CN115900459A - Method for relieving second insurance of airborne patrol missile fuze by means of weak environment information - Google Patents
Method for relieving second insurance of airborne patrol missile fuze by means of weak environment information Download PDFInfo
- Publication number
- CN115900459A CN115900459A CN202211593767.0A CN202211593767A CN115900459A CN 115900459 A CN115900459 A CN 115900459A CN 202211593767 A CN202211593767 A CN 202211593767A CN 115900459 A CN115900459 A CN 115900459A
- Authority
- CN
- China
- Prior art keywords
- fuse
- pitch angle
- missile
- angle
- environment information
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Pending
Links
Images
Classifications
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T90/00—Enabling technologies or technologies with a potential or indirect contribution to GHG emissions mitigation
Landscapes
- Aiming, Guidance, Guns With A Light Source, Armor, Camouflage, And Targets (AREA)
Abstract
The invention provides a method for relieving a second fuse of an airborne flying bomb fuze by means of weak environment information, wherein a pitch angle is used as environment information for relieving the second fuse of the fuze, when the pitch angle is larger than a set range, the flying bomb patrols in a gliding stage, and enters a judgment window, a counting signal is given out when the pitch angle in the judgment window meets the set range, and when the counting signal reaches a set number of times, the flying bomb patrols in a straight level flying stage, and the condition for relieving the second fuse of the fuze is met. The method provided by the invention has the advantages that even if long-time airflow interference occurs in the flying process, the normal release of the second insurance by the fuse is not influenced, and the method has certain anti-interference capability and has important practical application value for flying round.
Description
Technical Field
The invention belongs to the field of information detection and data processing, and particularly relates to a method for removing a second insurance of an airborne patrol missile fuze by means of weak environment information.
Background
The flying round is an important trend of the development of information ammunition in the world at present. The airborne patrol bomb is one of the core weapon types for the air force to accurately and effectively strike the ground target in the future informatization war. The launching and flying of the airborne patrol missile comprises the processes of throwing, rolling stabilization, missile wing popping, gliding flight, engine starting, linear level flight, tail end diving and the like. The patrol missile is different from a conventional guided missile, mainly distinguished from the fact that the patrol missile can simultaneously undertake tasks such as reconnaissance, attack, damage assessment and the like in the combat process, carries sensors such as a GPS (global positioning system), a radar, an accelerometer, a gyroscope and the like on the missile, and also carries a computing processor and a communication system which can receive instructions of an operation console and process data; based on the hardware and software bases, the patrol missile fuze can have a more flexible action mode.
The second safety insurance of the patrol missile detonator can be released in several ways of depending on minimum insurance distance (delay) information, continuation of journey engine ignition load, missile wing popup signals and gas power generated by propulsion, and the environmental information sources have weak characteristics and energy, cannot be directly utilized by the detonator and belong to weak environmental information. When the airborne patrol bomb pops up the missile wing, the missile wing is close to the thrown airplane, and the safety of the airplane is easily threatened; depending on the minimum insurance distance (delay) information, it cannot be accurately controlled, and the error becomes large when the interference occurs, so that it is not suitable for being used as an information source. The current technology which is mature and has high popularity depends on the recoil load generated when the endurance engine is ignited. However, the ignition signal of the endurance engine comes from a navigation system (judging the height) and a flight control system (judging the flight stability), when the cruise missile reaches the preset height and the flight is stable, the engine starts to ignite, and then the recoil load generated by fuze recognition is used for removing the second insurance.
The load generated by ignition of an airborne patrol missile engine is about 10g, the duration is long, the load is difficult to accurately identify for a traditional mechanical fuse, and a recoil mechanism needs to be additionally designed to identify the load; in addition, when the engine is ignited, the cruise missile is still in a gliding stage and is close to the launching airplane. At present, the second insurance of the fuse of the airborne patrol missile does not need to be relieved by directly depending on the weak environment information in flight, the processing steps are needed for identifying and extracting the weak environment information, the information is converted into a signal which can be used by the fuse, and a set of complete design method does not exist.
