CN111102886A - Gyro-free accurate guidance method for small micro aircraft - Google Patents
Gyro-free accurate guidance method for small micro aircraft Download PDFInfo
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F42—AMMUNITION; BLASTING
- F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
- F42B15/00—Self-propelled projectiles or missiles, e.g. rockets; Guided missiles
- F42B15/01—Arrangements thereon for guidance or control
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- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course or altitude of land, water, air, or space vehicles, e.g. automatic pilot
- G05D1/0088—Control of position, course or altitude of land, water, air, or space vehicles, e.g. automatic pilot characterized by the autonomous decision making process, e.g. artificial intelligence, predefined behaviours
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- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course or altitude of land, water, air, or space vehicles, e.g. automatic pilot
- G05D1/10—Simultaneous control of position or course in three dimensions
- G05D1/101—Simultaneous control of position or course in three dimensions specially adapted for aircraft
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- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course or altitude of land, water, air, or space vehicles, e.g. automatic pilot
- G05D1/10—Simultaneous control of position or course in three dimensions
- G05D1/107—Simultaneous control of position or course in three dimensions specially adapted for missiles
Abstract
The invention discloses a gyro-free precise guidance method for a small micro aircraft, which belongs to the technical field of flight navigation and guidance of aircrafts and is characterized in that a gyroscope or an inertia combined navigation device is not needed to measure attitude information of the aircrafts, only a seeker or a third-party measuring station is used for providing line-of-sight angle information relative to a target, approximate differentiation and integration are carried out on the line-of-sight angle information, a nonlinear anti-saturation differential signal and a nonlinear anti-saturation integral signal are obtained, and finally signal synthesis is carried out, so that a guidance signal can be formed to drive a stable-attitude tracking control loop of the aircrafts, and the precise guidance of the target is realized. The method can solve the problem of accurate guidance of the small micro aircraft under the condition that a gyroscope or other inertia components are inconvenient to install.
Description
Technical Field
The invention belongs to the field of flight control of aircrafts, and particularly relates to a method for realizing control and accurate guidance of small and miniature aircrafts without using a gyroscope.
Background
The precise guidance technology of the aircraft is a high-tech technology for military and civil use. The conventional methods include proportional guidance, pre-guidance, tracking guidance, etc., or modifications and variations of the above guidance methods. The large aircraft has enough space, so a gyroscope or an inertia combined navigation component can be installed, and by adopting the guidance method, the self attitude of the aircraft needs to be measured for comparison if the aircraft is guided in the front position, so that the final accurate guidance is realized. Small micro aircrafts such as small guided missiles, small electromagnetic cannons, police micro patrol helicopters and other micro aircrafts do not have the conditions for installing gyroscopes or inertia combined navigation components due to narrow space, but have the requirement of precise guidance, so the conventional guidance method cannot be adopted at all. Based on the background, the invention provides a method for forming the accurate guidance law only by adopting the sight angle information provided by the seeker or a third-party measuring station and performing simple computer calculation. Therefore, the method has theoretical innovativeness, and can solve the problem of guidance of a large and small type of micro aircrafts, so that the method has high economic value and engineering application value.
It is to be noted that the information invented in the above background section is only for enhancing the understanding of the background of the present invention, and therefore, may include information that does not constitute prior art known to those of ordinary skill in the art.
Disclosure of Invention
The invention aims to provide a gyro-free precise guidance method for a small micro aircraft, and further solves the problem that stable control and guidance can not be carried out due to the fact that inertial navigation equipment such as a gyroscope cannot be installed due to the narrow space of the small micro aircraft at least to a certain extent.
The invention provides a gyro-free accurate guidance method for a small micro aircraft, which comprises the following steps:
step S10: the method comprises the steps that a micro seeker is installed to measure the line-of-sight angle information between a small micro aircraft and a target, or a radio receiving device is installed to receive the line-of-sight angle information between the target and the small micro aircraft sent by a third-party navigation station measuring device;
step S20: constructing a differential equation according to the sight angle information of the aircraft and the target, and solving an approximate differential signal of the sight angle;
step S30: performing integral operation according to the line-of-sight angle signal to obtain a line-of-sight angle integral signal, and performing anti-saturation operation according to the line-of-sight angle approximate differential signal to obtain an anti-saturation differential signal of the line-of-sight angle;
step S40: according to the line-of-sight angle integral signal, carrying out nonlinear anti-saturation operation to obtain an integral anti-saturation signal of a line-of-sight angle;
step S50: designing a nonlinear transformation function according to the line-of-sight angle signal to obtain a line-of-sight angle nonlinear transformation signal, so that when the line-of-sight angle signal is large, turning is accelerated, the efficiency of a guidance law is increased, and the guidance precision is increased;
step S60: the line-of-sight angle signal, the anti-saturation differential signal and the anti-saturation integral signal are comprehensively synthesized with the line-of-sight angle nonlinear transformation signal and are directly input to an aircraft stabilizing system, so that accurate guidance between the aircraft and a target is realized.
