CN104062976B - A kind of is sinusoidal attitude of flight vehicle fast reserve method based on angular acceleration derivative - Google Patents

Abstract

The invention discloses a rapid attitude maneuvering method of an aircraft based on a sinusoidal angular acceleration derivative. According to the torque and angular momentum providing capabilities of the actuators of the control system, an attitude maneuvering path that undergoes three processes of acceleration, uniform velocity and deceleration is designed. During the acceleration and deceleration process, the derivative of the angular acceleration is guaranteed to be a standard sine curve, which ensures that the torque output during the entire maneuvering process is not only continuous, but also the first-order derivative is continuous, so that the torque output changes smoothly during the entire maneuvering process. The excitation effect of the flexible mode is small. After the attitude maneuver of the aircraft is in place, the attitude of the aircraft can be quickly stabilized due to the small vibration amplitude of the flexible mode, thereby realizing rapid maneuvering and rapid stability control. This method is especially suitable for fast maneuvering control of aircraft with severe coupling of flexible modes, and can realize fast and stable control requirements.

Description

A Rapid Attitude Maneuvering of Aircraft Based on Angular Acceleration Derivative as Sine Curve method

technical field

The invention relates to a fast maneuvering control method for a flexible spacecraft, in particular to a fast maneuvering method for an aircraft attitude based on a sinusoidal angular acceleration derivative, which belongs to the field of spacecraft attitude control.

Background technique

Modern spacecraft usually have flexible attachments with light structures such as large solar panels. During the flight of flexible spacecraft, it is often necessary to quickly perform large-angle attitude maneuvers to meet mission requirements. Theoretical analysis shows that due to the strong coupling between the dynamics of the central rigid body of the spacecraft and the vibration of the flexible attachment, the nonlinear dynamic characteristics are more significant when its attitude maneuvers to a large angle, which often leads to continuous strong vibration of the flexible attachment, and then Seriously affect the attitude movement, and even directly threaten the safety of the spacecraft structure. The attitude control and vibration suppression of the flexible spacecraft pose great challenges to the rapid maneuvering mode at a large angle.

For the attitude maneuver control and vibration suppression of flexible spacecraft, existing studies have shown that optimizing the maneuvering path is an effective measure to achieve large-angle rapid maneuvering and rapid stability. Current representative maneuver path planning methods include bang-bang-based trajectory planning methods and angular-acceleration sinusoidal-based planning methods. Among them, the bang-bang trajectory planning method has a short planning time, but the vibration excitation of the flexible mode is severe, and the stabilization time of the aircraft after maneuvering is long. The angular acceleration trajectory planning method based on the sinusoidal curve has little excitation to the flexible mode, but it takes into account the coupling relationship between the flexible mode and the attitude angular velocity of the spacecraft. Since the planning method only considers the continuation of the angular acceleration, the derivative of the angular acceleration will jump at the beginning and end of the planning, which still has a certain excitation effect on the flexible mode, which is not good for the stability of the attitude maneuver.

Aiming at the deficiencies of current maneuvering trajectory planning methods, the derivative of angular acceleration is considered for trajectory planning, and the final planned trajectory ensures the smoothness of angular acceleration. When the fast maneuvering trajectory planning curve based on the angular jerk sinusoidal curve is used for fast maneuvering, the flexible mode is basically not excited after the planned trajectory ends, and the aircraft has a high degree of stability after the attitude maneuver is in place.

Contents of the invention

The technical solution problem of the present invention is: overcome the deficiencies in the prior art, provide a kind of aircraft attitude rapid maneuvering method based on the angular acceleration derivative being a sinusoidal curve, this method can guarantee the quickness and smoothness of the attitude maneuvering process.

