Carrier rocket closed-loop test method based on real-time operation system

Abstract

A carrier rocket closed-loop test method based on a real-time operating system comprises the steps that 1, closed-loop test equipment is started, and closed-loop test software runs in the real-time operating system; 2, an external instruction controls closed-loop test software to read binding parameters, and kinetic model calculation is started; step 3, taking rocket launch information monitored by closed-loop test software MT (bus monitor) as input to the dynamic model, and calculating and outputting position, speed, attitude, angular velocity and acceleration information; 4, the inertial measurement unit model and the GNSS model calculate the positions and the speeds of the gyroscope, the meter adding pulse and the GNSS according to the dynamic model result; step 5, sending the gyro, the accelerometer pulse excitation signal and the GNSS position and speed to the inertial measurement unit equipment and the rocket aircraft through a serial port; and 6, after receiving the excitation signal, the inertial measurement unit superposes a self-sensitive error and outputs the superposed self-sensitive error to the rocket aircraft. The method is good in integration and high in practicability, and has efficient, safe and reliable engineering application effects.

Description

Technical Field

[0001] The present invention relates to a closed-loop test method for a launch vehicle based on a real-time operating system, and belongs to the technical field of launch vehicle experiments and tests. Specifically, the present invention relates to a simulated flight test method formed by input and output of closed-loop test software and launch vehicle flight software in an embedded real-time operating system. Background Art

[0002] Traditional launch vehicle testing methods mainly rely on ground integrated system testing and flight testing. Ground integrated system testing can often only verify the functional integrity of various stand-alone products. The flight software and control system equipment response are assessed through simulated flight tests. Usually, a 0.5g open-loop simulated flight test is used, which cannot truly simulate the flight environment. However, flight tests are expensive and risky, and cannot be used as a conventional launch vehicle testing method. Therefore, there is an urgent need for a method that can test launch vehicles more efficiently and safely. In addition, due to the high dynamics of launch vehicles, there are high requirements for system stability and real-time response speed. Summary of the invention

[0003] The technical problem to be solved by the present invention is to overcome the deficiencies of the prior art and solve the problem of simulating a closed-loop flight test of a launch vehicle on the ground.

[0004] The purpose of the present invention is achieved through the following technical solutions:

[0005] A closed-loop test method for a launch vehicle based on a real-time operating system comprises the following steps:

[0006] Step 1: Start the closed-loop test equipment and run the closed-loop test software in the real-time operating system;

[0007] Step 2: External command controls the closed-loop test software to read the binding parameters and start the dynamic model calculation;

[0008] Step 3: The arrow machine message monitored by the closed-loop test software MT (bus monitor) is used as input to the dynamic model to calculate the output position, speed, attitude, angular velocity, and acceleration information;

[0009] Step 4: The IMU model and GNSS model calculate the position and speed of the gyro, adder pulse and GNSS according to the results of the dynamic model;

[0010] Step 5: The gyro, pulse excitation signal, GNSS position and speed are sent to the IMU device and the rocket machine through the RS422 serial port;

[0011] Step 6: After receiving the excitation signal, the inertial group superimposes its own sensitive error and outputs it to the arrow machine.

[0012] In one embodiment of the present invention, in step 1:

[0013] 1) The closed-loop test equipment includes RS422 serial port peripherals, LAN port and 1553B bus.

[0014] 2) The real-time operating system has a latency of less than 1us and supports multi-task creation and memory management.

[0015] 3) The closed-loop test equipment runs the closed-loop test software, and the software initializes and starts the 1ms task, 10ms task, network command message receiving and sending task, and network telemetry message sending task.

[0016] In one embodiment of the present invention, in step 2:

[0017] 1) External commands are sent to the closed-loop test software via the LAN port to control the closed-loop test software to make corresponding actions and the test process of the entire control system.

[0018] 2) The binding parameters are in text format and are located in the main control microcomputer that sends instructions on the ground. The closed-loop software reads the parameters via FTP.

[0019] In one embodiment of the present invention, in step 3:

[0020] 1) The arrow machine 1553B bus works in BC (bus controller) mode, the closed-loop test software 1553B bus works in MT mode, and MT obtains the servo-rudder feedback message transmitted on the bus.

