CN105116905A - Aircraft attitude control method - Google Patents

Abstract

The invention discloses an aircraft attitude control method, and the method comprises the following steps: 1) building a translational motion equation of the mass center of an aircraft; 2) considering the impact on the attitude control from earth rotation, and obtaining a corresponding equation of rotation around the mass center according to the translational motion equation of the mass center of the aircraft, wherein the equation decides the rotating angle and angular rate of the mass center of the aircraft, and is mainly used for achieving the attitude control of the aircraft; (3) giving out a reference aerodynamic force model; 4) designing an adaptive controller based on a Terminal sliding form, wherein the design comprises the design of a buffer controller and the design of a fast loop controller.

Description

A kind of Spacecraft Attitude Control

Technical field

The present invention relates to a kind of fast aircraft manufacturing technology technical field, be specifically related to a kind of Spacecraft Attitude Control.

Background technology

The gesture stability of aircraft, mainly in order to meet the strict ignition operation condition of air suction type punching engine.In posture adjustment process, aircraft also will complete the actions such as air intake duct is opened, fuel injection, engine ignition simultaneously.Hypersonic aircraft is more and more extensive in current various application, and larger change occurs along with the difference of engine behavior its every dynamic coefficient, makes controlled device have very strong uncertainty.Due to the hypersonic properties of flow of aircraft, cause its aerodynamic characteristic and attitude angle strong coupling, aerodynamic modeling is complicated, is difficult to be directly used in systematic analysis and Controller gain variations.In simulation study, usually the aerodynamic coefficients fit utilizing numerical evaluation to obtain or interpolation set up aerodynamic model, therefore for Controller gain variations, Aerodynamic Coefficient is unknown uncertain parameter, and the attitude of hypersonic aircraft is interorbital strong coupling, model non-linear, requires that its attitude controller has very strong adaptability and robustness.

For the inaccurate problem of model, sliding formwork control mode provides the Systematization method solving and keep stable and consistent performance problem.The major advantage that sliding formwork controls is that system responses is insensitive with interference to the uncertainty of model.It is single-input single-output (SISO) nonlinear system that sliding formwork controls the most ripe field of research.Although sliding formwork controls there is outstanding robust property, pure sliding formwork controls also have shortcoming, as required large control and controlling flutter phenomenon.The performance that pure sliding formwork controls can be improved by the estimation that it is coupled with on-line parameter, and only has and can realize full state when feeding back, and sliding mode controller just can realize.

Summary of the invention

Technical matters to be solved by this invention is for the deficiencies in the prior art, provides a kind of Spacecraft Attitude Control.

In order to solve the problems of the technologies described above, the invention discloses a kind of Spacecraft Attitude Control, comprising the following steps:

1) the translation motion equation of aircraft barycenter is set up;

2) consider that earth rotation is on the impact of gesture stability, obtains corresponding rotation around center of mass equation according to the translation motion equation of aircraft barycenter.This equation determines angle and its angular speed of aircraft rotation around center of mass, is mainly used to the gesture stability realizing aircraft;

3) Aerodynamic Model of reference is provided;

4) adaptive controller based on Terminal sliding formwork (terminal sliding mode) designs, comprising buffer control unit design and fast loop Controller gain variations.

Abovementioned steps 1) the translation motion equation of aircraft barycenter obtain according to following condition,

Aircraft is considered as controlled particle, considers that Spherical Earth rotation is on the impact reentering motion, following Three Degree Of Freedom can be obtained and be loaded into motion model:

${\stackrel{\·}{r}}_e=v\sin \mathrm{\γ}---\left(1\right)$

$\stackrel{\·}{\mathrm{\θ}}=\frac{v}{r_e}\cos \mathrm{\γ}\cos x---\left(3\right)$

$\stackrel{\·}{v}=\frac{1}{m}(Y\sin \mathrm{\β}-D\cos \mathrm{\β})-g\sin \mathrm{\γ}+{\mathrm{\Ω}}^2{r}_e\cos \mathrm{\θ}(\sin \mathrm{\γ}\cos \mathrm{\θ}-\cos \mathrm{\γ}\sin \mathrm{\θ}\cos x)---\left(4\right)$