In an exemplary embodiment of the invention, constructing a differential equation according to the information of the line-of-sight angle of the aircraft from the target, and solving an approximate differential signal of the line-of-sight angle includes:
dq(n+1)＝dq(n)+ddq(n)*Δt；
where Δ t is the difference step, and is generally selected to be 0.001 per second. Wherein the initial values are set as: q. q.sb(0)＝qb(1) (ii) a dq (1) is 0. And q isb(1) Is selected according to qbIs determined by the 1 st data of (a), qb(n) is selected according to qbIs determined by the nth data of (a). The time interval between each datum is Δ t. Wherein T is1、T2For the filter time constant, dq (n +1) is the approximate differential signal of the solved line-of-sight angle, ddq (n) is the intermediate variable of the transition, dqb(n) is an input signal qb(n) digital differentiation, qbAnd (n) is a line-of-sight angle signal of the aircraft from the target.
In an exemplary embodiment of the present invention, the anti-saturation operation is performed according to the line-of-sight angle approximate differential signal to obtain an anti-saturation differential signal of the line-of-sight angle:
wherein dq is said line-of-sight angle approximately differential signal, k1、k2And epsilon1Is a constant positive parameter against saturation, udIs the anti-saturation differential signal of the line of sight angle.
In an exemplary embodiment of the present invention, the performing integration according to the line of sight angle to obtain a line of sight angle integrated signal, and then performing a nonlinear anti-saturation operation to obtain an anti-saturation integrated signal of the line of sight angle includes:
Sq＝∫qbdt；
wherein k is3、k4And epsilon2As a constant positive parameter against saturation, qbFor line-of-sight angle signals, SqFor line-of-sight angle integral signal, usThe signal is integrated for line-of-sight angles against saturation.
In an exemplary embodiment of the present invention, designing a nonlinear transformation function according to the line-of-sight angle signal, and obtaining the line-of-sight angle nonlinear transformation signal includes:
wherein k is5、k6And epsilon3For non-linear transformation of parameters, qbFor line-of-sight angle signals, uqThe signal is integrated for line-of-sight angles against saturation.
In an exemplary embodiment of the invention, the line-of-sight angle signal q is dependent on the first signalbAnti-saturation differential signal udAnti-saturation integrated signal usAngle of sight non-linear transformation signal uqPerforming comprehensive processing to obtain a final pilot law comprehensive signal, including:
u＝k7qb+k8ud+k9us+k10uq；
where u is the final pilot law combined signal, k7、k8、k9、k10For the guided constant, qbIs a line-of-sight angle signal udFor anti-saturation differential signal usIntegrating the signal u for anti-saturationqThe signal is transformed for line-of-sight non-linearity. And finally, transmitting the guidance law output signal u to an aircraft stable tracking loop, so that the aircraft can accurately guide the target.
The invention provides a gyro-free accurate guidance method for a small micro aircraft, which is mainly characterized in that no attitude signal of the aircraft is applied in a guidance signal, so that a gyro is not needed for measuring the attitude signal of the aircraft, and the method has great superiority for a plurality of simple aircrafts, or small guided cannonballs or electromagnetic cannonballs. Because the small micro aircraft cannot measure the attitude of the aircraft according to the inertial component no matter in space or cost, the gyro-free accurate guidance method provided by the invention is a very suitable choice. Therefore, the invention not only has novel characteristics in theory, but also can solve the problem of accurate guidance of the small micro aircraft under the condition that the inertia components are inconvenient to install. Therefore, the invention has high theoretical and engineering practical value and can be widely applied to the precise guidance of small and miniature aircrafts, small and miniature electromagnetic cannon weapons and small and miniature cannonballs.
It is to be understood that both the foregoing general description and the following detailed description are exemplary and explanatory only and are not restrictive of the invention, as claimed.
Drawings
The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments consistent with the invention and together with the description, serve to explain the principles of the invention. It is obvious that the drawings in the following description are only some embodiments of the invention, and that for a person skilled in the art, other drawings can be derived from them without inventive effort.
