WO2022007509A1 - 一种图像制导飞行器延时补偿方法及系统 - Google Patents

一种图像制导飞行器延时补偿方法及系统 Download PDF

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WO2022007509A1
WO2022007509A1 PCT/CN2021/094852 CN2021094852W WO2022007509A1 WO 2022007509 A1 WO2022007509 A1 WO 2022007509A1 CN 2021094852 W CN2021094852 W CN 2021094852W WO 2022007509 A1 WO2022007509 A1 WO 2022007509A1
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aircraft
filter
guidance
image
rate gyro
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PCT/CN2021/094852
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English (en)
French (fr)
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王辉
李涛
林德福
王江
王伟
袁亦方
宋韬
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北京理工大学
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Priority to JP2022576496A priority Critical patent/JP2023529492A/ja
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    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
    • G05D1/12—Target-seeking control
    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
    • G05D1/0094—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots involving pointing a payload, e.g. camera, weapon, sensor, towards a fixed or moving target
    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
    • G05D1/08—Control of attitude, i.e. control of roll, pitch, or yaw
    • G05D1/0808—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft
    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
    • G05D1/10—Simultaneous control of position or course in three dimensions
    • G05D1/101—Simultaneous control of position or course in three dimensions specially adapted for aircraft

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  • the invention relates to a delay compensation method for an image-guided aircraft, belonging to the field of guidance.
  • the aircraft system includes a launch unit and a command unit; the aircraft head is equipped with an infrared imaging seeker, which can ensure the stability of the aircraft's optical axis in space, and can complete functions such as searching, intercepting and tracking targets.
  • the aircraft transmits the image information in the field of view to the rear command unit through the radio data link in real time. And flexibly lock the target according to the current engagement scene, and then the shooter joystick controls the aircraft to track the target until it successfully hits.
  • the shooter's identification ability is especially suitable for target identification under complex natural environment interference, battlefield environment interference and modern camouflage technology. It also allows launching aircraft from limited space and has a wider target selection ability.
  • the extension of the imaging seeker It directly affects the bandwidth and tracking performance of the seeker. In the case of large delay, it will cause the time-domain response of the seeker to oscillate or even become unstable, which will eventually have a negative impact on the hit accuracy of the aircraft.
  • the delay includes: 1) the seeker signal output delay caused by limited onboard hardware computing resources; 2) the image transmission delay caused by image compression, decompression, and the shooter's responsiveness.
  • the traditional imaging seeker is a rate gyro platform seeker, and its control system is complex, bulky, expensive, and difficult to assemble and debug.
  • the prior art uses a strapdown seeker to stabilize the seeker platform through the information output by the strapdown inertial system.
  • the platform seeker rate gyroscope and other components are omitted, the seeker structure is simplified, and the guide
  • the algorithm for extracting the angular velocity of the optical axis in the inertial frame is relatively complicated, and it is necessary to reconstruct the field of view, and then differentiate to obtain the angular velocity of the line of sight.
  • the delay link will lead to an error in the estimation of the angular velocity of the line of sight, which may cause parasitic loops. Destabilize the guidance system.
  • the Smith predictor In the traditional time delay processing process, the Smith predictor is usually used. Smith predictor is a classic delay compensation technology. The control structure is simple and easy to debug. It is frequently used in industrial time delay systems.
  • the disadvantages of the estimator in engineering applications are: for a stable time-delay system, the controller has a slow response to disturbance rejection; for a time-delay system including an integral link, the controller has a steady-state error to the load disturbance; it cannot be applied to unstable time-delay systems. Therefore, when the Smith predictor is used to compensate the delay, it is difficult to improve the shortcomings of the strapdown seeker, and the guidance system still has great instability.
  • an image-guided aircraft delay compensation system is designed.
  • the aircraft is provided with a strapdown seeker, a guidance filter, an autopilot and an angular rate gyro ,
  • the strapdown seeker is used to measure the target angle of view, and its output signal is transmitted to the guidance filter;
  • the guidance filter is used to form the guidance command and transmit the command to the autopilot;
  • the autopilot controls the flight state of the aircraft according to the control commands.
  • the image-guided aircraft delay compensation system further includes an ⁇ - ⁇ filter, which estimates the transmitted parameters to obtain the line-of-sight angular velocity.
  • the image-guided aircraft delay compensation system further includes a first channel model and a second channel model,
  • the output signal of the strapdown seeker is corrected by the first channel model and transmitted to the ⁇ - ⁇ filter, and the output signal of the angular rate gyro is transmitted to the ⁇ - ⁇ filter after being compensated by the second channel model, and the ⁇ - ⁇ filter is solved.
  • the signal is transmitted to the guidance filter, and the guidance filter forms a guidance command according to the output signal of the alpha-beta filter, and transmits the command to the autopilot to control the flight state of the aircraft at the next moment.
  • the present invention also provides a delay compensation method for an image-guided aircraft, in which the measurement value of the target viewing angle is obtained by measuring the state of the aircraft at the previous moment by the strapdown seeker measure the target viewing angle The pitch rate at the last moment measured with the angular rate gyro The information is transmitted to the guidance filter, and the guidance filter solves the information to generate control instructions and transmit them to the autopilot, and the autopilot controls the aircraft state at the next moment according to the instruction information. The above process is repeated to continue to measure and control the state of the aircraft at the next moment, thereby forming a guidance loop.
  • the passed parameters are estimated by an alpha-beta filter before the guidance filter.
  • the delay link function G s in the strapdown seeker is:
  • ⁇ g are the natural frequency and damping ratio of the angular rate gyro, respectively
  • T SD is the signal processing time of the strapdown seeker.
  • the pitch angle is the pitch rate measured by the diagonal rate gyro Earn points.
