CN114237299A - Aircraft guidance method and device, electronic equipment and storage medium - Google Patents
Aircraft guidance method and device, electronic equipment and storage medium Download PDFInfo
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Abstract
The application provides an aircraft guidance method, an aircraft guidance device, electronic equipment and a storage medium, relates to the technical field of missile launching, and specifically comprises the following steps: acquiring the position, the flying speed and the trajectory inclination angle of the aircraft at the current moment; calculating a virtual control quantity in a polynomial form at the current moment according to the position, the flying speed and the ballistic inclination angle of the aircraft at the current moment; establishing a nonlinear function containing the attack angle parameters according to the virtual control quantity in the polynomial form at the current moment and the virtual control quantity containing the attack angle parameters; and solving the attack angle when the nonlinear function is 0, and taking the attack angle as the instruction attack angle between the current moment and the next moment. The method and the device can generate the optimal guidance instruction in real time according to the current state, do not depend on the nominal track, and have stronger adaptability.
Description
Technical Field
The application relates to the technical field of missile launching, in particular to an aircraft guidance method, an aircraft guidance device, electronic equipment and a storage medium.
Background
In the flight process of the aircraft from the climbing section to the cruise section with fixed height, because the climbing flight state needs to be switched to the cruise flight state, a control process of a transition section exists, the attack angle of the transition section is easy to generate large oscillation, the corresponding speed and the corresponding height also generate oscillation, and the transition process is not smooth. In addition, the common guidance method tracks a nominal trajectory which is designed in advance, and the adaptability is poor.
Disclosure of Invention
In view of the above, the present application provides an aircraft guidance method, an aircraft guidance device, an electronic device, and a storage medium, so as to solve the technical problem that in the prior art, a transition section exists in guidance law design of a climbing section and a cruise section of an aircraft, and an optimal guidance instruction can be generated in real time according to a current state, which is independent of a nominal trajectory and has stronger adaptability.
In one aspect, an embodiment of the present application provides an aircraft guidance method, including:
acquiring the position, the flying speed and the trajectory inclination angle of the aircraft at the current moment;
calculating a virtual control quantity in a polynomial form at the current moment according to the position, the flying speed and the ballistic inclination angle of the aircraft at the current moment;
establishing a nonlinear function containing the attack angle parameters according to the virtual control quantity in the polynomial form at the current moment and the virtual control quantity containing the attack angle parameters; and solving the attack angle when the nonlinear function is 0, and taking the attack angle as the instruction attack angle between the current moment and the next moment.
Further, calculating a polynomial-form virtual control quantity at the current moment according to the position, the flying speed and the ballistic inclination angle of the aircraft at the current moment; the method comprises the following steps:
the virtual control amount u (t) in the form of a polynomial at the present time is:
where t is the current time, vtSpeed of the aircraft at the present moment, γtThe trajectory inclination angle of the aircraft at the current moment; z is a radical oftThe z coordinate value of the aircraft at the current moment in the emission coordinate system; z is a radical offSpecified altitude for cruising flight, tfFor aircraft from the beginning to a given height zfTime of flight of (a).
Further, the aircraft flies from the beginning to a specified height zfTime of flight tfThe values of (A) are as follows:
wherein, tsetfTime of flight, t, of the aircraft being the desired guidance lawminAnd the value interval is 5 to 15 seconds for the residual flight threshold value of the cruise segment.
Further, establishing a nonlinear function containing the attack angle parameters according to the virtual control quantity in the polynomial form at the current moment and the virtual control quantity containing the attack angle parameters; the method comprises the following steps:
establishing a virtual control quantity u at the current momentcom(t) expression:
wherein T is engine thrust, alpha is an attack angle of a belt solution, rho is an atmospheric density function, and Cl(Ma, α) is a lift coefficient, and Mach number Ma is equal to vt-1/vair，vairThe local speed of sound; srefThe characteristic area of the aircraft, g is the gravity acceleration, and m is the mass of the aircraft;
the nonlinear function F (α) containing the angle of attack parameter is:
F(α)＝u(t)-ucom(t)。
further, solving the attack angle when the nonlinear function is 0, and taking the attack angle as the instruction attack angle from the current time to the next time includes:
solving alpha satisfying F (alpha) 0 by using a Newton iteration method, wherein the iteration process is as follows:
angle of attack alpha for the k +1 th iterationk+1Comprises the following steps:
wherein the content of the first and second substances,
angle of attack alpha for the k-th iteration for the non-linear functionkIs specifically expressed as
Wherein the content of the first and second substances,
the partial derivative of the lift coefficient to the angle of attack; the initial iteration value of the attack angle is the attack angle value at the last moment;
and when the preset iteration times are reached, taking the obtained attack angle as an instruction attack angle from the current moment to the next moment.