$\begin{array}{c}\stackrel{\·}{x}=\frac{1}{mv\cos \mathrm{\γ}}(L\sin u+D\sin \mathrm{\β}\cos u+Y\cos \mathrm{\β}\cos u)+\frac{v}{r_e}\cos \mathrm{\γ}\sin x\tan \mathrm{\θ}-\\ 2\mathrm{\Ω}(\tan \mathrm{\γ}\cos \mathrm{\θ}\cos x-\sin \mathrm{\θ})+\frac{{\mathrm{\Ω}}^2{r}_e}{v\cos \mathrm{\γ}}\sin \mathrm{\θ}\cos \mathrm{\θ}\sin x\end{array}---\left(5\right)$

$\begin{array}{c}\stackrel{\·}{\mathrm{\γ}}=\frac{1}{mv}(L\cos u-Dsin\mathrm{\β}\sin u-Ycos\mathrm{\β}\sin u)-(\frac{g}{v}-\frac{v}{r_e})cos\mathrm{\γ}+2\mathrm{\Ω}cos\mathrm{\θ}\sin x\\ +\frac{{\mathrm{\Ω}}^2{r}_e}{v}cos\mathrm{\θ}(cos\mathrm{\γ}cos\mathrm{\θ}+sin\mathrm{\γ}sin\mathrm{\θ}\cos x)\end{array}---\left(6\right)$

Wherein: state of flight r _e, θ, v, x, γ represent the earth's core distance, longitude, dimension, flying speed, course angle and flight-path angle respectively; M represents vehicle mass; represent gravitational acceleration, g ₀represent terrestrial gravitation constant; Ω represents rotational-angular velocity of the earth; L, D, Y represent that aircraft reenters lift, resistance and the side force received in process respectively.

Abovementioned steps 2) in rotation around center of mass equation mainly consider the impact of earth rotation on aircraft manufacturing technology, the Three Degree Of Freedom attitude motion model that can obtain under body axis system is

$\stackrel{\·}{p}=\frac{M_x}{I_{\mathrm{xx}}}+\frac{(I_{\mathrm{yy}}-I_{\mathrm{zz}})}{I_{\mathrm{xx}}}\mathrm{qr}---\left(10\right)$

$\stackrel{\·}{q}=\frac{M_y}{I_{\mathrm{yy}}}+\frac{(I_{\mathrm{zz}}-I_{\mathrm{xx}})}{I_{\mathrm{tt}}}\mathrm{pr}---\left(11\right)$

$\stackrel{\·}{r}=\frac{M_y}{I_{\mathrm{zz}}}+\frac{(I_{\mathrm{xx}}-I_{\mathrm{yy}})}{I_{\mathrm{zz}}}\mathrm{pq}---\left(12\right)$

Wherein: state p, q, r, α, β, u represent roll angle speed, pitch rate, yawrate, the angle of attack, yaw angle and pitch angle respectively; M _x, M _y, M _zbe respectively the control moment of rolling, pitching and jaw channel; I _ij(i=x, y, z, j=x, y, z) represents the moment of inertia of aircraft.

Abovementioned steps 3) in reference gas dynamic model be the aerodynamic data of hypersonic aircraft X-33 adopted, reenter lift L, the resistance D and side force Y that aircraft in process is subject to and be respectively:

L＝q _dSC _L(M _a，α(13)

D＝q _dSC _D(M _a，α)(14)

Y＝q _dSC _Y(M _a，α(15)

Wherein flight vehicle aerodynamic area of reference S=2690ft ², dynamic pressure q _d=0.5 ρ (r) v ², lift coefficient C _l(M _a, α), resistance coefficient C _d(M _a, α) and lateral force coefficient C _y(M _a, α) and be expressed as the function that angle of attack Q and Mach number Ma, Ma are defined as the ratio of flying speed and the velocity of sound.