  • the first channel model is set after the strapdown seeker, and the second channel model is set after the angular rate gyro to compensate the two channels where the strapdown seeker and the angular rate gyro are located, so that the strapdown seeker channel and the angular rate gyro are located.
  • the angular rate gyro channel bandwidth is the same.
  • the processing function performed in the first channel model is the same as the processing function performed in the angular rate gyro, and the processing function performed in the second channel model is the same as the delay link function performed in the strapdown seeker.
  • the first channel model performs the following function deal with:
  • the second channel model performs the following function deal with:
  • ⁇ g are the natural frequency and damping ratio of the angular rate gyro, respectively
  • T SD is the signal processing time of the strapdown seeker.
  • the image-guided aircraft delay compensation method provided according to the present invention has the following:
  • the strapdown seeker is adopted to stabilize the seeker platform through the information output by the strapdown inertial system, omitting the platform seeker rate gyroscope and other components, simplifying the seeker structure and reducing the amount of guidance Head volume, while reducing development and production costs;
  • FIG. 1 shows a schematic diagram of an image-guided aircraft system of a preferred embodiment
  • FIG. 2 shows a schematic diagram of an image-guided aircraft system of a preferred embodiment
  • FIG. 3 shows a schematic diagram of a method for estimating line-of-sight angular velocity of an image-guided aircraft in a preferred embodiment
  • FIG. 4 shows a schematic diagram of a delay compensation method for an image-guided aircraft based on line-of-sight angular velocity estimation in a preferred embodiment
  • FIG. 5 shows a schematic diagram of an image-guided aircraft system without delay compensation with an initial velocity pointing error in Embodiment 1;
  • FIG. 6 shows a schematic diagram of a time-delay compensation image-guided aircraft system with an initial velocity pointing error in Embodiment 2;
  • Fig. 7 shows the time-dependent change result of off-target amount in Experimental Example 1;
  • FIG. 8 shows the results of time-dependent changes in the pitch angular velocity in Experimental Example 1.
  • the present invention also provides an image-guided aircraft delay compensation system, including a strapdown seeker, a guidance filter, an autopilot and an angular rate gyro.
  • the strapdown seeker is used to measure the target angle of view, and its output signal is transmitted to the guidance filter;
  • the guidance filter is used to form the guidance command and transmit the command to the autopilot;
  • the autopilot manipulates the actuator (rudder surface) to deflect a certain angle according to the control command, and adjusts the flight attitude of the aircraft to change the acceleration, that is, the autopilot changes the speed of the aircraft by outputting the acceleration of the aircraft, and finally changes the position of the aircraft;
  • Angular rate gyroscopes are used to measure pitch angular velocity.
  • the image-guided aircraft delay compensation system further includes an alpha-beta filter for estimating the passed parameters.
  • ⁇ and ⁇ are undetermined parameters
  • T s is the update step size of the ⁇ - ⁇ filter
  • z is a variable in the discrete domain.
  • the output signal of the strapdown seeker is transmitted to the ⁇ - ⁇ filter
  • the solution signal of the ⁇ - ⁇ filter is transmitted to the guidance filter
  • the output signal of the angular rate gyro is transmitted to the guidance filter
  • the guidance filter is transmitted according to the ⁇ - ⁇ filter.
  • the filter and the output signal of the angular rate gyro form the guidance command, and the command is transmitted to the autopilot to control the flight state of the aircraft.
  • the angular rate gyro detects the pitch angular velocity of the aircraft to monitor its flight state; the strapdown seeker measures the target according to the flight state of the aircraft perspective.
  • the image-guided aircraft delay compensation system further includes a first channel model and a second channel model, and further, the first channel model and the second channel model are chips with solving capability .
  • the output signal of the strapdown seeker is corrected by the first channel model and then transmitted to the ⁇ - ⁇ filter, and the output signal of the angular rate gyro is transmitted to the ⁇ - ⁇ filter after being compensated by the second channel model, and the ⁇ - ⁇ filter
  • the solution signal is transmitted to the guidance filter, and the guidance filter forms the guidance command according to the output signal of the ⁇ - ⁇ filter, and transmits the command to the autopilot to control the flight state of the aircraft at the next moment.
  • the two channels where the strapdown seeker and the angular rate gyro are located are compensated by the first channel model and the second channel model, the field of view angle is reconstructed after compensation, and the reconstructed signal is transmitted to the ⁇ - ⁇ filter , the compensation of the two channels can be adjusted, and by setting the first channel model and the second channel model reasonably, the isolation degree after compensation can be zero, thereby reducing the disturbance of the strapdown seeker to the aircraft.
  • the following second-order transfer function G g processing is performed in the angular rate gyro:
  • the delay link G s in the strapdown seeker is:
  • the first channel model performs the following function deal with:
  • the second channel model performs the following function deal with:
  • ⁇ g are the natural frequency and damping ratio of the angular rate gyro, respectively
  • T SD is the signal processing time of the strapdown seeker.
  • the present invention provides a delay compensation method for an image-guided aircraft, which can effectively reduce the influence of delay on the guidance accuracy and improve many shortcomings in the guidance process of the strapdown seeker.
  • the measurement value of the target angle of view is obtained through the measurement of the aircraft state at the previous moment by the strapdown seeker measure the target viewing angle
  • the pitch rate at the last moment measured with the angular rate gyro The information is transmitted to the guidance filter, and the guidance filter solves the information to generate control instructions and transmit them to the autopilot, and the autopilot controls the aircraft state at the next moment according to the instruction information. Repeat the above process, continue to measure and control the state of the aircraft at the next moment, so as to form a guidance loop, as shown in FIG. 1 .
  • a filter preferably an alpha-beta filter, is also provided on the aircraft to estimate the passed parameters before the guidance filter.