[0021] 2) The message obtained by MT is input into the dynamic model for solution, and the position, velocity, attitude, angular velocity and acceleration information can be output as follows:

[0022] enter:

[0023] a) Engine structural characteristics binding: Engine installation quadrant angle ξ i , engine installation angle δ0, servo mechanism installation angle ζ i , Distance from engine swing point to theoretical tip x Pr , Distance from engine swing point to rocket body longitudinal axis

[0024] b) Servo system update: engine swing angle δ at each level ia and δ ib

[0025] c) Engine system update: thrust P of each engine i

[0026] d) Engine structural characteristics binding (structural deviation): engine thrust line lateral displacement Lateral direction of engine thrust line (The positive direction is the rotation around the ox1 axis, with the system oy1 axis as the starting point)

[0027] Output:

[0028] Thrust vector P b

[0029] Note: i = 8

[0030]

[0031] Solve the differential equation to obtain the apparent acceleration and angular velocity:

[0032] enter:

[0033] a) Mass characteristics update: total mass m, three components of inertia I z ,I y ,I x

[0034] b) Force model update: Aerodynamic force R b , aerodynamic moment M R , aerodynamic damping force F db , aerodynamic damping torque M d Thrust P b , nozzle force F gb , nozzle torque M g , control torque M c , inertial force F Ib , moment of inertia M I

[0035] Output:

[0036]

[0037]

[0038] The apparent acceleration (The three components in the emission system are ) and angular velocity (The three components in the emission system are ) is integrated to obtain the launch system position, velocity and attitude, and the dynamic model outputs this information.

[0039] In one embodiment of the present invention, in step 4:

[0040] 1) The launch system position and velocity output by the dynamic model are input to the GNSS model to calculate the GNSS position and velocity in the 84 coordinate system, as follows:

[0041] enter:

[0042] Position under the launch system The velocity of the launch system

[0043] Output:

[0044] Receiver output location Receiver output speed

[0045] The GNSS receiver mainly outputs time, velocity, position and 1PPS pulse, and the error model uses Gaussian white noise with zero mean. The time when the GNSS receiver sends the signal is recorded as t PPS The input of this model is the output of the kinetic model. The output is

[0046]

[0047] Where:

[0048] Current shooting position;

[0049] The current shooting speed;

[0050] The position of the transmitting system after adding noise, in m;

[0051] The velocity of the emission system after adding noise, in m / s.

[0052] dp: the vector of the rocket body's center of mass pointing to the differential GNSS receiver antenna, the coordinate in the rocket body coordinate system, the binding value, GNSSA and GNSSB are different;

[0053] The receiver output is converted from the transmitting coordinate system to the 84 coordinate system using the following formula:

[0054]

[0055] The position of the transmitting system after adding noise, in m;

[0056] The velocity of the transmitting system after adding noise, in m / s;

[0057] R0: The launch point's position in the geocentric launch coordinate system before takeoff, and initialization binding.

[0058] 84 The transformation matrix from the coordinate system to the emission coordinate system, initializes the binding.

[0059] The 84-series position output by the receiver, in meters;

[0060] 84 System velocity output by the receiver, in m / s;

[0061] δr GNSS : standard deviation of receiver position error, unit: m;

[0062] δV GNSS : Standard deviation of receiver velocity error, unit: m / s;

[0063] GussRand: Gaussian white noise with variance 1 and mean 0;

[0064] A0: Transmitting azimuth;

[0065] B0: latitude of the launch point;

[0066] λ0: longitude of the launch point.

[0067] 2) The angular velocity and acceleration output by the dynamic model are input to the inertial group model to calculate the pulse increment of the gyroscope and the adder. The total pulse amount is accumulated for each increment, as follows:

[0068] The input of the fiber-optic inertial group model is the apparent acceleration and angular velocity of the arrow system, based on which the apparent velocity increment and angular increment of the arrow system are calculated, and the output is a pulse digital quantity.

[0069] a) Incremental calculation

[0070] According to the input arrow system apparent acceleration and angular velocity, calculate the arrow system apparent velocity increment and angular increment.

[0071]

[0072] ΔW b(n) : Apparent acceleration increment in the current arrow shooting system, unit: m / s 2 ;

[0073] Δθ b(n) : The current angle increment of the arrow, in rad;

[0074] ω (n) : Current angular velocity of the arrow body, Unit: rad / s;

[0075] The apparent acceleration under the current arrow shooting system is Unit: m / s 2

[0076] T SAM ——Inertial group sampling period, constant value binding;

[0077] b) Pulse calculation

[0078] Gyro Pulse:

[0079]

[0080] Adding table pulse:

[0081]

[0082] After the above steps are completed, and Round towards 0.