In another aspect, an embodiment of the present application provides an aircraft guidance device, including:
the acquiring unit is used for acquiring the position, the flying speed and the trajectory inclination angle of the aircraft at the current moment;
the polynomial virtual control quantity calculating unit is used for calculating the polynomial virtual control quantity at the current moment according to the position, the flying speed and the ballistic inclination angle of the aircraft at the current moment;
the command attack angle calculation unit is used for establishing a nonlinear function containing attack angle parameters according to the virtual control quantity in the polynomial form at the current moment and the virtual control quantity containing the attack angle parameters; and solving the attack angle when the nonlinear function is 0, and taking the attack angle as the instruction attack angle between the current moment and the next moment.
In another aspect, an embodiment of the present application provides an electronic device, including: the aircraft guidance system comprises a memory, a processor and a computer program stored on the memory and capable of running on the processor, wherein the processor executes the computer program to realize the aircraft guidance method of the embodiment of the application.
In another aspect, the present application provides a computer-readable storage medium, on which a computer program is stored, where the computer program is executed by a processor to implement the aircraft guidance method of the present application.
The guidance method can generate the optimal guidance instruction in real time according to the current state, does not depend on the nominal track, and is high in adaptability; the method has the advantages of having good guidance effect, being capable of rapidly completing climbing and leveling tasks and guiding the video of the aircraft to cruise and fly, having strong robustness and being capable of smoothly completing the guidance tasks even under the condition of large deviation.
Drawings
In order to more clearly illustrate the detailed description of the present application or the technical solutions in the prior art, the drawings needed to be used in the detailed description of the present application or the prior art description will be briefly introduced below, and it is obvious that the drawings in the following description are some embodiments of the present application, and other drawings can be obtained by those skilled in the art without creative efforts.
FIG. 1 is a curve fitting lift coefficients provided by an embodiment of the present invention;
FIG. 2 is a response curve of a step signal for an angle of attack command according to an embodiment of the present invention;
FIG. 3 is a response curve of a sine wave signal for a commanded angle of attack according to an embodiment of the present invention;
FIG. 4 is a schematic diagram of a flight timing sequence provided by an embodiment of the present invention;
FIG. 5 is a flow chart of a method provided by an embodiment of the present application;
FIG. 6(a) is a graph of height versus longitudinal travel for a nominal case as provided by an embodiment of the present application;
FIG. 6(b) is a plot of height over time for a nominal case as provided by an embodiment of the present application;
FIG. 7(a) is a graph showing the Mach number as a function of time according to an embodiment of the present application;
figure 7(b) is a graph of ballistic inclination angle versus time provided by an embodiment of the present application;
FIG. 8(a) is a graph of Monte Carlo target height versus course provided in an embodiment of the present application;
FIG. 8(b) is a graph of Monte Carlo target height over time as provided by an embodiment of the present application;
FIG. 9 is a functional block diagram of an apparatus according to an embodiment of the present disclosure;
fig. 10 is a schematic hardware structure diagram of an electronic device according to an embodiment of the present application.
Detailed Description
In order to make the objects, technical solutions and advantages of the embodiments of the present application clearer, the technical solutions in the embodiments of the present application will be clearly and completely described below with reference to the drawings in the embodiments of the present application, and it is obvious that the described embodiments are some embodiments of the present application, but not all embodiments. The components of the embodiments of the present application, generally described and illustrated in the figures herein, can be arranged and designed in a wide variety of different configurations.
Thus, the following detailed description of the embodiments of the present application, presented in the accompanying drawings, is not intended to limit the scope of the claimed application, but is merely representative of selected embodiments of the application. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present application.
To facilitate a better understanding of the present application by those skilled in the art, a brief description of the technical terms involved in the present application will be given below.