  • the ⁇ - ⁇ filter is a filter that can be used for state estimation and data smoothing. It does not depend on the specific model of the system, is simple and effective, and is often used for distance, angle, and velocity estimation.
  • the target perspective of the measurement Estimate and obtain the angular velocity of the target viewing angle Angular velocity according to target viewing angle and pitch angular velocity Reconstruct the field of view to obtain the angular velocity between the reconstructed aircraft and the target line of sight will Passed to the guidance filter for solving.
  • ⁇ and ⁇ are undetermined parameters, and satisfy 0 ⁇ 1, 0 ⁇ 2, 0 ⁇ 4-2 ⁇ - ⁇ ;
  • T s is the update step size of the ⁇ - ⁇ filter, generally 0.01 to 0.05;
  • the above-mentioned ⁇ - ⁇ filter has a simple structure and good performance, and can effectively complete the estimation of the angular velocity.
  • the transfer function for guidance filter (s) treatment G 1, for the transfer function G 2 (s) in the autopilot processing is:
  • T g is the dynamic time constant of the guidance system and s is a complex variable.
  • the transfer function between the guidance filter and the autopilot is NV c , where N is the navigation coefficient, generally ranging from 4 to 6, and V c is the relative speed between the aircraft and the target; the autopilot and the angular rate
  • N is the navigation coefficient, generally ranging from 4 to 6
  • V c is the relative speed between the aircraft and the target
  • the transfer function between the gyroscopes is where V m is the speed of the aircraft, and T ⁇ is the time constant of the angle of attack of the aircraft.
  • the following continuous second-order system G g processing is performed in the angular rate gyroscope:
  • the delay link function G s in the strapdown seeker is:
  • ⁇ g are the natural frequency and damping ratio of the angular rate gyro, respectively;
  • T SD is the duration for the strapdown seeker to process the signal.
  • ⁇ ⁇ q ⁇ - ⁇ .
  • V c the relative speed of the aircraft and the target
  • T the final guidance time of the aircraft
  • t the flight time in the final guidance section of the aircraft
  • the Z axis is perpendicular to the horizontal plane, and the position projection Z t of the target on the Z axis is obtained by the imaging seeker.
  • the isolation degree is generally used to characterize the ability of the strapdown seeker to isolate the aircraft disturbance. The greater the isolation degree, the lower the ability of the strapdown seeker to isolate the aircraft disturbance.
  • the method of reconstructing the field of view forms an image guidance loop
  • the time delay T SD of the strapdown seeker and the dynamics of the ⁇ - ⁇ filter are inconsistent with the dynamics of the angular rate gyro, it will cause the seeker to malfunction. Oscillation and instability of the time domain response, resulting in extremely high isolation, will severely affect the performance of the guidance system.
  • a delay model is set to perform delay compensation on the image-guided aircraft, and the isolation degree after compensation can be reduced to zero through the delay model, so as to realize that the attitude motion of the aircraft has approximately no disturbance to the guidance command output by the strapdown seeker.
  • how to obtain the isolation before and after compensation is the difficulty of the present invention.
  • the target viewing angle and pitch angle are reconstructed, the angular motion of the aircraft is decoupled, and then estimated through the ⁇ - ⁇ filter, as shown in Figure 3.
  • the target viewing angle is measured by the strapdown seeker Pitch rate measured by diagonal rate gyroscope earn points
  • the alpha-beta filter will Filter to get the estimated line-of-sight angular velocity
  • the first channel model is set after the strapdown seeker Set the second channel model in the angular rate gyro channel Compensating the two channels, the isolation transfer function after compensation is:
  • the processing function performed in the first channel model is the same as the processing function performed in the angular rate gyro, and the processing function performed in the second channel model is the same as the delay link function in the strapdown seeker, The isolation degree after compensation is zero.
  • the processing function performed in the first channel model is set as:
  • the processing function performed in the second channel model is set to
  • ⁇ g are the natural frequency and damping ratio of the angular rate gyro, respectively
  • T SD is the signal processing time of the strapdown seeker.
  • the bandwidth of the strapdown seeker channel and the angular rate gyro channel are consistent, which solves the line-of-sight angular velocity estimation error caused by the delay link, and effectively compensates for the delay effect of the strapdown seeker.
  • the impact of guidance accuracy improves the accuracy of aircraft hits.
  • the output signal of the strapdown seeker is differentiated by the ⁇ - ⁇ filter to obtain the angular velocity of the target viewing angle
  • the angular rate gyro obtains the pitch rate by measuring The angular velocity of the aircraft according to the target viewing angle and pitch velocity
  • the field of view is reconstructed and calculated by the guidance filter.
  • the calculated information is transmitted to the autopilot, and the autopilot controls the aircraft according to the received information, and outputs information such as aircraft acceleration.
  • the aircraft acceleration information and transfer function Obtain the true pitch angular velocity of the aircraft, and perform one integration to obtain the aircraft pitch angle ⁇ ; perform two integrations according to the aircraft acceleration information and synthesize the image information to obtain the line-of-sight angle q ⁇ between the aircraft and the target, and then obtain the real target viewing angle ⁇ ⁇ .
  • the initial velocity pointing error V m ⁇ ⁇ of the aircraft is set in the guidance system model
  • the delay model is not set in the guidance loop, and the aircraft is not compensated, as shown in Figure 5.
  • the performance of the aircraft guidance system is represented by the variation of the missed target amount and the aircraft pitch angular velocity with the terminal guidance time.
  • the missed target amount is the deviation of the final hit point of the aircraft from the target caused by the error signal, which is an important parameter reflecting the performance of the aircraft guidance system, and the damage efficiency of the aircraft is a strong correlation function of the missed target amount;
  • the aircraft pitch angle speed can represent the stability of the aircraft during flight. The faster the aircraft pitch angle changes, the more unstable the aircraft flight.
