Method of removing the second safety of airborne cruise missile fuze by relying on weak environmental information

Abstract

The present invention proposes a method for releasing the second safety of an airborne cruise missile fuze by relying on weak environmental information, using the pitch angle as environmental information for releasing the second safety of the fuze. When the pitch angle is greater than a set range, it indicates that the cruise missile is in a gliding stage, and enters a judgment window. In the judgment window, when the pitch angle meets the set range, a counting signal is given. When the counting signal reaches a set number of times, it indicates that the cruise missile is in a straight and level flight stage, and the conditions for releasing the second safety of the fuze are met. The method provided by the present invention will not affect the normal release of the second safety of the fuze even if there is a long-term airflow interference during the flight. It has a certain anti-interference ability and has very important practical application value for cruise missiles.

Description

Technical Field

[0001] The invention belongs to the field of information detection and data processing, and in particular relates to a method for releasing the second safety of an airborne cruise missile fuze by relying on weak environmental information. Background Art

[0002] Cruise missiles are an important trend in the development of information-based ammunition in the world. Airborne cruise missiles are one of the core weapon types for the Air Force to carry out precise and effective strikes on ground targets in future information-based wars. The launch and flight of airborne cruise missiles include the processes of delivery, roll stabilization, ejection of wings, gliding flight, engine start, straight and level flight, and terminal dive. Cruise missiles are different from conventional guided munitions. The main difference is that they can simultaneously undertake reconnaissance, attack, and damage assessment tasks during combat. The missiles are equipped with sensors such as GPS, radar, accelerometers, and gyroscopes. They are also equipped with computing processors and communication systems that can receive console commands and process data. Based on these hardware and software foundations, cruise missile fuzes can have a more flexible mode of action.

[0003] The release of the second safety insurance of the cruise missile fuze can be divided into several ways, relying on the minimum insurance distance (delay) information, the ignition load of the endurance engine, the wing ejection signal, and the gas power generated by the propulsion. The characteristics and energy of these environmental information sources are weak and cannot be directly used by the fuze. They belong to weak environmental information. When the airborne cruise missile ejects the wing, it is close to the delivery aircraft, which is easy to threaten the safety of the aircraft; relying on the minimum insurance distance (delay) information, it cannot be accurately controlled, and the error will become larger when it is disturbed, so it is not suitable to be an information source. The more mature and popular technology at present is to rely on the recoil load generated when the endurance engine is ignited. However, the signal for the ignition of the endurance engine comes from the navigation system (to determine the altitude) and the flight control system (to determine the stability of the flight). When the cruise missile reaches the predetermined altitude and the flight is stable, the engine starts to ignite, and then the recoil load generated by the fuze is identified to release the second insurance.

[0004] The load generated by the engine ignition of an airborne cruise missile is about 10g and lasts for a long time. This is difficult for traditional mechanical fuzes to accurately identify, and an additional recoil mechanism needs to be designed to identify the load. In addition, when the engine is ignited, the cruise missile is still in the gliding stage and is close to the launching aircraft. At present, there is no direct way to release the second line of insurance of the airborne cruise missile fuze by relying on weak environmental information in flight. There is no complete design method for how to identify and extract weak environmental information, what processing steps are required, and convert the information into a signal that can be used by the fuze. Summary of the invention

[0005] The purpose of the present invention is to provide a method for releasing the second safety of an airborne cruise missile fuze by relying on weak environmental information. According to the method, the existing hardware and software resources on the cruise missile are used to identify and process the weak environmental information during the flight. From the fuze application level, a related method is designed to convert the information into a signal that can be recognized by the fuze circuit. Through this method, the airborne cruise missile fuze can directly use weak environmental information to release the second safety, without the need to design additional mechanisms, simplifying the intermediate steps and improving reliability. In addition, compared with relying on the engine ignition pulse load, this method has a farther safety distance from the delivery aircraft and is safer. It provides a method idea for the fuze to utilize weak environmental information.

[0006] The technical solution to realize the present invention is:

[0007] A method for releasing the second safety of an airborne cruise missile fuze by relying on weak environmental information, using the pitch angle as the environmental information for releasing the second safety of the fuze.