Disclosure of Invention
The invention aims to provide a method for relieving the second insurance of a fuse of an airborne patrol missile by relying on weak environment information, according to the method, the weak environment information in the flight process is identified and processed by utilizing the existing hardware and software resources on the patrol missile, and a related method is designed to convert the information into a signal which can be identified by a fuse circuit from the application level of the fuse; by the method, the airborne patrol cartridge fuse can directly use weak environment information to remove the second insurance, no additional mechanism is needed to be designed, intermediate steps are simplified, and reliability is improved. And a method idea is provided for the fuze to utilize weak environment information.
The technical solution for realizing the invention is as follows:
a method for removing the second fuse of airborne flying ammunition fuze by weak environmental information uses the pitch angle as the environmental information for removing the second fuse of the fuze,
when the pitch angle is larger than the set range, the patrol missile is in a gliding stage, a judgment window is entered, a counting signal is given out when the pitch angle in the judgment window meets the set range, and when the counting signal reaches the set times, the patrol missile is in a straight level flight stage and meets the condition of relieving the second fuse of the fuze.
Compared with the prior art, the invention has the following remarkable advantages:
the invention provides a set of design method, according to the method, the airborne patrol missile fuze can directly utilize weak environment information in the flight process to remove the second insurance through the existing resources on the missile, no additional mechanism is needed to be designed, the overall volume of the fuze can be reduced, the intermediate flow is simplified, and the action reliability is improved; furthermore, the method is a much greater safety distance from the delivering aircraft than by relying on engine firing pulse loading.
Drawings
Fig. 1 is a process for arming a fuze using weak environmental information.
Fig. 2 is a rotation relationship diagram of coordinate axes of the orbit coordinate system and coordinate axes of the projectile coordinate system.
Fig. 3 shows the change of the attitude angle gliding and straight-line level flying of the airborne cruise missile.
Fig. 4 is a flow of amplitude determination.
Fig. 5 is a counting circuit principle.
Fig. 6 shows the variation amplitudes of the yaw angle (a), the pitch angle (b) and the roll angle (c) in the straight-line level flight phase.
Detailed Description
The invention is further described with reference to the following figures and embodiments.
With reference to fig. 1, the method for removing the second insurance of the airborne patrol missile fuze by means of weak environment information comprises the following steps:
the selected weak environment information is aerodynamic force borne by the cruise missile in a gliding stage and a straight-line level flight stage. Since the installation of pressure sensors on the body or wings of the projectile destroys the aerodynamic properties of the projectile, the aerodynamic forces cannot be directly recognized by the projectile-mounted sensors. However, the missile body is unbalanced under the action of aerodynamic force, and the missile-borne controller controls the rudder wings to deflect to realize the rebalancing of the missile body, so that the change condition of the aerodynamic force can be reflected by the change of the attitude angle of the missile body. The projectile attitude angle is an aerodynamic environmental characteristic value.
Step 1.1, as can be seen from fig. 3, the environmental characteristic values (referred to as pitch angle and attack angle change) have significant changes in the gliding and straight level flight phases. According to the diagram, in the glide phase (t ≦ 230 s), the variation range of the characteristic value is ± 12 °; in the straight-line level flight stage (t >230 s), the characteristic value is in a stable gliding state.