1. Aircraft model
The aircraft adopts preset pneumatic data, the aircraft is a two-stage rocket, and the characteristic area of the full projectile is Sref m2. Both the mass characteristic and the thrust characteristic are functions that vary with time.
The aircraft mainly adopts an STT control mode, the particle trajectory only needs to consider the aerodynamic data when only an attack angle exists, and the range of the attack angle alpha is [ alpha ]min,αmax]Meanwhile, the pneumatic coefficient under the projectile coordinate system needs to be converted into pneumatic data under the speed coordinate system, and the conversion relationship is as follows:
Cl＝Cy cosα-Cx sinα
Cd＝Cy sinα+Cx cosα
wherein, CxIs the axial force coefficient, CyIs the normal force coefficient, ClIs the coefficient of lift in the velocity coordinate system, CdThe resistance coefficient in a speed coordinate system.
The main level flies in a controlled mode, so that the problem of moment balancing needs to be considered for mass point ballistic aerodynamic data of the main level, the main level also adopts an STT control mode, only longitudinal plane motion pneumatic parameters need to be considered, and therefore only a pitching moment coefficient needs to be balanced, and the balanced aerodynamic is selected to be used as three-degree-of-freedom mass point ballistic aerodynamic.
In order to ensure the smooth solution of the attack angle guidance instruction, it is necessary to perform polynomial fitting on the particle aerodynamic data of the aircraft to obtain a set of smooth aerodynamic functions, which have better smoothness when newton iteration is adopted for solution, and the lift coefficient of the aircraft is expressed in a polynomial form as follows:
Cl＝[ma3 ma2 ma 1]Clcoe[α3 α2 α 1]T
wherein, ClcoeFor the fitted lift matrix coefficients, m is the mass of the aircraft, the function is a function of the Mach number and the angle of attack, a partial derivative of the lift coefficients with respect to the angle of attack is required in the process of solving the angle of attack, and the expression can be expressed as
The fitting effect of the aerodynamic coefficient is shown in fig. 1, in the figure, a line represents a result obtained by calculating a fitting polynomial, and a node is the aerodynamic coefficient of the aircraft.
The delay influence of a flight control system is considered in the simulation process, a first-order link related to an attack angle is added into a dynamic equation, and the specific expression is
Wherein alpha iscomFor the command attack angle output by the guidance system, α is an attack angle state considering delay influence, in order to simulate the delay influence when the flight control system exists, a time constant is selected from Ts according to the aircraft characteristics and flight mission, the output response when the command attack angle is a step signal is shown in fig. 2, and the output response when the command attack angle is a sine wave signal is shown in fig. 3.
It can be seen that when the command attack angle is 10 degrees, it takes about 2 seconds to completely track, and when the command attack angle is a sinusoidal signal with a cycle of 6 seconds and an amplitude of 10 degrees, the delay error of the attack angle does not exceed 1.5 degrees. Therefore, the model can well represent the influence of the flight control system.
2. Flight dynamics modeling
Because the researched aircraft has the characteristics of short flight time and flight distance and the like, the dynamic model can be a plane geodetic three-degree-of-freedom particle flight dynamic model neglecting the influence of earth rotation and curvature, and the model is shown as follows
Wherein x, y and z are three coordinates under a launching coordinate system, v is the speed of the aircraft, gamma is the ballistic inclination angle of the aircraft, psi is the course angle of the aircraft, T is the thrust of the aircraft, alpha is the attack angle of the aircraft, sigma is the roll angle of the aircraft, D is the resistance of the aircraft, L is the lift force of the aircraft, g is the gravitational acceleration, and m is the mass of the aircraft.
The expressions of lift L and drag D are
Where ρ is an atmospheric density function, typically varying with altitude. Cl(Ma, α) and Cd(Ma, α) are lift and drag coefficients, which are functions of angle of attack and Mach number, Ma＝v/vair，vairThe local speed of sound, which is also a function of altitude, SrefIs the characteristic area. The expression of the gravity acceleration g is g ═ mu/(Re + z)2Wherein μ is the gravitational constant and Re is the earth's mean radius.
