Method and system for measuring the exit parameters of a launched carrier

Abstract

The present invention discloses a method and system for measuring the exit parameters of a launched carrier, and belongs to the field of sensor measurement and control technology. The method for measuring the exit parameters of a launched carrier disclosed by the present invention comprises: installing a micro-electromechanical inertial sensor system on the carrier; obtaining the direction change data and acceleration change data of the carrier through the micro-electromechanical inertial sensor system; and determining the real-time motion parameter information of the carrier according to the direction change data and the acceleration change data. The present invention is used to accurately measure the parameters such as the initial attitude, launch time, real-time position, real-time attitude and exit time of the launched carrier, judge the motion stage of the carrier according to these parameters, and implement the corresponding steering control strategy to achieve the purpose of stabilizing the carrier attitude and accurately controlling the motion direction.

Description

Technical Field

[0001] The invention belongs to the technical field of sensor measurement and control, and in particular relates to a method and system for measuring the exit parameters of a launched carrier. Background Art

[0002] Autonomous powered tubular launch vehicles usually include a control system, which is used to measure and control the motion attitude and trajectory to improve the control accuracy and efficiency of the vehicle. The vehicle control system generally includes an inertial measurement system and a steering gear control system. During the launch phase of a tubular launch vehicle, the forward overload usually reaches 100 to 200g, and the roll angular rate can reach more than 4000° / s.

[0003] During the high-overload launch process, a carrier control system is required that can accurately measure parameters such as the initial attitude, launch time, real-time position, real-time attitude and exit time of the launched carrier under a drastically changing mechanical environment. Summary of the invention

[0004] The object of the present invention is to provide a method and system for measuring the exit parameters of a launched vehicle, which are used to accurately measure parameters such as the initial attitude, launch time, real-time position, real-time attitude and exit time of the launched vehicle.

[0005] In order to achieve the above object, the present invention provides the following technical solutions:

[0006] A method for measuring the exit tube parameters of a launched carrier, comprising:

[0007] installing a micro-electromechanical inertial sensor system on a carrier;

[0008] Acquiring direction change data and acceleration change data of the carrier through the micro-electromechanical inertial sensor system;

[0009] The real-time motion parameter information of the carrier is determined according to the direction change data and the acceleration change data.

[0010] The micro-electromechanical inertial sensor system includes a three-axis gyroscope data collector and a three-axis acceleration data collector.

[0011] The three-axis gyroscope data collector includes two high-precision gyroscopes and a large-range gyroscope;

[0012] The three-axis acceleration data collector includes two high-precision accelerometers and a large-range accelerometer.

[0013] Determining the real-time motion parameter information of the carrier according to the direction change data and the acceleration change data is specifically:

[0014] Step a, obtaining first startup information, setting the information processing module to enter an alignment working mode;

[0015] Step b, continuously collecting direction change data and acceleration change data of the carrier;

[0016] Step c, detecting whether the acceleration change data is greater than a preset value, if the acceleration change data is less than the preset value, go to step d, if the acceleration change data is greater than the preset value, go to step e;

[0017] Step d, detecting whether the information processing module is in an alignment working mode or a navigation working mode; if the information processing module is in the alignment working mode, collecting the direction change data and the acceleration change data of the carrier at the first startup information acquisition moment, calculating the real-time position of the carrier, and going to step g; if the information processing module is in the navigation working mode, going to step f;

[0018] Step e, setting the information processing module end to the navigation working mode, and proceeding to step f, wherein the working mode conversion of the information processing module from the alignment working mode to the navigation working mode is irreversible;

[0019] Step f, iteratively calculating the real-time position and the real-time state of the carrier through the direction change data and the acceleration change data of the carrier;

[0020] Step g, output the real-time location information of the carrier and the real-time status of the carrier, and then go to step b.

[0021] The step a comprises:

[0022] Power on and start;

[0023] Port initialization;

[0024] The information processing module is set to an alignment working mode.

[0025] The step f comprises:

[0026] Launch point coordinate system attitude solution;

[0027] Calculation of the launch point coordinate system position;

[0028] The posture information of the carrier at the launch point and the position information of the carrier at the launch point are output.