In the straight-line flat flight stage, the projectile body is in a balanced state by external force, and the change of the attitude angle of the projectile body is caused by the consumption of projectile fuel, so that the weight of the projectile is changed, and the attitude angle is changed. According to the method, the attitude angle change amplitude of the LOCAAS cruise missile in the straight-line level flight stage is calculated by referring to the LOCAAS cruise missile related parameters. The calculation formula is as follows:
wherein Δ M is the change in the weight of the projectile, g is the gravitational acceleration value (unit: kg/N), H W Is the oil consumption (unit: mg/N s), T of the engine N Is engine thrust (in kN), Δ T L Is a time interval (unit: s), delta Y is the variation of the lifting force of the projectile (unit: N), delta alpha is the variation of the angle of attack (unit:degree),
is a lift coefficient caused by a unit angle of attack angle deflection>
Is the lift coefficient caused by the unit angle of deflection of the pitching rudder, and rho is the air density (unit kg/m) 3 )，v 2 Is the flying speed (unit: ma), S is the missile wing characteristic area (unit m) 2 )。
The oil consumption and thrust parameters of the engine refer to LOCAAS patrol missile TJ-50-12 type engine, H W ＝43.89，T N =0.267, take Δ T L =10, assuming that the height of the straight-line level flight phase is 1000m, the gravity acceleration is 9.8, the air density is 1.1111, the flight speed is 0.3, and the missile wing characteristic area is 0.052.
The calculations show that the projectile flight angle of attack varies by 0.02 ° over 10 s. In horizontal flight, the change of the attack angle of the projectile is the same as the change of the pitch angle. In summary, with 10s as the determination interval, the attitude angle change of the cruise missile reaches more than 10 degrees in the gliding stage, and the attitude angle change is 0.02 degree in the straight-line level flight stage, so that the change amplitude and the change frequency have obvious difference.
And step 1.2, the change of the attitude angle can be calculated through the acceleration and the angular speed of the projectile body, and the calculation precision of the attitude angle can be improved by introducing geomagnetic data. The identification parameters are thus determined as projectile acceleration, angular velocity and earth magnetic field.
And 2, selecting sensors for identification according to the physical parameters analyzed in the step 1, wherein the sensors are of different types, and the same detection principle is avoided.
2.1, selecting the existing MEMS IMU sensor in the flight control system: the MEMS triaxial accelerometer measures the acceleration of the projectile body, the MEMS triaxial gyroscope measures the angular velocity of the projectile body, and the MEMS triaxial magnetometer measures the magnetic field;
and 2.2, in order to increase the identification reliability, meet the requirement of redundancy safety of the fuze system and prevent the identification failure caused by the fault of a single sensor, establishing two independent sensor systems, wherein the sensors used in the two systems are the same.
The sensor system 1 adopts a sensor of an on-board flight control system, and the sensor system 2 adopts an MEMS integrated nine-axis sensor MPU-9250 (comprising a three-axis accelerometer, a three-axis gyroscope and a three-axis magnetometer), the volume of the sensor system is 3 multiplied by 1mm, and the power consumption is extremely low. In order to ensure the consistency of the measured data of the two systems, the installation position of the MPU-9250 and the position of the flight control sensor are symmetrical along the center of mass of the projectile body.
And 3, transmitting the acceleration, the angular velocity and the magnetic force data detected by the sensor to a missile-borne processor, and designing a data fusion algorithm to be Double extended Kalman filtering (Double-EKF). In order to avoid singularity phenomenon and calculate trigonometric function, an attitude quaternion is selected as an operation element, and the quaternion is solved and then converted into an Euler angle form, so that the attitude angle condition is conveniently expressed.
And 3.1, at least two coordinate systems are required to be established for determining the posture condition of the projectile body of the cruise projectile. Establishing a track coordinate system Ox 1 y 1 z 1 The origin of coordinates is the center of mass of the projectile, ox 1 The axis pointing from the origin in the direction of the earth's magnetic field, oz 1 The axis pointing from the origin to the centre of the earth, oy 1 The axis orientation is determined by the right hand rule.
In order to conveniently acquire data of acceleration, angular velocity and magnetic field, a projectile coordinate system Ox is established 2 y 2 z 2 The origin of coordinates is the center of mass of the projectile, ox 2 Axis pointing from origin to transverse axis of projectile, oy 2 The axis being perpendicular to the transverse plane of the projectile body and pointing in the direction of the longitudinal axis towards the ground, oz 2 The axial orientation is determined by the right-hand coordinate system.