The technical solution of the present invention is: a kind of aircraft attitude fast maneuvering method based on angular acceleration derivative is a sine curve, it is characterized in that the steps are as follows:

(1) Determine the maximum angular acceleration a _m of the attitude maneuver of the aircraft according to the configured actuators on the aircraft;

(2) Determine the maximum angular velocity of the aircraft attitude maneuver according to the angular momentum envelope of the actuator configured on the aircraft and the range of the sensor

(3) When the aircraft receives the attitude maneuver control command sent by the ground, the aircraft control system according to the received attitude maneuver angle θ _m and the determined maximum angular velocity of the attitude maneuver and attitude maneuvering maximum angular acceleration a _m to calculate the characteristic turning moment of aircraft attitude maneuvering flight trajectory;

(4) The aircraft control system uses the characteristic turning time calculated in step (3) to calculate the target attitude angular acceleration a _r and the target attitude angular velocity of the aircraft in real time when the angular acceleration derivative is a segmented sinusoidal curve and the target attitude angle θ _r ;

(5) The attitude maneuver target attitude angular acceleration a _r and the target attitude angular velocity calculated by the aircraft control system according to step (4) and the target attitude angle θ _r for aircraft attitude maneuver control.

The method for calculating the characteristic turning point is as follows: the attitude maneuvering flight trajectory of the aircraft is divided into three sections: an acceleration section, a constant velocity section and a deceleration section, wherein the time of the acceleration section is [0 t _m1 ], and the time of the constant velocity section is [t _m1 t _m2 ], the time of the deceleration section is [t _m2 t _m3 ], t _m1 , t _m2 , and t _m3 are the characteristic turning moments of the maneuvering flight trajectory, and they all take the initial moment of the aircraft attitude maneuvering flight as the starting point; The method is:

(a) when ${\mathrm{\θ}}_m>\frac{2\×{\stackrel{\·}{\mathrm{\θ}}}_m^2}{a_m}$ hour, $t_{m1}=\frac{2\×{\stackrel{\&Center\; Dot;}{\mathrm{\θ}}}_m}{a_m}{t}_{m1}=\frac{2\·{\stackrel{\·}{\mathrm{\θ}}}_m}{a_m},t_{m2}=\frac{2\&Center\; Dot;{\mathrm{\θ}}_m}{a_m\·t_{m1}},$ t _m3 =t _m1 +t _m2 ;

(b) when ${\mathrm{\θ}}_m\≤\frac{2\×{\stackrel{\&Center\; Dot;}{\mathrm{\θ}}}_m^2}{a_m}{\mathrm{\θ}}_m\≤\frac{2\&Center\; Dot;{\stackrel{\&Center\; Dot;}{\mathrm{\θ}}}_m^2}{a_m}$ hour, $t_{m1}=\sqrt{\frac{2\·{\mathrm{\θ}}_m}{a_m}},$ t _m2 =t _m1 , t _m3 =t _m1 +t _m2 .

The step (4) calculates in real time the target angular acceleration a _r and the target angular velocity during attitude maneuvering of the aircraft and the method of the target angle θ _r as:

(1) Obtain the angular frequency of the angular acceleration derivative sinusoidal curve

(2) During the attitude maneuvering process, the target attitude angular acceleration a _r =0.5·am·(1-cos(f· _t )), the target attitude angular velocity Target attitude angle θ _r =0.5·am·(0.5· _t ² -1/f+cos(f·t)/f ² ), where t is the maneuvering time;

(3) In the process of attitude maneuvering, the target attitude angular acceleration a _r = 0 in the constant speed segment, and the target attitude angular velocity ${\stackrel{\·}{\mathrm{\θ}}}_r=0.5\·a_m\·t_{m1},$ target attitude angle ${\mathrm{\θ}}_r=0.5\·a_m\·(t_{m1}\&Center\; Dot;t-0.5\·t_{m1}^2);$

