CN102393630A - Carrier aircraft landing guide and control system for inhibiting airflow disturbance of stern and control method for system - Google Patents

Abstract

The invention relates to a carrier-based aircraft landing guidance and control system and method for suppressing ship-tail airflow disturbance, and belongs to the technical field of carrier-based aircraft flight control. The system consists of a guidance subsystem mounted on a ship and a control subsystem mounted on an aircraft. The guidance subsystem includes a tracking radar, a radar stabilization platform, a high-speed general-purpose computer, a display platform, a data encoding transmitter, a data link monitor and Flight track recorder; control subsystem including autopilot, data link receiver, receiver decoder, autopilot coupler, autothrottle controller, and on-board radar equipment. The system introduces the altitude change rate feedback and lateral deviation rate feedback modules into the guidance law calculation sub-module of the high-speed general-purpose computer, which effectively suppresses the influence of the ship's wake disturbance and improves the accuracy of the lateral landing trajectory.

Description

Carrier aircraft landing guidance and control system and method for suppressing ship tail airflow disturbance

technical field

The invention relates to a carrier-based aircraft landing guidance and control system and method for suppressing ship-tail airflow disturbance, and belongs to the field of carrier-based aircraft flight control.

Background technique

Disturbance of ship tail airflow is the main factor that causes ship landing guidance error and affects ship landing safety. Pilots even referred to the complex airflow disturbance zone near the stern as entering "the gate of hell".

When the aircraft approaches and lands, and leaves the ship for the last 0.5 miles (800 meters), the MIL-F-8785C military specification regards the disturbance of the ship's tail airflow as a combination of four components. They are quantitatively described and specified to test the landing performance of the aircraft under airflow disturbance.

The turbulence of the tail air flow is composed of four components,

1) Free atmospheric turbulence component

2) Steady-state component of the wake (rooster wake)

3) The periodic component of the wake

4) The random component of the wake.

The characteristics of the free atmospheric turbulent components are independent of the position of the aircraft relative to the ship, and MIL-F-8785C specifies their spatial power spectrum.

The steady-state component of the ship's stern airflow is the main component of the ship's stern atmospheric disturbance. This airflow is caused by the air flowing out from the flat stern of the aircraft carrier when it is traveling against the wind. It is characterized in that in the vertical direction, it produces a unique rooster-tail-shaped wind force, and its wind direction is related to the distance from the stern. Near the stern is the effective downward wind, while away from the stern the downward wind decreases in proportion to the distance and is later changed to an upward wind. This is because during the actual landing process of the aircraft, because the aircraft carrier is sailing in the sea, the air at the stern of the ship is relatively thin, so the air behind is filled in, and coupled with the influence of the deck wind flow, the combined effect of these two winds forms a male Cocktail. Compared with the deck wind, the amplitude of the air to be filled later is very small, and it is actually superimposed on the deck wind in the form of incremental disturbance. The direction of the incremental disturbance is that the horizontal wind is tailwind, and the vertical wind is far away from the stern. It is an ascending segment, and it is a descending segment near the stern. You can refer to the rooster wake model provided by AIAA-79-1772. the

The periodic component of the ship's stern airflow is the wake generated by the ship's pitch, which is the wind force formed by the pitching motion of the deck. It varies with the pitch frequency of the ship, the magnitude of the pitch, the wind on the deck, and the distance of the aircraft from the ship. the

According to MIL-F-8785C, the random velocity component related to the ship is obtained by some form of white noise after the shaping filter.

Contents of the invention

The present invention proposes a carrier-based aircraft landing guidance and control system and method for suppressing the disturbance of the ship's tail airflow. By introducing the feedback information of the altitude change rate and the feedback information of the lateral deviation rate, the effect of well suppressing the disturbance of the ship's wake flow can be achieved. The purpose is to improve the landing accuracy of carrier-based aircraft.

The present invention adopts following technical scheme for solving its technical problem:

A carrier-based aircraft landing guidance and control system that suppresses the disturbance of the ship's tail airflow. It is composed of a guidance subsystem and a control subsystem. The guidance subsystem is loaded on the ship, including tracking radar, radar stabilization platform, high-speed general-purpose computer, and display platform. , data encoding transmitter, data link monitor and flight track recorder, among which, the high-speed general-purpose computer, tracking radar and flight track recorder are sequentially connected, the display platform, the high-speed general-purpose computer and the radar stabilization platform are sequentially connected, and the display platform and the high-speed general-purpose The computer is respectively connected to the data encoding transmitter, the data encoding transmitter, the data link monitor and the display platform are sequentially connected; the control subsystem is loaded on the aircraft, including the autopilot, the data link receiver, the receiving decoder, the autopilot Coupler, auto throttle controller and on-board radar equipment, among which, data link receiver, receiver decoder, autopilot coupler and autopilot are connected sequentially, and auto throttle controller and autopilot are bidirectionally connected; guidance subsystem The data encoding transmitter in the control subsystem is connected with the data link receiver in the control subsystem through radio waves, and the tracking radar in the guidance subsystem is connected with the on-board radar equipment in the control subsystem through Ka-band signals.