  • Example 1 The simulation results of the guidance system in Example 1 and Example 2 are shown in Figure 7 and Figure 8.
  • Example 1 the delay of the aircraft is not compensated, resulting in the variation of the missed target amount and the pitch angular velocity of the aircraft with the final guidance time.
  • the trend is that at the end of the terminal guidance/aircraft landing, the miss distance is about 10m, and the pitch angular velocity is about 12deg/s.
  • Embodiment 2 of compensating the aircraft delay can make the image-guided aircraft achieve a better terminal guidance effect.
  • Example 2 the amount of missed targets of the aircraft in the middle and late stages of terminal guidance is close to 0, that is, the guidance accuracy of the aircraft is higher.
  • Example 2 the pitch angular velocity of the aircraft in the middle and late stages of terminal guidance changes steadily and tends to 0, indicating that it is beneficial for the strapdown seeker to track the target continuously and stably.
  • orientation or positional relationship indicated by the terms “upper”, “lower”, “inner”, “outer”, “front”, “rear”, etc. is based on the working state of the present invention
  • the orientation or positional relationship is only for the convenience of describing the present invention and simplifying the description, rather than indicating or implying that the indicated device or element must have a specific orientation, be constructed and operated in a specific orientation, and therefore should not be construed as a limitation of the present invention .
  • first,” “second,” “third,” and “fourth” are used for descriptive purposes only and should not be construed to indicate or imply relative importance.

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Abstract

一种图像制导飞行器延时补偿方法及系统，在飞行器上设置有捷联导引头、制导滤波器、自动驾驶仪和角速率陀螺，在制导滤波器之前通过对目标视角和俯仰角进行视场角重构，在捷联导引头后设置第一通道模型，在角速率陀螺后设置第二通道模型，对捷联导引头和角速率陀螺所在的两个通道进行补偿，使得捷联导引头通道与角速率陀螺通道带宽一致，有效解决了捷联导引头延时对制导精度的影响，改善了飞行器由于延时导致的视线角速度估计误差，也改善了由于估计误差引起的寄生回路可能使制导系统不稳定的现象。

Description

一种图像制导飞行器延时补偿方法及系统 技术领域
本发明涉及一种图像制导飞行器延时补偿方法，属于制导领域。
背景技术
在图像制导飞行器系统中，包含发射单元与指挥单元；飞行器头部搭载红外成像导引头，成像导引头可以保证飞行器光轴在空间的稳定，同时能完成搜索、截获和跟踪目标等功能，使得飞行器在飞行过程中将视场内的图像信息实时通过无线电数据链路传输给后方的指挥单元，射手在指挥单元的图像显示屏前观察飞行器回传的图像，了解战场区域场景和目标情况，并依据当前交战场景灵活锁定目标，之后射手操纵手柄控制飞行器跟踪目标，直至成功命中。
利用射手的识别能力，特别适合于复杂自然环境干扰、战场环境干扰和现代伪装技术下的目标识别，也允许从有限空间发射飞行器且具有更宽广的目标选择能力，但是，成像导引头的延时直接影响导引头带宽的大小和跟踪性能，在大延时的情况下会导致导引头的时域响应振荡甚至失稳，最终对飞行器的命中精度产生不良影响。
具体地，该延时包括：1)有限的弹载硬件计算资源导致的导引头信号输出延时；2)图像压缩、解压和射手响应能力导致的图像传输延时。
此外，传统的成像导引头为速率陀螺式平台导引头，其控制系统复杂、体积较大、成本较高、装配调试难度较大。已有技术采用捷联导引头，通过捷联惯性系统输出的信息实现对导引头平台的 稳定，虽然省略了平台导引头速率陀螺等组件、简化了导引头结构、减小了导引头体积，但是其提取光轴在惯性系下转动角速度算法较为复杂，需要进行视场角重构，进而微分得到视线角速度，而延时环节将导致视线角速度估计误差，其引起的寄生回路可能使制导系统不稳定。