[0083] Where:

[0084] T SAM ——Inertial group sampling period, constant value binding;

[0085] Δθ b ——Rocket body angle increment, 5×1 matrix;

[0086] ΔN gyro ——Fiber optic inertial group gyro pulse increment, 5×1 matrix, the third and fourth rows are filled with 0;

[0087] ——Fiber optic inertial unit nth gyro pulse output, 5×1 matrix, initial value [0,0,0,0,0] T ;

[0088] ΔW b ——Increment of the apparent velocity of the arrow system, 3×1 matrix;

[0089] ΔN acce ——Fiber optic inertial group pulse increment, 5×1 matrix, fill the 3rd and 4th rows with 0;

[0090] ——The output of the nth beat of the fiber optic inertial group, 5×1 matrix, initial value is [0,0,0,0,0] T ;

[0091] δK: scale factor repeatability;

[0092] δE: gyro and table installation coefficient stability;

[0093] δE GX : Inertia group installation coefficient stability;

[0094] Gyro and table zero bias repeatability;

[0095] E gyro : gyro installation coefficient;

[0096] Eacce : Add meter installation coefficient;

[0097] K gyro : gyro scale factor;

[0098] K acce : Add the scale factor;

[0099] δE gyro : Gyro installation coefficient deviation;

[0100] δE acce : Add the installation coefficient deviation;

[0101] δK gyro : gyro scale factor deviation;

[0102] δK acce : Add the scale factor deviation.

[0103] In one embodiment of the present invention, in step 5:

[0104] 1) The full pulse excitation signal of the gyro and the meter is sent through the RS422 serial port every 10ms according to the output data frequency of the inertial group, and the inertial group receives the excitation signal.

[0105] 2) In the 10ms task of the closed-loop test software, the RS422 serial port is used to receive the actual value of the GNSS test site, subtract the local theoretical position and speed values ​​to obtain the GNSS error, and superimpose the 84 coordinate system position and speed calculated by the GNSS model as the output, which is transmitted to the arrow machine through the RS422 serial port to form a closed loop.

[0106] In one embodiment of the present invention, in step 6: after the inertial group receives the full pulse excitation signal of the gyro and the meter, the sensor error is superimposed to obtain the full pulse data of the inertial group with error, and the inertial group pulse data is transmitted to the 1553B bus for use by the rocket flight software, forming a closed loop.

[0107] Compared with the prior art, the present invention has the following beneficial effects:

[0108] (1) The closed-loop test method for a launch vehicle based on a real-time operating system of the present invention adopts an embedded real-time operating system with good stability, thereby ensuring the real-time and synchronization of the launch vehicle simulation flight test.

[0109] (2) The closed-loop test method for a launch vehicle based on a real-time operating system of the present invention can simulate the actual output values ​​of the inertial group equipment and the GNSS, making the simulated flight test of the launch vehicle more credible and improving the overall detection capability of the launch vehicle. BRIEF DESCRIPTION OF THE DRAWINGS

[0110] Figure 1It is a schematic diagram of the layout of the closed-loop test method of the present invention. DETAILED DESCRIPTION

[0111] To make the objectives, technical solutions and advantages of the present invention more clear, the embodiments of the present invention will be further described in detail below with reference to the accompanying drawings.

[0112] The layout of the closed-loop test is as follows Figure 1 As shown, a closed-loop test method for a launch vehicle based on a real-time operating system includes:

[0113] Step 1: Start the closed-loop test equipment and run the closed-loop test software in the real-time operating system.

[0114] 1) The closed-loop test equipment includes RS422 serial port peripherals, LAN port and 1553B bus.

[0115] 2) The real-time operating system has a latency of less than 1us and supports multi-task creation and memory management.

[0116] 3) The closed-loop test equipment runs the closed-loop test software, and the software initializes and starts the 1ms task, 10ms task, network command message receiving and sending task, and network telemetry message sending task.

