprise two control circuits for one of the control circuits to be associated with a portion of the sensors of the inertial unit of the control unit. the second type and the other control circuit 30 is associated with the other part of the sensors of the second type of inertial unit. Ideally, one of the control circuits is associated with six of the sensors of the second type of inertial unit (three of the angle sensors and three of the linear sensors) and the other of the control circuits is associated with the other two of the sensors. inertial unit of the second type. The control circuits are, preferably in this case, different technologies. As a variant, the sensitive axes of the sensors of the second type of inertial unit may comprise three axes coinciding with three canonical axes and an axis arranged along a trisector of these three axes. 10
    权利要求:
    Claims (6)
    [0001]
    REVENDICATIONS1. An inertial measurement system for an aircraft, comprising at least one processing unit connected to at least two inertial units, namely at least one first type inertial unit comprising at least three angular sensors and three linear sensors aligned on three distinct sensitive axes, and an inertial center of the second type, comprising at least four angular sensors and four linear sensors aligned on four distinct sensitive axes, the system being arranged to calculate an inertial data from the six sensors of the first type of inertial unit and four data items. inertial measurements from four distinct combinations of three angular sensors and three linear sensors selected from the sensors of the second type of inertial unit and the processing unit being arranged to compare the four inertial data from the measurements of the inertial unit. second type of to determine if the inertial unit of the second type is faulty.
    [0002]
    2. System according to claim 1, the four inertial data from the measurements of the inertial unit of the second type are calculated using the same algorithm.
    [0003]
    3. A system according to claim 1, wherein the second type of inertial unit comprises a first control circuit of a first set of sensors comprising three of the angular sensors and three of the linear sensors and a second control circuit of a second one. second set of sensors comprising the fourth angular sensor and the fourth linear sensor.
    [0004]
    4. System according to claim 1, wherein the processing unit is arranged to calculate a difference between the inertial data from measurements of the first type of inertial unit and one of the inertial data from the inertial unit of second type 5 and issue an alert when the deviation is greater than a predetermined value.
    [0005]
    5. System according to claim 1, wherein the sensitivity axes of the sensors of the second type of inertial unit comprise three axes coinciding with three canonical axes and an axis arranged along a trisector of these three axes.
    [0006]
    The system of claim 1, wherein the sensitivity axes of the sensors of the second type of inertial unit are aligned on diagonals of a cube. 20
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    同族专利:
    公开号 | 公开日
    CN107787441A|2018-03-09|
    CN107787441B|2022-03-01|
    EP3314209A1|2018-05-02|
    FR3038048B1|2017-07-07|
    EP3314209B1|2021-10-13|
    WO2016207029A1|2016-12-29|
    US10415995B2|2019-09-17|
    US20180195878A1|2018-07-12|
    引用文献:
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    WO2021261557A1|2020-06-24|2021-12-30|パナソニックＩｐマネジメント株式会社|Inertial force sensor|
    法律状态:
    2016-05-24| PLFP| Fee payment|Year of fee payment: 2 |
    2016-12-30| PLSC| Publication of the preliminary search report|Effective date: 20161230 |
    2017-05-23| PLFP| Fee payment|Year of fee payment: 3 |
    2018-05-25| PLFP| Fee payment|Year of fee payment: 4 |
    2018-06-15| CD| Change of name or company name|Owner name: SAFRAN ELECTRONICS & DEFENSE, FR Effective date: 20180515 |
    2020-05-20| PLFP| Fee payment|Year of fee payment: 6 |
    2021-05-19| PLFP| Fee payment|Year of fee payment: 7 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1555758A|FR3038048B1|2015-06-23|2015-06-23|INERTIAL MEASUREMENT SYSTEM FOR AN AIRCRAFT|FR1555758A| FR3038048B1|2015-06-23|2015-06-23|INERTIAL MEASUREMENT SYSTEM FOR AN AIRCRAFT|
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    PCT/EP2016/063649| WO2016207029A1|2015-06-23|2016-06-14|Inertial measurement system for an aircraft|
    CN201680036831.1A| CN107787441B|2015-06-23|2016-06-14|Inertial measurement system for aircraft|
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