In the simulation process, the influence of the thrust T of the engine on the atmospheric pressure needs to be considered, and the thrust expression is
T＝T0+S(P0-P)
Wherein S is the outlet area of the engine, P0Sea level atmospheric pressure, and P is local atmospheric pressure.
The aircraft that this application embodiment relates to is two-stage rocket, needs ignition many times in flight to have two sections unpowered gliding stages before and after the separation, in the simulation process, need carry out segmentation process with the simulation process according to specific control mode, whole simulation process is divided into four parts, and specific control time sequence is as shown in figure 4:
the working stage of the full-bomb engine is as follows: time is from t10To t1fSecond, where t10About 0.2 seconds, the launcher separation problem needs to be considered, t1fThe different working temperatures have different shutdown times for the shutdown time of the full-barrel engine. The aircraft is not controlled at the stage and flies at a zero-degree attack angle.
And (3) a full projectile gliding flight stage: time is from t1fTo t2fSecond, where t2fFor full projectile separation time, the flight phase thrust is zero. The aircraft in this phase is also uncontrolled and flies at zero angle of attack.
Main-stage separation gliding flight stage: time is from t2fTo t3fSecond, where t3fAt the primary firing time, the flight phase thrust is zero. The phase is also uncontrolled and flies at zero angle of attack.
And a main stage ignition climbing cruise stage: time is from t3fTo t4fSecond, where t3fThe working end time of the main-stage engine. The control is started at the stage, and the aircraft is guided to quickly climb to the height of 20 kilometers and pull according to the guidance law instructionFlat, and then maintain cruise flight for a long time.
After introducing the technical terms related to the present application, the design ideas of the embodiments of the present application will be briefly described below.
The application provides a boosting section climbing and cruising section flight integrated guidance method based on a polynomial guidance law design idea, and the method can obtain an analyzed guidance instruction according to an initial state, guide an aircraft to smoothly complete climbing and deviation pulling, and switch to a cruising flight state when the aircraft approaches a cruising height so as to complete horizontal equal-height flight. Meanwhile, in order to test the effectiveness of the method, a nominal guidance example test, a limit deviation guidance example test and a Monte Carlo target practice test are carried out, and simulation results show that the method has good guidance effect and robustness, and can smoothly complete guidance tasks even under the condition of large limit deviation.
When the aircraft enters a primary ignition climbing cruise phase, a guidance law needs to be designed to guide the aircraft to quickly enter a specified cruise altitude and ensure that the aircraft can horizontally cruise flight, so that the designed guidance law needs to ensure the terminal altitude and the trajectory inclination angle, and meanwhile, a corresponding guidance strategy needs to be designed to ensure that the aircraft can stably cruise for a long time.
The guidance law design cruises according to the judgment of a time condition, and an aircraft longitudinal particle motion equation taking time as a stop condition can be expressed as
Wherein T is engine thrust, and L is lift force of an aircraft; assuming that the ballistic inclination γ of the flight is small, sin γ ≈ γ, and considering the virtual control amount u ═ (T sin α + L)/m-g cos γ; the resulting equations of dynamics after linearization can be expressed as:
meanwhile, it is assumed that the control amount u has a polynomial form as follows
u＝a1(t-tf)2+a2(t-tf)
Wherein, a1And a2Are all polynomial coefficients, tfIs the time of flight of the aircraft to a specified altitude, then an integral operation on the control can be used to obtain an analytical expression for y and z
Meanwhile, according to the boundary conditions, the following four constraint relationships can be obtained
z(tf)＝zf,z(t)＝z；
y(tf)＝0,y(t)＝v sinγ；
Wherein z isfFor a specified altitude for cruising flight, when the aircraft reaches a specified altitude zfEntering a flat flight stage, the trajectory inclination angle is 0, namely y (t)f)＝0。
By substituting this into the expression, the following four parameters can be obtained:
the polynomial form of the virtual control amount can be expressed in the following form
Thus, a polynomial guidance law ensuring the altitude and trajectory inclination of the aircraft is designed, and it should be noted that the flight time and the distance of the aircraft of the guidance law need to be set by a user, and are uniformly set to be a specific value according to the initial overload condition in the simulation process.
