A Missile Rolling Attitude Control Method with Integrator Dynamic Limiting

Abstract

A missile rolling attitude control method with integrator dynamic limiting, belonging to the field of flight control technology. The present invention includes: a guidance command calculation unit calculates the rolling command of the missile and sends it to the rolling attitude controller, and a sensitive element unit measures the rolling angular velocity of the missile and sends it to the rolling attitude controller; the rolling attitude controller calculates the upper and lower bounds of the integrator output value according to the maximum available rudder deflection; limits the integrator output according to the upper and lower bounds; calculates the variable structure controller output according to the information after the limiting process; calculates the rolling rudder command according to the variable structure controller output and the rolling angular velocity; the rolling rudder command is sent to the rudder system to control the movement of the missile. The present invention solves the problem that when the missile rolls at a large angular velocity, due to the too large integrator output value, the rolling rudder deflection is in the limited state for a long time, and the rudder deflection needs a long time to get out of the saturation area, resulting in a large amplitude oscillation of the missile rolling angular velocity and unstable attitude.

Description

Technical Field

[0001] The present invention relates to a missile rolling attitude control method with an integrator dynamic limiter, belonging to the technical field of flight control. Background Art

[0002] Missile rolling attitude control means that after receiving a rolling attitude command, the rudder command is calculated by a rolling attitude controller and sent to the rudder system. The rudder system controls the deflection of the rudder surface to generate a control moment, thereby controlling the missile to roll to the desired attitude. By integrating the missile rolling angular velocity output by the sensitive element, the missile rolling attitude angle is obtained. After subtracting it from the rolling attitude command, it is input into the rolling attitude controller for the calculation of the rudder command. If the missile is subjected to a large interference during flight, resulting in the missile rolling at a large angular velocity, the rudder command output by the missile attitude controller will increase sharply and may reach the upper or lower limit of the rudder system. Thereafter, if the interference continues to exist, the rudder command will continue to increase, but the deflection angle of the rudder surface paid by the rudder system will no longer increase and will remain at the upper or lower limit position of the rudder system. At this time, the missile attitude control system exceeds the normal operating range and enters the saturation area. When the interference disappears, under the control of the rudder system, the missile rolling angular velocity gradually decreases. When the rolling angular velocity increases in the reverse direction, the output of the integrator begins to decrease, and the missile attitude control system gradually exits the saturation area. When the interference exists, the larger the value of the integrator, the longer the time to exit the saturation area. During the time of exiting the saturation area, the rudder surface remains at the limit position of the rudder system and cannot immediately make corresponding changes as the rolling direction of the missile changes, which will cause the rolling angular velocity to oscillate and even the missile to get out of control.

[0003] Currently, the most commonly used methods in industry to overcome the integral saturation phenomenon mainly include conditional integration method, integral separation method, and variable speed integration method. Among them, the conditional integration method is the most intuitive method. This method has a switch link. When the system controller has integral saturation, the integral action is stopped or limited. This method is straightforward and easy to implement, but its parameter selection is only for a certain fixed state. Once the system parameters of the controlled object change, this method cannot play a role in suppressing integral saturation. In the method provided by the present invention, the limited value output by the integrator changes dynamically according to the missile state. When the missile system parameters change, it can still play a role in suppressing integral saturation.

[0004] The integral separation method means that when the deviation between the controlled variable and the set value is large, the integral action is cancelled to prevent the system stability from decreasing and the overshoot from increasing due to the integral action. When the controlled variable approaches the set value, integral control is introduced to eliminate the static error and improve the control accuracy. The variable-speed integral method means changing the accumulation speed of the integral term to make it correspond to the deviation magnitude. When the deviation is large, the integral accumulation speed is slow; when the deviation is small, the integral accumulation speed is fast, and the integral action can change dynamically with the change of the system deviation. Both of these methods process the output of the integrator when it is not saturated. The output value of the integrator in the missile attitude controller reflects the rolling attitude of the missile. If the output of the integrator is changed when it is not saturated, the output of the integrator cannot truly reflect the rolling attitude, thus affecting the rolling attitude control accuracy. Therefore, these two methods cannot be used in this rolling attitude control method. Summary of the Invention

[0005] The technical problem solved by the present invention is to overcome the deficiencies of the prior art and provide a missile rolling attitude control method with dynamic amplitude limiting of the integrator, enabling the missile to quickly exit the saturation region when it is subjected to large disturbances and the integrator in the rolling attitude controller is saturated, promptly generating the corresponding rudder deflection, reducing the oscillation of the missile rolling angular velocity, and even the risk of missile out-of-control.