[0029] Compared with the prior art, in the method for measuring the exit parameters of a launched carrier provided by the present invention, a micro-electromechanical inertial sensor system is installed on the carrier to obtain direction change data and acceleration change data of the carrier, and finally the real-time motion parameter information of the carrier is determined according to the direction change data and the acceleration change data, so as to accurately measure the parameters such as the initial attitude, launch time, real-time position, real-time attitude and exit time of the launched carrier, judge the motion stage of the carrier according to these parameters, and implement the corresponding steering control strategy, so as to achieve the purpose of stabilizing the carrier attitude and accurately controlling the motion direction.

[0030] The present invention also provides a system for measuring the exit parameters of a launched carrier, the system comprising:

[0031] A processor, a communication driving circuit board having a plurality of communication interfaces and a data output interface, a plurality of gyroscopes and a plurality of accelerometers, wherein the processor is used to execute the above-mentioned method for measuring the outgoing tube parameters of the launched carrier;

[0032] The processor is electrically connected to the plurality of gyroscopes and the plurality of accelerometers respectively through a plurality of communication interfaces of the communication driving circuit board;

[0033] The processor is connected to the outside world through the data output interface.

[0034] The system for measuring the outgoing tube parameters of the launched carrier also includes a power module, and the power module is electrically connected to the processor and the communication drive circuit board respectively.

[0035] The plurality of gyroscopes include at least one large-range gyroscope and at least two high-precision gyroscopes;

[0036] The plurality of accelerometers include at least one large-range accelerometer and at least two high-precision accelerometers;

[0037] The launch vehicle exit parameter measurement system further comprises a cuboid, wherein at least one of the large-range gyroscopes and at least two of the high-precision gyroscopes are respectively arranged on three mutually perpendicular faces of the cuboid;

[0038] At least one of the large-range accelerometers and at least two of the high-precision accelerometers are respectively arranged on three mutually perpendicular faces of the rectangular parallelepiped.

[0039] The launch vehicle exit parameter measurement system further includes:

[0040] A housing, wherein the housing has a receiving cavity therein;

[0041] The cuboid is fixed at the bottom of the accommodating cavity, the communication driving circuit board is fixed on one side of the top of the accommodating cavity, the power module is fixed on the other side of the top of the accommodating cavity, and the processor is arranged on the communication driving circuit board.

[0042] Compared with the prior art, the beneficial effects of the system for measuring the exit parameters of a launched carrier provided by the present invention are the same as the beneficial effects of the method for measuring the exit parameters of a launched carrier described in the above technical solution, and will not be elaborated here. BRIEF DESCRIPTION OF THE DRAWINGS

[0043] Figure 1 It is a flowchart of the method for measuring the outgoing tube parameters of the launched carrier of the present invention;

[0044] Figure 2 Another flowchart of the method for measuring the exit tube parameters of the launched carrier of the present invention;

[0045] Figure 3 It is a structural block diagram of the system for measuring the exit parameters of the launched carrier of the present invention;

[0046] Figure 4 It is a cross-sectional view from a top view of the system for measuring the exit tube parameters of a launched carrier according to the present invention;

[0047] Figure 5 It is a cross-sectional view from a front view of the system for measuring the exit tube parameters of a launched carrier according to the present invention;

[0048] Figure 6 It is the internal circuit diagram of the emission parameter measurement system of the launched carrier of the present invention.

[0049] Figure 7 This is a diagram showing an example of a method for measuring the exit tube parameters of a launched carrier according to the present invention;

[0050] Figure numerals: 1. micro-electromechanical inertial sensor system; 2. large-range gyroscope; 3. large-range accelerometer; 4. high-precision gyroscope; 5. high-precision accelerometer; 6. communication driver circuit board; 7. processor; 8. SPI communication interface; 9. USART communication interface; 10. communication driver circuit; 11. data output interface; 12. power module; 13. 3.3V isolated power supply; 14. 5V isolated power supply; 15. shell; 16. mounting base; 17. cuboid; 18. SPI bus cable. DETAILED DESCRIPTION

[0051] In order to make the technical problems, technical solutions and beneficial effects to be solved by the present invention more clearly understood, the present invention is further described in detail below in conjunction with the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present invention and are not intended to limit the present invention.

[0052] The present invention is further described in detail below in conjunction with specific embodiments, which are intended to explain the present invention rather than to limit it.

[0053] Please refer to Figure 1 , which shows a flow chart of a method for measuring the outgoing tube parameters of a launched carrier according to the present invention, the method for measuring the outgoing tube parameters of a launched carrier comprises:

[0054] like Figure 4 As shown, in step 101, a micro-electromechanical inertial sensor system 1 is mounted on the carrier.