在传统的时延处理过程中，通常采用Smith预估器，Smith预估器为一种经典的延时补偿技术，控制结构简单且容易调试，在工业界时滞系统中应用频繁，但Smith预估器在工程应用时的缺点为：对于稳定时滞系统，控制器对扰动抑制响应较慢；对包含积分环节的时滞系统，控制器对负载干扰存在稳态误差；无法应用于不稳定时滞系统，因此，应用Smith预估器在对延时进行补偿时，难以对捷联导引头的缺点进行改善，制导系统仍然具有极大的不稳定性。
因此，亟需研究一种克服上述缺点的图像制导飞行器延时补偿方法。
发明内容
为了解决上述问题，本发明人进行了锐意研究，一方面，设计了一种图像制导飞行器延时补偿系统，在飞行器上设置有捷联导引头、制导滤波器、自动驾驶仪和角速率陀螺，
其中，捷联导引头用于测量目标视角，其输出信号传递至制导滤波器；
制导滤波器用于形成制导指令，将指令传递至自动驾驶仪；
自动驾驶仪根据控制指令控制飞行器飞行状态。
所述图像制导飞行器延时补偿系统还包括α-β滤波器，对传递的参数进行估计，从而得到视线角速度。
所述图像制导飞行器延时补偿系统还包括第一通道模型和第 二通道模型，
捷联导引头输出信号经过第一通道模型进行修正后传递至α-β滤波器，角速率陀螺输出信号经过第二通道模型补偿后传递至α-β滤波器，α-β滤波器解算信号传递至制导滤波器，制导滤波器根据α-β滤波器输出信号形成制导指令，将指令传递至自动驾驶仪以控制下一时刻飞行器飞行状态。
另一方面，本发明还提供了一种图像制导飞行器延时补偿方法，通过捷联导引头对上一时刻飞行器状态的测量获取目标视角测量值
Figure PCTCN2021094852-appb-000001
将目标视角测量值
Figure PCTCN2021094852-appb-000002
与角速率陀螺测量的上一时刻的俯仰角速度
Figure PCTCN2021094852-appb-000003
信息传递至制导滤波器，制导滤波器对信息进行解算生成控制指令传递至自动驾驶仪，自动驾驶仪根据指令信息控制下一时刻飞行器状态。重复上述过程，继续对下一时刻飞行器状态进行测量和对飞行器状态进行控制，从而形成制导回路。
在制导滤波器之前通过α-β滤波器对传递的参数进行估计。
所述角速率陀螺中进行以下二阶传递函数G g处理：
Figure PCTCN2021094852-appb-000004
捷联导引头中的延时环节函数G s为：
Figure PCTCN2021094852-appb-000005
其中，
Figure PCTCN2021094852-appb-000006
和ζ g分别为角速率陀螺的自然频率和阻尼比，T SD为捷联导引头处理信号的时长。
在制导滤波器之前通过对目标视角
Figure PCTCN2021094852-appb-000007
和俯仰角
Figure PCTCN2021094852-appb-000008
进行视场角重构，对飞行器角运动进行解耦，解耦后再通过α-β滤波器进行估算；
其中，俯仰角
Figure PCTCN2021094852-appb-000009
为通过对角速率陀螺测量得到的俯仰角速度
Figure PCTCN2021094852-appb-000010
进行积分获得。
在捷联导引头后设置第一通道模型，在角速率陀螺后设置第二通道模型，对捷联导引头和角速率陀螺所在的两个通道进行补偿，使得捷联导引头通道与角速率陀螺通道带宽一致。
第一通道模型中进行的处理函数与角速率陀螺中进行的处理函数相同，第二通道模型中进行的处理函数与捷联导引头中的延时环节函数相同。
所述第一通道模型中进行以下函数
Figure PCTCN2021094852-appb-000011
处理：
Figure PCTCN2021094852-appb-000012
所述第二通道模型中进行以下函数
Figure PCTCN2021094852-appb-000013
处理：
Figure PCTCN2021094852-appb-000014
其中，
Figure PCTCN2021094852-appb-000015
和ζ g分别为角速率陀螺的自然频率和阻尼比，T SD为捷联导引头处理信号的时长。
根据本发明提供的图像制导飞行器延时补偿方法，具有以下
有益效果：
(1)采用捷联导引头，通过捷联惯性系统输出的信息实现对导引头平台的稳定，省略了平台导引头速率陀螺等组件，简化了导引头结构，减小了导引头体积，同时降低了研制和生产成本；
(2)有效解决了捷联导引头延时对制导精度的影响；
(3)改善了飞行器由于延时导致的视线角速度估计误差，以及由于估计误差引起的寄生回路可能使制导系统不稳定的现象。
附图说明
图1示出一种优选实施方式的图像制导飞行器系统示意图；
图2示出一种优选实施方式的图像制导飞行器系统示意图；
图3示出一种优选实施方式的图像制导飞行器视线角速度估计方法示意图；
图4示出一种优选实施方式的图像制导飞行器基于视线角速度估计的延时补偿方法示意图；
图5示出实施例1中具有初始速度指向误差的未延时补偿图像制导飞行器系统示意图；
图6示出实施例2中具有初始速度指向误差的延时补偿图像制导飞行器系统示意图；
图7示出实验例1中脱靶量随时间变化结果；
图8示出实验例1中俯仰角速度随时间变化结果。
具体实施方式
下面通过对本发明进行详细说明，本发明的特点和优点将随着这些示例性说明而变得更为清楚、明确。
在这里专用的词“示例性”意为“用作例子、实施例或说明性”。这里作为“示例性”所说明的任何实施例不必解释为优于或好于其它实施例。尽管在附图中示出了实施例的各种方面，但是除非特别指出，不必按比例绘制附图。
一方面，本发明还提供了一种图像制导飞行器延时补偿系统，包括捷联导引头、制导滤波器、自动驾驶仪和角速率陀螺。
其中，捷联导引头用于测量目标视角，其输出信号传递至制导滤波器；
制导滤波器用于形成制导指令，将指令传递至自动驾驶仪；
自动驾驶仪根据控制指令操纵执行机构(舵面)偏转一定角度，调整飞行器飞行姿态以改变加速度，即自动驾驶仪通过输出飞行器加速度，改变飞行器速度，最终改变飞行器位置；
角速率陀螺用于测量俯仰角速度。
在一个优选的实施方式中，所述图像制导飞行器延时补偿系统还包括α-β滤波器，对传递的参数进行估算。
所述α-β滤波器中进行以下二阶离散传递函数G f(z)处理：
Figure PCTCN2021094852-appb-000016