Step 4) in the involved design of the adaptive controller based on Terminal sliding formwork comprise two parts: the design of buffer control unit and the design of fast loop control unit; The design of controller is based on following three hypothesis:

Suppose 1: ignore earth rotation impact;

Suppose 2: ignore the amount describing track in the attitude of flight vehicle equation of motion, that is:

Suppose 3: the impact considering Parameter uncertainties and external disturbance, and sin β=0, tan β=0, cos β=1 is set up;

Based on above-mentioned hypothesis, the controller model after being simplified is:

$\stackrel{\·}{\mathrm{\γ}}=\mathrm{J\ω}+\mathrm{\Δf}---\left(16\right)$

$\stackrel{\·}{\mathrm{\ω}}=f_f+g_{f}M+\mathrm{\Δd}---\left(17\right)$

Wherein: ω=[p, q, r] ^trepresent attitude angular rate vector γ=[α, β, the u] of Hypersonic Reentry Vehicles ^trepresent attitude angle vector, m=[M _x, M _y, M _z] ^tthe control moment of expression system, Δ f=[f ₁, f ₂, f ₃] ^trepresent that the impact of orbital motion item on attitude motion causes uncertain, Δ d=[d ₁, d ₂, d ₃] ^trepresent the disturbance of outer bound pair Systematical control moment, J ∈ R ^{3 × 3}, f _f∈ R ^{3 × 1}, g _f∈ R ^{3 × 3}, and have:

$J=\left[\begin{array}{ccc}0& 1& 0\\ \sin \mathrm{\α}& 0& -\cos \mathrm{\α}\\ -\cos \mathrm{\α}& 0& -\sin \mathrm{\α}\end{array}\right]---\left(18\right)$

$f_f={[\frac{(I_{\mathrm{yy}}-I_{\mathrm{zz}})}{I_{\mathrm{xx}}}\mathrm{qr},\frac{(I_{\mathrm{zz}}-I_{\mathrm{xx}})}{I_{\mathrm{yy}}}\mathrm{pr},\frac{(I_{\mathrm{xx}}-I_{\mathrm{yy}})}{I_{\mathrm{zz}}}\mathrm{pq}]}^T---\left(19\right)$

$g_f=\mathrm{diag}\{\frac{1}{I_{\mathrm{xx}}},\frac{1}{I_{\mathrm{yy}}},\frac{1}{I_{\mathrm{zz}}}\}---\left(20\right)$

Divide attitude mode based on Multiple Time Scales, in view of the dynamic response rate of inner ring is far faster than outer shroud, the design of controller be divided into fast, slow two parts:

4-1) our department is divided into design buffer control unit, guidances command ω for generation of fast loop _c; Choose following sliding-mode surface function: $\mathrm{\&sigma;}={\mathrm{\&gamma;}}_e+{\&Integral;}_0^{\mathrm{\&tau;}}(a_1{\mathrm{\&gamma;}}_e+b_1{\mathrm{\&gamma;}}_e^{q_1/p_1})\mathrm{dt},$ Wherein: γ _e=γ-γ _c, γ _cfor needing guidanceing command of tracking; q ₁, P ₁for positive odd number, and meet q ₁< P ₁< 2q ₁; a ₁, b ₁be positive definite diagonal matrix; Order,

$\stackrel{\&CenterDot;}{\mathrm{\&sigma;}}=-{\stackrel{\&CenterDot;}{\mathrm{\&gamma;}}}_e+a_1{\mathrm{\&gamma;}}_r+b_1{\mathrm{\&gamma;}}_r^{\frac{q_1}{p_1}}=\stackrel{\&CenterDot;}{\mathrm{\&gamma;}}-{\stackrel{\&CenterDot;}{\mathrm{\&gamma;}}}_c+a_1{\mathrm{\&gamma;}}_e+b_1{\mathrm{\&gamma;}}_e^{\frac{q_1}{p_1}}=\mathrm{J\&omega;}+\mathrm{\&Delta;f}-{\stackrel{\&CenterDot;}{\mathrm{\&gamma;}}}_c+a_1{\mathrm{\&gamma;}}_e+b_1{\mathrm{\&gamma;}}_e^{\frac{q_1}{p_1}}$

Get then design following controller:

$w_c=J^{-1}({\stackrel{\&CenterDot;}{\mathrm{\&gamma;}}}_c-a_1{\mathrm{\&gamma;}}_e-b_1{\mathrm{\&gamma;}}_e^{\frac{q_1}{p_1}}-\mathrm{\&zeta;}\mathrm{sgn}(\mathrm{\&sigma;}\left)\right)---\left(21\right)$

4-2) our department is divided into the fast loop control unit of design, for generation of rolling, pitching and driftage control moment M _c:

Choose following sliding-mode surface function:

$s={\mathrm{\&omega;}}_e+{\&Integral;}_0^t(a_2{\mathrm{\&omega;}}_e+b_2{\mathrm{\&omega;}}_e^{q_2/p_2})\mathrm{dt};$

Wherein ω _cfor inner ring needs guidanceing command of tracking;

Q ₂, p ₂for positive odd number, and meet q ₂< p ₂< 2q ₂; a ₂, b ₂be positive definite diagonal matrix,

Order,

$\begin{array}{c}\stackrel{\&CenterDot;}{s}=-{\stackrel{\&CenterDot;}{\mathrm{\&omega;}}}_e+a_2{\mathrm{\&omega;}}_e+b_2{\mathrm{\&gamma;}}_e^{\frac{q_2}{p_2}}=\stackrel{\&CenterDot;}{\mathrm{\&omega;}}-{\stackrel{\&CenterDot;}{\mathrm{\&omega;}}}_c+a_2{\mathrm{\&omega;}}_e+b_2{\mathrm{\&omega;}}_e^{\frac{q_2}{p_2}}=f_f+g_{f}M+\mathrm{\&Delta;d}-{\stackrel{\&CenterDot;}{\mathrm{\&omega;}}}_e+\\ a_2{\mathrm{\&omega;}}_e+b_2{\mathrm{\&omega;}}_e^{\frac{q_2}{p_2}}\end{array}---\left(22\right)$

Get then design following controller:

The following robustness analyzing the controller designed by this method according to Lyapunov function theory, mainly can be divided into two steps, analyze respectively to buffer control unit and fast loop control unit:

According to Liapunov Lyapunov function theory, robust analysis is carried out to the buffer control unit designed by this method:

If system (16) adopts control law (21), then when controller parameter meets time, buffering sliding formwork meets reaching condition, namely meets robustness demand.

Prove: get Lyapunov function v=1/2 σ ^tσ, designs institute's sliding surface function of getting and formula (21) by buffer control unit, asks Lie derivative, have it along system trajectory:

So, closed-loop system signal bounded, and sliding-mode surface σ meets Lyapunov asymptotically stability.

According to Lyapunov function theory, robust analysis is carried out to the fast loop control unit designed by this method:

If system (17) adopts control law (23), then when controller parameter meets η _i> | d _i| time (i=1,2,3), fast loop sliding formwork meets reaching condition, namely meets robustness demand.

Prove: get Lyapunov function v=1/2s ^ts, is got sliding surface function and formula (23) by fast loop Controller gain variations, asks Lie derivative, have it along system trajectory:

$\begin{array}{c}\stackrel{\&CenterDot;}{v}=s^T\stackrel{\&CenterDot;}{s}=s^T(\mathrm{\&Delta;d}-\mathrm{\&eta;sgn}\left(s\right))\&le;\underset{i=1}{\overset{i=3}{\mathrm{\&Sigma;}}}\left|d_i\right|\left|s_i\right|-\underset{i=1}{\overset{i=3}{\mathrm{\&Sigma;}}}\left|{\mathrm{\&eta;}}_i\right|\left|s_i\right|\\ =-{\mathrm{\&Sigma;}}_{i=1}^{i=3}({\mathrm{\&eta;}}_i-|d_i\left|\right)\left|s_i\right|\&le;0\end{array}---\left(25\right)$

So, closed-loop system signal bounded, and sliding-mode surface s meets Lyapunov asymptotically stability.

In the present invention, a point on variable represents that single order is led, and represent speed, it is exactly two points that second order is led, and represents acceleration.

Gesture stability problem when technical solution of the present invention reenters for hypersonic aircraft, devises the controller based on adaptive sliding modeling method, achieves the tenacious tracking to hypersonic aircraft attitude angle.The method is mainly divided into two ingredients: the Controller gain variations of buffering and the Controller gain variations in fast loop.Under Parameter uncertainties and Bounded Perturbations situation, respectively robust analysis is carried out to the controller designed by this method by Lyapunov function, in view of the characteristic of Terminal sliding formwork finite time convergence control, illustrate that the tracking error of controller at finite time convergence control to zero, and then can achieve quick tracking hypersonic aircraft being reentered to attitude angle instruction in process.