The track coordinate system and the missile coordinate system have the coordinate origin at the center of mass of the missile, and the rotation angles between the three coordinate axes of the missile coordinate system and the three axes of the track coordinate system can be obtained by rotating the track coordinate system to the missile coordinate system, and the angles can be used for describing the attitude information of the missile. As shown in FIG. 2, the pivoting sequence is Z → X → Y. From the orbital coordinate system Ox 1 y 1 z 1 Firstly, the phi angle is rotated around the Z axis to obtain an intermediate coordinate system Ox 'y' Z 1 (ii) a Then rotating around the X axis
Angle finding the intermediate coordinate system Ox' y 2 z', finally rotating theta angle around the Y axis to obtain a projectile coordinate system Ox 2 y 2 z 2 . Converting the matrix: />
step 3.2, setting the system state variable as quaternion q, q = [ q ] 1 ,q 2 ,q 3 ,q 4 ] T Wherein q is 4 Is a quaternion scalar part, [ q ] 1 ,q 2 ,q 3 ] T = ρ is quaternion vector part; and (3) obtaining a one-step estimation value of the system state variable q at the moment k +1 by the formula (5):
in the formula f (q) k+1 ) Is a system discrete state equation formed by the angular velocity and quaternion at the moment k, is a one-step estimation value of a system state variable q at the moment k +1,
is the measured value of angular velocity at time k (unit: rad/s), beta k For the moment k the gyroscope measures the drift (unit: rad/s), w k Is a true value of the angular velocity at instant k and->
Δ T is the sensor sampling interval, I 4×4 Is a fourth order identity matrix.
w＝[w x w y w z ] T vector output for three-axis angular velocity:
and (3) updating gyro measurement drift:
β k+1 ＝β k +β biass
β biass random drift for the gyroscope;
and (3) obtaining a priori estimated value of the system state variable q at the moment k +1 through formulas (6) and (7):
in the formula
Is the optimal estimation value of the quaternion vector part at the moment k, I 3×3 Is a third order identity matrix, q k+1|k Is a prior estimated value of a system state variable at a moment k +1, and the norm of the prior estimated value is as follows:
to q is k+1|k After normalization treatment, the product is obtained
F k Is the system state equation f (q) k+1 ) Of Jacobi matrix, η v Noise is measured for the gyroscope.
The prior prediction of the noise covariance matrix P of the extended Kalman filter at the moment k + 1:
P k+1|k is a 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.
Step 3.3, establishing a measurement updating equation of the system state variable according to the measured acceleration and magnetic force data at the moment k +1, and calculating the optimal estimation value of the system state variable at the moment k +1 through formulas (9) - (21):
due to the conversion matrix from the orbit coordinate system to the projectile coordinate system
The method is in an euler angle form, and is not beneficial to quaternion calculation, so that the method needs to convert the attitude conversion matrix into a quaternion form, and the A (q):
step 3.3.1, correcting the data of the acceleration sensor for one time:
estimating the acceleration of the projectile at the moment k + 1:
in the formula
Is determined by>
The estimated value of projectile acceleration at the moment k +1, [00 survival g] T Is to consider the influence of gravity acceleration on an estimated value in the vertical direction under natural conditions, an orbital coordinate system Ox 1 y 1 z 1 Is directed in the geocentric/gravitational direction, and therefore has a value of | g | only in the z-axis direction.
H k+1'1 Is that
The Jacobi matrix of (1).
In the formula K k+1'1 Is the Kalman gain value, R, of the acceleration data to the system state variable 1 Is the acceleration sensor measurement noise covariance matrix.
z k+1'1 Is the actual measurement value of the acceleration sensor at the time k + 1, z k+1'1 ＝[a k+1,x a k+1,y a k+1,z ] T And carrying out normalization treatment on the obtained product:
using estimated value of projectile acceleration at the moment K +1, actual measured value of acceleration sensor and Kalman gain K k+1'1 Calculating a quaternion correction q ε1 ：
In the formula q ε1 Is the quaternion correction quantity obtained according to the measurement data of the acceleration sensor,
is a quaternion value corrected by the acceleration data, I is a unit matrix, and->
And the state estimation error covariance matrix after the acceleration data is corrected.