Described high-speed general-purpose computer is equipped with deck motion compensation calculation submodule, ideal trajectory submodule, trajectory error signal calculation submodule, data stability processing submodule, aircraft dynamics information submodule and guidance law calculation submodule, wherein and The two-way connected deck motion compensation calculation submodule of the radar stabilization platform is respectively connected to the guidance law calculation submodule, the display platform and the data encoding transmitter through the trajectory error signal calculation submodule; the input terminals of the data stabilization processing submodule are respectively connected to the radar stabilization platform and The tracking radar is connected, the output terminal of the data stabilization processing submodule is connected with the input terminal of the trajectory error signal calculation submodule; the input terminal of the guidance law calculation submodule is connected with the trajectory error signal calculation submodule and the aircraft dynamics information submodule, The output end of the guidance law calculation sub-module is connected to the data encoding transmitter; the ideal trajectory sub-module is connected to the trajectory error signal calculation sub-module.

The control method of the carrier-based aircraft landing guidance and control system for suppressing the disturbance of the ship's tail airflow, including the longitudinal guidance control method and the lateral guidance control method:

(1) The longitudinal guidance control method includes a longitudinal height guidance control method and a longitudinal attitude control method,

为主反馈，有效地抑制舰尾气流扰动对飞机着舰的影响，具体方法是，依据引导律表达式，构建轨迹控制器，实现飞机飞行高度变化率

的反馈，达到抑制舰尾气流扰动对飞机着舰的影响，该轨迹控制器的引导律的表达式为： 1) The longitudinal altitude guidance method is to introduce the flight altitude change rate of the aircraft into the guidance law calculation sub-module of the high-speed general computer

The main feedback is to effectively suppress the influence of the ship's tail airflow disturbance on the aircraft's landing. The specific method is to construct a trajectory controller based on the guidance law expression to realize the flight altitude change rate of the aircraft.

Feedback to suppress the influence of ship tail airflow disturbance on aircraft landing, the expression of the guidance law of the trajectory controller is:

In the formula, K _P is the gain of the proportional term, K _i is the gain of the integral term, K _d is the gain of the differential term, K _dd is the gain of the second differential term, K ₀ is the total gain, H _com is the aircraft flight reference altitude command signal, H is the actual height feedback signal of the aircraft flight, s is a complex variable, where K _P , K _i , K _d , K _dd , K ₀ pass the actual height feedback signal gain of the aircraft flight △H to the aircraft flight reference height command signal gain △H _com The response is optimized to obtain;

反馈，构建纵向姿态控制器，该纵向姿态控制器的控制律表达式为： 2) The longitudinal attitude control method is to add the aircraft flight altitude change rate to the control law of the traditional flight control system, which is traditionally dominated by attitude control. Feedback and the second differential signal of the actual altitude signal of the aircraft

Feedback, build a longitudinal attitude controller, the control law expression of the longitudinal attitude controller is:

为升降舵回路传递函数，为姿态控制参数，通过根轨迹设计方法，来确定姿态控制参数，为飞机飞行实际高度变化率增量，

为飞机飞行实际俯仰角速率增量，

为飞机飞行实际高度变化率增量，为飞机飞行参考高度变化率增量，

为飞机飞行实际高度信号增量的二次微分； In the formula,

is the transfer function of the elevator loop, As the attitude control parameters, the attitude control parameters are determined by the root locus design method, is the increment of the actual altitude change rate of the aircraft,

is the actual pitch rate increment of the aircraft,

is the increment of the actual altitude change rate of the aircraft, is the aircraft flight reference altitude change rate increment,

is the quadratic differential of the signal increment of the actual flight height of the aircraft;

(2) The lateral guidance control method includes a lateral deviation guidance method and a lateral attitude control method:

(1) The lateral deviation guidance method is to introduce the feedback information of the lateral deviation rate into the guidance law sub-calculation module of the high-speed general-purpose computer to construct the lateral trajectory controller. The expression of the guidance law of the lateral trajectory controller is :

为比例项增益、

为积分项增益、

为微分项增益、

为总增益、s为复变量、y为与甲板中心线的侧偏离、y_com为期望的与甲板中心线侧偏离指令; In the formula,

is the proportional term gain,

is the integral term gain,

is the differential term gain,

is the total gain, s is a complex variable, y is the side deviation from the deck centerline, and y _com is the expected side deviation command from the deck centerline;

(2) The lateral attitude control method is to introduce the lateral deviation rate into the high-speed general-purpose computer guidance law calculation sub-module The feedback information of the lateral attitude controller is constructed, and the expression of the control law of the lateral attitude controller is:

为滚转角速度增量，

为偏航角速度增量，

为洗出网络，为迎角基准值，

为侧滑角增量，为控制参数，可以通过根轨迹方法获得， 

为常数，s为复变量，△y_com为期望的侧偏离速率指令。 In the formula, is the roll angle increment,

is the roll angular velocity increment,

is the yaw rate increment,

To wash out the network, is the reference value of the angle of attack,

is the sideslip angle increment, As the control parameter, it can be obtained by the root locus method,

is a constant, s is a complex variable, and △y _com is the expected side deviation speed command.