The present invention reduces the impact of outside high-frequency noise on system performance, makes Attitude Controller have very strong adaptability and robustness, can realize the attitude-adaptive adjustment that aircraft is entering.First the present invention utilizes Multiple Time Scales technology that attitude mode is divided into twin nuclei; Then respectively for each loop design Terminal sliding mode controller, and proved by Lyapunov theory and the stability of singular perturbation theory to system, ensure that the Globally asymptotic of whole control system, improve the robustness of aircraft manufacturing technology system to Parameters variation.

The present invention sets up adaptive sliding mode model on the basis of nonlinear model, and then realizes the robust control to attitude of flight vehicle.

This method reenters the feature of the non-linear and strong coupling of model in process for hypersonic aircraft, consider the impact of Parameter uncertainties and BOUNDED DISTURBANCES, based on quick sliding-mode method design adaptive controller, with realize to hypersonic aircraft reentry guidance instruction stable, follow the tracks of fast.

Accompanying drawing explanation

Below in conjunction with the drawings and specific embodiments the present invention to be done and further illustrate. above-mentioned and/or otherwise advantage of the present invention will become apparent.

Fig. 1 is principle of the invention block diagram.

Embodiment

A kind of Spacecraft Attitude Control, comprises the following steps:

1) the translation motion equation of aircraft barycenter is set up;

2) consider that earth rotation is on the impact of gesture stability, obtains corresponding rotation around center of mass equation according to the translation motion equation of aircraft barycenter.This equation determines angle and its angular speed of aircraft rotation around center of mass, is mainly used to the gesture stability realizing aircraft;

3) Aerodynamic Model of reference is provided;

4) adaptive controller based on Terminal sliding formwork (terminal sliding mode) designs, comprising buffer control unit design and fast loop Controller gain variations.

Abovementioned steps 1) the translation motion equation of aircraft barycenter obtain according to following condition, aircraft is considered as controlled particle, considers that Spherical Earth rotation is on the impact reentering motion, obtain following Three Degree Of Freedom and be loaded into motion model:

${\stackrel{\&CenterDot;}{r}}_e=v\sin \mathrm{\&gamma;}---\left(1\right)$

$\stackrel{\&CenterDot;}{\mathrm{\&theta;}}=\frac{v}{r_e}\cos \mathrm{\&gamma;}\cos x---\left(3\right)$

$\stackrel{\&CenterDot;}{v}=\frac{1}{m}(Y\sin \mathrm{\&beta;}-D\cos \mathrm{\&beta;})-g\sin \mathrm{\&gamma;}+{\mathrm{\&Omega;}}^2{r}_e\cos \mathrm{\&theta;}(\sin \mathrm{\&gamma;}\cos \mathrm{\&theta;}-\cos \mathrm{\&gamma;}\sin \mathrm{\&theta;}\cos x)---\left(4\right)$

$\begin{array}{c}\stackrel{\&CenterDot;}{x}=\frac{1}{\mathrm{mv}\cos \mathrm{\&gamma;}}(L\sin u+D\sin \mathrm{\&beta;}\cos u+Y\cos \mathrm{\&beta;}\cos u)+\frac{v}{r_e}\cos \mathrm{\&gamma;}\sin x\tan \mathrm{\&theta;}-2\mathrm{\&Omega;}\\ (\tan \mathrm{\&gamma;}\cos \mathrm{\&theta;csox}-\sin \mathrm{\&theta;})+\frac{{\mathrm{\&Omega;}}^2{r}_e}{v\cos \mathrm{\&gamma;}}\sin \mathrm{\&theta;}\cos \mathrm{\&theta;}\sin x\end{array}---\left(5\right)$

$\begin{array}{c}\stackrel{\&CenterDot;}{\mathrm{\&gamma;}}=\frac{1}{\mathrm{mv}}(L\cos u-D\sin \mathrm{\&beta;}\sin u-Y\cos \mathrm{\&beta;}\sin u)-(\frac{g}{v}-\frac{v}{r_e})\cos \mathrm{\&gamma;}+2\mathrm{\&Omega;}\cos \mathrm{\&theta;}\sin x\\ +\frac{{\mathrm{\&Omega;}}^2{r}_e}{v}\cos \mathrm{\&theta;}(\cos \mathrm{\&gamma;}\cos \mathrm{\&theta;}+\sin \mathrm{\&gamma;}\sin \mathrm{\&theta;}\cos x)\end{array}---\left(6\right)$