[0008] When the pitch angle is greater than the set range, it indicates that the cruise missile is in the gliding stage and enters the judgment window. When the pitch angle in the judgment window is within the set range, a counting signal is given. When the counting signal reaches the set number of times, it indicates that the cruise missile is in the straight and level flight stage and meets the conditions for releasing the second safety of the fuze.

[0009] Compared with the prior art, the present invention has the following significant advantages:

[0010] The present invention provides a set of design methods. According to this method, the fuze of an airborne cruise missile can use the existing resources on the missile and directly use the weak environmental information during the flight to release the second line of insurance. There is no need to design additional mechanisms. The overall volume of the fuze can be reduced, the intermediate process can be simplified, and the reliability of the action can be improved. In addition, compared with relying on the engine ignition pulse load, this method has a farther safety distance from the delivery aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0011] Figure 1 It is a process of using weak environmental information to disarm the fuse.

[0012] Figure 2 It is the rotation relationship diagram between the coordinate axes of the orbital coordinate system and the coordinate axes of the projectile coordinate system.

[0013] Figure 3 It is the changes of the attitude angle of the airborne cruise missile during the gliding and straight-level flight stages.

[0014] Figure 4 It is the amplitude judgment process.

[0015] Figure 5 It is the counting circuit principle.

[0016] Figure 6It is the variation range of yaw angle (a), pitch angle (b) and roll angle (c) during the straight and level flight phase. DETAILED DESCRIPTION

[0017] The present invention is further described below in conjunction with the accompanying drawings and specific embodiments.

[0018] Combined with Figure 1 The method of releasing the second safety of the airborne cruise missile fuze by relying on weak environmental information of the present invention comprises the following steps:

[0019] Step 1. The selected weak environmental information is the aerodynamic force that the cruise missile is subjected to during the gliding stage and the straight-level flight stage. Since installing a pressure sensor on the missile body or wings will destroy the aerodynamic performance of the missile body, the aerodynamic force cannot be directly identified by the on-missile sensor. However, the effect of aerodynamic force will cause the missile body to be unbalanced. The on-missile controller rebalances the missile body by manipulating the deflection of the rudder wing. Therefore, the change in aerodynamic force can be reflected by the change in the missile body attitude angle. The missile body attitude angle is the environmental characteristic value of aerodynamic force.

[0020] Step 1.1, by the attached Figure 3 It can be seen that the environmental characteristic values ​​(reference pitch angle, angle of attack changes) vary significantly during the gliding and straight-level flight stages. According to the figure, during the gliding stage (t≤230s), the characteristic value changes within a range of ±12°; during the straight-level flight stage (t>230s), the characteristic value is in a stable downward trend.

[0021] In the straight and level flight stage, the combined external force on the missile body is in a balanced state. The reason for the change in the missile body's attitude angle is the consumption of the missile fuel, which causes the weight of the missile to change, and the attitude angle also changes accordingly. The present invention refers to the relevant parameters of the LOCAAS cruise missile to calculate the change amplitude of its attitude angle in the straight and level flight stage. The calculation formula is as follows:

[0022]

[0023] ΔY=ΔG=ΔM×g (2)

[0024]

[0025] Where ΔM is the change in the weight of the projectile, g is the value of gravity acceleration (unit: kg / N), H W is the engine fuel consumption (unit: mg / Ns), T N is the engine thrust (in kN), ΔT L is the time interval (unit: s), ΔY is the change in lift of the projectile (unit: N), Δα is the change in angle of attack (unit: °), is the lift coefficient caused by a unit angle of attack deflection, is the lift coefficient caused by a unit angle of pitch rudder deflection, ρ is the air density (in kg / m 3 ), v 2 is the flight speed (unit: Ma), S is the characteristic area of ​​the wing (unit: m 2 ).

[0026] The fuel consumption and thrust parameters of the engine refer to the LOCAAS cruise missile TJ-50-12 engine, H W =43.89, T N =0.267, take ΔT L =10, assuming that the altitude of the straight and level flight stage is 1000m, the gravity acceleration is 9.8, the air density is 1.1111, the flight speed is 0.3, and the characteristic area of ​​the wing is 0.052.