其中，α、β为待定参数，T s为α-β滤波器的更新步长，z为离散域的变量。
优选地，捷联导引头输出信号传递至α-β滤波器，α-β滤波器解算信号传递至制导滤波器，角速率陀螺输出信号传递至制导滤波器，制导滤波器根据α-β滤波器和角速率陀螺输出信号形成制导指令，将指令传递至自动驾驶仪以控制飞行器飞行状态，角速率陀螺检测飞行器的俯仰角速度以监控其飞行状态；捷联导引头根据飞行器飞行状态测量目标视角。
在一个优选的实施方式中，所述图像制导飞行器延时补偿系统还包括第一通道模型和第二通道模型，进一步地，所述第一通道模型和第二通道模型为具有解算能力的芯片。
优选地，捷联导引头输出信号经过第一通道模型进行修正后传递至α-β滤波器，角速率陀螺输出信号经过第二通道模型补偿后传递至α-β滤波器，α-β滤波器解算信号传递至制导滤波器，制导滤波器根据α-β滤波器输出信号形成制导指令，将指令传递至自动驾驶仪以控制下一时刻飞行器的飞行状态。
通过第一通道模型和第二通道模型对捷联导引头和角速率陀螺所在的两个通道进行补偿，补偿后进行视场角重构，将重构后的信号传递至α-β滤波器，实现两个通道补偿可调，通过合理设置第一通道模型和第二通道模型，可实现补偿后隔离度为零，从而降低捷联导引头对飞行器扰动。
在一个优选的实施方式中，所述角速率陀螺中进行以下二阶传 递函数G g处理：
Figure PCTCN2021094852-appb-000017
捷联导引头中的延时环节G s为：
Figure PCTCN2021094852-appb-000018
所述第一通道模型中进行以下函数
Figure PCTCN2021094852-appb-000019
处理：
Figure PCTCN2021094852-appb-000020
所述第二通道模型中进行以下函数
Figure PCTCN2021094852-appb-000021
处理：
Figure PCTCN2021094852-appb-000022
其中，
Figure PCTCN2021094852-appb-000023
和ζ g分别为角速率陀螺的自然频率和阻尼比，T SD为捷联导引头处理信号的时长。
另一方面，本发明提供了一种图像制导飞行器延时补偿方法，有效减少延时对制导精度的影响，改善捷联导引头制导过程中的诸多缺点。
通过捷联导引头对上一时刻飞行器状态的测量获取目标视角测量值
Figure PCTCN2021094852-appb-000024
将目标视角测量值
Figure PCTCN2021094852-appb-000025
与角速率陀螺测量的上一时刻的俯仰角速度
Figure PCTCN2021094852-appb-000026
信息传递至制导滤波器，制导滤波器对信息进行解算生成控制指令传递至自动驾驶仪，自动驾驶仪根据指令信息控制下一时刻飞行器状态。重复上述过程，继续对下一时刻飞行器的状态进行测量和对飞行器状态进行控制，从而形成制导回路，如图1所示。
在一个优选的实施方式中，在飞行器上还设置有滤波器，优选为α-β滤波器，在制导滤波器之前对传递的参数进行估计。
所述α-β滤波器是一种可用于状态估计、数据平滑的滤波器，其不依赖系统的具体模型，简单有效，常用于对距离、角度、速度的估计。
在一个实施方式中，通过对测量的目标视角
Figure PCTCN2021094852-appb-000027
进行估计，获得目标视角的角速度
Figure PCTCN2021094852-appb-000028
根据目标视角的角速度
Figure PCTCN2021094852-appb-000029
与俯仰角速度
Figure PCTCN2021094852-appb-000030
重构视场角，获得重构后飞行器与目标视线的角速度
Figure PCTCN2021094852-appb-000031
将
Figure PCTCN2021094852-appb-000032
传递至制导滤波器进行解算。
进一步地，如图2所示，捷联导引头解算的目标视角σ θ＝q θ-θ，q θ为目标的视线角；重构后飞行器与目标视线的角速度
Figure PCTCN2021094852-appb-000033
更优选地，所述α-β滤波器中进行以下二阶离散传递函数G f(z)处理：
Figure PCTCN2021094852-appb-000034
其中，α、β为待定参数，且满足0＜α＜1，0＜β≤2,0＜4-2α-β；
T s为α-β滤波器的更新步长，一般为0.01～0.05；
z为离散域的变量,为采样信号在拉氏变换时引入，进一步地，z＝e Ts，其中T为采样周期，s为复变量；
上述α-β滤波器结构简单、性能良好，能够有效的完成角速度的估计。
在一个优选的实施方式中，制导滤波器中进行传递函数G 1(s)处理，自动驾驶仪中进行传递函数G 2(s)处理：
Figure PCTCN2021094852-appb-000035
其中，T g是制导系统动力学时间常数，s为复变量。
优选地，制导滤波器与自动驾驶仪之间传递函数为NV c，其中N为导航系数，一般在4～6之间取值，V c为飞行器与目标的相对速 度；自动驾驶仪与角速率陀螺之间传递函数为
Figure PCTCN2021094852-appb-000036
其中V m为飞行器速度，T α为飞行器攻角时间常数。
在一个优选的实施方式中，所述角速率陀螺中进行以下连续的二阶系统G g处理：
Figure PCTCN2021094852-appb-000037
捷联导引头中的延时环节函数G s为：
Figure PCTCN2021094852-appb-000038
其中，
Figure PCTCN2021094852-appb-000039
和ζ g分别为角速率陀螺的自然频率和阻尼比；
T SD为捷联导引头处理信号的时长。
在一个优选的实施方式中，真实的目标视角σ θ根据弹体运动学解算获得，即σ θ＝q θ-θ。具体地，将飞行器的加速度进行二次积分后，取其在Z轴上的位置投影Z m，结合目标在Z轴上的位置投影Z t，通过公式
Figure PCTCN2021094852-appb-000040
即可得获得飞行器与目标的视线角q θ，其中，V c为飞行器与目标的相对速度，T为飞行器末制导时间，t为飞行器末制导段中已经飞行的时间；
所述Z轴垂直于水平面，目标在Z轴上的位置投影Z t通过成像导引头获得。
