Sensor arrangement method based on flight mode identification requirement

Abstract

The invention discloses a sensor arrangement method based on flight mode identification requirements, which comprises the following steps: according to flight mode parameter identification requirements, analyzing various information of required sensors and types of parameters to be identified, setting the sensors and corresponding mounting positions according to the analyzed content, and aiming at a first-order beam mode of a non-boosted solid carrier rocket, determining a first-order beam mode of the non-boosted solid carrier rocket and a second-order beam mode of the non-boosted solid carrier rocket. The purpose of designing sensor arrangement according to identification requirements is achieved through an on-rocket sensor arrangement scheme, and the method is used for providing accurate and high-quality flight telemetry data.

Description

Technical Field

[0001] The invention relates to a sensor arrangement method based on flight mode identification requirements, and belongs to the technical field of elastic parameter design of solid launch vehicles. Background Art

[0002] In the process of elastic parameter design of solid launch vehicle, ground modal test combined with finite element simulation prediction is usually adopted. However, on the one hand, ground modal test will introduce the influence of additional tooling, and on the other hand, ground modal test conditions cannot simulate the influence of factors such as force and thermal environment on the elastic characteristics of the rocket during the actual flight of the launch vehicle. In addition, for ground modal test of solid launch vehicle, only modal tests of the engine full charge and empty shell can be carried out, and modal test of the intermediate state cannot be carried out to provide test data reference for elastic design. Therefore, in the elastic design of the intermediate state, only the modified finite element model can be used for prediction. It is worth noting that ground modal test is often only carried out in the development stage. After the payload of the launch vehicle changes, the modal parameters are given only by means of simulation analysis.

[0003] Using flight telemetry data to obtain rocket elastic parameters is an important means to make up for the shortcomings of ground tests. It can provide direction for studying the differences between modal characteristics and the earth, the accuracy of ground tests, and the reliability of simulation predictions, and provide theoretical support for improving performance indicators for subsequent launch missions of the same model. Accurate and high-quality flight telemetry data is the basis for obtaining flight modal parameters, and the sensor layout scheme on the rocket has an important impact on the signal quality. Therefore, in order to obtain high-quality flight modal identification data input, the sensor layout scheme must be designed according to the identification requirements. The existing sensor layout scheme is mainly designed based on the flight environment analysis requirements, and does not consider the special requirements of flight modal parameter identification for sensor accuracy, frequency, and position. At the same time, there is a lack of simulation verification of the sensor layout scheme. There are defects such as poor data quality and incomplete information when telemetry parameters are used for flight modal identification, which are difficult to overcome. Summary of the invention

[0004] The technical problem solved by the present invention is: in view of the defects of the existing technology in the telemetry parameters in the flight mode identification, such as poor data quality and incomplete information, which leads to the difficulty in the flight mode parameter identification and poor result credibility, a sensor arrangement method based on the flight mode identification needs is proposed.

[0005] The present invention solves the above technical problems by the following technical solutions:

[0006] A sensor arrangement method based on flight mode identification requirements includes:

[0007] Conduct demand analysis based on the flight mode identification task requirements and determine the modal parameters and types to be identified;

[0008] Determine the sensor type and sensor combination required for modal identification according to the modal parameters and types to be identified;

[0009] Conduct ground modal tests on the sensor combination to determine the installation position of each sensor on the arrow;

[0010] Design the measurement range and analysis frequency range of each sensor in the sensor combination;

[0011] After installing each sensor at the designated location, the sensor layout plan is tested. If the test results meet the requirements of the flight mode identification task, the sensor layout is completed; otherwise, the flight mode identification task requirement analysis is performed again.

[0012] The modal parameters and types to be identified include frequency identification only, frequency and damping identification, frequency and vibration mode identification, frequency and vibration mode and damping identification, and the maximum frequency to be identified; the sensor types include acceleration sensors, overload sensors, low-frequency vibration sensors, and rate gyroscopes; according to the determined modal parameters and types to be identified, a sensor combination that meets the requirements of the flight modal identification task and existing measuring elements on the arrow are selected from the sensor types to jointly realize the identification of the modal parameters.