Step 3.3.2, performing magnetometer data secondary correction:
estimating the earth magnetic field at the k +1 moment:
in the formula
Is determined by>
The derived estimated value of the magnetic field at the time k + 1, [ 10] T Considering the natural situation (assuming that the magnetic field strength is 1), the direction of the earth magnetic field is the north-south direction, and the orbit coordinate system Ox 1 y 1 z 1 The x-axis direction of (b) points in the direction of the geomagnetism (i.e., the north-south direction), and therefore the influence on the magnetic field data is considered only in the x-axis direction.
H k+1'2 Is that
The Jacobi matrix of (c).
K k+1'2 Is the Kalman gain value, R, of the magnetic data to the system state variable 2 Is the magnetometer measures the noise covariance matrix.
z k+1'2 Is the actual measurement value of the magnetometer at the time k + 1, z k+1'2 ＝[m k+1,x m k+1,y m k+1,z ] T And carrying out normalization treatment on the obtained product:
using the estimated value of the magnetic field at the moment K +1, the actual measured value of the magnetometer and the Kalman gain K k+1'2 Calculating a quaternion correction q ε2 ：
In the formula q ε2 Is a quaternion correction based on 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.
And 3.4, the quaternion can not directly describe the change condition of the attitude angle of the projectile body, and the change condition needs to be converted into an Euler angle form. The quaternion optimal estimation value at the moment of the above equation k +1 is converted into the euler angle equation as follows:
roll k+1 ,pitch k+1 ,yaw k+1 the roll angle, the pitch angle and the yaw angle of the projectile at the moment k +1 are respectively, wherein 180/pi is used for realizing unit conversion between radian and degree.
Step 5.1, obtaining a pitch angle change range 1 according to a pitch angle calculation result of the cruise missile in the straight level flight stage, summing the range 1 and the range 2 by combining an attack angle change range 2 caused by fuel consumption in the straight level flight stage to obtain a final range which is used as an amplitude basis for distinguishing the gliding stage and the straight level flight stage;
and 5.2, setting a time judgment interval to be t =1s by considering the resolving speed of the processor, the timeliness of the second insurance relief and the change curve of the attack angle, calculating the maximum amplitude difference of the attitude angle in the interval and storing the maximum amplitude difference into a storage unit, wherein the data acquisition interval is 0.1 s.
And 6, formulating an amplitude judgment flow according to the amplitude range set in the step 5 and the duration of the judgment interval.
Step 6.1, the amplitude judgment flow is shown in an attached figure 4, continuous 3 intervals are selected as a judgment period, if any amplitude in the period does not meet the set requirement, a judgment window is opened, and a high level is sent to a zero clearing end of the counter;
and 6.2, in the window, if all the amplitude values in the first period meet the set requirement, performing the step 7, and otherwise, continuing the step 6.1.
Step 7, when all the amplitudes in a period meet the requirements, a high level is sent to a counting end of the counter through the fuze control circuit, and the window is continued; if the amplitude value in the next period does not meet the set requirement according to the time sequence, ending the window, sending a high level to a zero clearing end of the counter, and returning to the step 6.1.
When the patrol missile encounters long-term airflow interference in the straight-line level flight stage, the attitude angle is changed violently, and the process returns to the step 6 again, so that the fuse circuit misjudges the stage of the patrol missile, and the second fuse of the fuse is released and delayed. However, the attitude angle is stabilized again with the adjustment of the flight control system on the missile, and the judgment process is continued.
And 8, setting the counting standard of each counter to be 10, wherein the principle of the counter is shown in figure 5, and when any one of the two counters meets the requirement, outputting a control signal and removing the second fuse of the fuse.