在飞行器中，一般采用隔离度表征捷联导引头对飞行器扰动隔离能力的大小，隔离度越大表示捷联导引头隔离飞行器扰动的能力越低。
发明人发现，通过获取目标视角的角速度
Figure PCTCN2021094852-appb-000041
与俯仰角速度
Figure PCTCN2021094852-appb-000042
对视场角重构的方式虽然形成了图像制导回路，但当捷联导引头的 时延T SD、α-β滤波器动力学与角速率陀螺动力学不一致时，会引起导引头的时域响应振荡、失稳，生产极高的隔离度，会剧烈影响制导系统的性能。
在本发明中，设置延时模型对图像制导飞行器进行延时补偿，可以通过延时模型将补偿后的隔离度降低到零，从而实现飞行器姿态运动对捷联导引头输出制导指令近似无扰动的效果，但如何获得补偿前后的隔离度是本发明的难点所在。
优选地，通过获取目标视角
Figure PCTCN2021094852-appb-000043
和俯仰角
Figure PCTCN2021094852-appb-000044
进行视场角重构，对飞行器角运动进行解耦，解耦后再通过α-β滤波器进行估计，如图3所示。
具体地，通过捷联导引头测量目标视角
Figure PCTCN2021094852-appb-000045
通过对角速率陀螺测量得到的俯仰角速度
Figure PCTCN2021094852-appb-000046
进行积分获得
Figure PCTCN2021094852-appb-000047
将
Figure PCTCN2021094852-appb-000048
与
Figure PCTCN2021094852-appb-000049
进行解耦，获得解耦后飞行器与目标的视线角
Figure PCTCN2021094852-appb-000050
α-β滤波器将
Figure PCTCN2021094852-appb-000051
滤波得到估计视线角速度
Figure PCTCN2021094852-appb-000052
通过视场角重构解耦飞行器与目标视线的角速度
Figure PCTCN2021094852-appb-000053
可获得对应的隔离度传递函数：
Figure PCTCN2021094852-appb-000054
进一步地，在捷联导引头通道与角速率陀螺通道中增加模型，对两个通道进行补偿，使得捷联导引头通道与角速率陀螺通道带宽一致，如图4所示。
具体地，在捷联导引头后设置第一通道模型
Figure PCTCN2021094852-appb-000055
在角速率陀螺通道设置第二通道模型
Figure PCTCN2021094852-appb-000056
对两个通道进行补偿，则补偿后的隔离度传递函数为：
Figure PCTCN2021094852-appb-000057
在本发明中，通过将第一通道模型中进行的处理函数与角速 率陀螺中进行的处理函数相同，第二通道模型中进行的处理函数与捷联导引头中的延时环节函数相同，即可实现补偿后的隔离度为零。
在一个优选的实时方式中，将所述第一通道模型中进行的处理函数设置为：
Figure PCTCN2021094852-appb-000058
将所述第二通道模型中进行的处理函数设置为
Figure PCTCN2021094852-appb-000059
其中，
Figure PCTCN2021094852-appb-000060
和ζ g分别为角速率陀螺的自然频率和阻尼比，T SD为捷联导引头处理信号的时长。
通过第一通道模型和第二通道模型，使得捷联导引头通道与角速率陀螺通道带宽一致，解决了延时环节带来的视线角速度估计误差，有效补偿了捷联导引头延时对制导精度的影响，提升了飞行器命中精度。
实施例
实施例1
捷联导引头输出信号经过α-β滤波器进行微分运算，获得目标视角的角速度
Figure PCTCN2021094852-appb-000061
角速率陀螺通过测量获得俯仰角速度
Figure PCTCN2021094852-appb-000062
飞行器根据目标视角的角速度
Figure PCTCN2021094852-appb-000063
与俯仰角速度
Figure PCTCN2021094852-appb-000064
进行视场角重构，经过制导滤波器解算，解算后的信息传递至自动驾驶仪，自动驾驶仪根据接收到的信息控制飞行器，并输出飞行器加速度等信息。根据飞行器加速度信息与传递函数
Figure PCTCN2021094852-appb-000065
得到飞行器真实的俯仰角速度，进行一次积分获得飞行器俯仰角θ；根据飞行器加速度信息进行两次积分并综合图像信息，得到飞行器与目标的视线角q θ，进而获得真实的目标视角σ θ。将q θ和θ反馈至捷联导引头，从而形成制导回 路，实现图像制导。
其中，在制导系统模型中设置飞行器初始速度指向误差V mε ν；
制导回路中不设置延时模型，不对飞行器进行补偿，如图5所示。
在本模拟实验中，通过飞行器初始速度指向误差描述飞行器初速的指向偏离理想轨迹的误差角，误差为ε v＝5deg。
实施例2
使用制导系统仿真软件进行模拟实验，在捷联导引头后设置第一通道模型
Figure PCTCN2021094852-appb-000066
对捷联导引头测量目标视角
Figure PCTCN2021094852-appb-000067
进行补偿，在角速率陀螺后设置第二通道模型
Figure PCTCN2021094852-appb-000068
对
Figure PCTCN2021094852-appb-000069
进行补偿，将补偿后的信号进行视场角重构，对飞行器角运动进行解耦，解耦后再通过α-β滤波器进行估计后进行制导滤波解算，解算出的信息传递至自动驾驶仪，其余设定与实施例1相同，其系统如图6所示。
其中，飞行器的参数设置如下：α-β滤波器中α＝0.99，β＝1.72，T s＝0.02s；制导系统参数为N＝4，T g＝0.4s，T α＝0.68s，V c＝V m＝160m/s；捷联导引头延时和速率陀螺参数为T SD＝40ms，
Figure PCTCN2021094852-appb-000070
ζ g＝0.7。
实验例1
通过脱靶量和飞行器俯仰角速度随末制导时间的变化来表现飞行器制导系统的性能。
所述脱靶量为误差信号导致飞行器最终命中点距离目标的偏差，是反映飞行器制导系统性能的重要参数，飞行器的毁伤效能是脱靶量的强相关函数；