[0013] The corresponding relationship between the installation position of each sensor on the arrow and the ground modal test is:

[0014] Sensors are classified according to the ground modal test, finite element simulation predicted vibration mode and vibration mode slope results. When the sensor type is an acceleration sensor, overload sensor, or low-frequency vibration sensor, it is installed at the peak or trough of the vibration mode predicted by the finite element simulation, or at both ends of the rocket; when the sensor type is a rate gyroscope, it is installed at the vibration mode node of the vibration mode predicted by the finite element simulation.

[0015] The range design method of each sensor is as follows:

[0016] Conduct trajectory simulation for each selected sensor, and design the corresponding range of each sensor based on the acceleration parameters, angular velocity parameters, and upper and lower bounds of the angular acceleration parameters obtained by analyzing the simulation results;

[0017] The analysis frequency range of each sensor is determined according to the lowest order frequency fmin and the maximum frequency fmax to be identified obtained in the ground modal test or finite element simulation.

[0018] The method for determining the maximum frequency fmax to be identified is:

[0019] According to the analysis results of flight mode identification requirements, the minimum value of the maximum frequency fmax is set to be no less than the first two modal frequencies of the entire missile in the analysis results to avoid the first two modes of the entire missile being excited by aerodynamic interference forces.

[0020] The sensor type and the sensor combination required for modal identification are determined according to the modal parameters and types to be identified. The corresponding relationship between the sensor type and the sensor combination required for modal identification and the modal parameters and types to be identified is:

[0021] When the modal parameters and types to be identified are frequency identification only or frequency and damping identification, the sensors are the existing measuring elements and rate gyroscopes on the arrow; if the current sensor type and the range of the sensor combination required for modal identification meet the requirements of the flight modal identification task after design, the current sensor combination is maintained, otherwise other sensors are added;

[0022] When the modal parameters and types to be identified are frequency and vibration mode identification, frequency and vibration mode and damping identification, the sensors use the existing measuring elements and rate gyroscopes on the arrow, and new acceleration sensors and rate gyroscopes specially used for flight mode identification are added.

[0023] The installation positions of the acceleration sensor, overload sensor, and low-frequency vibration sensor are selected at the peaks and troughs of the vibration modes predicted by finite element simulation or at the two ends of the rocket where the vibration modes meet the requirements; the installation position of the velocity gyroscope is selected at the nodes of each order vibration mode; when all types of sensors are installed, the current installation position is used as the preliminary design plan for the sensor installation position.

[0024] The sensor analysis frequency range is set to the lowest order frequency fmin to the maximum frequency to be identified fmax. Each sensor performs ground modal tests or finite element simulations to obtain the first two order frequencies of the entire arrow, which are f1 and f2 respectively. The lowest order frequency fmin is set less than 0.5f1, and the maximum frequency fmax to be identified is set greater than f2.

[0025] The method for testing the sensor arrangement scheme is as follows:

[0026] According to the ground modal test results, a model is established and dynamic response simulation is carried out to obtain the response values ​​at each measuring point of the whole arrow. According to the obtained response value F(t), judgment is made:

[0027]

[0028] In the formula, M, C, and K are the model stiffness matrix, damping matrix, and stiffness matrix respectively. x is the acceleration, velocity, and displacement response, respectively, and F(t) is the external excitation; when the response value meets the requirements of the flight mode identification task, the current sensor installation position is reasonable, and the sensor layout plan inspection is completed after traversing all sensors.

[0029] Before checking the sensor layout, determine the impossible installation positions and delete them according to the installation restrictions of the rocket body structure, and use the remaining positions to check the sensor layout by measuring the response values ​​of the corresponding measurement points;

[0030] If the corresponding measurement point response values ​​of the remaining positions all meet the requirements of the flight mode identification task, the sensor layout plan is tested; if the corresponding measurement point response values ​​of the installation positions do not meet the requirements of the flight mode identification task, the flight mode identification task requirement analysis is performed again;

[0031] Among them, if the modal parameters and types to be identified cannot be identified in each sensor installation position, it is deemed that the flight mode identification task requirements are not met.