Example 1
A group of static 14h MEMS IMU data is selected and solved through a Double-EKF algorithm to obtain the absolute error of the attitude angle (pitch angle) estimation value, and the amplitude judgment range is helped to be determined. The sampling interval of the sensor data is set to 0.1s, and the initial value q of the quaternion is set 0 ＝[0 0 0 T Initial value of gyroscope bias
β biass ＝1e-19，η v =1e-13, system noise covariance matrix Q = [0.010.010.010.01 =] T The acceleration sensor measures the covariance matrix R of the noise 1 ＝[0.10.10.1] T Magnetometer measures the noise covariance matrix R 2 ＝[0.0010.0010.001] T . And (4) settling attitude angle data in 14h through formulas (4) to (24), as shown in figure 6 ((a) is a yaw angle calculation result, (b) is a pitch angle calculation result, and (c) is a roll angle calculation result). The pitch angle is used as a judgment reference (the attack angle is equal to the pitch angle in horizontal flight), and the variation range of the pitch angle is +/-0.04 degrees; the variation range of the roll angle is-0.25-0.15 degrees, and the variation range of the yaw angle is-0.2-0.5 degrees. The attitude angle amplitude standard was set to ± 0.1 ° (0.08 ° +0.02 °).
Those matters not described in detail in the present specification are well known in the art to which the skilled person pertains.
Claims (6)
1. A method for releasing the second fuse of an airborne flying bomb fuse by means of weak environment information is characterized in that a pitch angle is used as the environment information for releasing the second fuse of the fuse,
when the pitch angle is larger than the set range, the patrol missile is in a gliding stage, a judgment window is entered, a counting signal is given out when the pitch angle in the judgment window meets the set range, and when the counting signal reaches the set times, the patrol missile is in a straight level flight stage and meets the condition of relieving the second fuse of the fuze.
2. The method for releasing the second insurance of the airborne patrol bomb fuze according to the weak environment information in claim 1, wherein the change range of the pitch angle in the straight-line level flight stage is used as a setting range: the pitch angle variation range of the straight-line level flight stage is determined by two aspects: on one hand, the calculation accuracy range 1 of the pitch angle is obtained, and on the other hand, the variation range 2 of the attitude angle caused by fuel consumption of the cruise missile is obtained; the final range is obtained by summing range 1 and range 2;
the pitch angle calculation formula is as follows:
wherein the pitch k+1 The value of the pitch angle at the moment k +1 is normalized,
the attitude angle change caused by fuel consumption is calculated by the formula:
wherein, delta alpha is the change of an attack angle (the change of the attack angle is equal to the change of a pitch angle during horizontal flight), delta Y is the change of a lifting force of the projectile body,
is a lift coefficient caused by a unit angle of attack angle deflection>
Is the lift coefficient caused by the unit angle of deflection of the pitching rudder, rho is the air density, v 2 The flying speed is S, and the characteristic area of the missile wing is S.
3. The method for relieving second insurance of the airborne patrol missile fuze by relying on weak environment information according to claim 2, wherein the pitch angle calculation is obtained by the following steps:
3.1, firstly, acquiring the acceleration, the angular velocity and the data of the earth magnetic field of the flying projectile;
3.2, obtaining the attitude information of the projectile body through the conversion of coordinate axes between the track coordinate system and the projectile body coordinate system;
3.3, setting a quaternion as a system variable, solving a priori estimation value of the quaternion at the moment k +1 by using the optimal estimation value of the quaternion at the moment k and projectile body angular velocity data at the moment k, and correcting the priori estimation value by using acceleration data and magnetic force data to obtain the optimal estimation value of the quaternion at the moment k + 1;
and 3.4, obtaining the pitch angle value at the k +1 moment by using the optimal estimation value of the quaternion at the k +1 moment.
4. The method for releasing the second insurance of the fuse of the airborne patrol missile according to the weak environment information in claim 2, wherein the mass change value of the patrol missile caused by fuel consumption in the straight-line level flight stage can be obtained by the following formula:
ΔY＝ΔM×g
where Δ M is the projectile body weight change, g is the gravitational acceleration value, H W For engine oil consumption, T N As engine thrust, Δ T L Is a time interval.