[0117] For example, the closed-loop test equipment used has a total of 8 serial ports, and the baud rates of the inertial group RS422 serial port and the GNSS RS422 serial port are both 115200; the closed-loop test equipment selects the VxWorks real-time operating system, and the closed-loop test software runs in the VxWorks real-time operating system, starting the 1ms task, 10ms task, network command message receiving and sending task, and network telemetry message sending task. The usage of the 1ms task and the 10ms task is as follows:

[0118] a) The programs running in the 1ms task include MT bus monitor receiving messages and decoding processing, dynamic model calculation, inertial group model calculation, GNSS model calculation, and inertial group data RS422 serial port sending data;

[0119] b) The program running in the 10ms task includes GNSS RS422 serial port receiving and sending processing.

[0120] Step 2: External command controls the closed-loop test software to read the binding parameters and start the dynamic model calculation.

[0121] 1) External commands are sent to the closed-loop test software via the LAN port to control the closed-loop test software to make corresponding actions and the test process of the entire control system.

[0122] 2) The binding parameters are in text format and are located in the main control microcomputer that sends instructions on the ground. The closed-loop software reads the parameters via FTP.

[0123] For example: the launch vehicle main control microcomputer test software is connected to the closed-loop test software as a TCP client, and then the main control microcomputer test software sends network instructions to control the closed-loop test software to make corresponding actions and the test process of the entire control system; because real-time model calculation is required in the closed-loop test, and the model calculation depends on flight trajectory parameters and other hardware equipment such as the zero bias and installation error of the inertial group, these data need to be changed because the trajectory of each launch vehicle is different and the basic parameters of the inertial group equipment are also different. Therefore, in order to facilitate the use of text format for parameter binding, files are uploaded through FTP.

[0124] Step 3: The closed-loop test software MT (bus monitor) monitors the arrow machine message as input to the dynamic model to calculate the output position, velocity, attitude, angular velocity, and acceleration information.

[0125] 1) The arrow machine 1553B bus works in BC (bus controller) mode, the closed-loop test software 1553B bus works in MT mode, and MT obtains the servo-rudder feedback message transmitted on the bus.

[0126] 2) The message obtained by MT is input into the dynamic model for solution, and the position, velocity, attitude, angular velocity and acceleration information can be output as follows:

[0127] enter:

[0128] a) Engine structural characteristics binding: Engine installation quadrant angle ξ i , engine installation angle δ0, servo mechanism installation angle ζ i , Distance from engine swing point to theoretical tip x Pr , the distance from the engine swing point to the longitudinal axis of the rocket body l Pi

[0129] b) Servo system update: engine swing angle δ at each level ia and δ ib

[0130] c) Engine system update: thrust P of each engine i

[0131] d) Engine structural characteristics binding (structural deviation): engine thrust line lateral displacement Lateral direction of engine thrust line (The positive direction is the rotation around the ox1 axis, with the system oy1 axis as the starting point)

[0132] Output:

[0133] Thrust vector P b

[0134] Note: i = 8

[0135]

[0136] Solve the differential equation to obtain the apparent acceleration and angular velocity:

[0137] enter:

[0138] a) Mass characteristics update: total mass m, three components of inertia I z ,I y ,I x

[0139] b) Force model update: Aerodynamic force R b , aerodynamic moment M R , aerodynamic damping force F db , aerodynamic damping torque M d Thrust P b , nozzle force F gb , nozzle torque M g , control torque M c , inertial force F Ib , moment of inertia M I

[0140] Output:

[0141]

[0142]

[0143] The apparent acceleration (The three components in the emission system are ) and angular velocity (The three components in the emission system are ) is integrated to obtain the launch system position, velocity and attitude, and the dynamic model outputs this information.

[0144] For example, the flight software running in the arrow machine has its 1553 bus working in BC mode. When the BC scheduler obtains the servo telemetry message, the closed-loop test software MT can monitor this message, and then parse the rudder feedback value from the message, use the rudder feedback value as the dynamic model input, and calculate the engine thrust vector P b , then the combined aerodynamic force R b , aerodynamic moment M R , aerodynamic damping force F db , aerodynamic damping torque M d , nozzle force F gb , nozzle torque M g , control torque M c , inertial force F Ib , moment of inertia MI Calculate the apparent acceleration and angular velocity Finally, integrate the apparent acceleration to get the velocity, integrate again to get the position, integrate the angular velocity to get the attitude angle, and complete the calculation of the dynamic model. Use the position, velocity, attitude, apparent acceleration and angular velocity as the output of the dynamic model for the next step.