[0055] The above-mentioned micro-electromechanical inertial sensor system 1 includes a three-axis gyroscope data collector and a three-axis acceleration data collector. The three-axis gyroscope data collector includes two high-precision gyroscopes 4 and a large-range gyroscope 2; the three-axis acceleration data collector includes two high-precision accelerometers 5 and a large-range accelerometer 3. The high-precision gyroscope 4, the large-range gyroscope 2, the high-precision accelerometer 5 and the large-range accelerometer 3 can all adopt micro-electromechanical sensor modules.

[0056] The large-range gyroscope 2 is used to measure the roll angular rate, and the range of the large-range gyroscope 2 is 2000-9000° / s. In one example, the large-range gyroscope 2 can be an MGS4000 gyroscope, and its range is 4000° / s.

[0057] The large-range accelerometer 3 is used to measure the forward acceleration, and the range of the large-range accelerometer 3 is 100 to 200 g. In one example, the large-range accelerometer 3 can be a MAX200 accelerometer, and its range is 200 g.

[0058] The high-precision gyroscope 4 is used to measure angular velocity, and the measurement accuracy of the high-precision gyroscope 4 is in the range of 0.5 to 3° / h. In one example, the high-precision gyroscope 4 can be an MGS300 gyroscope, and its measurement accuracy is 3° / h.

[0059] The high-precision accelerometer 5 is used to measure the acceleration of the yaw and pitch axes, and its measurement accuracy is better than 300 μg. In one example, the high-precision accelerometer 5 can be a MAX030 accelerometer with a measurement accuracy of 300 μg.

[0060] Step 102, obtaining direction change data and acceleration change data of the carrier through the micro-electromechanical inertial sensor system 1;

[0061] Step 103: Determine the real-time motion parameter information of the carrier according to the direction change data and the acceleration change data. Figure 2 As shown, step 103 is specifically as follows:

[0062] Step a, obtain the first startup information and set the information processing module to enter the alignment working mode; in this step, the first startup information is used to wake up the information processing module. After the first startup information wakes up the information processing module, the information processing module executes, powers on, initializes the port, and sets the information processing module to the alignment working mode.

[0063] Step b, continuously collecting direction change data and acceleration change data of the carrier;

[0064] Step c, detecting whether the acceleration change data is greater than a preset value, if the acceleration change data is less than the preset value, go to step d, if the acceleration change data is greater than the preset value, go to step e;

[0065] Step d, detecting whether the information processing module is in an alignment working mode or a navigation working mode; if the information processing module is in the alignment working mode, collecting the direction change data and the acceleration change data of the carrier at the first startup information acquisition moment, calculating the real-time position of the carrier, and going to step g; if the information processing module is in the navigation working mode, going to step f;

[0066] Step e, setting the information processing module end to the navigation working mode, and going to step f, wherein the working mode conversion of the information processing module from the alignment working mode to the navigation working mode is irreversible; it should be noted that when the information processing module itself executes the commands stored in itself, it does not have the authority to convert the navigation working mode into the alignment working mode. However, the external master control system can issue instructions to instruct the information processing module to convert the navigation working mode into the alignment working mode.

[0067] Step f, iteratively calculating the real-time position and real-time state of the carrier through the direction change data and the acceleration change data of the carrier; this step f is used to send the attitude solution result of the launch point coordinate system, the position solution result of the launch point coordinate system, and output the attitude information of the carrier at the launch point and the position information of the carrier at the launch point.

[0068] Step g, output the real-time location information of the carrier and the real-time status of the carrier, and then go to step b.

[0069] In a specific embodiment, if Figure 7 An example flow chart of the measurement software is shown.

[0070] Step (1): After the measurement system of the carrier is powered by DC5V, the processor 7 is powered on and started;

[0071] Step (2): Initialize the measurement system's GPIO input and output ports (status indication and chip selection), SPI port (data acquisition) and USART serial port (data transmission). Also initialize various timing modules;

[0072] Step (3): After the measurement system is initialized, the status flag ST is first set to 1, indicating that the measurement system enters the alignment working mode (alignment working mode, i.e. ST = 1, i.e. alignment working mode). At this time, the carrier remains stationary and the measurement system does not output data. The time is 1 second. After the alignment is completed, ST = 1, the measurement system starts to output data, and can also maneuver to aim at the target.