[0006] The technical solution of the present invention is: a missile rolling attitude control method with dynamic amplitude limiting of the integrator, where the missile includes a guidance command resolution unit, a sensitive element unit, and a rudder system, and includes:

[0007] Step 1: The guidance command resolution unit calculates the rolling command of the missile and sends it to the rolling attitude controller, and the sensitive element unit measures the rolling angular velocity of the missile and sends it to the rolling attitude controller;

[0008] Step 2: The rolling attitude controller calculates the upper and lower bounds of the output value of the integrator based on the maximum available rudder deflection of the missile and the rolling angular velocity;

[0009] Step 3: Limit the output of the integrator according to the upper and lower bounds of the output value of the integrator calculated in Step 2;

[0010] Step 4: Calculate the output of the variable structure controller based on the output value of the integrator after amplitude limiting processing in Step 3 and information such as the rolling command and the rolling angular velocity;

[0011] Step 5: Calculate the rolling rudder command based on the output of the variable structure controller calculated in Step 4 and the rolling angular velocity;

[0012] Step 6: Send the rolling rudder command to the rudder system to control the movement of the missile.

[0013] Further, the calculation formulas for the upper and lower bounds of the output value of the integrator in Step 2 are:

[0014]

[0015] Among them, Wx is the rolling angular velocity, Gma is the integrator output value, Gmac is the rolling command, and U Rmax is the preset maximum available rudder deflection of the missile, and a1, a2, β, g1, μ, Kr1, and Kr2 are preset fixed gain values, and Er and Ur bjg are intermediate calculation variables, and Ur low is the upper bound of the integrator output, and Ur high is the lower bound of the integrator output.

[0016] Furthermore, the formula for limiting the integrator output in step three is:

[0017]

[0018] Furthermore, the calculation formula for the output of the variable structure controller in step four:

[0019]

[0020] Er = Gma - Gmac

[0021]

[0022] Among them, Br is the output of the variable structure controller.

[0023] Furthermore, the calculation formula for the rolling rudder command in step five is:

[0024] Qr = Kr1 × Br

[0025] Zr = Kr2 × Wx

[0026] Dr = -Qr - Zr

[0027] Among them, Qr and Zr are intermediate calculation variables, and Dr is the rolling rudder command.

[0028] A missile rolling attitude control system with integrator dynamic amplitude limiting includes: a guidance command calculation unit, a sensitive element unit, and a rudder system; the guidance command calculation unit calculates the rolling command of the missile and sends it to the rolling attitude controller, and the sensitive element unit measures the angular velocity of the missile and sends it to the rolling attitude controller. The rolling attitude controller calculates the rolling rudder command based on the rolling command and the rolling angular velocity; the rolling attitude controller sends the rolling rudder command to the rudder system, and the rudder system controls the movement of the missile, which is used to perform a missile rolling attitude control method with integrator dynamic amplitude limiting.

[0029] A computer-readable storage medium stores a computer program, and when the computer program is executed by a processor, the steps of the missile rolling attitude control method with an integrator dynamic limit are implemented.

[0030] A missile rolling attitude control device with an integrator dynamic limit includes a memory, a processor, and a computer program stored in the memory and executable on the processor. When the processor executes the computer program, the steps of the missile rolling attitude control method with an integrator dynamic limit are implemented.