[0145] Step 4: The IMU model and GNSS model calculate the position and velocity of the gyro, adder pulse and GNSS according to the results of the dynamic model.

[0146] 1) The launch system position and velocity output by the dynamic model are input to the GNSS model to calculate the GNSS position and velocity in the 84 coordinate system, as follows:

[0147] enter:

[0148] Position under the launch system The velocity of the launch system

[0149] Output:

[0150] Receiver output location Receiver output speed

[0151] The GNSS receiver mainly outputs time, velocity, position and 1PPS pulse, and the error model uses Gaussian white noise with zero mean. The time when the GNSS receiver sends the signal is recorded as t PPS The input of this model is the output of the kinetic model. The output is

[0152]

[0153] Where:

[0154] Current shooting position;

[0155] The current shooting speed;

[0156] The position of the transmitting system after adding noise, in m;

[0157] The velocity of the emission system after adding noise, in m / s.

[0158] dp: the vector of the rocket body's center of mass pointing to the differential GNSS receiver antenna, the coordinate in the rocket body coordinate system, the binding value, GNSSA and GNSSB are different;

[0159] The receiver output is converted from the transmitting coordinate system to the 84 coordinate system using the following formula:

[0160]

[0161] The position of the transmitting system after adding noise, in m;

[0162] The velocity of the transmitting system after adding noise, in m / s;

[0163] R0: The launch point's position in the geocentric launch coordinate system before takeoff, and initialization binding.

[0164] 84 The transformation matrix from the coordinate system to the emission coordinate system, initializes the binding.

[0165] The 84-series position output by the receiver, in meters;

[0166] 84 System velocity output by the receiver, in m / s;

[0167] δr GNSS : standard deviation of receiver position error, unit: m;

[0168] δV GNSS : Standard deviation of receiver velocity error, unit: m / s;

[0169] GussRand: Gaussian white noise with variance 1 and mean 0;

[0170] A0: Transmitting azimuth;

[0171] B0: latitude of the launch point;

[0172] λ0: longitude of the launch point.

[0173] 2) The angular velocity and acceleration output by the dynamic model are input to the inertial group model to calculate the pulse increment of the gyroscope and the adder. The total pulse amount is accumulated for each increment, as follows:

[0174] The input of the fiber-optic inertial group model is the apparent acceleration and angular velocity of the arrow system, based on which the apparent velocity increment and angular increment of the arrow system are calculated, and the output is a pulse digital quantity.

[0175] a) Incremental calculation

[0176] According to the input arrow system apparent acceleration and angular velocity, calculate the arrow system apparent velocity increment and angular increment.

[0177]

[0178] ΔW b(n) : Apparent acceleration increment in the current arrow shooting system, unit: m / s 2 ;

[0179] Δθ b(n) : The current angle increment of the arrow, in rad;

[0180] ω (n) : Current angular velocity of the arrow body, Unit: rad / s;

[0181] The apparent acceleration under the current arrow shooting system is Unit: m / s 2

[0182] T SAM ——Inertial group sampling period, constant value binding;

[0183] b) Pulse calculation

[0184] Gyro Pulse:

[0185]

[0186] Adding table pulse:

[0187]

[0188] After the above calculation steps are completed, and Round towards 0.

[0189] Where:

[0190] T SAM ——Inertial group sampling period, constant value binding;

[0191] Δθ b ——Rocket body angle increment, 5×1 matrix;

[0192] ΔN gyro ——Fiber optic inertial group gyro pulse increment, 5×1 matrix, the third and fourth rows are filled with 0;

[0193] ——Fiber optic inertial unit nth gyro pulse output, 5×1 matrix, initial value [0,0,0,0,0] T ; ΔW b ——Increment of the apparent velocity of the arrow system, 3×1 matrix;

[0194] ΔN acce ——Fiber optic inertial group pulse increment, 5×1 matrix, fill the 3rd and 4th rows with 0;

[0195] ——The output of the nth beat of the fiber optic inertial group, 5×1 matrix, initial value is [0,0,0,0,0] T ; δK: scale factor repeatability;

[0196] δE: gyro and table installation coefficient stability;

[0197] δE GX : Inertia group installation coefficient stability;

[0198] Gyro and table zero bias repeatability;

[0199] E gyro : gyro installation coefficient;

[0200] E acce : Add meter installation coefficient;

[0201] K gyro : gyro scale factor;

[0202] K acce : Add the scale factor;

[0203] δE gyro : Gyro installation coefficient deviation;

[0204] δE acce : Add the installation coefficient deviation;

[0205] δK gyro : gyro scale factor deviation;

[0206] δK acce : Add the scale factor deviation.