[0032] The advantages of the present invention compared with the prior art are:

[0033] The present invention provides accurate and high-quality flight telemetry data for flight mode identification through a sensor arrangement method that meets the requirements of flight mode identification, and provides a data support basis for studying the sky-ground differences in modal characteristics, the accuracy of ground tests, and the reliability of simulation predictions. At the same time, it can make full use of existing measurement components on the arrow, reduce the number of sensors required for flight mode identification, and take into account economic requirements. The sensor arrangement scheme that has been arranged is verified by simulation in advance, and the scheme has high feasibility and strong practicality. BRIEF DESCRIPTION OF THE DRAWINGS

[0034] Figure 1 A schematic diagram of the sensor arrangement process based on the flight mode identification requirements provided by the present invention;

[0035] Figure 2 The ground test or finite element simulation provided by the present invention obtains the schematic diagram of the first two vibration modes of the whole arrow. DETAILED DESCRIPTION

[0036] A sensor arrangement method based on flight mode identification requirements analyzes various types of sensor information and parameter types to be identified according to the flight mode parameter identification requirements, sets sensors and corresponding installation positions according to the analyzed contents, and realizes the purpose of designing sensor arrangement according to identification requirements through the on-rocket sensor arrangement scheme for the first-order beam mode of an unboosted solid launch vehicle, so as to provide accurate and high-quality flight telemetry data.

[0037] The sensor layout scheme based on the flight mode identification requirements provides a specific layout, analysis, and design process from the aspects of sensor selection, range design, and location selection;

[0038] According to the requirements of flight modal parameter identification, the required sensor information is analyzed. For the first-order beam mode of unboosted solid launch vehicle, there are two nodes that require at least three sensor positions to identify the vibration mode, and three nodes that identify the second-order vibration mode require at least four sensor installation positions to identify the vibration mode. The first three-order elastic frequency parameters of solid launch vehicle are usually within 50Hz, and its first-order frequency is sometimes even around 5Hz. Therefore, the effective frequency range of the sensor must include 0~50Hz. Usually, overload sensors can be used to measure the vibration acceleration data of the flight projectile, which can meet the requirements of flight modal identification. In special cases, inertial groups, rate gyroscopes, etc. can also be used;

[0039] The actual locations available for installing sensors in solid-propellant launch vehicles are relatively limited, mainly concentrated in the fairing, satellites, interstages, and tail sections of each stage. When installing sensors, it is necessary to select the peaks and troughs of the modal vibration type of interest and avoid the vibration type nodes. In addition, the sensor installation location should avoid points that are more affected by flight noise, equipment brackets, and bulkheads that are easily affected by local modes, and select locations where the bulkhead is reinforced and close to the engine column section.

[0040] During the sensor range design process, the dynamic response simulation of the missile body is carried out according to the ground wind tunnel pressure pulsation test and aerodynamic numerical simulation, and the response magnitude at the sensor installation position can be obtained. Based on this, the appropriate sensor range can be selected; when selecting the sensor range, the amplitude resolution of the sampling data must be guaranteed.

[0041] The following is further described in conjunction with the accompanying drawings and preferred embodiments of the specification:

[0042] In the current embodiment, the sensor arrangement scheme based on the flight mode recognition requirement is as follows: Figure 1 As shown, the specific steps are:

[0043] (1) Conduct flight mode identification requirements analysis and determine the modal parameters to be identified, including: frequency identification only, frequency + damping identification, frequency + vibration mode identification, frequency + vibration mode + damping identification, and the maximum frequency fmax to be identified;

[0044] (2) Select sensor type. According to the flight modal parameter identification requirement analysis results in step (1), determine the type of sensor, which generally includes: acceleration sensor, overload sensor, low-frequency vibration sensor, rate gyroscope, and make full use of the existing measurement components on the arrow, while taking into account the economic requirements;

[0045] (3) Based on the sensor selection results in (2), complete the design (or selection) of the sensor installation location. The sensor location selection is determined based on the vibration mode and vibration mode slope results predicted by the ground modal test and finite element simulation. Acceleration, overload, and low-frequency vibration sensors are installed at the vibration mode peak, trough, or both ends of the rocket. Rate gyro sensors are installed at the vibration mode nodes.

[0046] (4) Based on the sensor results in step (3), the sensor range design is completed. The range design is determined based on the acceleration parameters, angular velocity parameters, and upper and lower limits of the angular acceleration parameters obtained from the trajectory simulation and attitude control target analysis; the sensor analysis frequency range design is determined based on the lowest order frequency fmin obtained from the ground test or finite element simulation and the fmax in step (1);

[0047] (5) The sensor layout plan is tested to confirm whether it meets the requirements in step (1) and whether it meets the installation restrictions of the rocket body structure. If so, the sensor layout plan is established. Otherwise, it jumps back to (2) to design the next round of this sensor layout plan.