(4) During the attitude maneuvering process, the target attitude angular acceleration a _r = -0.5· _{am ·(1-cos(f(tt m2 )} )), the target attitude angular velocity ${\stackrel{\&Center\; Dot;}{\mathrm{\θ}}}_r=-0.5\&Center\; Dot;a_m\&Center\; Dot;(t-t_{m3}-\sin \left(f(t-t_{m2})\right)/f),$ target attitude angle ${\mathrm{\θ}}_r=-0.5a_m(0.5t^2-t_{m3}\&Center\; Dot;t+\frac{\cos \left(f(t-t_{m2})\right)}{f^2})-\frac{1}{f^2}+t_{m1}^2+0.5t_{m3}(t_{m2}-t_{m1}).$

Compared with the prior art, the present invention has the advantages that: the traditional aircraft attitude control method based on Bang-Bang trajectory planning will excite the large-scale vibration of the flexible mode due to the jump of the moment, and is not suitable for the rapid maneuvering and rapid stability control of the flexible aircraft . According to the torque and angular momentum providing ability of the actuator of the control system, the present invention designs an attitude maneuvering path that experiences three processes of acceleration, constant velocity and deceleration. During the acceleration and deceleration processes, the derivative of the angular acceleration is guaranteed to be a standard sinusoidal curve, ensuring The torque output in the whole maneuvering process is not only continuous, but also the first order derivative is continuous, which makes the torque output change smoothly in the whole maneuvering process, and the excitation effect on the flexible mode in the attitude maneuvering process is small. After the attitude maneuver of the aircraft is in place, the attitude of the aircraft can be quickly stabilized due to the small vibration amplitude of the flexible mode, thereby realizing rapid maneuvering and rapid stability control. This method is especially suitable for fast maneuvering control of aircraft with severe coupling of flexible modes, and can realize fast and stable control requirements. In addition, compared with the aircraft attitude control method based on angular acceleration sinusoidal trajectory planning, the present invention has a smaller excitation effect on the flexible mode of the aircraft, and the method has obvious advantages in occasions requiring higher stability.

Description of drawings

Fig. 1 is the block flow diagram of the inventive method;

Fig. 2 adopts the maneuver path that the inventive method obtains;

Fig. 3 is the aircraft attitude angle error curve when adopting this method control;

Fig. 4 is the aircraft attitude angular velocity error curve when adopting this method control;

Fig. 5 adopts the aircraft windshield flexible modal coordinate displacement that the method of the present invention obtains;

Fig. 6 is the coordinate displacement of the flexible modal coordinates of the sailboard of the aircraft under the sinusoidal maneuvering path;

Fig. 7 shows the coordinate displacement of the aircraft sailboard flexible mode under the bang-bang maneuvering path.

detailed description

In order to improve the efficiency of on-orbit use, the aircraft is paying more and more attention to the ability to maneuver quickly and stably at large angles. For example, the aircraft that uses the high-torque control moment gyroscope for attitude control requires the maneuvering speed to be more than 10 times the maneuvering speed of the traditional wheel-controlled aircraft. Since aircraft generally need to carry large flexible solar panels, the fundamental frequency of the panels is low, which makes the aircraft easily cause vibration in the flexible mode during rapid maneuvering and affects the attitude stability of the aircraft body, which cannot meet the working environment of the load. Therefore, a practical attitude maneuver control method must not only ensure that the maneuvering speed of the aircraft is fast, but also ensure that the attitude is fast and stable after the maneuver is in place. For a flexible aircraft, one of the ways to achieve rapid stabilization is to ensure that the vibration of the flexible mode is not excited as much as possible during the maneuvering process. The method of the present invention proceeds from this idea, and by optimizing the maneuvering path, it is ensured that the attitude maneuvering angle trajectory, angular velocity trajectory and angular acceleration trajectory all change smoothly during the entire rapid maneuvering process, not only without moment jumps, but also ensuring a constant moment. The order derivative is continuous, so the flexible mode vibration is hardly excited during the whole maneuvering process, and the fast and stable attitude maneuvering control is realized.