The beneficial effects of the present invention are as follows:

1. A new longitudinal guidance and control system has been established, which introduces the feedback information of the altitude change rate. When the ship's wake acts on the carrier-based aircraft, since the altitude change rate is equivalent to the track inclination angle, the input signal of the guidance loop Directly control the inclination angle of the track, thereby quickly correcting the landing track, and effectively suppressing the influence of the ship's wake disturbance.

2. Established a new lateral guidance and control system, which introduced lateral deviation rate feedback information. Since the introduction of lateral deviation rate feedback is equivalent to introducing track deflection angle feedback, the control of roll angle is transformed into direct The control of the deflection angle of the track improves the accuracy of the lateral landing track.

Description of drawings

Figure 1 is a schematic diagram of the composition and structure of the automatic landing guidance and control system.

Figure 2 is a schematic diagram of the structure of the longitudinal trajectory controller.

Fig. 3 is a schematic diagram of the slope response of the actual altitude feedback signal increment △H of the aircraft flight to the aircraft flight reference altitude command signal increment △ _Hcom when there is no ship exhaust flow disturbance.

时飞机飞行的实际高度反馈信号增量△H对飞机飞行参考高度指令信号增量△H_com的斜坡响应示意图。 Figure 4 is the introduction of flight altitude change rate

Schematic diagram of the slope response of the actual altitude feedback signal increment △H of the aircraft flight to the aircraft flight reference altitude command signal increment △H _com .

Figure 5 is a schematic diagram of the structure of the longitudinal attitude controller.

Fig. 6 is a schematic diagram of the comparison of the frequency response characteristics of the traditional flight control system and the improved longitudinal flight control system after introducing the altitude change rate.

Figure 7 is a schematic diagram of the structure of the lateral trajectory controller.

Fig. 8 is a schematic diagram of the anti-crosswind effect of the traditional lateral landing guidance control system and the lateral landing guidance control system after introducing the lateral deviation rate.

Figure 9 is a schematic diagram of the structure of the lateral attitude controller.

为升降舵回路传递函数，

为姿态控制参数，通过轨迹设计方法，来确定姿态控制参数，

为飞机飞行实际高度变化率增量，

为飞机飞行实际俯仰角速率增量，

为飞机飞行实际高度变化率增量，

为飞机飞行参考高度变化率增量，

为飞机飞行实际高度信号增量的二次微分；y为与甲板中心线的侧偏离、y_com为期望的与甲板中心显得侧偏离指令，

为滚转角增量，

为滚转角速度增量，

为偏航角速度增量，为洗出网络，

为迎角基准值，为侧滑角增量，

为控制参数，可以通过根轨迹方法获得， 

为常数，s为复变量， △y_com为期望的侧偏离速率指令。；

为传统飞行控制系统，

为引入高度变化率后的纵向改进型飞行控制系统，

为传统侧向着舰引导控制系统，

为引入侧向偏离速率后的侧向着舰引导控制系统。 Among them: H _com is the aircraft flight reference height command signal, H is the actual flight height feedback signal of the aircraft, K _P is the gain of the proportional term, K _i is the gain of the integral term, K _d is the gain of the differential term, and K _dd is the gain of the second differential term , K _o is the total gain, s is the complex variable,

is the transfer function of the elevator loop,

As the attitude control parameters, the attitude control parameters are determined by the trajectory design method,

is the increment of the actual altitude change rate of the aircraft,

is the actual pitch rate increment of the aircraft,

is the increment of the actual altitude change rate of the aircraft,

is the aircraft flight reference altitude change rate increment,

is the quadratic differential of the actual altitude signal increment of the aircraft; y is the lateral deviation from the deck centerline, and y _com is the expected lateral deviation command from the deck center,

is the roll angle increment,

is the roll angular velocity increment,

is the yaw rate increment, To wash out the network,

is the reference value of the angle of attack, is the sideslip angle increment,

As the control parameter, it can be obtained by the root locus method,

is a constant, s is a complex variable, and △y _com is the expected lateral deviation speed command. ;

For traditional flight control systems,

For the longitudinal improved flight control system after introducing the altitude change rate,

It is a traditional lateral landing guidance control system,

It is the lateral landing guidance control system after introducing the lateral deviation rate.

Detailed ways

The invention will be described in further detail below in conjunction with the accompanying drawings.

The automatic landing guidance and control system is divided into a guidance subsystem and a control subsystem. The guidance subsystem is loaded on the ship, and the control subsystem is loaded on the aircraft. The Automatic Landing Guidance and Control System (ACLS) consists of two parts, the ship guidance subsystem and the aircraft control subsystem, as shown in Figure 1.

The onboard guidance subsystem includes Ka-band tracking radar, radar stabilization platform, high-speed general-purpose computer, display platform, data encoding transmitter, data link monitor, and flight track recorder. The information after the trajectory error signal is processed by the guidance law is transmitted from the aircraft carrier to the aircraft in the form of Ka-band signal. The tracking radar of the ship's guidance subsystem scans with a conical scanning antenna. The radar system continuously tracks the trajectory of the aircraft and compares the actual tracked trajectory with the ideal trajectory.

