Aircraft guidance control system and method accommodating directional gyro failure

Abstract

This invention discloses an aircraft guidance and control system and method that takes into account the failure of the directional gyroscope. In this system, data collected by seven accelerometers is used to calculate the data of the alternative directional gyroscope in real time. At the same time, a sliding mode adaptive guidance law is used to guide and control the aircraft, so that the overall hit accuracy of the aircraft is kept within the ideal range. This has important practical significance and engineering application value.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of aircraft guidance control, and particularly relates to an aircraft guidance control system and method considering failure of directional gyro. BACKGROUND

[0002] The conventional aircraft uses a space directional gyro to identify the roll angle of the aircraft, to realize signal decomposition and gravity orientation. During flight, the space directional gyro generally has the following problems. First, the space directional gyro is limited by the range. When the gyro hits the frame, that is, the space directional gyro reaches the critical stability, the microcomputer will output an incorrect guidance instruction, causing the aircraft to deviate. Since further energy loss is generated during the deviation correction process, the range and accuracy of the aircraft are ultimately affected. In addition, although there are some schemes to replace the directional gyro in the prior art, the actual application effect of these schemes is poor, and the final accuracy is affected to different degrees. Therefore, when researching the directional gyro replacement scheme, the overall guidance control process of the aircraft needs to be considered.

[0003] Due to the above reasons, the present inventors have made in-depth research on the existing guidance control scheme and directional gyro replacement scheme in order to design a new control system and method that can solve the above problems. SUMMARY

[0004] In order to overcome the above problems, the present inventors have made intensive research and designed an aircraft guidance control system and method considering failure of directional gyro. In this system, the data of the directional gyro is replaced by the data calculated in real time by the data of seven accelerometers, and the aircraft is guided and controlled by the corresponding sliding mode adaptive guidance law, so that the overall hit accuracy of the aircraft is maintained within the ideal range, which has important practical significance and engineering application value, thereby completing the present application.

[0005] Specifically, the purpose of the present application is to provide an aircraft guidance control system considering failure of directional gyro. The system includes a required overload calculation module:

[0006] In the terminal guidance segment, the required overload calculation module obtains the required overload of the aircraft in real time by the following formula (I):

[0007]

[0008] Wherein, a M represents the required overload;

[0009] represents the relative motion speed of the projectile and the target;

[0010] represents the line-of-sight angular rate of the projectile and the target;

[0011] r represents the relative distance between the missile and the target;

[0012] λ represents the line-of-sight angle of the missile and the target;

[0013] α, b and κ all represent design parameters;

[0014] e1 and e2 both represent state variables;

[0015] S represents the terminal sliding mode surface of the terminal sliding mode control system;

[0016] ρ represents an adaptive parameter;

[0017] γ M represents the velocity inclination angle of the aircraft.

[0018] The terminal sliding mode surface is obtained by the following formula (II):

[0019] s = e2 + α|e1 |sgn(e1) (II)

[0020] wherein the state variable e1 = λ - λ f ,

[0021] The adaptive parameter ρ is obtained by the following formula (III):

[0022] ρ = ω3|e2 | (III)

[0023] wherein ω3 represents the roll angular velocity;

[0024] |e2| represents the absolute value of the line-of-sight angle rate of the missile and the target.

[0025] The system further comprises seven accelerometers and a rotation speed calculation module arranged on the aircraft,

[0026] wherein the accelerometer one, the accelerometer two and the accelerometer three are all used for measuring the acceleration in the x-axis direction of the body coordinate system, and the outputs thereof are A a , A b , A g , respectively;

[0027] The accelerometer four and the accelerometer five are both used for measuring the acceleration in the y-axis direction of the body coordinate system, and the outputs thereof are A c , A d , respectively;

[0028] The accelerometer six and the accelerometer seven are both used for measuring the acceleration in the z-axis direction of the body coordinate system, and the outputs thereof are A e , A f , respectively;

[0029] The output information A a , Ab , A g , A c , A d , A e , A f All are transmitted to the rotation speed solving module in real time.