[0073] Step (4): 1ms timing judgment, the software uses the timer interrupt service program 1ms timing cycle to perform main functional modules such as data acquisition, attitude and position solution and data output;

[0074] Step (5): 6-axis inertial sensor data acquisition, the 6-axis inertial sensor includes 3-axis gyroscope data acquisition and 3-axis acceleration data acquisition (large-range gyroscope 2, large-range accelerometer 3, two high-precision gyroscopes 4, two high-precision accelerometers 5), the SPI acquisition module with SPI port in the system sends a read command through the DIN line according to the SPI protocol of each sensor, and receives data through the DOUT line;

[0075] Step (6): The sensor data is decoded and weighted according to the protocol to obtain the carrier angular velocity and acceleration values, and error compensation is performed according to the installation error and temperature error coefficients;

[0076] Step (7): Calculate the combined acceleration ΣA of the three axes of the carrier according to formula (1) and determine whether the combined acceleration ΣA is greater than 2g. If ΣA>2g, go to step (11);

[0077] Where: Ax, Ay, Az are the acceleration measurement values ​​of the X, Y, and Z axes respectively.

[0078] Step (8): If the combined acceleration ΣA≤2g, determine whether the current state flag ST is equal to 1; otherwise, the carrier is operating in the measurement mode, and go to step (12);

[0079] Step (9): If ST = 1, the carrier is in the alignment stage, and the attitude is calculated according to the coordinate system at the time of power-on. Since the carrier is in a stationary state at this time, the pitch angle θ and the roll angle φ are calculated using the acceleration measurement value for inverse trigonometric function calculation (see formulas (2) and (3)). The algorithms of formulas (2) and (3) are commonly used calculation formulas in the inertial navigation industry. The coordinate system at the time of power-on uses the current carrier orientation as the initial yaw angle ψ0 = 0°, and the actual measured values ​​of the carrier pitch angle and roll angle at that time as the initial pitch angle θ0 and roll angle φ0;

[0080]

[0081]

[0082] Step (10): During the alignment phase, the measurement system needs to output data including time (ms), angular velocity, acceleration and attitude angle through the RS422 serial port according to the protocol. The time value is a negative ms and the position output is 0. Go to step (15);

[0083] Step (11): If ΣA>2g, it means that the carrier is ignited in the launch tube. The measurement software uses this time as the measurement timestamp 0, and the subsequent time is a positive number of ms; ST=2 is set (ST=2 means that the measurement system enters the navigation working mode), and the measurement system works in the measurement mode;

[0084] Step (12): The attitude solution after launch is different from the attitude solution before launch. After launch, the pitch angle and roll angle are switched to quaternion solution using angular rate measurement values, and the acceleration measurement values ​​are no longer used for inverse trigonometric function solution. This is because the acceleration value after launch changes dynamically, resulting in a large error in the attitude angle solution. The definition of quaternion is shown in formula (4). The new q value obtained by quaternion solution is used for the attitude transformation matrix Perform iterative calculation (see formula (5)) to obtain the new posture transformation matrix Formula (6) is the Euler angle representation of the attitude conversion matrix. The new attitude angle can be obtained by performing inverse trigonometric calculations according to formula (6). Formulas (4) to (6) are all referenced to commonly used calculation formulas to illustrate the software calculation process.

[0085] q=q0+q1i+q2j+q3k (4)

[0086]

[0087]

[0088] Step (13): Position solution after launch. As shown in formula (7), after launch, the three-axis accelerometer measurement value is integrated in the navigation coordinate system to obtain the speed, and the current speed of the navigation coordinate system is integrated twice to obtain the position increment of the acquisition and solution cycle of 1ms, and then the current position is obtained. The algorithm belongs to the commonly used calculation formula of inertial navigation and will not be discussed here. The position solution is based on the size of the launch device and the position solution result. The moment when the forward (X-direction) position value exceeds the length of the launch tube is the moment when the carrier leaves the tube (in ms). At this time, the control system can safely steer according to the control strategy to control the carrier attitude.

[0089]

[0090] Step (14): During the measurement phase, the measurement system needs to output data including time (in ms), angular velocity, acceleration, attitude angle and position through the RS422 serial port in a protocol format.

[0091] Step (15): The measurement system outputs the data packet time (in ms), three angular rates (Wx WyWz), three accelerations (Ax Ay Az), yaw angle ψ, pitch angle θ, roll angle φ and three positions (Sx Sy Sz) through the RS422 serial port, and the output is 0. Go to step (3).