1. Tracking radar

When the aircraft enters the radar intercept window, the tracking radar locks the aircraft, tracks the flight trajectory of the aircraft, and obtains the flight distance, azimuth, and pitch angle of the aircraft relative to the tracking radar measurement coordinate system until the aircraft lands or goes around.

The radar used for ACLS is a high-precision guided landing radar, which emits left-right and up-down scanning beams from the heading and glide antennas in the direction of aircraft landing. Tracking radar acquisition tracks the target aircraft after the aircraft passes through the landing window and measures the aircraft in spherical coordinates with the tracking radar antenna as the coordinate origin. The high-speed general-purpose computer converts the aircraft data measured by the tracking radar into a Cartesian coordinate system composed of distance, height and lateral position, and sets the origin of the coordinates at the position of the scheduled landing point. The radar-stabilized platform, which senses the movement of the carrier deck, approaches the tracking radar antenna and feeds its measured carrier deck movement information into a high-speed general-purpose computer in order to finally establish the aircraft's position in a stable horizontal coordinate system (inertial coordinate system).

The tracking radar in the radar system has a 4-foot diameter parabolic antenna, a 0.5° beamwidth, a cone-scanning radar operating in the Ka band (33.2GHZ), and a pulse repetition frequency of 2000 pulses/second. Peak power is 40 watts. The operating range is 8 nautical miles to 300 feet.

The angle sensor in the radar measurement system uses an optical incremental shaft encoder with a resolution of 14 bits. These encoders generate pulse trains as the radar's gimbal rotates around the vertical axis and the orientation angle, giving the absolute angle through a buffer counter. A high-resolution high-speed counter is used to measure the time between the transmitted pulse and the received pulse.

Before finding the target, use the computer in the radar system to control the elevation angle and azimuth angle of the tracking radar antenna, so that it scans with a rectangular search pattern according to the heading given by the ship's traffic control computer.

After the target is found, the control of the antenna is realized by the radar tracking system.

2. Radar stabilization platform

It transforms the actually detected aircraft position information into an inertial coordinate system with the ideal landing point on the deck as the origin, which can eliminate the influence of deck motion.

The radar stabilized platform (including accelerometer) is a two-axis gyro stabilized platform that measures the pitch angle, roll angle and heave motion of the ship. The rotation of the gimbal of the platform is a pulse train generated by an axial incremental encoder, which is then measured by a buffer.

A single-axis accelerometer is fixed on the radar stable platform, and the vertical acceleration is measured in the form of a DC signal, which is converted into a digital signal by an analog-to-digital converter with a multi-way switch. This measures the heave motion of the ship.

The main function of the radar stabilization platform is to provide the ship's motion information to the computer, so that the aircraft motion can be measured in the inertial space coordinates. Additionally deck motion compensation commands are provided.

3. High-speed general-purpose computer

It is used to establish the inertial stable landing measurement coordinate system, filter the landing error signal, and calculate the guidance law.

There are two computers in the system . Each performs all calculation tasks for the two incoming aircraft at a rate of 20 operations per second. The computer also calculates the on-line diagnosis and off-line monitoring of the two redundancy systems. And the guidance error and command are modulated into VHF carrier. All are sent to the aircraft at 10 times/second. The following discrete messages are also sent: Landing Check, ACL Lock on, Autopilot coupler available, Command Control, Voice, 10 Seconds, Go Around, etc.

The main task of the high-speed general-purpose computer is to calculate the height given signal according to the distance of the aircraft and the glide slope requirements, and compare it with the actual height to form a height deviation signal. In addition, in the lateral channel, the measured lateral position of the aircraft is compared with the centerline position of the aircraft carrier deck to form a lateral lateral deviation signal. According to the requirements of trajectory guidance dynamic characteristics and factors such as anti-deck motion and anti-radar electronic noise, the above two errors are processed by filtering, limiting differential, and integral according to a certain guidance law. Then form a data link and send it to the aircraft.

4. Data link monitor

The data link monitor continuously detects the flight path error transmitted by the data link. If the error does not meet the requirements, the system will transfer to Mode II or Mode III or generate a go-around signal. Mode II refers to the working mode of the instrument landing system, that is, in the cockpit, use the indication of the pointer instrument or the head-up display, and use the error information provided by the automatic landing guidance system to perform manual landing and guide the aircraft to About 3/4n mile away from the ship. Mode III refers to the ship's control approach system, that is, the pilot completes the landing task through the command information given by the ship's console operator.

5. Display platform

Monitor and control various functions of the system.

The automatic approach is monitored and Mode III - a talk-down is performed by the operator of the display device in the event of data link monitor, autopilot coupler, autopilot failure. While executing Mode III, the operator notes the aircraft type, data link address, etc. And under certain circumstances, perform a go-around operation.

The display device can also record the landing track, flight speed, descent speed, ship motion, and ship collision speed of each aircraft, so as to record the driving results.

The control subsystem on the aircraft includes the autopilot (automatic flight control system), data link receiver, receiver decoder, autopilot coupler, autothrottle controller and radar booster.

1. Autopilot

The autopilot is installed on the aircraft and it is the interface between the data link and the aircraft control panel. Pilots use it to select the automatic landing guidance subsystem. In the autopilot can provide state transitions and the state of the signal, turn on the logic circuit, control the signal limit. In addition, it is used to process the signals sent by the data link to control the pitch angle and roll angle of the aircraft, and couple it to the flight control system.