FIG. 1 is a flow chart of a design implementation of a gyro-free precision guidance method for a small micro aircraft provided by the invention;
FIG. 2 is a graph of the relative motion of an aircraft and a target in a course plane (in meters) according to the method provided by the embodiment of the invention;
FIG. 3 is a miss-measure curve (in meters) for a method provided by an embodiment of the invention;
FIG. 4 is a graph of end-of-range magnification (in meters) for the amount of miss in a method provided by an embodiment of the invention;
FIG. 5 is a graph of actual yaw angle versus desired yaw angle (in degrees) for a method provided by an embodiment of the present invention;
FIG. 6 is a graph of the output of the steering law in degrees/second according to the method of the present invention;
FIG. 7 is a graph of aircraft sideslip angle (in degrees/second) for a method provided by an embodiment of the present invention;
FIG. 8 illustrates an aircraft yaw rudder deflection angle curve (in degrees) for a method provided by an embodiment of the present invention;
FIG. 9 is a graph of the change in the angle of sight of an aircraft and a target according to the method of the present invention (in meters);
FIG. 10 shows a non-linear transformation of the signal (unit: degree) from the line of sight according to the method of the present invention;
FIG. 11 shows a line-of-sight approximately differential signal (in degrees/second) of a method provided by an embodiment of the invention;
FIG. 12 is a graph of line-of-sight angular anti-saturation differential signal (in degrees/sec) of a method provided by an embodiment of the present invention;
FIG. 13 is a graph of the line-of-sight anti-saturation integrated signal (in degrees/sec) of the method provided by an embodiment of the present invention;
FIG. 14 is a graph of the line-of-sight angle integrated signal (in degrees/sec) of the method provided by the embodiment of the present invention.
Detailed Description
Example embodiments will now be described more fully with reference to the accompanying drawings. Example embodiments may, however, be embodied in many different forms and should not be construed as limited to the examples set forth herein; rather, these embodiments are provided so that this disclosure will be thorough and complete, and will fully convey the concept of example embodiments to those skilled in the art. The described features, structures, or characteristics may be combined in any suitable manner in one or more embodiments. In the following description, numerous specific details are provided to provide a thorough understanding of embodiments of the invention. One skilled in the relevant art will recognize, however, that the invention may be practiced without one or more of the specific details, or with other methods, components, devices, steps, and so forth. In other instances, well-known technical solutions have not been shown or described in detail to avoid obscuring aspects of the invention.
The invention relates to a gyro-free precise guidance method for a small micro aircraft, which is characterized in that sight angle information of a target is provided only through a seeker or a third-party measuring station, approximate differentiation and integration are carried out on the sight angle information, a nonlinear anti-saturation differential signal and a nonlinear anti-saturation integral signal are obtained, and finally signal synthesis is carried out, so that a guidance signal can be formed to drive a stable tracking control loop of the aircraft, and the precise guidance of the target is realized. The method is characterized in that in all the information of guidance, the attitude information of the aircraft does not need to be acquired, so compared with the traditional preposed guidance, tracking guidance and the like, the method does not need to install a gyroscope or an inertia combined navigation device to measure the attitude information of the aircraft, is particularly suitable for small and micro aircrafts, can save a large amount of space, reduces the volume of the aircraft, reduces the cost of the aircraft, better utilizes the miniaturization and the microminiaturization of the aircraft, and has very wide application prospect.
Hereinafter, a gyro-less precision guidance method for a small micro-aircraft according to an exemplary embodiment of the present invention will be explained and explained with reference to the drawings. Referring to fig. 1, a gyro-less precision guidance method for a small micro aircraft may include the following steps:
step S10: and the micro seeker is arranged to measure the sight angle information between the small micro aircraft and the target, or the radio receiving equipment is arranged to receive the sight angle information between the target and the small micro aircraft sent by the third-party navigation station measuring equipment.
Specifically, the invention provides the following two ways to obtain the line-of-sight angle information between the aircraft and the target. The first is a slightly larger aircraft, if space permits, a miniature seeker can be mounted to measure the line-of-sight angular rate of relative motion between the aircraft and the target, denoted as
The viewing angle is then integrated from the viewing angle rate to obtain the viewing angle, which is denoted as qbWhich satisfies
Where dt represents the integration of the time signal.
The second type is a small aircraft, and because the size is small, measurement equipment such as a guide head cannot be installed on the aircraft, wireless receiving equipment can be installed, and guide information sent by third-party navigation station measurement equipment can be received. Namely from the thirdMeasuring the sight angle q of the aircraft relative to the target by the square navigation stationbAnd then sent to the small aircraft.