[0030] Wherein, the rotation speed solving module obtains the rotation speed of the aircraft in real time through the following formula (four):

[0031]

[0032] Wherein, ω1 represents the pitch angle velocity;

[0033] ω2 represents the yaw angle velocity;

[0034] ω3 represents the roll angle velocity;

[0035] C1, C2, C3 all represent process quantities, which are obtained through the following formula (five):

[0036]

[0037] Wherein, the rotation speed solving module transmits the rotation speed information of the aircraft to the required overload solving module in real time;

[0038] The required overload solving module receives the information transmitted by the rotation speed solving module in real time, and also receives the information transmitted by the directional gyro in real time,

[0039] Preferably, when the roll angle velocity ω3 in the information transmitted by the directional gyro reaches 55 degrees, the adaptive parameter is obtained through the information transmitted by the rotation speed solving module.

[0040] The application also provides a kind of aircraft guidance control method considering the failure of directional gyro, in which:

[0041] In the terminal guidance section, the required overload of the aircraft is obtained in real time through the following formula (one):

[0042]

[0043] Wherein, a M represents the required overload;

[0044] represents the relative motion speed of missile and target;

[0045] represents the line-of-sight angular rate of missile and target;

[0046] r represents the relative distance of missile and target;

[0047] λ represents the line-of-sight angle of missile and target;

[0048] a, b and k all represent design parameters;

[0049] e1 and e2 all represent state variables;

[0050] s represents a terminal sliding mode surface;

[0051] p represents an adaptive parameter;

[0052] gamma M represents the velocity inclination angle of the aircraft.

[0053] wherein the terminal sliding mode surface is obtained by the following formula (II):

[0054] s = e2 + a |e1| b sgn (e1) (II)

[0055] wherein the state variable e1 = l - l f ,

[0056] wherein the adaptive law of the adaptive parameter p is the following formula (III):

[0057] p = w3 |e2| (III)

[0058] wherein w3 represents the roll angle;

[0059] |e2| represents the absolute value of the line-of-sight angular rate of the missile.

[0060] The present application has the beneficial effects including:

[0061] (1) The aircraft guidance control system and method considering the failure of the directional gyro provided by the present application can ensure the guidance control precision of the aircraft in the case of failure of the directional gyro, and improve the adaptability of the aircraft to the harsh environment.

[0062] (2) The aircraft guidance control system and method considering the failure of the directional gyro provided by the present application can eliminate the dependence of the aircraft on the directional gyro with a small amount of additional cost or no additional cost, and can improve the hit accuracy and adaptability of the aircraft compared with the traditional aircraft. BRIEF DESCRIPTION OF DRAWINGS

[0063] Figure 1 show the flight trajectories of two aircraft in Embodiment 1 of the present application;

[0064] Figure 2 show the flight trajectories of two aircraft in Embodiment 1 of the present application. DETAILED DESCRIPTION

[0065] The application will be further described in details by the following drawings and examples. The features and advantages of the application will become more apparent through these descriptions.

[0066] The term "exemplary" is used herein to mean "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other implementations. Unless specifically indicated otherwise, the drawings shown in the Figures are not necessarily to scale.

[0067] The application provides a flight vehicle guidance control system considering the failure of directional gyro, which comprises a required overload solution module:

[0068] The required overload solution module obtains the required overload of the flight vehicle in real time through the following formula (I) in the terminal guidance stage:

[0069]

[0070] wherein, a M represents the required overload;

[0071] represents the relative velocity between the projectile and the target, which is obtained through the following formula:

[0072]

[0073] V T represents the velocity of the target, which is obtained in real time by the guidance device;

[0074] γ T represents the velocity inclination angle of the target, which is obtained in real time by the guidance device; the guidance device can be an image guidance device or a laser guidance device;

[0075] V M represents the velocity of the flight vehicle, which is obtained in real time by the satellite receiver or the inertial navigation system;

[0076] γ M represents the velocity inclination angle of the flight vehicle, which is obtained in real time by the satellite receiver or the inertial navigation system;

[0077] represents the angular rate of the line of sight between the projectile and the target;

[0078] represents the relative distance between the projectile and the target, and λ represents the angular rate of the line of sight between the projectile and the target, which are obtained in real time according to the position of the target and the position of the flight vehicle;

[0079] α, b and κ all represent design parameters, and are preferably taken as α = 10, b = 0, and κ = 0.

[0080] e1 and e2 both represent state variables, and are preferably defined as e1 = λ - λf ,

[0081] S represents a terminal sliding surface;

[0082] ρ represents an adaptive parameter;

[0083] In a preferred embodiment, the terminal sliding surface is obtained by the following formula (II):

[0084] s = e2 + a |e1| b sgn(e1) (II)

[0085] wherein the state variable e1 = λ - λ f ,

[0086] In a preferred embodiment, the adaptive parameter ρ is obtained by the following formula (III):

[0087] ρ = ω3 |e2| (III)

[0088] wherein ω3 represents a roll angular velocity;

[0089] |e2| represents the absolute value of the line-of-sight angular rate of the projectile.

[0090] In a preferred embodiment, the system further comprises 7 accelerometers and a rotation speed calculation module arranged on the aircraft,

[0091] Among the 7 accelerometers, the first accelerometer, the second accelerometer and the third accelerometer are used to measure the acceleration in the x-axis direction of the projectile coordinate system, and their outputs are A a , A b , A g ;

[0092] The fourth accelerometer and the fifth accelerometer are used to measure the acceleration in the y-axis direction of the projectile coordinate system, and their outputs are A c , A d ;

[0093] The sixth accelerometer and the seventh accelerometer are used to measure the acceleration in the z-axis direction of the projectile coordinate system, and their outputs are A e , A f ;

[0094] The output information of the 7 accelerometers is transmitted to the rotation speed calculation module in real time, and the roll angular velocity information is obtained by the rotation speed calculation module in real time.

[0095] In a preferred embodiment, the rotational speed calculation module obtains the rotational speed of the aircraft in real time using the following formula (iv): the roll speed is the rotational speed of the aircraft only around the main axis, and does not include rotational offset in other directions;

[0096]

[0097] Where ω1 represents the pitch angular velocity;

[0098] ω2 represents the yaw rate;

[0099] ω3 represents the roll angular velocity;

[0100] C1, C2, and C3 all represent process quantities, obtained through the following formula (V):

[0101]

[0102] Wherein, the χ a The x-axis represents the projection of the line connecting accelerometer one and the center of mass of the spacecraft in the projectile coordinate system onto the x-axis; b This represents the projection of the line connecting accelerometer 2 and the spacecraft's center of mass onto the X-axis in the projectile coordinate system; the χ... g The x-axis represents the projection of the line connecting accelerometer 3 and the center of mass of the spacecraft in the projectile coordinate system onto the x-axis; c The x-axis represents the projection of the line connecting accelerometer four and the center of mass of the spacecraft onto the Y-axis in the projectile coordinate system; d The x-axis represents the projection of the line connecting accelerometer 2 and the center of mass of the spacecraft onto the Y-axis in the projectile coordinate system; e The x-axis represents the projection of the line connecting accelerometer 5 and the center of mass of the spacecraft onto the Z-axis in the projectile coordinate system; f This represents the projection of the line connecting the accelerometer six and the center of mass of the aircraft onto the Z-axis in the projectile coordinate system. Preferably, the origin O of the projectile coordinate system is the center of mass of the aircraft, the X-axis is the axis of symmetry of the aircraft shell, and the positive direction of the X-axis points towards the nose of the aircraft; the Y-axis is located in the longitudinal plane of symmetry of the aircraft, the Y-axis is perpendicular to the X-axis, and the positive direction of the Y-axis points upward; the Z-axis is perpendicular to the longitudinal plane of symmetry of the aircraft, and when viewed along the launch direction at the moment of launch, the positive direction of the Z-axis points to the right.

[0103] Preferably, the rotational speed calculation module is connected to the required overload calculation module, and transmits the aircraft's rotational speed information to the required overload calculation module in real time;

[0104] The required overload calculation module receives information from the rotation speed calculation module in real time, and also receives information from the directional gyroscope in real time.

[0105] Preferably, when the roll angle obtained in three consecutive sampling periods in the information delivered by the directional gyro is above 55 degrees, the adaptive parameter is obtained by solving the information delivered by the rotation speed solving module.

[0106] The application also provides a flight vehicle guidance control method considering the failure of directional gyro, in which:

[0107] In the terminal guidance stage, the required overload of the flight vehicle is obtained in real time by the following formula (I):

[0108]

[0109] The terminal sliding surface is obtained by the following formula (II):

[0110] s = e2 + a |e1| b sgn(e1) (II)

[0111] wherein the state variable e1 = λ - λ f 、

[0112] The adaptive law of the adaptive parameter p is the following formula (III):

[0113] p = ω3 |e2 | (III)

[0114] wherein ω3 represents the roll angular velocity;

[0115] |e2| represents the absolute value of the line-of-sight angular rate of the missile and target.

[0116] Embodiment

[0117] Embodiment 1: In a digital-analog simulation software, two identical flight vehicles are called, and are launched towards the same target place 30 km away under the same weather conditions at the same place;

[0118] In the first flight vehicle, a flight vehicle guidance control system considering the failure of directional gyro is added, the directional gyro in the flight vehicle fails after 3 seconds of entering the terminal guidance stage, i.e. the roll angle obtained in three consecutive sampling periods after 3 seconds of entering the terminal guidance stage is above 55 degrees; the roll information is provided in real time by the directional gyro before the failure of the directional gyro, and the roll information is provided in real time by the rotation speed solving module after the failure of the directional gyro

[0119] In the first flight vehicle, in the terminal guidance stage, the required overload of the flight vehicle is obtained in real time by the required overload solving module by the following formula (I):

[0120]

[0121] a = 10, b = 0.5, κ = 0.2, e1 = λ - λ f ,

[0122] The terminal sliding surface is obtained by the following formula (II):

[0123] s = e2 + a |e1| b sgn(e1) (II)

[0124] The adaptive parameter p is obtained by the following formula (III):

[0125] p = w3 |e2| (III)

[0126] w3 represents the roll angular velocity, i.e. in the first 3 seconds of entering the terminal guidance stage, the value is obtained by the directional gyro on the aircraft, and after 3 seconds, it is obtained by the following formula (IV):

[0127]

[0128] wherein C1, C2, C3 represent process quantities, which are obtained by the following formula (V):

[0129]

[0130] The final flight trajectory of the first aircraft is shown by the solid line in Figure 1 .

[0131] The proportional guidance guidance rate is used for guidance control in the second aircraft, i.e. the overload A directional gyro is also installed, and the directional gyro in the second aircraft is set to stop working after 3 seconds of entering the terminal guidance stage, and no longer provides roll information.

[0132] The final flight trajectory of the second aircraft is shown by the dashed line in Figure 1 .

[0133] Example 2:

[0134] The same two aircraft as in Example 1 are used, and the flight target is changed to a target 35 km away, and the final flight trajectories of the two aircraft are shown in Figure 2 , wherein the solid line represents the trajectory of the first aircraft, and the dashed line represents the trajectory of the second aircraft.

[0135] It can be seen from Figure 1 and Figure 2 that the aircraft guidance control system considering the failure of the directional gyro can ensure the hit accuracy of the aircraft under the condition of failure of the directional gyro of the aircraft, and if the proportional guidance guidance rate control scheme encounters the failure of the directional gyro, a large deviation of the landing point will be caused, resulting in miss.

[0136] The application has been described above with reference to preferred embodiments. However, these embodiments are merely exemplary and are intended to be illustrative only. Various substitutions and alterations are possible in view of the disclosure of this application without departing from the spirit and scope of the application.

Claims

1. An aircraft guidance and control system that takes into account the failure of directional gyroscopes, characterized in that, The system comprises a required overload solving module: In the terminal guidance phase, the required overload solving module obtains the required overload of the vehicle in real time through the following formula (I): wherein a M represents the required overload; V represents the velocity of the projectile relative to the target; represents the ballistic line-of-sight angular rate; r represents the relative distance between the missile and the target; λ represents the line-of-sight angle between the missile and the target; α, b and κ all represent design parameters; e1 and e2 both represent state variables; s represents a terminal sliding mode surface; ρ represents an adaptive parameter; gamma M denotes the speed inclination angle of the aircraft; The terminal sliding mode surface is obtained through the following formula (II): s = e2+ a |e1| b sgn(e1) (two) where the state variable e1= λ - λ f , The adaptive parameter ρ is obtained through the following formula (III): ρ = ω3|e2| (III) wherein ω3 represents the roll angular velocity; |e2| represents the absolute value of the line-of-sight angular rate between the missile and the target.

2. The vehicle guidance control system considering the failure of the directional gyroscope according to claim 1, wherein the system further comprises seven accelerometers arranged on the vehicle and a rotation speed solving module, 3. The vehicle guidance control system considering the failure of the directional gyroscope according to claim 2, wherein the rotation speed solving module obtains the rotation speed of the vehicle in real time through the following formula (IV): Among them, the first accelerometer, the second accelerometer and the third accelerometer are used for measuring the acceleration in the x-axis direction of the projectile coordinate system, and the outputs are A a ,A b ,A g ; Accelerometer four and accelerometer five are used to measure the acceleration in the y-axis direction of the projectile coordinate system, and their outputs are A c , A d ; Accelerometer six and accelerometer seven are used to measure the acceleration in the z-axis direction in the body coordinate system, and their outputs are A e , A f ; The output information A of the 7 accelerometers a b g c d e f are all transmitted to the rotation speed calculation module in real time.​​​​​​ wherein ω1 represents the pitch angular velocity; ω2 represents the yaw angular velocity; ω3 represents the roll angular velocity; C1, C2 and C3 all represent process quantities, which are obtained through the following formula (V):

4. The vehicle guidance control system considering the failure of the directional gyroscope according to claim 3, wherein the rotation speed solving module transmits the rotation speed information of the vehicle to the required overload solving module in real time; The required overload solving module receives the information transmitted by the rotation speed solving module in real time, and also receives the information transmitted by the directional gyroscope in real time, 。 When the roll angular velocity ω3 in the information transmitted by the directional gyroscope reaches 55 degrees, the adaptive parameter is solved through the information transmitted by the rotation speed solving module. In the method: In the terminal guidance phase, the required overload of the vehicle is obtained in real time through the following formula (I): r represents the relative distance between the missile and the target; 5. A guidance control method for an aircraft which takes into account the failure of a directional gyroscope, characterized in that, λ represents the line-of-sight angle between the missile and the target; α, b and κ all represent design parameters; wherein a M represents the required overload; represents the velocity of the relative motion of the projectile and the target; represents the ballistic line-of-sight angular rate; e1 and e2 both represent state variables; s represents a terminal sliding mode surface; ρ represents an adaptive parameter; The terminal sliding mode surface is obtained through the following formula (II): The adaptive parameter ρ is obtained through the following formula (III): ρ = ω3|e2| (III) gamma M denotes the speed inclination angle of the aircraft; wherein ω3 represents the roll angular velocity; s = e2+ a |e1| b sgn(e1)(two) where the state variable e1= λ - λ f 、 |e2| represents the absolute value of the line-of-sight angular rate between the missile and the target. ​ ​ ​ ​

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