[0027] The calculation results show that within 10 seconds, the missile body's flight angle of attack changes by 0.02°. In horizontal flight, the change in the missile body's angle of attack is the same as the change in the pitch angle. In summary, with 10 seconds as the judgment interval, the attitude angle of the cruise missile changes by more than 10° during the gliding phase, and changes by 0.02° during the straight and level flight phase, and there are significant differences in the amplitude and frequency of change.

[0028] Step 1.2: The change of attitude angle can be calculated by the acceleration and angular velocity of the projectile. The accuracy of attitude angle calculation can be improved by introducing geomagnetic data. Therefore, the identification parameters are determined to be the acceleration, angular velocity and earth magnetic field.

[0029] Step 2: According to the physical parameters analyzed in step 1, select the sensor for identification. The sensors used should be of different types to avoid the same detection principle.

[0030] Step 2.1, select the existing MEMS IMU sensors in the flight control system: MEMS three-axis accelerometer to measure the acceleration of the projectile, MEMS three-axis gyroscope to measure the angular velocity of the projectile, and MEMS three-axis magnetometer to measure the magnetic field;

[0031] Step 2.2: To increase the reliability of identification, meet the redundant safety requirements of the fuze system, and prevent identification failure due to a single sensor failure, two independent sensor systems are established, and the sensors used in the two systems are the same.

[0032] Sensor system 1 uses the sensors of the missile's flight control system, and sensor system 2 uses the MEMS integrated nine-axis sensor MPU-9250 (including a three-axis accelerometer, a three-axis gyroscope, and a three-axis magnetometer), which has a volume of 3×3×1mm and extremely low power consumption. To ensure the consistency of the dual-system measurement data, the installation position of the MPU-9250 is symmetrical with the flight control sensor position along the missile's center of mass.

[0033] Step 3: The acceleration, angular velocity, and magnetic data detected by the sensor will be transmitted to the onboard processor, and the data fusion algorithm is designed to be a double extended Kalman filter (Double-EKF). In order to avoid singularity phenomena and calculate trigonometric functions, attitude quaternions are selected as operands, and the quaternions are converted into Euler angles to facilitate the expression of attitude angle conditions.

[0034] Step 3.1: To determine the attitude of the cruise missile, at least two coordinate systems need to be established. Establish the orbital coordinate system Ox1y1z1, with the origin of the coordinate system being the center of mass of the missile, the Ox1 axis pointing from the origin to the direction of the earth's magnetic field, the Oz1 axis pointing from the origin to the center of the earth, and the Oy1 axis direction determined by the right-hand rule.

[0035] In order to obtain acceleration, angular velocity and magnetic field data, the projectile coordinate system Ox2y2z2 is established. The origin of the coordinate system is the center of mass of the projectile. The Ox2 axis points from the origin to the horizontal flight direction of the projectile. The Oy2 axis is perpendicular to the cross section of the projectile and points to the ground along the longitudinal axis. The direction of the Oz2 axis is determined by the right-hand coordinate system.

[0036] The coordinate origins of the orbital coordinate system and the missile body coordinate system are both at the center of mass of the missile body. By rotating the orbital coordinate system toward the missile body coordinate system, the rotation angle between the three coordinate axes of the missile body coordinate system and the three axes of the orbital coordinate system can be obtained. This angle can be used to describe the attitude information of the missile body. Figure 2 As shown, the rotation order around the axis is Z→X→Y. The orbital coordinate system Ox1y1z1 first rotates around the Z axis by an angle ψ to obtain the intermediate coordinate system Ox'y'z1; then rotates around the X axis The intermediate coordinate system Ox'y2z' is obtained by rotating the Y axis by an angle θ to obtain the projectile coordinate system Ox2y2z2. Transformation matrix:

[0037]

[0038] is the transformation matrix between the two coordinate systems. ψ, θ is the Euler angle, ψ represents the yaw angle, represents the roll angle, and θ represents the pitch angle. The computational relationship between the coordinate vectors of the two coordinate systems is:

[0039]

[0040] Step 3.2: Assume that the system state variable is quaternion q, q = [q1, q2, q3, q4] T , where q4 is the quaternion scalar part, [q1,q2,q3] T =ρ is the quaternion vector part; the one-step estimate of the system state variable q at time k+1 is obtained by formula (5):

[0041]

[0042] Where f(q k+1 ) is the discrete state equation of the system consisting of the angular velocity and quaternion at time k, and is the one-step estimate of the system state variable q at time k+1. is the measured value of angular velocity at time k (unit: rad / s), β k is the gyro measurement drift at time k (unit: rad / s), w k is the true value of the angular velocity at time k, and ΔT is the sensor sampling interval, I 4×4 is the fourth-order identity matrix.