[0092] In summary, in the method for measuring the exit parameters of a launched carrier provided by the present invention, a micro-electromechanical inertial sensor system is installed on the carrier to obtain the direction change data and the acceleration change data of the carrier, and finally the real-time motion parameter information of the carrier is determined according to the direction change data and the acceleration change data, so as to accurately measure the parameters such as the initial attitude, launch time, real-time position, real-time attitude and exit time of the launched carrier, judge the motion stage of the carrier according to these parameters, and implement the corresponding steering control strategy, so as to achieve the purpose of stabilizing the attitude of the carrier and accurately controlling the motion direction.

[0093] Please refer to Figure 3 , Figure 4 and Figure 5 The present invention also discloses a system for measuring the parameters of a launched carrier. The system for measuring the parameters of a launched carrier comprises:

[0094] A processor 7, a communication driving circuit board 6 having a plurality of communication interfaces and a data output interface 11, a plurality of gyroscopes and a plurality of accelerometers, wherein the processor 7 is used to execute the above-mentioned method for measuring the outgoing tube parameters of the launched carrier;

[0095] The processor 7 is electrically connected to the plurality of gyroscopes and the plurality of accelerometers respectively through the plurality of communication interfaces of the communication driving circuit board 6;

[0096] The processor 7 is connected to the outside world through the data output interface 11 for communication.

[0097] Among them, Figure 4 As shown, the plurality of gyroscopes include at least one large-range gyroscope 2 and at least two high-precision gyroscopes 4;

[0098] The plurality of accelerometers include at least one large-range accelerometer 3 and at least two high-precision accelerometers 5;

[0099] The launch vehicle exit parameter measurement system further comprises a cuboid 17, wherein at least one of the large-range gyroscopes 2 and at least two of the high-precision gyroscopes 4 are respectively arranged on three mutually perpendicular faces of the cuboid 17;

[0100] At least one of the large-range accelerometers 3 and at least two of the high-precision accelerometers 5 are respectively arranged on three mutually perpendicular faces of the cuboid 17 .

[0101] The launch vehicle exit parameter measurement system further includes:

[0102] The housing 15 has a receiving cavity inside and is fixedly connected to the launched carrier.

[0103] The cuboid 17 is fixed to the bottom of the accommodating cavity, the communication driving circuit board 6 is fixed to one side of the top of the accommodating cavity, the power module 12 is fixed to the other side of the top of the accommodating cavity, and the processor 7 is arranged on the communication driving circuit board 6.

[0104] The system for measuring the outgoing tube parameters of the launched carrier further includes a power module 12 , and the power module 12 is electrically connected to the processor 7 and the communication driving circuit board 6 , respectively.

[0105] In the above example, the shell 15 has an outer dimension of 70x55mm and is made of free-cutting structural steel. The data output interface 11 is installed on the top; the mounting base 16 is the main body of the measurement system structure, with an outer dimension of 68x68x33mm and a material of A12. There are 4 3.5-diameter through holes on the bottom flange of the mounting base, which are used for positioning and screwing the product on the carrier. A 8x8x0.5mm square platform is left at the bottom of each 3.5-diameter through hole to prevent warping deformation errors caused by full-surface contact.

[0106] A sensor installation cuboid 17 is designed at the center of the upper end of the mounting and fixing base 16. When designing the sensor installation cuboid 17, requirements are put forward for the mutual parallelism and verticality of the six faces, namely, the front, back, left, right, upper end face and the mounting table surface of the mounting and fixing base 16 on the carrier. It is required that the verticality of two adjacent faces of the cuboid 17 is 0.01, and the parallelism is 0.01. Gyroscopes and accelerometers are installed on the five faces of the sensor installation cuboid 17, wherein the upper end face is bonded with a large-range gyroscope 2 (Wx) and a high-precision accelerometer 5 (Az), the front end face is bonded with a large-range accelerometer 3 (Ax), the left end face is bonded with a high-precision gyroscope 4 (Wz); the right end face is bonded with a high-precision accelerometer 5 (Ay); and the rear end face is bonded with a high-precision gyroscope 4 (Wy).

[0107] The large-range gyroscope 2 has a range of 2000-9000° / s and a measurement accuracy better than 20° / h. In one example, a domestically produced MGS4000 gyroscope with a range of 4000° / s, a measurement accuracy of 20° / h and a package of LCC44 is used to measure the carrier roll angular rate.