Because the controlled variable u is a virtual controlled variable, after the virtual controlled variable is obtained, the controlled variable u needs to be converted into an attack angle controlled variablecomThe expression of (a) is:
it can be seen that the equation is a nonlinear function about the attack angle α, and an analytic solution cannot be directly obtained, so that a corresponding numerical solution needs to be obtained by means of a numerical solution technology, and firstly, the pneumatic parameter C needs to be calculatedlPerforming a polynomial fit to obtain a smooth function with respect to the aerodynamic parameter, the non-linear function F (α) may be expressed as:
F(α)＝u-ucom
an attack angle satisfying F (α) ═ 0 can be obtained by newton iteration, and this is taken as the command attack angle at the current time.
According to the method, the climbing and cruising integrated polynomial guidance law of the primary level of the boosting section is designed according to the design thought of the polynomial guidance law, the guidance instruction which can guide the aircraft to quickly climb to the designated height and flatten the trajectory inclination angle can be provided according to the initial state of the primary level ending of the active section, and when the aircraft approaches the cruising height, the guidance method can be switched to the stable cruising stage to guide the aircraft to cruise at the same height. Due to the adoption of an integrated design idea, the guidance instructions among all the sections can be seamlessly connected. The nominal guidance example, the extreme deviation guidance example and the Monte Carlo target practice test are used for verifying the effectiveness and robustness of the guidance method, and simulation results show that the guidance method not only has a good guidance effect, can rapidly complete a climbing and leveling task and guide an aircraft to cruise and fly, but also has strong robustness, and can successfully complete the guidance task even under the condition of large deviation.
After introducing the application scenario and the design concept of the embodiment of the present application, the following describes a technical solution provided by the embodiment of the present application.
As shown in fig. 5, an embodiment of the present application provides an aircraft guidance method, including the following steps:
step 101: acquiring the position, the flying speed and the trajectory inclination angle of the aircraft at the current moment;
step 102: calculating a virtual control quantity in a polynomial form at the current moment according to the position, the flying speed and the ballistic inclination angle of the aircraft at the current moment;
when the aircraft approaches the cruising altitude, the cruise control system needs to maintain the high-altitude cruising flight for a long time at the cruising altitude, and in order to obtain a better guidance effect and avoid control jump caused by guidance law switching, the guidance law of the climbing section and the guidance law of the cruising section are designed in a combined manner to obtain the guidance effect of the cruising section. When the remaining flight time approaches zero, a specific remaining flight time is specified so as to maintain a long cruising flight, the aircraft flying from the beginning to a specified altitude zfTime of flight tfThe method comprises the following steps:
wherein, tsetfTime of flight, t, of the aircraft being the desired guidance lawminAnd selecting the residual flight threshold value of the cruise segment according to a specific task, wherein the value interval is 5 to 15 seconds. Aircraft time of flight t using the above guidance lawfThe seamless connection of the guidance law of the climbing section and the cruise section can be realized, so that the continuous climbing and cruise guidance effects are obtained.
The virtual control amount u (t) in the form of a polynomial at the present time may be expressed in the form:
where t is the current time, vtSpeed of the aircraft at the present moment, γtThe trajectory inclination angle of the aircraft at the current moment; z is a radical oftThe z coordinate value of the aircraft at the current moment in the emission coordinate system; z is a radical offSpecified altitude for cruising flight, tfFor aircraft from the beginning to a given height zfTime of flight of (a).
Step 103: establishing a nonlinear function containing the attack angle parameters according to the virtual control quantity in the polynomial form at the current moment and the virtual control quantity containing the attack angle parameters; and solving the attack angle when the nonlinear function is 0, and taking the attack angle as the instruction attack angle between the current moment and the next moment.
Firstly, establishing an expression of a virtual control quantity:
establishing a nonlinear function F (alpha) containing an attack angle parameter:
F(α)＝u(t)-ucom(t)
solving the attack angle satisfying F (alpha) as 0 by using a Newton iteration method, and taking the attack angle as an instruction attack angle alphacom(ii) a The iterative process is as follows:
wherein the content of the first and second substances,
the partial derivative of the nonlinear function to the attack angle is expressed by
Wherein the content of the first and second substances,
the partial derivative of the lift coefficient with respect to the angle of attack. Finally, the solution meeting the nonlinear equation can be obtained through cyclic iteration, and in the simulation process, the initial value of the attack angle is selected as the attack angle of the previous step of the resolving periodThe initial value of the initial attack angle is zero, the iteration times are limited in the simulation process in order to ensure the stability of calculation, and each calculation does not exceed five iterations. The simulation results later show that the method can quickly obtain the command attack angle.