[0048] In step (1), the condition for selecting the maximum frequency fmax to be identified is that its minimum value is not less than the first two modal frequencies of the entire missile, because usually the first two orders can be excited by aerodynamic interference forces.

[0049] In step (2), the specific method for selecting the sensor type is as follows:

[0050] (a) For frequency identification only or frequency + damping identification, the sensor can make full use of the onboard inertial group, rate gyro, etc. If the range and other requirements are met, there is no need to add a sensor specifically used for flight mode identification;

[0051] (b) For frequency + mode identification or frequency + mode + damping identification, in addition to making full use of the onboard inertial group and rate gyro, additional acceleration and rate gyro sensors specifically used for flight mode identification should be added;

[0052] In step (3), the specific method for selecting the sensor installation location is as follows:

[0053] According to ground tests or finite element simulation, the first two vibration modes of the whole arrow are obtained. As shown in the figure below, the installation position of the acceleration sensor is selected at the peak, trough or the larger vibration mode at both ends of the rocket of each vibration mode (☆), and the installation position of the rate gyro sensor is selected at the node of each vibration mode (△). The preliminary design of the sensor installation position is completed. Figure 2 As shown;

[0054] In step (4), the specific method for designing the sensor analysis frequency range fmin to fmax is as follows:

[0055] The first two frequencies of the entire arrow are obtained based on ground tests or finite element simulations, which are f1 and f2 respectively. Then fmin should be less than 0.5f1 and fmax should be greater than f2.

[0056] In step (5), the specific method for testing the sensor arrangement scheme is as follows:

[0057] First, the numerical model is established and modified according to the ground test results, and then the dynamic response simulation is carried out according to the following formula to obtain the response at each measuring point:

[0058]

[0059] Among them, M, C, and K are the system stiffness matrix, damping matrix, and stiffness matrix respectively. x is the acceleration, velocity, and displacement response, respectively, and F(t) is the external excitation.

[0060] According to the installation space restriction requirements of the rocket body structure, the installation positions that cannot be realized in step (4) are deleted, and then the modal parameter identification work is carried out using the measurement point responses corresponding to the remaining positions. If the modal parameters to be identified determined in step (1) can be identified and meet the flight modal identification requirements, the sensor layout plan is tested. Otherwise, it fails and the flight modal identification requirements analysis is carried out again.

[0061] 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.

[0062] The contents not described in detail in the specification of the present invention belong to the common knowledge of those skilled in the art.

Claims

1. A sensor arrangement method based on flight mode identification requirements, characterized in that include: Conduct demand analysis based on the flight mode identification task requirements and determine the modal parameters and types to be identified; Determine the sensor type and sensor combination required for modal identification according to the modal parameters and types to be identified; Conduct ground modal tests on the sensor combination to determine the installation position of each sensor on the arrow; Design the measurement range and analysis frequency range of each sensor in the sensor combination; After installing each sensor at the designated location, the sensor layout plan is tested. If the test results meet the requirements of the flight mode identification task, the sensor layout is completed; otherwise, the flight mode identification task requirement analysis is performed again.

2. The sensor arrangement method based on flight mode identification requirements according to claim 1 is characterized in that: The modal parameters and types to be identified include frequency identification only, frequency and damping identification, frequency and vibration mode identification, frequency and vibration mode and damping identification, and the maximum frequency to be identified; the sensor types include acceleration sensors, overload sensors, low-frequency vibration sensors, and rate gyroscopes; according to the determined modal parameters and types to be identified, a sensor combination that meets the requirements of the flight modal identification task and existing measuring elements on the arrow are selected from the sensor types to jointly realize the identification of the modal parameters.

3. The sensor arrangement method based on flight mode identification requirements according to claim 1 is characterized in that: The corresponding relationship between the installation position of each sensor on the arrow and the ground modal test is: Sensors are classified according to the ground modal test, finite element simulation predicted vibration mode and vibration mode slope results. When the sensor type is an acceleration sensor, overload sensor, or low-frequency vibration sensor, it is installed at the peak or trough of the vibration mode predicted by the finite element simulation, or at both ends of the rocket; when the sensor type is a rate gyroscope, it is installed at the vibration mode node of the vibration mode predicted by the finite element simulation.