[0108] The large-range accelerometer 3 has a range of 100-200g and a measurement accuracy better than 1mg. In one example, a domestically produced MAX200 accelerometer with a range of 200g, a measurement accuracy of 1mg, and a package of LCC44 is used to measure the acceleration of the carrier's forward axis.

[0109] The high-precision gyroscope 4 has a measurement accuracy better than 300 μg. In one example, a domestically produced MGS300 gyroscope is selected, with a range of 300° / s, a measurement accuracy of 3° / h, and a package of LCC44, for measuring the pitch and yaw angular rates of the carrier.

[0110] The high-precision accelerometer 5 is a domestically produced MAX030 accelerometer with a measuring range of 30g, a measuring accuracy of 300μg, and a package of LCC44, and is used to measure the yaw and pitch axial acceleration of the carrier.

[0111] The communication drive circuit board 6 at least includes: a processor 7; and an RS422 communication drive circuit 10. The processor 7 is the core of the entire measurement system, and the measurement software is downloaded into the program memory of the processor through the SWD port. The SPI communication interface 8 and the USART communication interface 9 are integrated on the processor 7.

[0112] The processor 7 uses the domestic MCU series CS32F103CB, packaged as LQFP48, with an operating voltage of 3.3V and an internal clock of not less than 72MHz. The processor 7 collects the carrier motion parameters in real time according to the software process, obtains the current attitude angle and position parameters of the carrier relative to the launch point coordinate system through the attitude position solution algorithm, and can accurately obtain the time value (ms) relative to the launch time, with an error of no more than 3ms, which is used for measuring and judging the carrier's tube exit position and time. The designed collection and solution cycle is 1ms

[0113] The SPI communication interface 8 is the host end of the SPI bus network, and the six sensors, namely, the large-range gyroscope 2, the large-range accelerometer 3, the high-precision gyroscope 4 and the high-precision accelerometer 5, are the slave ends. The processor 7 collects the carrier angular velocity and acceleration data measured by the gyroscope and the accelerometer in real time through the SPI communication interface 8.

[0114] The USART communication interface 9 is the slave end of the RS422 bus network. The RS422 communication network is composed of the USART communication interface 9, the RS422 communication drive circuit 10 and the data interface 11. The processor 7 uploads the posture position and other data to the carrier control computer through the USART communication interface 9 at a frequency of 1KHz through the data interface 11 connector.

[0115] The RS422 communication drive circuit 10 uses MAX3488ESA, full-duplex serial communication, SO-8 package, and an operating voltage of 3.3 V. The data interface 11 uses a J30J-15ZK connector.

[0116] The power supply module 12 at least comprises: a power supply module 12 having a 3.3V isolated power supply 13 and a 5V isolated power supply 14, and the power supply module 12 provides 3.3V and 5V working power supplies for the measurement system.

[0117] It should be noted that when an element is referred to as being "fixed to" or "disposed on" another element, it can be directly on the other element or indirectly on the other element. When an element is referred to as being "connected to" another element, it can be directly connected to the other element or indirectly connected to the other element.

[0118] In addition, the terms "standby" and "navigation" are used for descriptive purposes only and should not be understood as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Therefore, the features defined as "standby" and "navigation" may explicitly or implicitly include one or more of the features. In the description of the present invention, the meaning of "multiple" is two or more, unless otherwise clearly and specifically defined. The meaning of "several" is one or more, unless otherwise clearly and specifically defined.

[0119] In the description of the present invention, it is necessary to understand that the directions or positional relationships indicated by the terms "up", "down", "front", "back", "left", "right", etc. are based on the directions or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the present invention and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific direction, be constructed and operated in a specific direction, and therefore cannot be understood as a limitation on the present invention.

[0120] In the description of the present invention, it should be noted that, unless otherwise clearly specified and limited, the terms "installed", "connected", and "connected" should be understood in a broad sense, for example, it can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection or an indirect connection through an intermediate medium, it can be the internal connection of two elements or the interaction relationship between two elements. For ordinary technicians in this field, the specific meanings of the above terms in the present invention can be understood according to specific circumstances.

[0121] The above shows and describes the basic principles and main features of the present invention and the advantages of the present invention. It should be understood by those skilled in the art that the present invention is not limited to the above embodiments. The above embodiments and descriptions are only for explaining the principles of the present invention. Without departing from the spirit and scope of the present invention, the present invention may have various changes and improvements, which fall within the scope of the present invention to be protected. The scope of protection of the present invention is defined by the attached claims and their equivalents.