5. The method for releasing the second insurance of the fuse of the airborne patrol missile according to the weak environment information of the claim 1, wherein the data of the acceleration, the angular velocity and the earth magnetic field of the patrol missile are collected by an existing MEMSIMU sensor in the flight control system.
6. The method for releasing the second insurance of the fuse of the airborne patrol missile according to the weak environment information of claim 5, wherein the data of the acceleration, the angular velocity and the earth magnetic field of the patrol missile are collected by an integrated nine-axis sensor.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202211593767.0A CN115900459A (en) | 2022-12-13 | 2022-12-13 | Method for relieving second insurance of airborne patrol missile fuze by means of weak environment information |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202211593767.0A CN115900459A (en) | 2022-12-13 | 2022-12-13 | Method for relieving second insurance of airborne patrol missile fuze by means of weak environment information |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN115900459A true CN115900459A (en) | 2023-04-04 |
Family
ID=86485329
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202211593767.0A Pending CN115900459A (en) | 2022-12-13 | 2022-12-13 | Method for relieving second insurance of airborne patrol missile fuze by means of weak environment information |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN115900459A (en) |
-
2022
- 2022-12-13 CN CN202211593767.0A patent/CN115900459A/en active Pending
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| Siouris | Missile guidance and control systems | |
| US11898828B1 (en) | Closed, self-contained ballistic apogee detection module and method | |
| Kain et al. | Rapid transfer alignment for tactical weapon applications | |
| US6883747B2 (en) | Projectile guidance with accelerometers and a GPS receiver | |
| CN110017830B (en) | Method for resolving aircraft attitude by using geomagnetic information and gravity sensor | |
| CN111238474A (en) | Strapdown seeker non-singular line-of-sight angular velocity extraction method based on inclined coordinate system | |
| Roh et al. | Dynamic accuracy improvement of a MEMS AHRS for small UAVs | |
| JP5338464B2 (en) | Inertial navigation device, flying object, and navigation data calculation method | |
| CN110017808B (en) | Method for resolving aircraft attitude by using geomagnetic information and accelerometer | |
| Ilg | Guidance, navigation, and control for munitions a thesis submitted to the faculty of | |
| Parfiriev et al. | Model of inertial navigation system for unmanned aerial vehicle based on MEMS | |
| Pamadi et al. | Assessment of a GPS guided spinning projectile using an accelerometer-only IMU | |
| CN115900459A (en) | Method for relieving second insurance of airborne patrol missile fuze by means of weak environment information | |
| CN115615261B (en) | Fusion method for rocket elastic identification with large slenderness ratio and line-of-sight angular rate extraction | |
| Kim et al. | Guidance and control system design for impact angle control of guided bombs | |
| Ohlmeyer et al. | Guidance, navigation and control without gyros: A gun-launched munition concept | |
| de Celis et al. | Neural Network-Based Controller For Terminal Guidance Applied In Short-Range Rockets | |
| CN112937832A (en) | Air-drop type unmanned aerial vehicle and throwing method thereof | |
| Milinović et al. | Experimental and simulation testing of flight spin stability for small caliber cannon projectile | |
| Lu et al. | Loitering Missile Environment Information analysis and application in Fuze S&A | |
| Pingan et al. | Attitude measurement of special aircraft based on geomagnetic and angular velocity sensors | |
| Zhang et al. | Attitude Measurement of Special Aircraft Based on Geomagnetic and Angular Velocity Sensors | |
| Xu et al. | Influence of random disturbance on hit accuracy of Rocket Launcher | |
| Pingan et al. | Research Article Attitude Measurement of Special Aircraft Based on Geomagnetic and Angular Velocity Sensors | |
| Li et al. | Research on aerial alignment method of guided projectiles based on INS/radar command combination |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination |