Missile large static instability control method based on swing nozzle

Abstract

The invention relates to a missile large static instability control method based on a swing nozzle. The swing nozzle is adopted to control a pitching channel and a yaw channel, and an air rudder is adopted to control a rolling channel; the control method of the rolling channel comprises the steps that a rolling channel control instruction is obtained through a variable structure control method, and the rolling channel control instruction is output to the rudder system after being subjected to amplitude limiting; according to the thrust obtained through online identification, the missile flight process is divided into three stages including a thrust establishing stage, a thrust stabilizing stage and a thrust ending stage; in three stages of a missile flight process, a composite structure method is adopted, a pitching channel control instruction and a yaw channel control instruction are calculated, and the pitching channel control instruction and the yaw channel control instruction are output to a rudder system after being subjected to filtering and amplitude limiting; the control parameters of the thrust establishing section and the thrust stabilizing section are the same; and after entering the thrust ending section, correcting control parameters of the thrust establishing section and the thrust stabilizing section, and eliminating control errors caused by thrust uncertainty of the thrust ending section.

Description

Technical Field

[0001] The invention belongs to the technical field of tactical weapon flight control, and in particular relates to a missile large static instability control method based on a swing nozzle. Background Art

[0002] The center of mass of a missile that uses a swinging nozzle as an actuator is further back than that of an ordinary missile or launch vehicle. The static instability of the missile may be above 6%, or even more than 20%. Therefore, the pitch channel has large static instability and weak control characteristics. In view of the above characteristics, taking a launch vehicle as an example, the load reduction control technology + PID control is usually used to control the rocket attitude. Introducing overload feedback for the statically indeterminate rocket body is equivalent to increasing the aerodynamic torque coefficient, and the rocket body state can be changed from statically unstable to statically stable. This method has been successfully used in engineering practice, but it also has the following disadvantages:

[0003] (1) The load reduction effect and attitude stability control are contradictory;

[0004] (2) It requires multiple sets of inertial measurement equipment to obtain the flight status of the rocket body, which is not suitable for missile control;

[0005] (3) This control scheme has poor robustness and is suitable for launch vehicles with relatively fixed trajectories, but not for missiles with larger flight envelopes.

[0006] In addition, missiles usually use solid rocket engines, and the engine working time and thrust size have large discrete characteristics under high and low temperature working conditions; the control torque drops sharply in the thrust tail section, affecting the separation interference suppression performance; the aileron rudder effect is low in the subsonic stage, which puts higher requirements on the initial interference suppression and roll control of the roll. Summary of the invention

[0007] The technical problem solved by the present invention is: to overcome the shortcomings of the prior art, to provide a large static instability control method for a missile with a swing nozzle, to solve the control problems of large static instability, weak control stability ratio, critical static stability separation and rolling disturbance suppression under uncertainty, and to improve control accuracy.

[0008] The technical solution of the present invention is: a method for controlling the large static instability of a missile based on a swing nozzle, the method comprising the following steps: using the swing nozzle to control the pitch channel and the yaw channel, and using the air rudder to control the roll channel;

[0009] The control method of the rolling channel is as follows: a variable structure control method is used to obtain a rolling channel control command, and the rolling channel control command is output to the rudder system after being limited;

[0010] The pitch channel and yaw channel control method is as follows: according to the thrust obtained by online identification, the missile flight process is divided into three stages: thrust establishment stage, thrust stabilization stage and thrust termination stage; in the three stages of the missile flight process, a composite structure method is used to calculate the pitch channel control instructions and the yaw channel control instructions, and the pitch channel control instructions and the yaw channel control instructions are output to the rudder system after filtering and limiting; the control parameters of the thrust establishment stage and the thrust stabilization stage are the same; after entering the thrust termination stage, the control parameters of the thrust establishment stage and the thrust stabilization stage are corrected to eliminate the control error caused by the thrust uncertainty in the thrust termination stage.