2. Data link receiver

The data link receiver receives the data link signal sent by the ship, and sends the signal to the flight control system after filtering.

3. Receive decoder

The receiving decoder obtains the glide track error signal from the tracking radar of the ship's guidance subsystem, and displays this signal on the cross pointer of the instrument. In mode I, the driver monitors the autopilot with the help of the indication of the instrument, and mode I refers to the fully automatic landing mode; The plane lands.

4. Autopilot coupler

After the autopilot coupler is coupled with the autopilot, the automatic control of the trajectory movement of the aircraft is completed.

5. Auto throttle controller

The automatic throttle controller can automatically adjust the throttle to ensure that the flight angle of attack and flight speed are constant during the landing process. It uses information from angle-of-attack sensors, accelerometers, yoke displacement, and onboard signals to automatically control the motor servos connected to the engine throttles.

6. On-board radar equipment

The radar equipment on the aircraft is used to receive the Ka-band signal sent by the tracking radar, and then send the aircraft's position data to the aircraft carrier in the form of X-band signal.

According to the system structure diagram, the working mechanism of the automatic landing guidance and control system is explained. When the aircraft enters the radar interception window, the tracking radar continuously tracks the aircraft, and digitally encodes the angle information and distance information of the tracking antenna into the high-speed general-purpose computer. At the same time, the deck motion information measured by the radar stabilization platform is sent to a high-speed general-purpose computer. After data processing, the tracking information of the tracking radar eliminates the influence of the ship's roll, pitch, heading and fluctuation, thereby obtaining The precise position of the aircraft in the inertial space coordinate system. The origin of the measurement coordinate system of this inertial space is set at the expected landing point of the aircraft, the X-axis is along the centerline of the runway, and the Z-axis is along the vertical direction of the aircraft carrier.

Comparing the coordinate information in the inertial space of the aircraft with the optimized ideal trajectory stored in the high-speed general-purpose computer, two kinds of instruction information are generated:

One is the trajectory error command information, which is sent to the aircraft through the ground-air data link. The error information includes the vertical landing height error and the lateral deviation of the aircraft relative to the ship's measurement coordinate system, that is, the centerline of the flight deck. The aircraft receives the error signal and displays it to the pilot through the pointer instrument or HUD.

The second is flight control command information, or autopilot information, which is also sent to the aircraft through the ground-air data link. The guidance command of the trajectory includes longitudinal and lateral channels, which are respectively formed by calculating the longitudinal and lateral guidance errors through their respective guidance laws. Under the effect of vertical and lateral guidance commands, the flight control system continuously corrects its own flight path, so that the aircraft tries to fly according to the set ideal trajectory, that is, vertically according to the glide path of about 3.5°, and laterally according to the middle line of the runway.

The trajectories of different aircraft are stored in the high-speed general-purpose computer to meet the guidance requirements of different aircraft. The ideal trajectory stored in the high-speed general computer can be temporarily changed according to the situation. For example, it can be used for a constant glide angle approach, a land flare approach, or a large angle approach like a V/STOL aircraft, and a helicopter hover approach, etc.

In order to reduce the spread error of landing and improve the accuracy of landing, deck motion compensation is performed 12 seconds before landing, that is, the deck motion information sensed by the radar stabilization platform is introduced into a high-speed general-purpose computer for calculation of compensation information, and then calculated with the error The information is sent to the aircraft together, so that the aircraft can follow the deck movement and make corresponding maneuvers, so as to reduce the landing error caused by the deck movement.

The control method of the present invention mainly introduces the altitude change rate feedback information and the lateral deviation rate feedback information respectively in the shipboard aircraft landing guidance and control system, effectively suppresses the influence of the ship tail airflow interference, and improves the shipboard aircraft landing guidance and control system. aircraft landing accuracy. It includes a longitudinal guidance control method and a lateral guidance control method. The carrier-based aircraft mainly refers to aircraft, which are hereinafter referred to as aircraft.

(1) The longitudinal guidance control method also includes the longitudinal height guidance method and the longitudinal attitude control method;

的反馈信息，达到抑制舰尾气流扰动对飞机着舰的影响。舰尾气流是造成飞机着舰准确度的主要因素，舰尾气流的气流扰动作用飞机时，直接影响飞机的航迹倾斜角γ，由于

，

为飞机飞行实际高度变化率增量，

为飞机飞行速度，

为航迹倾斜角增量，因此，采用引入飞机飞行高度变化率

作为纵向高度引导控制系统的反馈信息，当舰尾气流作用于飞机时，由于

相当于

，所以从导引控制系统的输入信号直接控制飞机的航迹倾斜角增量

，能迅速纠正飞机飞行航迹，有效地抑制了舰尾气流扰动对飞机着舰准确度的影响。具体方法是，依据引导控制律表达式，构建纵向轨迹控制器，该轨迹控制器的引导律表达式: (1) The longitudinal altitude guidance method is to introduce the flight altitude change rate of the aircraft into the guidance law calculation sub-module of the high-speed general computer