[0043] Represents the optimal estimate of the quaternion at time k and has been normalized:

[0044]

[0045] w=[w x w y w z ] T Output vector for the three-axis angular velocity:

[0046]

[0047] Gyro measurement drift update:

[0048] β k+1 =β k +β biass

[0049] β biass is the random drift of the gyroscope;

[0050] The prior estimate of the system state variable q at time k+1 is obtained by using formulas (6) and (7):

[0051]

[0052] In the formula is the optimal estimate of the vector part of the quaternion at time k, I 3×3 is the third-order identity matrix, is the prior estimate of the system state variable at time k+1, and its norm is:

[0053]

[0054] right After normalization, we get

[0055]

[0056] F k is the system state equation f(q k+1 )’s Jacobi matrix, η v Measure noise for the gyroscope.

[0057]

[0058] The prior prediction of the noise covariance matrix P of the extended Kalman filter at time k+1 is:

[0059]

[0060] P k+1|k is the one-step state estimation error covariance matrix at time k+1, P k|k is the state estimation error covariance matrix at time k, and Q is the system noise covariance matrix.

[0061] Step 3.3: Based on the measured acceleration and magnetic data at time k+1, establish the measurement update equation of the system state variables, and calculate the optimal estimated value of the system state variables at time k+1 using formulas (9)-(21):

[0062] Since the transformation matrix from the orbital coordinate system to the projectile coordinate system is It is in the form of Euler angles, which is not conducive to quaternion calculation, so it needs to be converted into a quaternion attitude transformation matrix. After the conversion, A(q) is obtained:

[0063]

[0064] Step 3.3.1: Correct the acceleration sensor data once:

[0065] Estimate the acceleration of the projectile at time k+1:

[0066]

[0067] In the formula Is The estimated value of the projectile acceleration at time k+1 is derived as [0 0 |g|] T This is to consider the influence of gravity acceleration in the vertical direction on the estimated value under natural conditions. The z-axis direction of the orbital coordinate system Ox1y1z1 points to the center of the earth / gravity direction, so there is only a |g| value in the z-axis direction.

[0068]

[0069] H k+1'1 yes The Jacobi matrix of .

[0070]

[0071] Where K k+1'1 is the Kalman gain value of the acceleration data to the system state variable, and R1 is the acceleration sensor measurement noise covariance matrix.

[0072] z k+1'1 is the actual measurement value of the acceleration sensor at time k+1, z k+1'1 =[a k+1,x a k+1,y a k+1,z ] T , normalize it:

[0073]

[0074] is the normalized acceleration sensor measurement value, ‖‖z k+1'1 ‖‖ is the norm of the acceleration sensor measurement value, and the calculation formula is as follows:

[0075]

[0076] Using the estimated value of the projectile acceleration at time k+1, the actual measurement value of the acceleration sensor and the Kalman gain K k+1'1 , calculate the quaternion correction q ε1 :

[0077]

[0078] Where q ε1 It is the quaternion correction value obtained based on the acceleration sensor measurement data. is the quaternion value corrected by the acceleration data, I is the unit matrix, is the state estimation error covariance matrix after the acceleration data is corrected.

[0079] Step 3.3.2: Second correction of magnetometer data:

[0080] Estimate the Earth's magnetic field at time k+1:

[0081]

[0082] In the formula Is The estimated value of the magnetic field at time k+1 is derived,

[100] T Considering the natural situation (assuming the magnetic field strength is 1), the direction of the earth's magnetic field is north-south, and the x-axis direction of the orbital coordinate system Ox1y1z1 points to the geomagnetic direction (ie, north-south), so the impact on the magnetic field data is considered only in the x-axis direction.

[0083]

[0084] H k+1'2 yes The Jacobi matrix of .

[0085]

[0086] K k+1'2 is the Kalman gain value of the magnetic data to the system state variables, and R2 is the magnetometer measurement noise covariance matrix.

[0087] z k+1'2 is the actual measurement value of the magnetometer at time k+1, z k+1'2 =[m k+1,x m k+1,y m k+1,z ] T , normalize it:

[0088]

[0089] is the normalized magnetometer measurement value, ‖‖z k+1'2 ‖‖ is the norm of the magnetometer measurement value, calculated as follows:

[0090]

[0091] Using the magnetic field estimation value at time k+1, the actual measurement value of the magnetometer and the Kalman gain K k+1'2 , calculate the quaternion correction q ε2 :

[0092]

[0093] Where q ε2 is the quaternion correction obtained based on the magnetometer measurement data, is the optimal estimate of the quaternion at time k+1, P k+1|k+1 is the state estimation error covariance matrix of the system at time k+1.

[0094] Step 3.4: Quaternions cannot directly describe the change of the missile's attitude angle, so they need to be converted into Euler angles. The formula for converting the optimal estimate of the quaternion at time k+1 into Euler angles is as follows:

[0095]

[0096] roll k+1 ,pitch k+1 ,yaw k+1They are the roll angle, pitch angle, and yaw angle values ​​of the projectile at time k+1, where 180° / π is used to achieve the unit conversion between 'radians' and 'degrees'.

[0097] Step 4. According to the principle of state feedback utilization of fuze environmental information, the transmission mode of attitude angle data is selected as IIC, and the transmitting parties are the on-board processor and the fuze control circuit; the data acquisition frequency is set to 10Hz, and the data transmission frequency is 20Hz; the data storage logic is transmission, temporary storage MCU to merged storage.

[0098] Step 5: Take the change of the pitch angle during the straight and level flight stage as the judgment object, and use its amplitude change range as the setting range; analyze the factors that affect the pitch angle change range during this stage, and set the duration of the judgment interval.

[0099] Step 5.1, according to the pitch angle solution result of the cruise missile in the straight and level flight stage, the pitch angle change range 1 is obtained, combined with the attack angle change range 2 caused by fuel consumption in the straight and level flight stage, the range 1 and the range 2 are summed to obtain the final range, which is used as the amplitude basis for distinguishing the gliding and straight and level flight stages;

[0100] Step 5.2, the data collection interval is 0.1s. Considering the processor solution speed, the timeliness of the second safety release, and the change curve of the angle of attack, the time judgment interval is set to t=1s, and the maximum amplitude difference of the attitude angle within the interval is calculated and stored in the storage unit.

[0101] Step 6: According to the amplitude range and the duration of the judgment interval set in step 5, formulate an amplitude judgment process.

[0102] Step 6.1, see the attached figure for the amplitude judgment process Figure 4 , select three consecutive intervals as a judgment cycle. If any value in the cycle does not meet the set requirements, open the judgment "window" and send a high level to the counter reset terminal;

[0103] Step 6.2: In the "window", if all amplitudes in the first cycle meet the set requirements, proceed to step 7, otherwise continue with step 6.1.

[0104] Step 7: Whenever all amplitudes meet the requirements in a cycle, a high level is sent to the counter counting end through the fuse control circuit to continue the "window"; if the amplitude in the next cycle does not meet the set requirements in chronological order, the "window" is ended, and a high level is sent to the counter clearing end, returning to step 6.1.

[0105] When the cruise missile encounters long-term airflow interference during the straight and level flight phase, the attitude angle will change dramatically, returning to step 6, causing the fuze circuit to misjudge the stage of the cruise missile, delaying the release of the second fuse. However, with the adjustment of the flight control system on the missile, the attitude angle will stabilize again, and the judgment process will continue.

[0106] Step 8: Set the counting standard of each counter to 10. See the attached Figure 5 When any one of the two counters meets the requirements, a control signal is output to release the second fuse safety.