[0207] For example, the transmitting system position and velocity output after the dynamic model calculation are used as the input of the GNSS model calculation. The position is calculated through the receiver antenna vector compensation and coordinate system matrix conversion to obtain the 84 coordinate system position. The velocity is calculated through the coordinate system matrix conversion to obtain the velocity in the 84 coordinate system, thus completing the GNSS model calculation. The apparent acceleration and angular velocity output after the dynamic model calculation are used as the input of the inertial group model. The apparent velocity increment and angular increment are calculated in conjunction with the inertial group sampling period. Then, the pulse increment is calculated based on the binding data of the inertial group (scale factor repeatability, gyro, plus meter installation coefficient stability, inertial group installation coefficient stability and gyro, plus meter zero bias repeatability). Finally, the calculation results of each beat are accumulated to obtain the inertial group model output - the full amount of gyro plus meter pulse.

[0208] Step 5: The gyro, pulse excitation signal, GNSS position and speed are sent to the IMU device and the rocket machine through the RS422 serial port.

[0209] 1) The full pulse excitation signal of the gyro and the meter is sent through the RS422 serial port every 10ms according to the output data frequency of the inertial group, and the inertial group receives the excitation signal.

[0210] 2) In the 10ms task of the closed-loop test software, the RS422 serial port is used to receive the actual value of the GNSS test site, subtract the local theoretical position and speed values ​​to obtain the GNSS error, and superimpose the 84 coordinate system position and speed calculated by the GNSS model as the output, which is transmitted to the arrow machine through the RS422 serial port to form a closed loop.

[0211] For example, in the closed-loop test software, the dynamic model, IMU model and GNSS model calculations are all performed in the 1ms task. Therefore, the IMU pulse and the 84-frame position and velocity of the GNSS are updated every 1ms and counted in the 1ms task. When the calculation is accumulated ten times, the IMU pulse data is organized into a byte data frame, which is an excitation signal and is sent to the IMU through the RS422 serial port of the IMU. The GNSS receiver data is updated every 1s. Therefore, in order to slow down the access to the RS422 register, the received data is placed in the 10ms task. When the GNSS data is received and parsed successfully, the local theoretical position and velocity are deducted, and then the 84-frame position and velocity calculated by the GNSS model are superimposed to obtain the position and velocity values ​​of the actual receiver error. Finally, the results are organized into a byte data frame and sent to the rocket through the RS422 serial port, and the GNSS data forms a closed loop.

[0212] Step 6: After receiving the excitation signal, the inertial group superimposes its own sensitive error and outputs it to the arrow machine.

[0213] After the inertial group receives the full pulse excitation signal from the gyro and the meter, it superimposes the sensor error to obtain the full pulse data of the inertial group with error, and transmits the inertial group pulse data to the 1553B bus for use by the rocket flight software, forming a closed loop.

[0214] For example, the total pulse amount calculated by the inertial group model is the theoretical value of the launch vehicle's flight trajectory. This data is sent to the inertial group as an excitation. After receiving it, the inertial group superimposes the static pulse result to obtain the total pulse amount with the real inertial group error. The hardware connection output of the inertial group uses the 1553 bus. When the BC of the rocket flight software schedules the inertial group pulse data in a 10ms period, the inertial group sends the data to the rocket as an RT (remote terminal) device, so the inertial group pulse full amount data also forms a closed loop.

[0215] Compared with software running directly on a chip and running in a real-time operating system, the present invention has stronger real-time performance, faster multi-task switching, can greatly reduce the difficulty of software development, and is more convenient and efficient in memory management. The closed-loop test method provided by the present invention obtains the servo feedback 1553 bus message as the input of the model for calculation, obtains the inertial group and GNSS data required by the flight software, and then provides the data to the flight software to form a closed-loop circuit, which is more realistic than the open-loop simulated flight test.

[0216] In the field of launch vehicle testing, the present invention has the characteristics of high testing accuracy, high reliability and strong feasibility, and has high engineering application value.