[0031] The advantages of the present invention compared with the prior art are as follows:

[0032] By adding an integrator limit to the missile rolling attitude controller, the problem that when the missile rolls at a large angular velocity, due to the excessive output value of the integrator, the rolling rudder deflection is in the limit state for a long time, and the rudder deflection needs a long time to get out of the saturation area, resulting in a large oscillation of the missile rolling angular velocity and unstable attitude is solved. At the same time, the integrator limit method is a dynamic limit, and the upper and lower limits of the integrator output are calculated in real time according to the missile state, which can play a role in suppressing integral saturation in different flight states of the missile. The calculation of the present invention is simple, the performance is stable, and it is easy to be implemented in engineering. BRIEF DESCRIPTION OF THE DRAWINGS

[0033] Figure 1 is a schematic flow chart of the missile rolling attitude control method considering integrator dynamic limit provided by the present invention;

[0034] Figure 2 is a schematic principle diagram of the missile rolling attitude control method considering integrator dynamic limit provided by the present invention. DETAILED DESCRIPTION OF THE INVENTION

[0035] In order to better understand the above technical solutions, the technical solutions of the present application will be described in detail below through the accompanying drawings and specific embodiments. It should be understood that the specific features in the embodiments of the present application and the embodiments are detailed descriptions of the technical solutions of the present application, rather than limitations on the technical solutions of the present application. Without conflict, the technical features in the embodiments of the present application and the embodiments can be combined with each other.

[0036] The following further describes in detail a missile rolling attitude control method with an integrator dynamic limit provided by the embodiments of the present application with reference to the accompanying drawings of the specification. The specific implementation manner may include (as Figures 1-2 shown):

[0037] Step 1: The guidance command resolution unit calculates the rolling command of the missile and sends it to the rolling attitude controller, and the sensitive element unit measures the rolling angular velocity of the missile and sends it to the rolling attitude controller;

[0038] Step 2: The roll attitude controller calculates the upper and lower bounds of the integrator output value based on the maximum available rudder deflection of the missile and the roll angular velocity;

[0039] Step 3: Limit the integrator output according to the upper and lower bounds of the integrator output value calculated in Step 2;

[0040] Step 4: Calculate the variable structure controller output based on the integrator output value after amplitude limiting in Step 3 and information such as the roll command and the roll angular velocity;

[0041] Step 5: Calculate the roll rudder command based on the variable structure controller output calculated in Step 4 and the roll angular velocity;

[0042] Step 6: Send the roll rudder command to the rudder system to control the movement of the missile.

[0043] Furthermore, the calculation formulas for the upper and lower bounds of the integrator output value in Step 2 are:

[0044] Gma = ∫Wx

[0045] Er = Gma - Gmac

[0046]

[0047] where Wx is the roll angular velocity, Gma is the integrator output value, Gmac is the roll command, U [[ID=2⑨]] Rmax is the maximum available rudder deflection of the missile, is a preset fixed value, a1, a2, β, g1, μ, Kr1, and Kr2 are preset fixed value gains, Er and Ur bjg are intermediate calculation variables, Ur low is the upper bound of the integrator output, Ur high is the lower bound of the integrator output.

[0048] Furthermore, the formula for amplitude limiting the integrator output in Step 3 is:

[0049]

[0050] Furthermore, the calculation formula for the variable structure controller output in Step 4:

[0051]

[0052] Er = Gma - Gmac

[0053]

[0054] where Br is the variable structure controller output.

[0055] Further, the calculation formula for the rolling rudder command in Step 5 is as follows:

[0056] Qr = Kr1 × Br

[0057] Zr = Kr2 × Wx

[0058] Dr = -Qr - Zr

[0059] Where Qr and Zr are intermediate calculation variables, and Dr is the rolling rudder command.

[0060] The present invention also provides a guidance command calculation unit, a sensitive element unit, and a rudder system for implementing the missile rolling attitude control method with integrator dynamic limiting. The guidance command calculation unit calculates the rolling command of the missile and sends it to the rolling attitude controller. The sensitive element unit measures the angular velocity of the missile and sends it to the rolling attitude controller. The rolling attitude controller calculates the rolling rudder command based on the rolling command and the rolling angular velocity. The rolling attitude controller sends the rolling rudder command to the rudder system, and the rudder system controls the movement of the missile.