In order to verify the guidance effect of the designed guidance law, the guidance effects for different nominal conditions are considered below, the initial conditions of the aircraft being as follows:
| x0 | y0 | z0 | v0 | ψ0 | Hcruise control system | |
| EXAMPLES-1 | 0 | 0 | 1215 | 25 | 0 | 18000 |
| EXAMPLES-2 | 0 | 0 | 1215 | 25 | 0 | 20000 |
| EXAMPLES-3 | 0 | 0 | 1215 | 25 | 0 | 22000 |
The guidance results are shown in fig. 6(a) and 6(b), the guidance law can be well adapted to the cruise altitudes under different calculation examples, the climbing period reaches the specified cruise altitude at the specified time, and then the guidance law guides the aircraft to keep equal-altitude cruise flight.
As shown in fig. 7(a) and 7(b), the speed of the aircraft has a certain difference for different cruising altitudes, when the cruising altitude of the aircraft is low, the air density is high, so that the deceleration is obvious, the speed and the mach number are continuously reduced, the speed difference of the aircraft is small at 20 km and 22 km, the thrust and the air resistance are approximately balanced in the cruising section, and the speed and the mach number do not change greatly.
The designed polynomial guidance law can smoothly complete the tasks of rapid entering of a climbing section into a cruising height, leveling of a trajectory, stable equal-height flight of the cruising section and the like, is simple in structure and easy to realize, generates a guidance instruction which is an analytic function, can ensure smooth handover of two flight stages and is suitable for different cruising heights.
In order to further consider the guidance effect of the designed guidance law under the comprehensive pull bias, in this section, a monte carlo target practice test is performed, the considered bias includes an initial bias, a pneumatic bias, an atmospheric environment bias and a thrust under different temperatures, the target practice times are 3000 times, and the initial pull bias is set as follows at 0 ℃, 20 ℃ and 40 ℃ for 1000 times respectively:
| pulling bias item | Value of pull bias | Distribution of | Pulling bias item | Value of pull bias | Distribution of |
| |
100 | 3σ | Δψ | 0 | - |
| Δy | 0 | - | ΔCl | ±15 | 3σ |
| Δz | |||||
| 100 | 3σ | ΔCd | ±10 | 3σ | |
| Δv | |||||
| 20 | 3σ | Δρ | ±15 | 3σ | |
| Δγ | |||||
| 5 | 3σ |
The simulation results are shown in fig. 8(a) and 8(b), and it can be seen from the figures that comprehensive deviation leads to great dispersion of flight trajectory, but the guidance law can smoothly complete the tasks of high-speed flight of climbing segment, leveling trajectory, stable cruise segment and the like.
Based on the above embodiments, an aircraft guidance device is provided in the embodiments of the present application, and referring to fig. 9, an aircraft guidance device 200 provided in the embodiments of the present application at least includes:
an obtaining unit 201, configured to obtain a position, a flight speed, and a trajectory inclination of an aircraft at a current time;
a polynomial virtual control quantity calculation unit 202, configured to calculate a polynomial virtual control quantity at the current time according to the position, the flight speed, and the trajectory inclination of the aircraft at the current time;
the instruction attack angle calculation unit 203 is used for establishing a nonlinear function containing attack angle parameters according to the virtual control quantity in the polynomial form at the current moment and the virtual control quantity containing the attack angle parameters; and solving the attack angle when the nonlinear function is 0, and taking the attack angle as the instruction attack angle between the current moment and the next moment.
As a possible implementation manner, the polynomial virtual control quantity calculating unit 202 is specifically configured to:
the virtual control amount u (t) in the form of a polynomial at the present time is:
where t is the current time, vtSpeed of the aircraft at the present moment, γtThe trajectory inclination angle of the aircraft at the current moment; z is a radical oftThe z coordinate value of the aircraft at the current moment in the emission coordinate system; z is a radical offSpecified altitude for cruising flight, tfFor aircraft from the beginning to a given height zfTime of flight of (a).