Step S20: and constructing a differential equation according to the sight angle information of the aircraft and the target, and solving an approximate differential signal of the sight angle.
Since the resolution of the precise differential signal is not easy to implement, the present invention uses an approximate differential approach to implement the precise pilot signal. Of course, if the line of sight angular rate signal can be provided directly from the seeker signal measurement, then the solution of the exact differential here is not required. But the signal measured by the seeker is also actually an approximately differential signal.
For the second mode, i.e. the third-party navigation station transmits the transmitted line-of-sight information qbAfter the small flyer is fed, the small micro-aircraft carries out approximate differential signal solving according to the following mode. The approximately differentiated signal of the line of sight angle is denoted dq here, and its calculation is performed according to the following difference equation:
dq(n+1)＝dq(n)+ddq(n)*Δt；
where Δ t is the difference step, and is generally selected to be 0.001 per second. Wherein the initial values are set as: q. q.sb(0)＝qb(1) (ii) a dq (1) is 0. And q isb(1) Is selected according to qbIs determined by the 1 st data of (a), qb(n) is selected according to qbIs determined by the nth data of (a). The time interval between each datum is Δ t. Wherein T is1、T2For the filtering time constant, see the following example implementation.
Step S30: and performing integral operation according to the line-of-sight angle signal to obtain a line-of-sight angle integral signal, and performing anti-saturation operation according to the line-of-sight angle approximate differential signal to obtain an anti-saturation differential signal of the line-of-sight angle.
Specifically, firstly, according to the line-of-sight angle information, the line-of-sight angle integral information is solved by an accumulation method, defined as Sq, and the integral equation is solved to satisfy the following conditions: sq＝∫qbdt, dt represents integration over time.
Next, in order to avoid the influence of the excessive output of the differential signal on the steering accuracy, the differential signal after the anti-saturation processing is subjected to the anti-saturation processing by using the following nonlinear function, and the differential signal after the anti-saturation processing is referred to as an anti-saturation differential signal and is recorded as ud. The anti-saturation algorithm proceeds as follows:
wherein k is1、k2And epsilon1For the anti-saturation parameters, the physical meaning and selection principle are illustrated in the examples hereinafter.
Step S40: and carrying out nonlinear anti-saturation operation according to the line-of-sight angle integral signal to obtain an integral anti-saturation signal of the line-of-sight angle.
If the integrated signal is directly applied to the pilot law, the problem that the signal is too large due to saturation exists. Therefore, the anti-saturation processing is also performed by using the following nonlinear function, and the integrated signal after the anti-saturation processing is referred to as an anti-saturation integrated signal and is denoted as us. The anti-saturation algorithm proceeds as follows:
wherein k is3、k4And epsilon2For the anti-saturation parameters, the physical meaning and selection principle are illustrated in the examples hereinafter.
Step S50: and designing a nonlinear transformation function according to the line-of-sight angle signal, so that when the line-of-sight angle signal is large, turning is accelerated, the efficiency of a guidance law is increased, and the guidance precision is increased.
The line-of-sight angle signal can be applied directly to the lead law, while the following non-linear processing can be used, indicating if the line-of-sight angle signal | qb| is greater than ε3Then, k is superimposed on the basis of the original pilot law signal5And (4) quickening the turning speed. Specifically, the anti-saturation processing may be performed by using a nonlinear function, and the line-of-sight angle signal after the anti-saturation processing is referred to as a line-of-sight nonlinear conversion signal and is denoted as uq. Although the same anti-saturation function is used for the path signal, the effect is that non-linear transformation is performed in addition to anti-saturation, and thus the path signal is referred to as a line-of-sight non-linear transformation signal. The nonlinear transformation algorithm is carried out according to the following equation:
wherein k is5、k6And epsilon3The non-linear transformation parameters are selected in detail as will be described in the following examples.
Step S60: the line-of-sight angle signal, the anti-saturation differential signal and the anti-saturation integral signal are comprehensively synthesized with the line-of-sight angle nonlinear conversion signal and are directly input to an aircraft stabilizing system, and accurate guidance between the aircraft and a target is realized.
Specifically, for the above-mentioned line-of-sight angle signal qbAnti-saturation differential signal udAnti-saturation integrated signal usAngle of sight non-linear transformation signal uqAnd carrying out comprehensive processing to obtain a final guide law comprehensive signal. The linear synthesis is carried out as follows:
u＝k7qb+k8ud+k9us+k10uq；
wherein the parameter k7、k8、k9、k10The detailed selection is described in the examples below.