[0061] In the solution provided in the embodiment of the present application, with reference to Figure 1 , Step 1: The guidance command calculation unit calculates the rolling command of the missile and sends it to the rolling attitude controller. The sensitive element unit measures the rolling angular velocity of the missile and sends it to the rolling attitude controller. Step 2: The rolling attitude controller calculates the upper and lower bounds of the integrator output value based on the maximum available rudder deflection. Step 3: Limit the integrator output according to the upper and lower bounds of the integrator output value calculated in Step 2. Step 4: Calculate the variable structure controller output based on the integrator output value after the limit processing in Step 3, as well as information such as the rolling command and the rolling angular velocity. Step 5: Calculate the rolling rudder command based on the variable structure controller output calculated in Step 4 and the rolling angular velocity. Step 6: Send the rolling rudder command to the rudder system to control the movement of the missile.

[0062] In an embodiment of the present invention, the gains a1, a2, β, g1, μ, Kr1, and Kr2 are determined in advance according to the flight state. For example, when the missile speed is 1000 m / s and the altitude is 5 km, a1 = 30, a2 = 1, β = 0.1, g1 = 200, μ = 0.5, Kr1 = 0.1, and Kr2 = 0.05.

[0063] In this embodiment, the specific working process is described as follows:

[0064] Calculate the upper and lower bounds of the integrator output value according to the following formula:

[0065] Gma = ∫Wx

[0066] Er = Gma - Gmac

[0067]

[0068] Among them, Wx is the rolling angular velocity, Gma is the output value of the integrator, Gmac is the rolling command, and U Rmax is the maximum available rudder deflection of the missile, which is a preset fixed value. a1, a2, β, g1, μ, Kr1, and Kr2 are preset fixed value gains. Er and Ur bjg are intermediate calculation variables, and Ur low is the upper bound of the output of the integrator, and Ur high is the lower bound of the output of the integrator.

[0069] The output of the integrator is limited according to the following formula:

[0070]

[0071] The output of the variable structure controller is calculated according to the following formula:

[0072]

[0073] Among them, Br is the output of the variable structure controller.

[0074] The rolling rudder command is calculated according to the following formula:

[0075] Qr = Kr1 × Br

[0076] Zr = Kr2 × Wx

[0077] Dr = -Qr - Zr

[0078] Among them, Qr and Zr are intermediate calculation variables, and Dr is the rolling rudder command.

[0079] In the present invention, by adding integrator limiting in the missile rolling attitude controller, the problem that when the missile rolls at a large angular velocity, due to the too large output value of the integrator, the rolling rudder deflection is in the limited state for a long time, and the rudder deflection needs a long time to get out of the saturation area, resulting in a large oscillation of the missile rolling angular velocity and unstable attitude is solved.

[0080] This application provides a computer-readable storage medium, and the computer-readable storage medium stores computer instructions. When the computer instructions run on a computer, the computer is made to execute Figure 1 the method described above.

[0081] Those skilled in the art should understand that the embodiments of the present application can be provided as a method, a system, or a computer program product. Therefore, the present application can take the form of a complete hardware embodiment, a complete software embodiment, or an embodiment combining software and hardware aspects. Moreover, the present application can take the form of a computer program product implemented on one or more computer-usable storage media (including but not limited to disk memories and optical memories, etc.) that contain computer-usable program code.

[0082] The present application is described with reference to the flowcharts and / or block diagrams of methods, apparatuses (systems), and computer program products according to the embodiments of the present application. It should be understood that each flow and / or block in the flowchart and / or block diagram, as well as the combination of flows and / or blocks in the flowchart and / or block diagram, can be realized by computer program instructions. These computer program instructions can be provided to the processor of a general-purpose computer, a special-purpose computer, an embedded processor, or other programmable data processing devices to generate a machine, such that the instructions executed by the processor of the computer or other programmable data processing devices generate means for realizing the functions specified in Figure 1 one flow or multiple flows and / or blocks Figure 1 one block or multiple blocks.

[0083] These computer program instructions can also be stored in a computer-readable memory that can direct a computer or other programmable data processing device to work in a specific manner, such that the instructions stored in the computer-readable memory generate a manufactured article including instruction means that realizes the functions specified in Figure 1 one flow or multiple flows and / or blocks Figure 1 one block or multiple blocks.