飞行器俯仰角速度能够表示飞行器飞行过程中稳定性，飞行 器俯仰角变化速度越快，则飞行器飞行越不稳定。
实施例1和实施例2中制导系统仿真后结果如图7、图8所示，实施例1中未对飞行器延时进行补偿，导致飞行器脱靶量与俯仰角速度随末制导时间的变化呈现振荡发散的趋势，在末制导结束/飞行器着陆时，脱靶量约为10m,俯仰角速度约为12deg/s。相比之下，对飞行器延时进行补偿的实施例2可以使图像制导飞行器达到较好的末制导效果。
通过图7可知，实施例2中飞行器在末制导中后期的脱靶量趋近于0，即飞行器制导精度更高。
通过图8可知，实施例2中飞行器在末制导中后期的俯仰角速度变化平稳，且趋于0，说明其有利于捷联导引头持续稳定地跟踪目标。
在本发明的描述中，需要说明的是，术语“上”、“下”、“内”、“外”、“前”、“后”等指示的方位或位置关系为基于本发明工作状态下的方位或位置关系，仅是为了便于描述本发明和简化描述，而不是指示或暗示所指的装置或元件必须具有特定的方位、以特定的方位构造和操作，因此不能理解为对本发明的限制。此外，术语“第一”、“第二”、“第三”、“第四”仅用于描述目的，而不能理解为指示或暗示相对重要性。
以上结合具体实施方式和范例性实例对本发明进行了详细说明，不过这些说明并不能理解为对本发明的限制。本领域技术人员理解，在不偏离本发明精神和范围的情况下，可以对本发明技术方案及其实施方式进行多种等价替换、修饰或改进，这些均落入本发明的范围内。本发明的保护范围以所附权利要求为准。

Claims (10)

  1. 一种图像制导飞行器延时补偿系统，其特征在于，在飞行器上设置有捷联导引头、制导滤波器、自动驾驶仪和角速率陀螺，
    其中，捷联导引头用于测量目标视角，其输出信号传递至制导滤波器；
    制导滤波器用于形成制导指令，将指令传递至自动驾驶仪；
    自动驾驶仪根据控制指令调整飞行器飞行姿态，从而改变飞行器加速度，进而改变飞行器速度和飞行器位置。
  2. 根据权利要求1所述的图像制导飞行器延时补偿系统，其特征在于，
    所述图像制导飞行器延时补偿系统还包括α-β滤波器，对传递的参数进行估计，从而得到视线角速度。
  3. 根据权利要求1所述的图像制导飞行器延时补偿系统，其特征在于，
    所述图像制导飞行器延时补偿系统还包括第一通道模型和第二通道模型，
    捷联导引头输出信号经过第一通道模型进行修正后传递至α-β滤波器，角速率陀螺输出信号经过第二通道模型补偿后传递至α-β滤波器，α-β滤波器解算信号传递至制导滤波器，制导滤波器根据α-β滤波器输出信号形成制导指令，将指令传递至自动驾驶仪以控制下一时刻飞行器飞行状态。
  4. 一种图像制导飞行器延时补偿方法，其特征在于，
    通过捷联导引头对上一时刻飞行器状态的测量获取目标视角测量值
    Figure PCTCN2021094852-appb-100001
    将目标视角测量值
    Figure PCTCN2021094852-appb-100002
    与角速率陀螺测量的上一时刻的俯仰角速度
    Figure PCTCN2021094852-appb-100003
    信息传递至制导滤波器，制导滤波器对信息进行解算生成控制指令传递至自动驾驶仪，自动驾驶仪根据指令信息控制下一时刻飞行器状态，重复上述过程，继续对下一 时刻飞行器状态进行测量和对飞行器状态进行控制，从而形成制导回路。
  5. 根据权利要求4所述的图像制导飞行器延时补偿方法，其特征在于，
    在制导滤波器之前通过α-β滤波器对传递的参数进行估计，从而得到视线角速度。
  6. 根据权利要求4所述的图像制导飞行器延时补偿方法，其特征在于，
    所述角速率陀螺中进行以下二阶传递函数G g处理：
    Figure PCTCN2021094852-appb-100004
    捷联导引头中的延时环节函数G s为：
    Figure PCTCN2021094852-appb-100005
    其中，
    Figure PCTCN2021094852-appb-100006
    和ζ g分别为角速率陀螺的自然频率和阻尼比，T SD为捷联导引头处理信号的时长。
  7. 根据权利要求4所述的图像制导飞行器延时补偿方法，其特征在于，
    在制导滤波器之前通过对目标视角
    Figure PCTCN2021094852-appb-100007
    和俯仰角
    Figure PCTCN2021094852-appb-100008
    进行视场角重构，对飞行器角运动进行解耦，解耦后再通过α-β滤波器进行估算；
    其中，俯仰角
    Figure PCTCN2021094852-appb-100009
    为通过对角速率陀螺测量得到的俯仰角速度
    Figure PCTCN2021094852-appb-100010
    进行积分获得。
  8. 根据权利要求4所述的图像制导飞行器延时补偿方法，其特征在于，
    在捷联导引头后设置第一通道模型，在角速率陀螺后设置第二通道模型，对捷联导引头和角速率陀螺所在的两个通道进 行补偿，使得捷联导引头通道与角速率陀螺通道带宽一致。
  9. 根据权利要求8所述的图像制导飞行器延时补偿方法，其特征在于，
    第一通道模型中进行的处理函数与角速率陀螺中进行的处理函数相同，第二通道模型中进行的处理函数与捷联导引头中的延时环节函数相同。
  10. 根据权利要求9所述的图像制导飞行器延时补偿方法，其特征在于，
    所述第一通道模型中进行以下函数
    Figure PCTCN2021094852-appb-100011
    处理：
    Figure PCTCN2021094852-appb-100012
    所述第二通道模型中进行以下函数
    Figure PCTCN2021094852-appb-100013
    处理：
    Figure PCTCN2021094852-appb-100014
    其中，
    Figure PCTCN2021094852-appb-100015
    和ζ g分别为角速率陀螺的自然频率和阻尼比，T SD为捷联导引头处理信号的时长。
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