As a possible embodiment, the aircraft flies from the beginning to a specified height zfTime of flight tfThe values of (A) are as follows:
wherein, tsetfTime of flight, t, of the aircraft being the desired guidance lawminAnd the value interval is 5 to 15 seconds for the residual flight threshold value of the cruise segment.
As a possible implementation, the command angle of attack calculation unit 203 is specifically configured to:
establishing a nonlinear function containing the attack angle parameters according to the virtual control quantity in the polynomial form at the current moment and the virtual control quantity containing the attack angle parameters; the method comprises the following steps:
establishing a virtual control quantity u at the current momentcom(t) expression:
wherein T is engine thrust, alpha is an attack angle of a belt solution, rho is an atmospheric density function, and Cl(Ma, α) is a lift coefficient, and Mach number Ma is equal to vt-1/vair，vairThe local speed of sound; srefThe characteristic area of the aircraft, g is the gravity acceleration, and m is the mass of the aircraft;
the nonlinear function F (α) containing the angle of attack parameter is:
F(α)＝u(t)-ucom(t)；
solving alpha satisfying F (alpha) 0 by using a Newton iteration method, wherein the iteration process is as follows:
angle of attack alpha for the k +1 th iterationk+1Comprises the following steps:
wherein the content of the first and second substances,
angle of attack alpha for the k-th iteration for the non-linear functionkIs specifically expressed as
Wherein the content of the first and second substances,
the partial derivative of the lift coefficient to the angle of attack; the initial iteration value of the attack angle is the attack angle value at the last moment;
and when the preset iteration times are reached, taking the obtained attack angle as an instruction attack angle from the current moment to the next moment.
Based on the foregoing embodiments, an embodiment of the present application further provides an electronic device, and referring to fig. 10, an electronic device 300 provided in an embodiment of the present application at least includes: the aircraft guidance system comprises a processor 301, a memory 302 and a computer program stored on the memory 302 and capable of running on the processor 301, wherein the processor 301 realizes the aircraft guidance method provided by the embodiment of the application when executing the computer program.
The electronic device 300 provided by the embodiment of the present application may further include a bus 303 connecting different components (including the processor 301 and the memory 302). Bus 303 represents one or more of any of several types of bus structures, including a memory bus, a peripheral bus, a local bus, and so forth.
The Memory 302 may include readable media in the form of volatile Memory, such as Random Access Memory (RAM) 3021 and/or cache Memory 3022, and may further include Read Only Memory (ROM) 3023.
The memory 302 may also include a program tool 3024 having a set (at least one) of program modules 3025, the program modules 3025 including, but not limited to: an operating subsystem, one or more application programs, other program modules, and program data, each of which, or some combination thereof, may comprise an implementation of a network environment.
It should be noted that the electronic device 300 shown in fig. 10 is only an example, and should not bring any limitation to the functions and the scope of the application of the embodiments.
Embodiments of the present application further provide a computer-readable storage medium, which stores computer instructions, and the computer instructions, when executed by a processor, implement the aircraft guidance method provided by embodiments of the present application.
It should be noted that although several units or sub-units of the apparatus are mentioned in the above detailed description, such division is merely exemplary and not mandatory. Indeed, the features and functions of two or more units described above may be embodied in one unit, according to embodiments of the application. Conversely, the features and functions of one unit described above may be further divided into embodiments by a plurality of units.
Further, while the operations of the methods of the present application are depicted in the drawings in a particular order, this does not require or imply that these operations must be performed in this particular order, or that all of the illustrated operations must be performed, to achieve desirable results. Additionally or alternatively, certain steps may be omitted, multiple steps combined into one step execution, and/or one step broken down into multiple step executions.
While the preferred embodiments of the present application have been described, additional variations and modifications in those embodiments may occur to those skilled in the art once they learn of the basic inventive concepts. Therefore, it is intended that the appended claims be interpreted as including preferred embodiments and all alterations and modifications as fall within the scope of the application.
Finally, it should be noted that: the above embodiments are only used for illustrating the technical solutions of the present application, and not for limiting the same; although the present application has been described in detail with reference to the foregoing embodiments, it should be understood by those of ordinary skill in the art that: the technical solutions described in the foregoing embodiments may still be modified, or some or all of the technical features may be equivalently replaced; and the modifications or the substitutions do not make the essence of the corresponding technical solutions depart from the scope of the technical solutions of the embodiments of the present application.