As shown in Figure 1, it is a flowchart of the method of the present invention. In the present invention, when performing attitude fast maneuvering, fast and stable control, a maneuvering trajectory is first planned, and the maneuvering trajectory includes the attitude angle, attitude angular velocity and angular acceleration of the whole maneuvering process, and then a corresponding control law is constructed to control the aircraft to run according to the planned trajectory , so as to complete the attitude maneuver control. The planning of the maneuvering path includes two steps. One is to calculate the key turning points t _m1 , t _m2 , and t _m3 of the maneuvering trajectory according to the attitude maneuvering capability of the aircraft before maneuvering; The angular acceleration a _r of the real-time maneuvering process, the target angular velocity Target angle θ _r . Specific steps are as follows:

(1) Determine the maximum angular acceleration a _m of the attitude maneuver of the aircraft according to the configured actuators on the aircraft;

(2) Determine the maximum angular velocity of the aircraft attitude maneuver according to the angular momentum envelope of the actuator configured on the aircraft and the range of the sensor

(3) When the aircraft receives the attitude maneuver control command sent by the ground, the aircraft control system according to the received attitude maneuver angle θ _m and the determined maximum angular velocity of the attitude maneuver and attitude maneuvering maximum angular acceleration a _m to calculate the characteristic turning moment of aircraft attitude maneuvering flight trajectory;

(4) The aircraft control system uses the characteristic turning time calculated in step (3) to calculate the target attitude angular acceleration a _r and the target attitude angular velocity of the aircraft in real time when the angular acceleration derivative is a segmented sinusoidal curve and the target attitude angle θ _r ;

(5) The attitude maneuver target attitude angular acceleration a _r and the target attitude angular velocity calculated by the aircraft control system according to step (4) and the target attitude angle θ _r for aircraft attitude maneuver control.

The maneuvering trajectory is generally divided into three processes, which are the acceleration segment, the constant velocity segment, and the deceleration segment. However, the constant velocity segment is not included in every maneuver process. For attitude maneuvers with small maneuvering angles, there are only the acceleration segment and the deceleration segment.

(1) The determination method of the maximum maneuvering angular acceleration a _m : assume that the maximum moment that the aircraft can provide in a certain direction of the body is T _mi , and the moment of inertia of the aircraft in the corresponding direction is I _Si , then the aircraft can provide in this direction maximum angular velocity For different maneuvering directions, the maximum angular acceleration in the corresponding direction can be calculated separately, and finally a _m is taken as a _m <min{a _mi }, and i is for all possible maneuvering directions;

(2) Maximum maneuvering angular velocity The determination method of : similar to the determination method of the maximum angular acceleration a _m , respectively calculate the maximum angular momentum H _mi that can be provided by the possible maneuvering direction, and the maximum angular velocity in this direction finally take as i is for all possible maneuvering directions.

(3) When the maximum maneuvering angular acceleration a _m and the maximum maneuvering angular velocity are determined After that, the maneuverability of the aircraft has been determined. When the aircraft receives the maneuver command, the acceleration maneuver angular velocity is θ _m , and trajectory planning can be performed. During trajectory planning, the time turning points are first calculated. The maneuvering flight trajectory of the aircraft is divided into three sections, which are respectively called the acceleration section, the constant _{velocity section} and the _{deceleration section} . _m3 ]. t _m1 , t _m2 , and t _m3 are all counted from the starting moment of the maneuver, which is called the characteristic turning moment of the maneuvering path. The method of obtaining the turning point of the feature is:

(11) when , the maneuvering angle of the aircraft is relatively large at this time, and the aircraft includes a constant speed segment when performing attitude maneuvers. t _m3 =t _m1 +t _m2 .

(12) when , the maneuvering angle of the aircraft is small at this time, and the aircraft only has two processes of acceleration and deceleration, where the acceleration time Since there is no constant velocity section, t _m2 =t _m1 , and the deceleration section is symmetrical to the acceleration section, t _m3 =t _m1 +t _m2 .

(4) During maneuver control, each control cycle is based on the planned maneuver trajectory, according to the maneuver moment and the key time turning point of the maneuver trajectory to determine the current specific stage of maneuver, and calculate the specific angular acceleration and angular velocity according to the trajectory planning method and maneuvering angle.

(41) Calculate the angular frequency of the angular acceleration derivative sinusoidal curve: It can be seen that the angular velocity derivative of the maneuvering acceleration section is a complete cycle of the sine curve, and the angular acceleration is the integral of the sine curve.

(42) When the maneuvering time t≤t _m1 , the maneuvering process is in the acceleration stage. Target angular acceleration a _r =0.5a _m (1-cos(f t)), target angular velocity Target angle θ _r =0.5a _m (0.5t ² -1/f+cos(f·t)/f ² ), where t is the maneuvering time. From the expression of a _r we know that, ${\stackrel{\&CenterDot;}{a}}_r\left(0\right)=0,{\stackrel{\&Center\; Dot;}{a}}_r\left(t_{m1}\right)=0,$ That is, the derivative of angular acceleration is continuous.

(43) When the maneuvering time t _m1 <t≤t _m2 , the maneuvering process is in the constant velocity section, the target attitude angular acceleration a _r =0, the target attitude angular velocity Corresponding to the maximum maneuvering angular velocity, the target attitude angle ${\mathrm{\&theta;}}_r=0.5\&CenterDot;a_m\&Center\; Dot;(t_{m1}\&Center\; Dot;t-0.5\&CenterDot;t_{m1}^2);$

(44) When the maneuvering time t _m2 <t≤t _m3 , the maneuvering process is in the deceleration stage, the target attitude angular acceleration a _r =-0.5a _m (1-cos(f(tt _m2 ))), the target attitude angular velocity ${\stackrel{\&CenterDot;}{\mathrm{\&theta;}}}_r=-0.5a_m(t-$ $t_{m3}-\sin \left(f(t-t_{m2})\right)/f),$ target attitude angle ${\mathrm{\&theta;}}_r=-0.5a_m(0.5t^2-t_{m3}\&CenterDot;t+\frac{\cos \left(f(t-t_{m2})\right)}{f^2})-\frac{1}{f^2}+t_{m1}^2+0.5t_{m3}(t_{m2}-t_{m1}).$ The angular acceleration of the maneuvering deceleration section and the acceleration section is antisymmetric, so the deceleration section also has the characteristic of continuous angular acceleration derivative.

The method of the invention can also be popularized and applied to attitude maneuvers of aircrafts, missiles and other aircrafts.

Example: Take a typical flexible aircraft rolling axis maneuvering process of 25° as an example, assuming that the rotational inertia of the aircraft is 3000kg.m ² , the fundamental frequency of the flexible sailboard is 0.8Hz, and the control period is 0.125s. Assume that the maximum torque that the aircraft actuator can provide is 55Nm, and the maximum angular momentum envelope is 210Nms. First, according to the maximum output torque and the angular momentum envelope, the maximum angular acceleration of the aircraft maneuvering is determined as a _m = 1.0°/s, and the maximum maneuvering angular velocity is Then the key rotor time points are calculated, and t _m1 =7.071s, t _m2 =7.071s, t _m3 =14.142s are obtained, and thus the maneuvering trajectory curve obtained by this method is obtained, as shown in Fig. 2 . Fig. 3 and Fig. 4 are curves of attitude angle error and angular velocity error of the aircraft when the method is used to control respectively. Statistics show that the time for the aircraft to achieve 0.001°/s (3σ) stability is 28.375s. Further analysis shows that the aircraft attitude control method based on angular acceleration sine trajectory planning can achieve 0.001°/s (3σ) stability in 31.250s, and the control method based on Bang-Bang trajectory planning can achieve the attitude stability of the aircraft within 50s Still greater than 0.001°/s, indicating that the attitude of the aircraft is not yet stable. Fig. 5 shows the modal coordinate displacement curve of the flexible sailboard when adopting this method, and Fig. 6 shows the modal coordinate displacement curve of the flexible sailboard when adopting the attitude control method based on angular acceleration sinusoidal trajectory planning, Fig. 7 shows the modal coordinate displacement curve of the flexible sailboard when the Bang-Bang trajectory planning attitude control method is adopted. By comparing Fig. The vibration displacement of the linear mode is very small, and the modal vibration is severe when the control method of the Bang-Bang trajectory is used, and the modal vibration amplitude is between the two when the angular acceleration sinusoidal trajectory planning control method is used. It can be seen that this method can effectively improve the excitation effect of the attitude maneuver on the flexible mode, and the time taken for the maneuver to stabilize is the shortest, thereby improving the maneuverability of the aircraft.

The content that is not described in detail in the description of the present invention belongs to the well-known technology of those skilled in the art.

Claims (2)

1. one kind is sinusoidal attitude of flight vehicle fast reserve method based on angular acceleration derivative, it is characterised in that step is such as Under:

(1) attitude of flight vehicle motor-driven maximum angular acceleration a is determined according to the executing agency configured on aircraft_m；

(2) determine that attitude of flight vehicle is motor-driven according to the executing agency's angular momentum envelope configured on aircraft and sensor range Big angular speed

(3) when aircraft receives the attitude maneuver control instruction that ground sends, flight control system is according to receiving Attitude maneuver angle, θ_mWith fixed attitude maneuver maximum angular rateWith attitude maneuver maximum angular acceleration a_mCalculate and fly The feature turnover moment of row device attitude maneuver flight path；

(4) flight control system utilizes the feature turnover moment that step (3) calculated, according to angular acceleration derivative be segmentation just Chord curve calculate in real time attitude of flight vehicle motor-driven time targeted attitude angular acceleration a_r, targeted attitude angular speedWith target appearance State angle, θ_r；

(5) flight control system is according to step (4) calculated attitude maneuver targeted attitude angular acceleration a_r, targeted attitude Angular speedWith targeted attitude angle θ_rCarry out attitude of flight vehicle maneuver autopilot；

Described calculating feature turnover the moment method be: set attitude of flight vehicle maneuverable path be divided into accelerating sections, at the uniform velocity section and Braking section three sections, wherein the time of accelerating sections is [0, t_m1], the at the uniform velocity time of section is [t_m1,t_m2], the time of braking section is [t_m2, t_m3], t_m1、t_m2、t_m3Feature for maneuverable path is transferred the moment, all with attitude of flight vehicle maneuvering flight initial time for rising Point timing；Obtaining the method in feature turnover moment is:

(a) whenTime,t_m3=t_m1+t_m2；

(b) whenTime,t_m2=t_m1, t_m3=t_m1+t_m2。

One the most according to claim 1 is sinusoidal attitude of flight vehicle fast reserve side based on angular acceleration derivative Method, it is characterised in that: target angular acceleration a during described step (4) calculating aircraft attitude maneuver in real time_r, target angular velocityWith angle on target θ_rMethod be:

(1) the sinusoidal angular frequency of angular acceleration derivative is asked for

(2) accelerating sections targeted attitude angular acceleration a during attitude maneuver_r=0.5 a_m(1-cos (f t)), targeted attitude Angular speedTargeted attitude angle θ_r=0.5 a_m·(0.5·t²-1/f+ cos(f·t)/f²), wherein t is the time kept in reserve；

(3) at the uniform velocity section targeted attitude angular acceleration a during attitude maneuver_r=0, targeted attitude angular speedTargeted attitude angle

(4) braking section targeted attitude angular acceleration a during attitude maneuver_r=-0.5 a_m·(1-cos(f(t-t_m2))), target Attitude angular velocityTargeted attitude angle