And finally, transmitting the guidance law output signal u to an aircraft stable tracking loop, so that the aircraft can accurately guide the target. Because the aircraft stable tracking loop has various modes, such as an attitude stable tracking loop, an overload stable tracking loop, a bus stable tracking loop and the like, the design of the aircraft stable tracking loop has published teaching materials or book descriptions, and the design is not the protection content of the invention, so the aircraft stable tracking loop is not described in detail herein.
Case implementation and computer simulation result analysis
First, the aircraft speed is set to about 220 m/s. Setting the three-dimensional coordinates of the initial position of the aircraft to be (0, 0, 0) and the three-dimensional coordinates of the initial position of the target to be (3500, 1, -500). Where 1 is the altitude of the target, -500 is the initial deviation of the target, and 3500 is the initial distance of the aircraft from the target. The target speed is set to move in the horizontal plane, the speed is 18m/s, and the direction is-22 degrees to the direction of the x axis.
And (5) the line-of-sight angle signal acquisition method in the step one does not need to be additionally described. Selecting T in the second step1＝0.4、T20.1. The integration in the third step is performed by adopting a general Euler cumulative method, and k is selected1＝0.3、k20.2 and ε125. Selecting k from the fourth step3＝0.5、k40.1 and ε20.5. Selecting k in step five5＝0.2、k60.1 and ε35. Selecting k in the sixth step7＝3、k8＝1、k9＝1、k10＝0.2。
The final case implementation results are shown in fig. 2 to 14 below. Fig. 2 is a projection of the equivalent motion curve of the aircraft and the target in the horizontal plane. Fig. 3 and 4 are guide miss-hits and enlarged curves, and it can be seen that the final miss-hits are less than 1.5 m, so the present invention can accurately hit targets with sizes greater than 2 m, both in motion and delicate. Fig. 5 is a comparison curve of the actual yaw angle and the expected yaw angle of the method provided by the present invention, and fig. 6 is an output curve of the guidance law provided by the present invention, so that it can be seen that the output of the guidance law is smooth, and the amplitude can meet the tracking requirement of the attitude stabilization loop.
Fig. 7 and 8 are actual sideslip angle and rudder deflection angle generated in the process of implementing the above guidance of the aircraft, respectively, and it can be seen that the problem of excessive sideslip angle and excessive rudder deflection caused by the guidance law output provided by the present invention is not caused. Fig. 9 is a line-of-sight angle curve during guidance of the aircraft, showing the rapid tip enlargement. Fig. 10 is a nonlinear transformation curve of the view angle, and it can be seen that the final output has a smaller amplitude after nonlinear transformation, and has an anti-saturation function. Fig. 11 is a view angle approximately differential signal, fig. 12 is a view angle anti-saturation differential signal, and it can be seen by comparison that the approximately differential signal still has the problem of too large amplitude, and the anti-saturation differential signal has a bounded output. Fig. 13 and 14 are respectively a line-of-sight angle-integrated signal and a line-of-sight angle-integrated anti-saturation signal, and it can be seen that the output of the integrated anti-saturation signal is also bounded. Although the amplitude of the signals is small, the signals are combined to achieve fine adjustment of the guidance precision of the target, and therefore high-precision guidance is achieved finally. It is worth to be noted that the method is completed by adopting a six-degree-of-freedom nonlinear model of the aircraft which is almost close to a real model, so that a curve of dynamic transformation of the sideslip angle and the rudder deflection angle of the aircraft can be reflected. Therefore, the case result is closer to the result of the real guide, and the reliability is high. In conclusion, the invention provides an accurate guidance method without a gyro or inertial navigation combination, and the method has great innovativeness and high engineering value.
Other embodiments of the invention will be apparent to those skilled in the art from consideration of the specification and practice of the invention disclosed herein. This application is intended to cover any variations, uses, or adaptations of the invention following, in general, the principles of the invention and including such departures from the present disclosure as come within known or customary practice within the art to which the invention pertains. It is intended that the specification and examples be considered as exemplary only, with a true scope and spirit of the invention being indicated by the following claims.