Feedback information to suppress the impact of ship tail airflow disturbance on aircraft landing. The airflow at the ship's tail is the main factor that causes the aircraft's landing accuracy. When the airflow disturbance of the ship's tail airflow affects the aircraft, it directly affects the aircraft's track inclination angle γ, because

,

is the increment of the actual altitude change rate of the aircraft,

is the flight speed of the aircraft,

is the increment of the flight path inclination angle, therefore, the rate of change of the flight altitude of the aircraft is introduced

As the feedback information of the longitudinal altitude guidance control system, when the ship's exhaust airflow acts on the aircraft, due to

equivalent to

, so the input signal from the guidance control system directly controls the aircraft's track inclination angle increment

, can quickly correct the flight path of the aircraft, and effectively suppress the influence of the disturbance of the ship's tail airflow on the accuracy of the aircraft's landing. The specific method is to construct a longitudinal trajectory controller according to the expression of the guidance control law, and the expression of the guidance law of the trajectory controller is:

The structure of the longitudinal trajectory controller constructed by this formula is shown in Figure 2.

，

，

，

，

，以上述这些寻得值作为初值，再采用梯度下降法继续二次寻优，可得最终优化值为：

In the formula, K _P is the gain of the proportional term, K _i is the gain of the integral term, K _d is the gain of the differential term, K _dd is the gain of the second differential term, K ₀ is the total gain, H _com is the aircraft flight reference altitude command signal, H is the actual height feedback signal of the aircraft flight, s is a complex variable, where K _P , K _i , K _d , K _dd , K ₀ pass the actual height feedback signal gain of the aircraft flight △H to the aircraft flight reference height command signal gain △H _com The response can be optimized; the optimization toolbox in MATLAB can be used to optimize the multivariate parameters by using the gradient descent method to obtain the above five incremental control parameters, which can be obtained

,

,

,

,

, using the above found values as the initial values, and then using the gradient descent method to continue the second optimization, the final optimized value can be obtained as:

The values of the five incremental control parameters when there is no ship exhaust flow disturbance are:

Substituting the above-mentioned five incremental control parameter values in the absence of ship exhaust disturbance into Fig. 2, the slope response of the actual flight height feedback signal increment △H to the aircraft flight reference altitude command signal increment △H _com can be obtained as shown in the figure 3.

Substituting the optimized values of the above-mentioned five incremental control parameters into the trajectory controller in Figure 2, the slope of the actual altitude feedback signal increment △H of the aircraft flight to the aircraft flight reference altitude command signal increment △H _com can be obtained The response is shown in Figure 4.

Comparing Fig. 3 and Fig. 4, it can be seen that the guidance control system with feedback of flight height change rate has higher longitudinal landing trajectory accuracy than the traditional guidance control system with pitch angle feedback.

反馈和飞机飞行实际高度信号二次微分信号

反馈，构建纵向姿态改进型飞行控制方法，具体方法是，在高速通用计算机的导引律计算子模块中，引入飞机飞行高度变化率

反馈和飞机飞行实际高度信号二次微分信号

反馈，构建纵向姿态控制器，该纵向姿态控制器的引导控制律表达式为： (2) The longitudinal attitude control method is to add the aircraft flight altitude change rate to the control law of the traditional flight control system, which is traditionally based on attitude control.

Feedback and the second differential signal of the actual altitude signal of the aircraft

Feedback, to build a longitudinal attitude improved flight control method, the specific method is to introduce the flight altitude change rate of the aircraft into the guidance law calculation sub-module of the high-speed general computer

Feedback and the second differential signal of the actual altitude signal of the aircraft

Feedback, to construct a longitudinal attitude controller, the expression of the guidance control law of the longitudinal attitude controller is:

The longitudinal attitude controller constructed by this expression is shown in Fig. 5.

为升降舵回路传递函数，

为姿态控制参数，通过轨迹设计方法，来确定姿态控制参数，

为飞机飞行实际高度变化率增量，

为飞机飞行实际俯仰角速率增量，为飞机飞行实际高度变化率增量，为飞机飞行参考高度变化率增量，

为飞机飞行实际高度信号增量的二次微分；由于

，

，

为飞行速度，

为航迹倾斜角增量，

为参考航迹倾斜角增量，由此可知，纵向姿态改进型飞行控制方法相当于对传统飞行控制系统进行了航迹倾斜角增量

和参考航迹倾斜角增量

反馈校正，它相当于把俯仰姿态角的控制转为直接对航迹角的控制，展宽了飞行控制系统的频带，有利于对舰尾气流作用下的扰动运动，能快速纠偏。 In the formula,

is the transfer function of the elevator loop,

As the attitude control parameters, the attitude control parameters are determined by the trajectory design method,

is the increment of the actual altitude change rate of the aircraft,

is the actual pitch rate increment of the aircraft, is the increment of the actual altitude change rate of the aircraft, is the aircraft flight reference altitude change rate increment,

is the quadratic differential of the actual altitude signal increment of the aircraft; because

,

,

is the flight speed,

is the track inclination angle increment,

In order to refer to the increment of track inclination angle, it can be seen that the improved flight control method of longitudinal attitude is equivalent to the traditional flight control system.

and the reference track inclination angle increment

Feedback correction, which is equivalent to converting the control of the pitch attitude angle to the direct control of the track angle, broadens the frequency band of the flight control system, which is beneficial to the disturbance movement under the action of the ship's tail airflow, and can quickly correct the deviation.

。已知传统飞行控制系统的姿态控制器参数为

，

，通过轨迹设计方法确定的改进型纵向飞行控制系统的姿态控制器参数

，

。将上述参数值代入如图5所示的纵向姿态控制器中，得到如图6所示的传统飞行控制系统与引入高度变化率

后的改进型纵向飞行控制系统的频率响应特性比较示意图。由图6可知，改进型纵向飞行控制系统的带宽约为3.47rad/s，在2rad/s处相移为

；而传统飞行控制系统的带宽约为2.5rad/s，在2rad/s处相移为，可见改进后的飞行控制系统的带宽明显增加。 Determine the parameters of the attitude controller by trajectory design method

. It is known that the attitude controller parameters of the traditional flight control system are

,

, the attitude controller parameters of the improved longitudinal flight control system determined by the trajectory design method

,

. Substituting the above parameter values into the longitudinal attitude controller shown in Figure 5, the traditional flight control system and the introduced altitude change rate shown in Figure 6 are obtained

Schematic diagram of the comparison of the frequency response characteristics of the improved longitudinal flight control system. It can be seen from Figure 6 that the bandwidth of the improved longitudinal flight control system is about 3.47rad/s, and the phase shift at 2rad/s is

; while the bandwidth of the traditional flight control system is about 2.5rad/s, the phase shift at 2rad/s is , it can be seen that the bandwidth of the improved flight control system increases significantly.

(2) Lateral guidance control method

When an aircraft lands, under the disturbance of lateral airflow, it will have a great impact on the landing accuracy, causing the aircraft to deviate from the centerline of the runway. To this end, the lateral deviation information is introduced into the lateral flight control system to broaden the bandwidth of the lateral landing guidance system and overcome the influence of lateral airflow disturbances on aircraft deviation from the centerline of the runway when landing. Because the flight control system is composed of two channels of aileron and rudder, the aileron channel is the roll angle control system, which receives the roll angle guidance command from the aircraft carrier, and realizes the track control of the aircraft by controlling the roll of the aircraft, while the rudder The channel plays a role in coordinating the turn, trying to make the sideslip angle of the aircraft equal to zero during the roll process.

Therefore, the lateral guidance control method includes the lateral deviation guidance method and the lateral attitude control method:

(1) The lateral deviation guidance control method is to introduce the feedback information of the lateral deviation rate into the guidance law calculation sub-module of the high-speed general-purpose computer to construct the lateral trajectory controller. The guidance law of the lateral trajectory controller is The expression is:

为比例项增益、

为积分项增益、

为微分项增益、

为总增益、s为复变量、y为与甲板中心线的侧偏离、y_com为期望的与甲板中心线侧偏离指令。 Through this guidance law expression, the constructed lateral trajectory controller is shown in Fig. 7. In the formula,

is the proportional term gain,

is the integral term gain,

is the differential term gain,

is the total gain, s is a complex variable, y is the side deviation from the deck centerline, and y _com is the expected side deviation command from the deck centerline.

Since the rudder channel has sideslip angle feedback, it has a certain anti-sidewind effect. Due to the introduction of the lateral deviation rate feedback, it is equivalent to the introduction of the feedback of the track deflection angle. The control method of the lateral flight control system is changed from the control of the roll angle to the direct control of the track deflection angle, so that the performance of the lateral landing guidance system is significantly improved, as shown in Figure 8. In the figure, is the response curve of the traditional lateral landing guidance control system under the disturbance of lateral airflow, is the response curve of the lateral landing guidance control system under the disturbance of lateral airflow after introducing the lateral deviation rate. It can be seen from the figure that the control of the lateral landing guidance system under the traditional lateral flight control system will still cause obvious lateral landing deviation under the disturbance of the lateral airflow, while the lateral landing guidance system under the improved lateral flight control system The lateral landing deviation controlled under the disturbance of lateral airflow has been significantly improved.

(2) The lateral attitude control method is to introduce the feedback information of the lateral deviation rate in the high-speed general-purpose computer guidance law calculation sub-module to construct the lateral attitude controller. The control law of the lateral attitude controller is expressed as:

Through the above two formulas, the constructed lateral attitude controller is shown in Figure 9.

为滚转角增量，为滚转角速度增量，

为偏航角速度增量，

为洗出网络，

为迎角基准值，

为侧滑角增量，

为控制参数，可以通过根轨迹方法获得，

为常数，s为复变量， △y_com为期望的侧偏离速率指令。  In the two formulas,

is the roll angle increment, is the roll angular velocity increment,

is the yaw rate increment,

To wash out the network,

is the reference value of the angle of attack,

is the sideslip angle increment,

As the control parameter, it can be obtained by the root locus method,

is a constant, s is a complex variable, and △y _com is the expected lateral deviation speed command.

Claims (3)

1. a carrier-borne aircraft that suppresses the stern flow perturbation warship guiding and control system; It is characterized in that forming by guiding subsystem and RACS; The guiding subsystem is loaded on the warship; Comprise tracking radar, radar stable platform, High Speed General computing machine, display platform, digital coding transmitter, data chainning watch-dog and flight path registering instrument, wherein, High Speed General computing machine, tracking radar and flight path registering instrument are linked in sequence; Display platform, High Speed General computing machine and radar stable platform are linked in sequence; Display platform is connected with the digital coding transmitter respectively with the High Speed General computing machine, the digital coding transmitter, and data chainning watch-dog and display platform are linked in sequence; RACS loads aboard; Comprise radar equipment on robot pilot, data chainning receiver, receiver decoder, autopilot coupler, auto-throttle controller and the machine; Wherein, data chainning receiver, receiver decoder, autopilot coupler and robot pilot be linked in sequence, the auto-throttle controller is connected with robot pilot is two-way; The digital coding transmitter of guiding in the subsystem is connected through radiowave with data chainning receiver in the RACS, guides that radar equipment is connected through the Ka-band signal on the machine in tracking radar and the RACS in the subsystem.

2. the carrier-borne aircraft of inhibition stern flow perturbation according to claim 1 warship guiding and control system; It is characterized in that being provided with in the described High Speed General computing machine deck motion compensation calculations submodule, ideal trajectory submodule, trajectory error calculated signals submodule, data stabilization processing sub, aircraft dynamics information submodule and guidance law calculating sub module, wherein be connected guidance law calculating sub module and display platform and digital coding transmitter respectively through trajectory error calculated signals submodule with the two-way continuous deck motion compensation calculations submodule of radar stable platform; The input end of data stabilization processing sub links to each other with tracking radar with the radar stable platform respectively, and the output terminal of data stabilization processing sub is connected with the input end of trajectory error calculated signals submodule; The input end of guidance law calculating sub module is connected in trajectory error calculated signals submodule and aircraft dynamics information submodule, and the output terminal of guidance law calculating sub module is connected in the digital coding transmitter; The ideal trajectory submodule is connected in trajectory error calculated signals submodule.

3. the carrier-borne aircraft based on the described inhibition stern of claim 1 flow perturbation the control method of warship guiding and control system, it is characterized in that, comprises longitudinal guide control method and side direction guidance control method:

(1) said longitudinal guide control method comprises vertical elevation guidance control method and longitudinal attitude control method,

1) vertically the elevation guidance method is in the guidance law calculating sub module of High Speed General computing machine; Introducing aircraft flight altitude rate

is primary feedback; Suppress the stern flow perturbation effectively and aircraft the influence of warship; Concrete grammar is; According to guiding rule expression formula; Make up tracking controller; Realize the feedback of aircraft flight altitude rate

; Reach and suppress the influence that the stern flow perturbation warship to aircraft, the expression formula of the guiding rule of this tracking controller is:

In the formula, K _PBe proportional term gain, K _iBe integral term gain, K _dBe differential term gain, K _DdBe the gain of second differential item, K ₀Be full gain, H _ComBe aircraft flight reference altitude command signal, H is an aircraft flight true altitude feedback signal, and s is complex variable, wherein K _P, K _i, K _d, K _Dd, K ₀True altitude feedback signal through aircraft flight gains △ H to aircraft flight reference altitude command signal gain △ H _ComResponse carry out optimizing and obtain;

2) the longitudinal attitude control method is in the control law that is controlled to be main traditional flight control system traditionally with attitude, to add aircraft flight altitude rate

feedback and aircraft flight true altitude signal second differential signal

feedback; Make up the longitudinal attitude controller, the control law expression formula of this longitudinal attitude controller is:

In the formula;

is the elevating rudder return transfer function;

is the attitude controlled variable; Through the root locus method for designing; Confirm the attitude controlled variable;

is aircraft flight true altitude rate of change increment;

is the actual angle of pitch rate increment of aircraft flight;

is aircraft flight true altitude rate of change increment;

is aircraft flight reference altitude rate of change increment, and

is the second differential of aircraft flight true altitude signal increment;

(2) said side direction guidance control method comprises lateral deviation bootstrap technique and side direction attitude control method:

(1) the lateral deviation bootstrap technique is in the sub-computing module of the guidance law of High Speed General computing machine, introduces the feedback information of lateral deviation speed, makes up the side track controller, and the expression formula of the guiding rule of this side track controller is:

In the formula,

For proportional term gain,

For integral term gain,

For differential term gain,

For full gain, s are that complex variable, y are for the lateral deviation with the center deck line leaves, y _ComFor instructing leaving of expectation with center deck line lateral deviation;

(2) the side direction attitude control method is in High Speed General computer-guided rule calculating sub module; Introduce the feedback information of lateral deviation speed ; Make up the side direction attitude controller, the expression formula of the control law of this side direction attitude controller:

In the formula,

Be the roll angle increment, Be the angular velocity in roll increment,

Be the yaw rate increment, For washing out network,

Be angle of attack reference value, Be the yaw angle increment,

Be controlled variable, can obtain through the root locus method,

Be constant, s is a complex variable, △ y _ComFor the lateral deviation of expectation is instructed from speed.

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