4. The sensor arrangement method based on flight mode identification requirements according to claim 2 is characterized in that: The range design method of each sensor is as follows: Conduct trajectory simulation for each selected sensor, and design the corresponding range of each sensor based on the acceleration parameters, angular velocity parameters, and upper and lower bounds of the angular acceleration parameters obtained by analyzing the simulation results; The analysis frequency range of each sensor is determined according to the lowest order frequency fmin and the maximum frequency fmax to be identified obtained in the ground modal test or finite element simulation.

5. The sensor arrangement method based on flight mode identification requirements according to claim 4 is characterized in that: The method for determining the maximum frequency fmax to be identified is: According to the analysis results of flight mode identification requirements, the minimum value of the maximum frequency fmax is set to be no less than the first two modal frequencies of the entire missile in the analysis results to avoid the first two modes of the entire missile being excited by aerodynamic interference forces.

6. The sensor arrangement method based on flight mode identification requirements according to claim 2 is characterized in that: The sensor type and the sensor combination required for modal identification are determined according to the modal parameters and types to be identified. The corresponding relationship between the sensor type and the sensor combination required for modal identification and the modal parameters and types to be identified is: When the modal parameters and types to be identified are frequency identification only or frequency and damping identification, the sensors are the existing measuring elements and rate gyroscopes on the arrow; if the current sensor type and the range of the sensor combination required for modal identification meet the requirements of the flight modal identification task after design, the current sensor combination is maintained, otherwise other sensors are added; When the modal parameters and types to be identified are frequency and vibration mode identification, frequency and vibration mode and damping identification, the sensors use the existing measuring elements and rate gyroscopes on the arrow, and new acceleration sensors and rate gyroscopes specially used for flight mode identification are added.

7. The sensor arrangement method based on flight mode identification requirements according to claim 2 is characterized in that: The installation positions of the acceleration sensor, overload sensor, and low-frequency vibration sensor are selected at the peaks and troughs of the vibration modes predicted by finite element simulation or at the two ends of the rocket where the vibration modes meet the requirements; the installation position of the velocity gyroscope is selected at the nodes of each order vibration mode; when all types of sensors are installed, the current installation position is used as the preliminary design plan for the sensor installation position.

8. The sensor arrangement method based on flight mode identification requirements according to claim 7 is characterized in that: The sensor analysis frequency range is set to the lowest order frequency fmin to the maximum frequency to be identified fmax. Each sensor performs ground modal tests or finite element simulations to obtain the first two order frequencies of the entire arrow, which are f1 and f2 respectively. The lowest order frequency fmin is set less than 0.5f1, and the maximum frequency fmax to be identified is set greater than f2.

9. The sensor arrangement method based on flight mode identification requirements according to claim 1, characterized in that: The method for testing the sensor arrangement scheme is as follows: According to the ground modal test results, a model is established and dynamic response simulation is carried out to obtain the response values ​​at each measuring point of the whole arrow. According to the obtained response value F(t), judgment is made: In the formula, M, C, and K are the model stiffness matrix, damping matrix, and stiffness matrix respectively. x is the acceleration, velocity, and displacement response, respectively, and F(t) is the external excitation; when the response value meets the requirements of the flight mode identification task, the current sensor installation position is reasonable, and the sensor layout plan inspection is completed after traversing all sensors.

10. The sensor arrangement method based on flight mode identification requirements according to claim 1, characterized in that: Before checking the sensor layout, determine the impossible installation positions and delete them according to the installation restrictions of the rocket body structure, and use the remaining positions to check the sensor layout by measuring the response values ​​of the corresponding measurement points; If the corresponding measurement point response values ​​of the remaining positions all meet the requirements of the flight mode identification task, the sensor layout plan is tested; if the corresponding measurement point response values ​​of the installation positions do not meet the requirements of the flight mode identification task, the flight mode identification task requirement analysis is performed again; Among them, if the modal parameters and types to be identified cannot be identified in each sensor installation position, it is deemed that the flight mode identification task requirements are not met.

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