[0011] Preferably, the scroll channel control instruction is:

[0012] G R =g3×S(Δγ,ω x )+g2×ω x

[0013] Among them, Δγ is the roll angle command deviation, ω x is the rolling angular velocity, S(Δγ,ω x ) is the transfer function of the variable structure controller, g2 and g3 are the control parameters of the roll channel;

[0014] S(Δγ,ω x )=(α1|Δγ|+α2|ω x +β|)×sign(g1Δγ+ω x )

[0015] Wherein, α1, α2, β, and g1 are constants greater than zero.

[0016] Preferably, the pitch channel and the yaw channel are recorded as controlled channels. In the thrust establishment phase and the thrust stabilization phase, the control instruction G of the controlled channel is p for:

[0017]

[0018] Among them, N yc is the overload instruction of the controlled channel; N y is the real-time overload of the controlled channel; ω z is the angular velocity of the controlled channel; K i ,K g ,K sf ,K a is the control parameter of the controlled channel; t is the flight time of the missile.

[0019] Preferably, in the thrust establishment stage and the thrust stabilization stage, the control parameters of the controlled channel are

[0020]

[0021] in,

[0022]

[0023] f sf1 、f i1 、f g1 、f sf2 、f g2 Determine from the following table:

[0024] Composite structure method control parameter table

[0025] MA1 MA2 MA3 MA4 <![CDATA[f sf1 ]]> <![CDATA[f1(x 11 ,y 11 ,z 11 )]]> <![CDATA[f1(x 12 ,y 12 ,z 12 )]]> <![CDATA[f1(x 13 ,y 13 ,z 13 )]]> <![CDATA[f1(x 14 ,y 14 ,z 14 )]]> <![CDATA[f i1 ]]> <![CDATA[f1(x 21 ,y 21 ,z 21 )]]> <![CDATA[f1(x 22 ,y 22 ,z 22 )]]> <![CDATA[f1(x 23 ,y 23 ,z 23 )]]> <![CDATA[f1(x 24 ,y 24 ,z 24 )]]> <![CDATA[f g1 ]]> <![CDATA[f1(x 31 ,y 31 ,z 31 )]]> <![CDATA[f1(x 32 ,y 32 ,z 32 )]]> <![CDATA[f1(x 31 ,y 31 ,z 31 )]]> <![CDATA[f1(x 34 ,y 34 ,z 34 )]]>

[0026] f1(x ij ,y ij ,z ij )=(x ij ×Ma+y ij )×(1+z ij ×ALP) / Q

[0027] Preferably, in the thrust tailing section, the controlled channel control parameters are:

[0028]

[0029]

[0030] Among them, N3, N4, M3, and M4 are constants.

[0031] Preferably, the thrust building phase, the stabilization phase, and the tail phase are divided according to time:

[0032] 0≤t≤Ta is the thrust establishment stage, Ta≤t≤Tb is the thrust stabilization stage, and t≥Tb is the thrust tail stage;

[0033] Ta is the moment when the thrust identification estimated value Thrust rises to 0.9×thrust; thrust is the thrust stable value; Tb is the moment when the thrust identification estimated value Thrust drops to 0.9×thrust.

[0034] Preferably, the thrust stability value thrust is determined by the following method:

[0035] Obtain the curve F of ground test thrust changing with time under high temperature conditions, normal temperature conditions, and low temperature conditions th_h (t), F th_n (t), F th_l (t);

[0036] The thrust identification estimated value Thrust is compared with the ground test thrust under high temperature conditions, normal temperature conditions and low temperature conditions at the corresponding missile flight time, and the value closest to the thrust identification estimated value Thrust is determined as the thrust stability value thrust.

[0037] Preferably, the thrust identification estimate is:

[0038] Thrust = N x -F x

[0039] N x is the real-time axial overload, F x is the aerodynamic resistance.

[0040] Preferably, the aerodynamic resistance F x Related to altitude, Mach number, and resultant angle of attack ALP.

[0041] The beneficial effects of the present invention compared with the prior art are:

[0042] (1) The present invention is based on a missile that uses a swing nozzle as an actuator and its large static instability and weak control characteristics, and selects a control method of "swing nozzle control pitch channel + air rudder control roll channel" to adapt to the control problems caused by the large static instability and weak control characteristics and the sharp decrease in control torque in the thrust trailing stage, thereby improving the control performance of the roll channel in the subsonic stage;

[0043] (2) The present invention aims to solve the problem of rolling interference suppression under strong uncertainty. The rolling loop adopts a variable structure control method based on the inner loop of the roll angular velocity instruction, which effectively suppresses the roll angular velocity interference on the basis of stability, ensures three-channel decoupling, improves the robustness of the control system, and has a strong ability to suppress oblique blowing interference;

[0044] (3) The present invention addresses the problem of large discreteness of engine thrust at high and low temperatures, and adopts thrust online identification technology to identify engine thrust, thereby suppressing the fitting error of swing jet control efficiency caused by thrust discreteness;

[0045] (4) The pitch channel of the present invention adopts a composite control loop structure to improve the adaptability to static instability changes, and the control parameters of the stable control system are designed by compromise through bandwidth coordination technology, thereby improving the robustness of the control system and having the adaptability to static instability changes of the missile;

[0046] (5) The present invention adopts the amplitude-phase correction technology and the adaptive control technology to design the filter, which can suppress the resonance and weaken the influence of high-frequency vibration on the performance of the swing-jet servo mechanism;

[0047] (6) The present invention obtains the real-time thrust of the engine through the thrust online identification technology, and compensates the pitch control channel control system, which can reduce the discreteness of the attitude angle and attitude angular velocity during the thrust tailing process before the two-stage separation, and solves the problem of decreased control ability in the thrust tailing section. BRIEF DESCRIPTION OF THE DRAWINGS

[0048] Figure 1 The figure is a schematic diagram of the structure of a rolling channel control system provided in an embodiment of the present invention.

[0049] Figure 2 This is a schematic diagram of the structure of a pitch channel control system provided in an embodiment of the present invention. DETAILED DESCRIPTION

[0050] The present invention is further described in detail below with reference to the accompanying drawings and specific implementation modes.

[0051] like Figure 1 , Figure 2 As shown, the present invention provides a method for controlling the large static instability of a missile with a swing nozzle. Based on a missile using a swing nozzle as an actuator and its large static instability weak control characteristics, a control strategy of a swing nozzle controlling a pitch channel and an air rudder controlling a roll channel is adopted. The strategy can effectively suppress the interference of the roll angular velocity and ensure the decoupling of the three channels, ensure sufficient stability of rigid body motion and elastic motion, weaken the influence of high-frequency vibration on the performance of the swing nozzle servo mechanism, and reduce the discreteness of the attitude angle and attitude angular velocity in the thrust tailing process. The implementation steps are:

[0052] The pitch channel and yaw channel are controlled by a swing nozzle, and the roll channel is controlled by an air rudder;

[0053] The control method of the rolling channel is as follows: a variable structure control method is used to obtain a rolling channel control command, and the rolling channel control command is output to the rudder system after being limited;

[0054] The pitch channel and yaw channel control method is as follows: according to the thrust obtained by online identification, the missile flight process is divided into three stages: thrust establishment stage, thrust stabilization stage and thrust termination stage; in the three stages of the missile flight process, a composite structure method is used to calculate the pitch channel control instructions and the yaw channel control instructions, and the pitch channel control instructions and the yaw channel control instructions are output to the rudder system after filtering and limiting; the control parameters of the thrust establishment stage and the thrust stabilization stage are the same; after entering the thrust termination stage, the control parameters of the thrust establishment stage and the thrust stabilization stage are corrected to eliminate the control error caused by the thrust uncertainty in the thrust termination stage.

[0055] The present invention is described in detail from the following aspects:

[0056] 1. Roll channel control design; 2. Thrust online identification; 3. Pitch channel control design; 4. Filter; 5. Thrust trailing section control compensation.

[0057] The flight state quantities involved are: Ma is the flight Mach number; H is the flight altitude; ALP is the synthetic angle of attack; Vm is the flight speed; Q is the dynamic pressure; N x Real-time axial overload feedback, etc.

[0058] 1. Roll channel control design

[0059] like Figure 1 The figure shows a large static instability control strategy for a missile with a swinging nozzle, in which a variable structure control method is adopted for the roll channel.

[0060] Roll channel control command G R for:

[0061] G R =g3×S(Δγ,ω x )+g2×ω x

[0062] Among them, Δγ is the roll angle command deviation, ω x Roll angular velocity, S(Δγ,ω x ) is the transfer function of the variable structure controller, g2 and g3 are the control parameters of the roll channel;

[0063] S(Δγ,ω x )=(α1|Δγ|+α2|ω x +β|)×sign(g1Δγ+ω x )

[0064] Wherein α1, α2, β, and g1 are constants greater than zero.

[0065] As shown in Table 1, the control parameters are obtained according to the linear interpolation of Mach number (MA1, MA2, MA3, MA4) according to Table 1, where a ij ,b ij It is related to the aerodynamic characteristics and index requirements of the missile, and technical personnel in this field can design it according to actual conditions and needs.

[0066] g(a ij ,b ij )=(a ij ×Ma+b ij ) / Q

[0067] Table 1 Control parameters at characteristic points

[0068] MA1 MA2 MA3 MA4 <![CDATA[g2]]> <![CDATA[g(a 11 ,b 11 )]]> <![CDATA[g(a 12 ,b 12 )]]> <![CDATA[g(a 13 ,b 13 )]]> <![CDATA[g(a 14 ,b 14 )]]> <![CDATA[g3]]> <![CDATA[g(a 21 ,b 21 )]]> <![CDATA[g(a 22 ,b 22 )]]> <![CDATA[g(a 23 ,b 23 )]]> <![CDATA[g(a 24 ,b 24 )]]>

[0069] 2. Thrust Online Identification

[0070] According to the ground test data, the thrust is divided into the following types according to time:

[0071] 0≤t≤Ta is the thrust establishment stage, Ta≤t≤Tb is the thrust stabilization stage, and t≥Tb is the thrust tail stage;

[0072] Ta is the moment when Thrust rises to 0.9×thrust; Tb is the moment when Thrust drops to 0.9×thrust. thrust is the thrust stability value; Tb is the moment when the thrust identification estimate value Thrust drops to 0.9×thrust.

[0073] The thrust stability value thrust is determined by the following method:

[0074] Obtain the curve F of ground test thrust changing with time under high temperature conditions, normal temperature conditions, and low temperature conditions th_h (t), F th_n (t), F th_l (t);

[0075] The thrust identification estimated value Thrust is compared with the ground test thrust under high temperature conditions, normal temperature conditions and low temperature conditions at the corresponding missile flight time, and the value closest to the thrust identification estimated value Thrust is determined as the thrust stability value thrust.

[0076] Based on the ground aerodynamic test data, the aerodynamic drag is estimated using the missile flight state obtained in real time, and finally the thrust identification is completed. The thrust identification estimate value Thrust is:

[0077] Thrust=Nx-F x

[0078] Among them, F x is the aerodynamic drag, and is related to the aerodynamic drag F x , height H, Mach number Ma, synthetic angle of attack ALP related variables, the parameter table is shown in Table 2, the parameters are obtained according to the linear interpolation of the height (H1, H2, H3, H4) according to Table 2, where n i ,m i It is related to the aerodynamic drag characteristics of the missile, and technical personnel in this field can design it according to actual conditions and needs.

[0079] f x (n i ,m i )=n i ×Ma+m i ×ALP

[0080] Table 2 Parameters of feature points

[0081] H1 H2 H3 H4 <![CDATA[F x ]]> <![CDATA[fx(n1,m1)]]> <![CDATA[fx(n2,m2)]]> <![CDATA[fx(n3,m3)]]> <![CDATA[fx(n4,m4)]]>

[0082] 3. Pitch channel control design

[0083] The pitch channel and yaw channel are recorded as controlled channels. In the thrust establishment phase and thrust stabilization phase, the control command G p for:

[0084]

[0085] Among them, N yc is the overload instruction of the controlled channel; N y is the real-time overload of the controlled channel; ω z is the angular velocity of the controlled channel; K i ,K g ,K sf ,K a is the control parameter of the controlled channel; t is the flight time of the missile.

[0086] This step mainly designs the thrust building section and thrust stabilization section control systems.

[0087] In the thrust building phase and thrust stabilization phase, the control parameters of the controlled channel are:

[0088]

[0089] in,

[0090]

[0091] f sf1 、f i1 、f g1 、f sf2 、f g2 Determine from the following table:

[0092] Table 3 Control parameters at characteristic points

[0093]

[0094]

[0095] As shown in Table 3, the control parameters are obtained by linear interpolation according to the Mach number (MA1, MA2, MA3, MA4) according to Table 3; N1, N2, M1, M2 are constants; where x ij ,y ij ,z ij It is related to the aerodynamic characteristics and mission indicators of the missile, and technical personnel in this field can design it according to actual conditions and needs.

[0096] f1(x ij ,yij ,z ij )=(x ij ×Ma+y ij )×(1+z ij ×ALP) / Q

[0097] Filter Design

[0098] The filter scheme adopts the first-order + high-frequency bias scheme. The filter center frequency and filtering depth are designed according to the first-order elastic modal frequency and high-order elastic frequency of the projectile obtained from ground tests.

[0099]

[0100] Among them, Gf1(s) and Gf2(s) are the first-order and second-order filter transfer functions respectively; T1 and T2 are the filter time constants respectively; ζ1 and ζ2 are the structural filter damping coefficients, T3T4 are the filter time constants respectively; ζ3 and ζ4 are the structural filter damping coefficients.

[0101] 5. Thrust Trailing Section Control Compensation

[0102] When the engine enters the thrust end stage (t>Tb), thrust tail stage compensation is performed.

[0103] At this time, the control parameters are designed as:

[0104]

[0105]

[0106] Among them, N3, N4, M3, and M4 are constants.

[0107] In summary, the missile static instability control method based on the swing nozzle provided by the present invention has the following characteristics:

[0108] In order to achieve the above object, the present invention is implemented by the following technical solutions:

[0109] (1) To solve the problem of rolling disturbance suppression under strong uncertainty, the present invention adopts a variable structure control method based on the inner loop of the roll angular velocity instruction in the rolling loop, which effectively suppresses the rolling angular velocity interference on the basis of stability, improves the missile's ability to suppress oblique blowing interference and its adaptability to model uncertainty, and the aileron control efficiency of the entire missile is extremely low in the initial stage of leaving the orbit, so a damping loop is used for stabilization to ensure three-channel decoupling.

[0110] (2) The present invention aims at the problem of low-mode large static instability and weak handling stability ratio control. The pitch channel adopts a composite control method to improve the adaptability to changes in static instability. The dynamic performance of the swing-jet servo mechanism and the rapid response of the projectile are comprehensively considered in the design of rigid body control parameters. The control parameters of the stability control system are compromised through the bandwidth coordination technology to improve the adaptability to changes in static instability.

[0111] (3) The present invention uses amplitude-phase correction technology and adaptive control technology to optimize the design of elastic vibration suppression correction network and loop control parameters to ensure sufficient stability of rigid body motion and elastic motion, and uses digital filtering technology to achieve high-frequency resonance suppression, thereby weakening the impact of high-frequency vibration on the performance of the swing-jet servo mechanism.

[0112] (4) To address the problem of large discreteness of engine thrust under high and low temperature conditions, the present invention uses thrust online identification technology to identify engine thrust based on ground engine test data and ground aerodynamic blowing data, through online measurement of real-time acceleration feedback, flight altitude speed and other flight state quantities, so as to suppress the swing jet control efficiency fitting error caused by thrust discreteness.

[0113] (5) In view of the decrease in control capability of the thrust trailing section, the pitch channel of the present invention compensates for the control parameters of the trailing section based on online thrust identification + longitudinal overload auxiliary judgment, thereby eliminating the control error caused by thrust uncertainty.

[0114] Although the present invention has been disclosed as above in the form of a preferred embodiment, it is not intended to limit the present invention. Any person skilled in the art may make possible changes and modifications to the technical solution of the present invention by using the methods and technical contents disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and modifications made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solution of the present invention shall fall within the protection scope of the technical solution of the present invention.

Claims

1. A missile static instability control method based on a swing nozzle, characterized in that The method comprises the following steps: using a swing nozzle to control a pitch channel and a yaw channel, and using an air rudder to control a roll channel; The control method of the rolling channel is as follows: a variable structure control method is used to obtain a rolling channel control command, and the rolling channel control command is output to the rudder system after being limited; The pitch channel and yaw channel control method is as follows: according to the thrust obtained by online identification, the missile flight process is divided into three stages: thrust establishment stage, thrust stabilization stage and thrust termination stage; in the three stages of the missile flight process, a composite structure method is used to calculate the pitch channel control instructions and the yaw channel control instructions, and the pitch channel control instructions and the yaw channel control instructions are output to the rudder system after filtering and limiting; the control parameters of the thrust establishment stage and the thrust stabilization stage are the same; after entering the thrust termination stage, the control parameters of the thrust establishment stage and the thrust stabilization stage are corrected to eliminate the control error caused by the thrust uncertainty in the thrust termination stage.

2. A missile large static instability control method based on a swing nozzle according to claim 1, characterized in that: The scroll channel control instructions are: G R =g3×S(Δγ,ω x )+g2×ω x Among them, Δγ is the roll angle command deviation, ω x is the rolling angular velocity, S(Δγ,ω x ) is the transfer function of the variable structure controller, g2, g 3 It is the roll channel control parameter; S(Δγ,ω x )=(α1|Δγ|+α2|ω x +β|)×sign(g1Δγ+ω x ) Wherein, α1, α2, β, and g1 are constants greater than zero.

3. The method for controlling the large static instability of a missile based on a swing nozzle according to claim 1, characterized in that: The pitch channel and yaw channel are recorded as controlled channels. In the thrust establishment phase and thrust stabilization phase, the control command G p for: Among them, N yc is the overload instruction of the controlled channel; N y is the real-time overload of the controlled channel; ω z is the angular velocity of the controlled channel; K i ,K g ,K sf ,K a is the control parameter of the controlled channel; t is the flight time of the missile.

4. A missile large static instability control method based on a swing nozzle according to claim 3, characterized in that: In the thrust building stage and thrust stabilization stage, the control parameters of the controlled channel are: in, f sf1 、f i1 、f g1 、f sf2 、f g2 Determine from the following table: Composite structure method control parameter table f1(x ij ,y ij ,z ij )=(x ij ×Ma+y ij )×(1+z ij ×ALP) / Q。 5. A missile large static instability control method based on a swing nozzle according to claim 4, characterized in that: In the thrust tailing stage, the controlled channel control parameters are: Among them, N3, N4, M3, and M4 are constants.

6. A missile large static instability control method based on a swing nozzle according to claim 3, characterized in that: The thrust building phase, stabilization phase, and tail phase are divided according to time: 0≤t≤Ta is the thrust establishment stage, Ta≤t≤Tb is the thrust stabilization stage, and t≥Tb is the thrust tail stage; Ta is the moment when the thrust identification estimated value Thrust rises to 0.9×thrust; thrust is the thrust stable value; Tb is the moment when the thrust identification estimated value Thrust drops to 0.9×thrust.

7. A missile large static instability control method based on a swing nozzle according to claim 6, characterized in that: The thrust stability value thrust is determined by the following method: Obtain the curve F of ground test thrust changing with time under high temperature conditions, normal temperature conditions, and low temperature conditions th_h (t), F th_n (t), F th_l (t); The thrust identification estimated value Thrust is compared with the ground test thrust under high temperature conditions, normal temperature conditions and low temperature conditions at the corresponding missile flight time, and the value closest to the thrust identification estimated value Thrust is determined as the thrust stability value thrust.

8. The method for controlling the large static instability of a missile based on a swing nozzle according to claim 3 is characterized in that: The thrust identification estimate is: Thrust=N x -F x N x is the real-time axial overload, F x is the aerodynamic resistance.

9. The method for controlling the large static instability of a missile based on a swing nozzle according to claim 3 is characterized in that: The aerodynamic resistance F x Related to altitude, Mach number, and resultant angle of attack ALP.

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