Claims (6)
1. A gyro-free precise guidance method for a small micro aircraft is characterized by comprising the following steps:
step S10: the method comprises the steps that a micro seeker is installed to measure the line-of-sight angle information between a small micro aircraft and a target, or a radio receiving device is installed to receive the line-of-sight angle information between the target and the small micro aircraft sent by a third-party navigation station measuring device;
step S20: constructing a differential equation according to the sight angle information of the aircraft and the target, and solving an approximate differential signal of the sight angle;
step S30: performing integral operation according to the line-of-sight angle signal to obtain a line-of-sight angle integral signal, and performing anti-saturation operation according to the line-of-sight angle approximate differential signal to obtain an anti-saturation differential signal of the line-of-sight angle;
step S40: according to the line-of-sight angle integral signal, carrying out nonlinear anti-saturation operation to obtain an integral anti-saturation signal of a line-of-sight angle;
step S50: designing a nonlinear transformation function according to the line-of-sight angle signal to obtain a line-of-sight angle nonlinear transformation signal, so that when the line-of-sight angle signal is large, turning is accelerated, the efficiency of a guidance law is increased, and the guidance precision is increased;
step S60: the line-of-sight angle signal, the anti-saturation differential signal and the anti-saturation integral signal are comprehensively synthesized with the line-of-sight angle nonlinear transformation signal and are directly input to an aircraft stabilizing system, so that accurate guidance between the aircraft and a target is realized.
2. The gyroscopic-free precision guidance method for the small micro aircraft as claimed in claim 1, wherein a differential equation is constructed according to the line-of-sight angle information of the aircraft from the target, and solving the approximate differential signal of the line-of-sight angle comprises:
dq(n+1)＝dq(n)+ddq(n)*Δt；
where Δ t is the difference step, and is generally selected to be 0.001 per second. Wherein the initial values are set as: q. q.sb(0)＝qb(1) (ii) a dq (1) is 0. And q isb(1) Is selected according to qbIs determined by the 1 st data of (a), qb(n) is selected according toqbIs determined by the nth data of (a). The time interval between each datum is Δ t. Wherein T is1、T2For the filter time constant, dq (n +1) is the approximate differential signal of the solved line-of-sight angle, ddq (n) is the intermediate variable of the transition, dqb(n) is an input signal qb(n) digital differentiation, qbAnd (n) is a line-of-sight angle signal of the aircraft from the target.
3. The gyroscopic-free precise guidance method for the small micro aircraft as claimed in claim 2, characterized in that anti-saturation operation is performed according to the line-of-sight angle approximate differential signal to obtain an anti-saturation differential signal of the line-of-sight angle:
wherein dq is said line-of-sight angle approximately differential signal, k1、k2And epsilon1Is a constant positive parameter against saturation, udIs the anti-saturation differential signal of the line of sight angle.
4. The gyroscopic-free precise guidance method for the small micro aircraft as claimed in claim 1, wherein the integration is performed according to the line-of-sight angle to obtain a line-of-sight angle integrated signal, and then the nonlinear anti-saturation operation is performed to obtain an anti-saturation integrated signal of the line-of-sight angle, comprises:
Sq＝∫qbdt；
wherein k is3、k4And epsilon2As a constant positive parameter against saturation, qbFor line-of-sight angle signals, SqFor line-of-sight angle integral signal, usThe signal is integrated for line-of-sight angles against saturation.
5. The gyroscopic-free precision guidance method for the small micro aircraft according to claim 1, wherein designing a nonlinear transformation function according to the line-of-sight angle signal to obtain a line-of-sight nonlinear transformation signal comprises:
wherein k is5、k6And epsilon3For non-linear transformation of parameters, qbFor line-of-sight angle signals, uqThe signal is integrated for line-of-sight angles against saturation.
6. The gyroscopic-free precision guidance method for small micro-aircraft according to claim 1, characterized in that it consists in said line-of-sight angle signal qbAnti-saturation differential signal udAnti-saturation integrated signal usAngle of sight non-linear transformation signal uqPerforming comprehensive processing to obtain a final pilot law comprehensive signal, including:
u＝k7qb+k8ud+k9us+k10uq；
where u is the final pilot law combined signal, k7、k8、k9、k10For the guided constant, qbIs a line-of-sight angle signal udFor anti-saturation differential signal usIntegrating the signal u for anti-saturationqThe signal is transformed for line-of-sight non-linearity.
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| CN112034874A (en) * | 2020-09-10 | 2020-12-04 | 中国人民解放军海军航空大学 | Aircraft attitude stability tracking control method based on nonlinear sliding mode |
| CN112082549A (en) * | 2020-09-10 | 2020-12-15 | 中国人民解放军海军航空大学 | Aircraft simple mass center control method only measuring acceleration |
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