[0107] Example 1

[0108] A set of 14h static MEMSIMU data is selected and solved by Double-EKF algorithm to obtain the absolute error of the attitude angle (pitch angle) estimation value to help determine the amplitude judgment range. The sensor data sampling interval is set to 0.1s, and the quaternion initial value q0 and the gyroscope offset initial value β0, β biass =1e-19,η v =1e-13, system noise covariance matrix Q = [0.010.010.010.01] T , acceleration sensor measurement noise covariance matrix R1 = [0.10.10.1] T , magnetometer measurement noise covariance matrix R2 = [0.0010.0010.001] T By using formulas (4)-(24), the attitude angle data within 14 hours is calculated, as shown in the attached figure. Figure 6 ((a) is the result of yaw angle calculation, (b) is the result of pitch angle calculation, and (c) is the result of roll angle calculation). The pitch angle is used as a reference for judgment (the angle of attack is equal to the pitch angle during horizontal flight), and its range is ±0.04°; the roll angle range is -0.25° to 0.15°, and the yaw angle range is -0.2° to 0.5°. The attitude angle amplitude standard is set to ±0.1° (0.08°+0.02°).

[0109] The contents not described in detail in the specification of the present invention belong to the prior art known to the professional and technical personnel in this field.

Claims

1. A method for releasing the second safety of an airborne cruise missile fuze by relying on weak environmental information, characterized in that: The elevation angle is used as the environmental information to release the second line of defense of the fuze. When the pitch angle is greater than the set range, it indicates that the cruise missile is in the gliding stage and enters the judgment window. In the judgment window, if the pitch angle meets the set range, a counting signal is given. When the counting signal reaches the set number of times, it indicates that the cruise missile is in the straight and level flight stage, and the conditions for releasing the second fuse safety are met. The pitch angle variation range in the straight-and-level flight phase is used as the setting range: the pitch angle variation range in the straight-and-level flight phase is determined by two aspects: on the one hand, it comes from the pitch angle calculation accuracy range 1, and on the other hand, it comes from the attitude angle variation range 2 caused by the fuel consumption of the cruise missile; the final range is obtained by summing range 1 and range 2; The pitch angle calculation formula is: Where pitch k+1 is the value of the pitch angle at time k+1, which has been normalized. Represents the optimal estimate of the quaternion at time k+1; The calculation formula for attitude angle change caused by fuel consumption is: Where Δα is the change in angle of attack, ΔY is the change in lift of the projectile, is the lift coefficient caused by a unit angle of attack deflection, is the lift coefficient caused by a unit angle of pitch rudder deflection, ρ is the air density, v 2 is the flight speed, and S is the characteristic area of ​​the wing.

2. The method for releasing the second safety of an airborne cruise missile fuze by relying on weak environmental information according to claim 1 is characterized in that: The pitch angle calculation is obtained by the following steps: 3.

1. First, collect the acceleration, angular velocity and earth magnetic field data of the cruise missile; 3.

2. Obtain the attitude information of the projectile by converting the coordinate axes between the orbital coordinate system and the projectile coordinate system; 3.

3. Set the quaternion as the system variable, use the optimal estimate of the quaternion at time k and the angular velocity data of the projectile at time k to solve the prior estimate of the quaternion at time k+1, and then correct the prior estimate through the acceleration data and magnetic data to obtain the optimal estimate of the quaternion at time k+1; 3.

4. Use the optimal estimate of the quaternion at time k+1 to obtain the value of the pitch angle at time k+1.

3. The method for releasing the second safety of an airborne cruise missile fuze by relying on weak environmental information according to claim 1 is characterized in that: The mass change of the cruise missile caused by fuel consumption during the straight-line flight phase can be obtained by the following formula: ΔY=ΔM×g Where ΔM is the change in the weight of the projectile, g is the value of gravity acceleration, H W is the engine fuel consumption, T N is the engine thrust, ΔT L For the time interval.

4. The method for releasing the second safety of an airborne cruise missile fuze by relying on weak environmental information according to claim 1 is characterized in that: The acceleration, angular velocity and earth magnetic field data of the cruise missile are collected through the existing MEMS IMU sensors in the flight control system.

5. The method for releasing the second safety of an airborne cruise missile fuze by relying on weak environmental information according to claim 4 is characterized in that: The cruise missile’s acceleration, angular velocity, and Earth’s magnetic field data are also collected through an integrated nine-axis sensor.

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