[0217] The contents not described in detail in the specification of the present invention belong to the common knowledge of those skilled in the art.

[0218] Although the present invention has been disclosed as above in the form of a preferred embodiment, it is not intended to limit the present invention. Any person skilled in the art may make possible changes and modifications to the technical solution of the present invention by using the methods and technical contents disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and modifications made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solution of the present invention shall fall within the protection scope of the technical solution of the present invention.

Claims

1. A closed-loop test method for a launch vehicle based on a real-time operating system, characterized in that: include: Step 1: Start the closed-loop test equipment and run the closed-loop test software in the real-time operating system; Step 2: External command controls the closed-loop test software to read the binding parameters and start the dynamic model calculation; Step 3: The closed-loop test software MT monitors the arrow machine message as input to the dynamic model to calculate the output position, speed, attitude, angular velocity, and acceleration information; Step 4: The IMU model and GNSS model calculate the position and speed of the gyro, adder pulse and GNSS according to the results of the dynamic model; Step 5: The gyro, meter pulse, GNSS position and speed are sent to the IMU device and the rocket machine through the RS422 serial port; Step 6: After receiving the excitation signal, the inertial group superimposes its own sensitive error and outputs it to the arrow machine.

2. The launch vehicle closed-loop test method according to claim 1, characterized in that: In step 1, the closed-loop test equipment includes RS422 serial port peripherals, LAN port and 1553B bus; the real-time operating system has a delay time of less than 1us and supports multi-task creation and memory management; the closed-loop test equipment runs the closed-loop test software, and the software initializes and starts 1ms tasks, 10ms tasks, network command message receiving and sending tasks, and network telemetry message sending tasks.

3. The launch vehicle closed-loop test method according to claim 2, characterized in that: The programs running in the 1ms task include MT bus monitor receiving messages and decoding processing, dynamic model calculation, inertial group model calculation, GNSS model calculation, and inertial group data RS422 serial port sending data; The program running in the 10ms task includes GNSS RS422 serial port receiving and sending processing.

4. The launch vehicle closed-loop test method according to claim 1, characterized in that: In step 2, the external command is sent to the closed-loop test software via the LAN port, which is used to control the closed-loop test software to make corresponding actions and the test process of the entire control system; The binding parameters are in text format and are located in the main control microcomputer that sends instructions on the ground. The closed-loop test software reads the parameters via FTP.

5. The launch vehicle closed-loop test method according to claim 1, characterized in that: In step 3, the arrow machine 1553B bus works in BC mode, and the closed-loop test software 1553B bus works in MT mode. MT obtains the servo rudder feedback message transmitted on the bus; the message obtained by MT is input into the dynamic model for solution, and can output position, speed, attitude, angular velocity, and acceleration information.

6. The launch vehicle closed-loop test method according to claim 1, characterized in that: In step 4, the transmitting system position and velocity output by the dynamic model are input to the GNSS model to calculate the GNSS position and velocity in the 84 coordinate system. The angular velocity and acceleration output by the dynamic model are input to the inertial group model to calculate the pulse increment of the gyroscope and the adding table. The increment of each beat is accumulated to obtain the total pulse amount.

7. The launch vehicle closed-loop test method according to claim 6, characterized in that: The input of the fiber-optic inertial group model is the apparent acceleration and angular velocity of the arrow system, based on which the apparent velocity increment and angular increment of the arrow system are calculated, and the output is a pulse digital quantity.

8. The launch vehicle closed-loop test method according to claim 1, characterized in that: In step 5, the full-scale pulse excitation signal of the gyro and the meter is sent through the RS422 serial port every 10 ms according to the output data frequency of the inertial group, and the inertial group receives the excitation signal; in the 10 ms task of the closed-loop test software, the RS422 serial port is used to receive the actual value of the GNSS test site, subtract the local theoretical position and speed values ​​to obtain the GNSS error, and superimpose the 84 coordinate system position and speed calculated by the GNSS model as the output, which is transmitted to the rocket through the RS422 serial port to form a closed loop.

9. The launch vehicle closed-loop test method according to claim 1, characterized in that: In step 6, after the inertial group receives the full pulse excitation signal from the gyro and the meter, it superimposes the sensor error to obtain the full pulse data of the inertial group with error, and transmits the inertial group pulse data to the 1553B bus for use by the rocket flight software, forming a closed loop.

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