Claims

1. A method for measuring the exit parameters of a launched carrier, characterized in that: include: installing a micro-electromechanical inertial sensor system on a carrier; Acquiring direction change data and acceleration change data of the carrier through the micro-electromechanical inertial sensor system; Determine the real-time motion parameter information of the carrier according to the direction change data and the acceleration change data; Determining the real-time motion parameter information of the carrier according to the direction change data and the acceleration change data is specifically: Step a, obtaining first startup information, setting the information processing module to enter an alignment working mode; Step b, continuously collecting direction change data and acceleration change data of the carrier; Step c, detecting whether the acceleration change data is greater than a preset value, if the acceleration change data is less than the preset value, go to step d, if the acceleration change data is greater than the preset value, go to step e; Step d, detecting whether the information processing module is in an alignment working mode or a navigation working mode; if the information processing module is in the alignment working mode, collecting the direction change data and the acceleration change data of the carrier at the first startup information acquisition moment, calculating the real-time position of the carrier, and going to step g; If the information processing module is in the navigation working mode, go to step f; Step e, setting the information processing module end to the navigation working mode, and proceeding to step f, wherein the working mode conversion of the information processing module from the alignment working mode to the navigation working mode is irreversible; Step f, iteratively calculating the real-time position and the real-time state of the carrier through the direction change data and the acceleration change data of the carrier; Step g, output the real-time location information of the carrier and the real-time status of the carrier, and then go to step b.

2. The method for measuring the exit parameters of a launched carrier according to claim 1, characterized in that: The step a comprises: Power on and start; Port initialization; The information processing module is set to an alignment working mode.

3. The method for measuring the exit parameters of a launched carrier according to claim 1, characterized in that: The step f comprises: Launch point coordinate system attitude solution; Calculation of the launch point coordinate system position; The posture information of the carrier at the launch point and the position information of the carrier at the launch point are output.

4. The method for measuring the exit parameters of a launched carrier according to claim 1, characterized in that: The micro-electromechanical inertial sensor system includes a three-axis gyroscope data collector and a three-axis acceleration data collector.

5. The method for measuring the exit parameters of a launched carrier according to claim 4, characterized in that: The three-axis gyroscope data collector includes two high-precision gyroscopes and a large-range gyroscope; The three-axis acceleration data collector includes two high-precision accelerometers and a large-range accelerometer.

6. A system for measuring the outgoing tube parameters of a launched carrier based on the method for measuring the outgoing tube parameters of a launched carrier according to any one of claims 1 to 5, characterized in that: The launch vehicle exit parameter measurement system comprises: A processor, a communication driver circuit board having a plurality of communication interfaces and a data output interface, a plurality of gyroscopes and a plurality of accelerometers, wherein the processor is used to execute the method for measuring the outgoing tube parameters of the launched carrier according to any one of claims 1 to 5; The processor is electrically connected to the plurality of gyroscopes and the plurality of accelerometers respectively through a plurality of communication interfaces of the communication driving circuit board; The processor is connected to the outside world through the data output interface.

7. The system for measuring the exit parameters of a launched carrier according to claim 6, characterized in that: The system for measuring the outgoing tube parameters of the launched carrier also includes a power module, and the power module is electrically connected to the processor and the communication drive circuit board respectively.

8. The system for measuring the exit parameters of a launched carrier according to claim 7, characterized in that: The plurality of gyroscopes include at least one large-range gyroscope and at least two high-precision gyroscopes; The plurality of accelerometers include at least one large-range accelerometer and at least two high-precision accelerometers; The launch vehicle exit parameter measurement system further comprises a cuboid, wherein at least one of the large-range gyroscopes and at least two of the high-precision gyroscopes are respectively arranged on three mutually perpendicular faces of the cuboid; At least one of the large-range accelerometers and at least two of the high-precision accelerometers are respectively arranged on three mutually perpendicular faces of the rectangular parallelepiped.

9. The system for measuring the exit parameters of a launched carrier according to claim 8, characterized in that: The launch vehicle exit parameter measurement system further includes: A housing, wherein the housing has a receiving cavity therein; The cuboid is fixed at the bottom of the accommodating cavity, the communication driving circuit board is fixed on one side of the top of the accommodating cavity, the power module is fixed on the other side of the top of the accommodating cavity, and the processor is arranged on the communication driving circuit board.

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