Claims (8)
1. An aircraft guidance method, comprising:
acquiring the position, the flying speed and the trajectory inclination angle of the aircraft at the current moment;
calculating a virtual control quantity in a polynomial form at the current moment according to the position, the flying speed and the ballistic inclination angle of the aircraft at the current moment;
establishing a nonlinear function containing the attack angle parameters according to the virtual control quantity in the polynomial form at the current moment and the virtual control quantity containing the attack angle parameters; and solving the attack angle when the nonlinear function is 0, and taking the attack angle as the instruction attack angle between the current moment and the next moment.
2. The aircraft guidance method according to claim 1, wherein the virtual control amount in the form of a polynomial at the present time is calculated from the position, the flight speed, and the ballistic inclination angle of the aircraft at the present time; the method comprises the following steps:
the virtual control amount u (t) in the form of a polynomial at the present time is:
where t is the current time, vtSpeed of the aircraft at the present moment, γtThe trajectory inclination angle of the aircraft at the current moment; z is a radical oftThe z coordinate value of the aircraft at the current moment in the emission coordinate system; z is a radical offSpecified altitude for cruising flight, tfFor aircraft from the beginning to a given height zfTime of flight of (a).
3. The aircraft guidance method of claim 2 wherein the aircraft flies from the beginning to a specified altitude zfTime of flight tfThe values of (A) are as follows:
wherein, tsetfTime of flight, t, of the aircraft being the desired guidance lawminAnd the value interval is 5 to 15 seconds for the residual flight threshold value of the cruise segment.
4. The aircraft guidance method according to claim 3, wherein the nonlinear function including the angle of attack parameter is established based on the virtual control quantity in the polynomial form at the present time and the virtual control quantity including the angle of attack parameter; the method comprises the following steps:
establishing a virtual control quantity u at the current momentcom(t) expression:
wherein T is engine thrust, alpha is an attack angle of a belt solution, rho is an atmospheric density function, and Cl(Ma, α) is a lift coefficient, and Mach number Ma is equal to vt-1/vair，vairThe local speed of sound; srefThe characteristic area of the aircraft, g is the gravity acceleration, and m is the mass of the aircraft;
the nonlinear function F (α) containing the angle of attack parameter is:
F(α)＝u(t)-ucom(t)。
5. the aircraft guidance method of claim 4, wherein solving for the angle of attack at which the non-linear function is 0 as the commanded angle of attack from the current time to the next time comprises:
solving alpha satisfying F (alpha) 0 by using a Newton iteration method, wherein the iteration process is as follows:
angle of attack alpha for the k +1 th iterationk+1Comprises the following steps:
wherein the content of the first and second substances,
angle of attack alpha for the k-th iteration for the non-linear functionkThe specific expression of the partial derivative of (c) is:
wherein the content of the first and second substances,
the partial derivative of the lift coefficient to the angle of attack; the initial iteration value of the attack angle is the attack angle value at the last moment;
and when the preset iteration times are reached, taking the obtained attack angle as an instruction attack angle from the current moment to the next moment.
6. An aircraft guidance device, comprising:
the acquiring unit is used for acquiring the position, the flying speed and the trajectory inclination angle of the aircraft at the current moment;
the polynomial virtual control quantity calculating unit is used for calculating the polynomial virtual control quantity at the current moment according to the position, the flying speed and the ballistic inclination angle of the aircraft at the current moment;
the command attack angle calculation unit is used for establishing a nonlinear function containing attack angle parameters according to the virtual control quantity in the polynomial form at the current moment and the virtual control quantity containing the attack angle parameters; and solving the attack angle when the nonlinear function is 0, and taking the attack angle as the instruction attack angle between the current moment and the next moment.
7. An electronic device, comprising: memory, a processor and a computer program stored on the memory and executable on the processor, the processor implementing the aircraft guidance method according to any one of claims 1 to 5 when executing the computer program.
8. A computer-readable storage medium, characterized in that the computer-readable storage medium has stored thereon a computer program which, when being executed by a processor, implements the aircraft guidance method according to any one of claims 1 to 5.
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