姿控火箭发动机脉冲推力测试系统研究
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摘要
姿控火箭发动机已经成为现代空间飞行器的重要组成部分,广泛应用于卫星等航天器的姿态调整、精密定位、中段修正、对接、交会、分离和制动中。姿控火箭发动机的力值较小,且为脉冲力。脉冲推力的准确测量对改进发动机性能、提高航天器的控制精度、节省有限的星上能源具有重要意义。现有的推力测试装置和方法已不能满足当前姿控火箭发动机脉冲推力的测量要求。
本文以姿控火箭发动机的脉冲推力测量为研究对象,以弥补现有测力装置在动态小推力测量方面的不足为出发点,从理论分析、仿真和实验等方面对动态推力测试系统的设计、分析进行了深入研究,研制出了基于压电石英传感器的姿控火箭发动机脉冲推力测试系统。
首先根据姿控火箭发动机脉冲推力的测量要求,从频率响应和瞬态响应对推力测试系统进行分析,研究了影响测试系统动态响应误差的因素和规律。固有频率和阻尼比是影响频率响应误差的主要参数,而瞬态响应误差则与固有频率和被测推力的上升沿时间的乘积有关。该乘积与超调量的关系曲线在低阻尼下会产生振荡,在曲线振荡范围内,通过调整固有频率与上升沿时间的乘积接近整数,可有效降低低阻尼系统的超调误差。针对上升沿幅值误差反映输出与输入差异方面的不足,采用跟踪误差代替上升沿幅值误差确定测试系统的动态设计参数。根据确定出的测试系统的固有频率和阻尼比的设计要求,选择压电石英力传感器作为力敏元件对姿控火箭发动机的脉冲推力进行测量。
研究了晶片数量和切型对压电传感器的固有频率、灵敏度以及零漂的影响,确定了传感器的晶片数量和切型。设计了具有一体化结构的壳体,建立了壳体的刚度数学模型,利用数学模型和有限元结合的方法,对壳体进行了刚度固有频率分析,确定出壳体参数。两对剪切型压电传感器对称布置在壳体弹性环中,依靠传感器与弹性环结合面的摩擦力实现推力传递,从而完成了测试系统的核心一测力平台的设计。采用锤击实验确定出测试系统的幅频特性曲线,研究了附加质量和质心变化对测试系统固有频率的影响。建立了测试系统的静态和动态标定系统,进行了静态和动态标定实验。该测试系统的固有频率和静态性能指标均满足设计要求,但动态标定输出波形上有叠加振荡,且输出幅值小于输入。
为了改进测试系统的动态测量性能、消除输出波形上的叠加振荡,首先建立测试系统传递函数模型。在已知幅频和相频特性曲线的条件下,测试系统的传递函数只与阻尼比有关,采用半功率法确定出阻尼比,得到测试系统的传递函数模型。基于该模型,在不改变固有频率的前提下,建立了测试系统的阻尼补偿传递函数模型,解决了阻尼补偿中出现的发散问题。利用阻尼补偿传递函数模型对测试系统的动态标定数据进行了阻尼补偿,验证了这种补偿方法的有效性。采用跟踪误差和冲量误差对补偿前后的推力波形进行了评定,结果表明,该测试系统经补偿后的动态测量误差能够满足姿控火箭发动机脉冲推力的测量要求。
As an important component of modern spacecraft, attitude-control rocket engines have been widely used in satellites, space shuttle and aircrafts. They play a very importance role in attitude adjustment, precision orientation, midcourse correction, docking, rendezvous, separation, braking. The thrust of attitude-control rocket engines is a small pulsed thrust. Accurately measuring the pulsed thrust has important significance to improving engine performance, promoting spacecraft control precision, saving limited fuel carried by small-satellites. The present thrust measurement devices and methods cannot satisfy the requirements of the pulsed thrust.
In order to make up for the deficiency of the studies existing, this paper takes the pulsed thrust measurement of attitude-control rocket engines as research object, makes further study on the design and analysis of pulsed thrust measurement system from theoretical simulation to experiment analysis. A pulsed thrust measurement system based on piezoelectric quartz sensors has been developed for attitude-control rocket engine.
According to the measurement requirements of the pulsed thrust, the factors which effect dynamic reponse error of the measurement system are studied by analysis of frequency response and transient response. The natural frequency and damping ratio are the main parameters affecting frequency response errors. Transient response errors are related to the product of natural frequency and the rising edge time of the measured thrust. Under low damping conditions, the ralationship curve of the product and the overshoot generates oscillation. The overshoot error of the low damping system can be effectively reduced by adjusting the product near integer within oscillation range. Aiming at the deficiency of the amplitude error of the rising edge affecting the difference between the output and input waveforms, tracking error is used to substitute the amplitude error of the rising edge to determine the dynamic design parameters of the thrust measurement system. According to design requirements of determined natural frequency and damping ratio of the system, piezoelectric quartz force sensors are selected as the force-sensitive element to measure the pulsed thrust of attitude-control rocket engines.
The influences of the quantity and type of wafer on the natual frequency, sensitivity and zero drift of piezoelectric sensors are studied, and the quantity and type of wafer of the sensors are determined. An integrated shell is designed, and its stiffness model is established. The parameters of the shell are determined by analyzing its stiffness and natural frequency using stiffness model and finite element method. Two shear-type piezoelectric quartz sensors are symmetrically distributed inside of the elastic rings of the shell. The thrust is transferred by means of friction of fitting surfaces. As the core of the thrust measurement system, the thrust-measuring platform is established. The amplitude-frequency characteristic curve of the system is obtained by hammer experiment. The influences of the additional mass and centroid position on the natural frequency of the system are studied. The static and dynamic calibration systems are established, and calibration experiments are carried out. The natural frequency and static indexes can meet the design requirements, but the output waveforms of the dynamic calibration generate additonal damped oscillation, the amplitude of the output waveform is also less than of the input waveform.
The transfer function model of the system is established in order to improving the dynamic measurement performance and eliminating the additional damped oscillation. If the amplitude-frequency and phase-frequency response curves are known, then the transfer function of the system is just related to damping ratios. The transfer function model can be established by determining the damping ratios with half-power method. Based on the model, damping compensation transfer function model of the system is established under the premise of remaining frequency unchanged. The divergence problem of the damping compensation is resolved. The dynamic calibration data is compensated by using damping compensation transfer function model, and the validity of the compensation method is verified. The thrust waveforms before and after damping compensation are evaluated by using tracking error and impulse error. The experimental results indicate that the dynamic measurement error of the system after damping compensation can satisfy the measurement requirements of pulsed thrust of the attitude-control rocket engines.
本文以姿控火箭发动机的脉冲推力测量为研究对象,以弥补现有测力装置在动态小推力测量方面的不足为出发点,从理论分析、仿真和实验等方面对动态推力测试系统的设计、分析进行了深入研究,研制出了基于压电石英传感器的姿控火箭发动机脉冲推力测试系统。
首先根据姿控火箭发动机脉冲推力的测量要求,从频率响应和瞬态响应对推力测试系统进行分析,研究了影响测试系统动态响应误差的因素和规律。固有频率和阻尼比是影响频率响应误差的主要参数,而瞬态响应误差则与固有频率和被测推力的上升沿时间的乘积有关。该乘积与超调量的关系曲线在低阻尼下会产生振荡,在曲线振荡范围内,通过调整固有频率与上升沿时间的乘积接近整数,可有效降低低阻尼系统的超调误差。针对上升沿幅值误差反映输出与输入差异方面的不足,采用跟踪误差代替上升沿幅值误差确定测试系统的动态设计参数。根据确定出的测试系统的固有频率和阻尼比的设计要求,选择压电石英力传感器作为力敏元件对姿控火箭发动机的脉冲推力进行测量。
研究了晶片数量和切型对压电传感器的固有频率、灵敏度以及零漂的影响,确定了传感器的晶片数量和切型。设计了具有一体化结构的壳体,建立了壳体的刚度数学模型,利用数学模型和有限元结合的方法,对壳体进行了刚度固有频率分析,确定出壳体参数。两对剪切型压电传感器对称布置在壳体弹性环中,依靠传感器与弹性环结合面的摩擦力实现推力传递,从而完成了测试系统的核心一测力平台的设计。采用锤击实验确定出测试系统的幅频特性曲线,研究了附加质量和质心变化对测试系统固有频率的影响。建立了测试系统的静态和动态标定系统,进行了静态和动态标定实验。该测试系统的固有频率和静态性能指标均满足设计要求,但动态标定输出波形上有叠加振荡,且输出幅值小于输入。
为了改进测试系统的动态测量性能、消除输出波形上的叠加振荡,首先建立测试系统传递函数模型。在已知幅频和相频特性曲线的条件下,测试系统的传递函数只与阻尼比有关,采用半功率法确定出阻尼比,得到测试系统的传递函数模型。基于该模型,在不改变固有频率的前提下,建立了测试系统的阻尼补偿传递函数模型,解决了阻尼补偿中出现的发散问题。利用阻尼补偿传递函数模型对测试系统的动态标定数据进行了阻尼补偿,验证了这种补偿方法的有效性。采用跟踪误差和冲量误差对补偿前后的推力波形进行了评定,结果表明,该测试系统经补偿后的动态测量误差能够满足姿控火箭发动机脉冲推力的测量要求。
As an important component of modern spacecraft, attitude-control rocket engines have been widely used in satellites, space shuttle and aircrafts. They play a very importance role in attitude adjustment, precision orientation, midcourse correction, docking, rendezvous, separation, braking. The thrust of attitude-control rocket engines is a small pulsed thrust. Accurately measuring the pulsed thrust has important significance to improving engine performance, promoting spacecraft control precision, saving limited fuel carried by small-satellites. The present thrust measurement devices and methods cannot satisfy the requirements of the pulsed thrust.
In order to make up for the deficiency of the studies existing, this paper takes the pulsed thrust measurement of attitude-control rocket engines as research object, makes further study on the design and analysis of pulsed thrust measurement system from theoretical simulation to experiment analysis. A pulsed thrust measurement system based on piezoelectric quartz sensors has been developed for attitude-control rocket engine.
According to the measurement requirements of the pulsed thrust, the factors which effect dynamic reponse error of the measurement system are studied by analysis of frequency response and transient response. The natural frequency and damping ratio are the main parameters affecting frequency response errors. Transient response errors are related to the product of natural frequency and the rising edge time of the measured thrust. Under low damping conditions, the ralationship curve of the product and the overshoot generates oscillation. The overshoot error of the low damping system can be effectively reduced by adjusting the product near integer within oscillation range. Aiming at the deficiency of the amplitude error of the rising edge affecting the difference between the output and input waveforms, tracking error is used to substitute the amplitude error of the rising edge to determine the dynamic design parameters of the thrust measurement system. According to design requirements of determined natural frequency and damping ratio of the system, piezoelectric quartz force sensors are selected as the force-sensitive element to measure the pulsed thrust of attitude-control rocket engines.
The influences of the quantity and type of wafer on the natual frequency, sensitivity and zero drift of piezoelectric sensors are studied, and the quantity and type of wafer of the sensors are determined. An integrated shell is designed, and its stiffness model is established. The parameters of the shell are determined by analyzing its stiffness and natural frequency using stiffness model and finite element method. Two shear-type piezoelectric quartz sensors are symmetrically distributed inside of the elastic rings of the shell. The thrust is transferred by means of friction of fitting surfaces. As the core of the thrust measurement system, the thrust-measuring platform is established. The amplitude-frequency characteristic curve of the system is obtained by hammer experiment. The influences of the additional mass and centroid position on the natural frequency of the system are studied. The static and dynamic calibration systems are established, and calibration experiments are carried out. The natural frequency and static indexes can meet the design requirements, but the output waveforms of the dynamic calibration generate additonal damped oscillation, the amplitude of the output waveform is also less than of the input waveform.
The transfer function model of the system is established in order to improving the dynamic measurement performance and eliminating the additional damped oscillation. If the amplitude-frequency and phase-frequency response curves are known, then the transfer function of the system is just related to damping ratios. The transfer function model can be established by determining the damping ratios with half-power method. Based on the model, damping compensation transfer function model of the system is established under the premise of remaining frequency unchanged. The divergence problem of the damping compensation is resolved. The dynamic calibration data is compensated by using damping compensation transfer function model, and the validity of the compensation method is verified. The thrust waveforms before and after damping compensation are evaluated by using tracking error and impulse error. The experimental results indicate that the dynamic measurement error of the system after damping compensation can satisfy the measurement requirements of pulsed thrust of the attitude-control rocket engines.
引文
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