[0084] These computer program instructions can also be loaded onto a computer or other programmable data processing device, such that a series of operation steps are executed on the computer or other programmable device to generate a computer-implemented process, so that the instructions executed on the computer or other programmable device provide steps for realizing the functions specified in Figure 1 one flow or multiple flows and / or blocks Figure 1 one block or multiple blocks.

[0085] It should be noted that in this text, relational terms such as first and second are only used to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any actual relationship or order between these entities or operations. Moreover, the term "comprising", "including" or any other variant thereof is intended to cover non-exclusive inclusion, so that a process, method, article or device comprising a series of elements not only includes those elements, but also includes other elements not expressly listed, or elements inherent to such process, method, article or device.

[0086] Obviously, those skilled in the art can make various changes and modifications to this application without departing from the spirit and scope of this application. Thus, if these modifications and variations of this application fall within the scope of the claims of this application and their equivalent technologies, this application is also intended to include these modifications and variations.

[0087] The content not described in detail in the specification of the present invention belongs to the well-known technology of those skilled in the art.

Claims

1. A missile rolling attitude control method with an integrator dynamic amplitude limiting is provided. The missile includes a guidance command resolution unit, a sensitive element unit, and a rudder system, and is characterized in that Including: Step 1: The guidance command resolution unit calculates the rolling command of the missile and sends it to the rolling attitude controller. The sensitive element unit measures the rolling angular velocity of the missile and sends it to the rolling attitude controller; Step 2: The rolling attitude controller calculates the upper and lower bounds of the integrator output value based on the maximum available rudder deflection of the missile and the rolling angular velocity; Step 3: Limit the integrator output according to the upper and lower bounds of the integrator output value calculated in Step 2; Step 4: Calculate the variable structure controller output based on the integrator output value after amplitude limiting in Step 3, as well as information such as the rolling command and rolling angular velocity; Step 5: Calculate the rolling rudder command based on the variable structure controller output calculated in Step 4 and the rolling angular velocity; Step 6: Send the rolling rudder command to the rudder system to control the movement of the missile; The formulas for the upper and lower bounds of the integrator output value in Step 2 are: Gma = ∫Wx Er = Gma - Gmac Among them, Wx is the rolling angular velocity, Gma is the integrator output value, Gmac is the rolling command, and U Rmax is the preset maximum available rudder deflection of the missile, a1, a2, β, g1, μ, Kr1, and Kr2 are preset constant gains, Er and Ur bjg are intermediate calculation variables, and Ur low is the upper bound of the integrator output, and Ur high is the lower bound of the integrator output.

2. A missile rolling attitude control method with integrator dynamic amplitude limiting according to claim 1, characterized in that, The formula for amplitude limiting the integrator output in Step 3 is:

3. A missile rolling attitude control method with integrator dynamic amplitude limiting according to claim 1, characterized in that The formula for calculating the variable structure controller output in Step 4: Er = Gma - Gmac where Br is the variable structure controller output.

4. A missile rolling attitude control method with integrator dynamic amplitude limiting according to claim 3, characterized in that, The formula for calculating the rolling rudder command in Step 5 is: Qr = Kr1 × Br Zr = Kr2 × Wx Dr = -Qr - Zr where Qr and Zr are intermediate variables in the calculation, and Dr is the rolling rudder command.

5. A missile rolling attitude control system with an integrator dynamic amplitude limiter, comprising: The guidance command resolution unit, the sensitive element unit, the rudder system; The guidance command resolution unit calculates the rolling command of the missile and sends it to the rolling attitude controller. The sensitive element unit measures the angular velocity of the missile and sends it to the rolling attitude controller. The rolling attitude controller calculates the rolling rudder command based on the rolling command and the rolling angular velocity; The rolling attitude controller sends the rolling rudder command to the rudder system, and the rudder system controls the movement of the missile. It is characterized in that it is used to perform a missile rolling attitude control method with integrator dynamic amplitude limiting provided in any one of claims 1 to 4.

6. A computer-readable storage medium storing a computer program, characterized in that, When the computer program is executed by a processor, it implements the steps of the method described in any one of claims 1 to 4.

7. A missile rolling attitude control device with an integrator dynamic amplitude limiting, comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, characterized in that: When the processor executes the computer program, it implements the steps of the method described in any one of claims 1 to 4.

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