Wherein: state of flight r _e, θ, v, x, γ represent the earth's core distance, longitude, dimension, flying speed, course angle and flight-path angle respectively; M represents vehicle mass; represent gravitational acceleration, g ₀represent terrestrial gravitation constant; Ω represents rotational-angular velocity of the earth; L, D, Y represent that aircraft reenters lift, resistance and the side force received in process respectively.

Abovementioned steps 2) in rotation around center of mass equation mainly consider the impact of earth rotation on aircraft manufacturing technology, the Three Degree Of Freedom attitude motion model that can obtain under body axis system is

$\stackrel{\&CenterDot;}{p}=\frac{M_x}{I_{\mathrm{xx}}}+\frac{(I_{\mathrm{yy}}-I_{\mathrm{zz}})}{I_{\mathrm{xx}}}\mathrm{qr}---\left(10\right)$

$\stackrel{\&CenterDot;}{q}=\frac{M_y}{I_{\mathrm{yy}}}+\frac{(I_{\mathrm{zz}}-I_{\mathrm{xx}})}{I_{\mathrm{tt}}}\mathrm{pr}---\left(11\right)$

$\stackrel{\&CenterDot;}{r}=\frac{M_y}{I_{\mathrm{zz}}}+\frac{(I_{\mathrm{xx}}-I_{\mathrm{yy}})}{I_{\mathrm{zz}}}\mathrm{pq}---\left(12\right)$

Wherein: state p, q, r, α, β, u represent roll angle speed, pitch rate, yawrate, the angle of attack, yaw angle and pitch angle respectively; M _x, M _y, M _zbe respectively the control moment of rolling, pitching and jaw channel; I _ij(i=x, y, z, j=x, y, z) represents the moment of inertia of aircraft.

Abovementioned steps 3) in reference gas dynamic model be the aerodynamic data of hypersonic aircraft X-33 adopted, reenter lift L, the resistance D and side force Y that aircraft in process is subject to and be respectively:

L＝q _dSC _L(M _a，α(13)

D＝q _dSC _D(M _a，α)(14)

Y＝q _dSC _Y(M _a，α(15)

Wherein flight vehicle aerodynamic area of reference S=2690ft ², dynamic pressure q _d=0.5 ρ (r) v ², lift coefficient C _l(M _a, α), resistance coefficient C _d(M _a, α) and lateral force coefficient C _y(M _a, α) and be expressed as angle of attack Q and Mach number M _a, M _abe defined as the function of the ratio of flying speed and the velocity of sound.

Step 4) in the involved design of the adaptive controller based on Terminal sliding formwork comprise two parts: the design of buffer control unit and the design of fast loop control unit; The design of controller is based on following three hypothesis:

Suppose 1: ignore earth rotation impact;

Suppose 2: ignore the amount describing track in the attitude of flight vehicle equation of motion, that is:

Suppose 3: the impact considering Parameter uncertainties and external disturbance, and sin β=0, tan β=0, cos β=1 is set up;

Based on above-mentioned hypothesis, the controller model after being simplified is:

$\stackrel{\&CenterDot;}{\mathrm{\&gamma;}}=\mathrm{J\&omega;}+\mathrm{\&Delta;f}---\left(16\right)$

$\stackrel{\&CenterDot;}{\mathrm{\&omega;}}=f_f+g_{f}M+\mathrm{\&Delta;d}---\left(17\right)$

Wherein: ω=[p, q, r] ^trepresent attitude angular rate vector γ=[α, β, the u] of Hypersonic Reentry Vehicles ^trepresent attitude angle vector, m=[M _x, M _y, M _z] ^tthe control moment of expression system, Δ f=[f ₁, f ₂, f ₃] ^trepresent that the impact of orbital motion item on attitude motion causes uncertain, Δ d=[d ₁, d ₂, d ₃] ^trepresent the disturbance of outer bound pair Systematical control moment, J ∈ R ^{3 × 3}, f _f∈ R ^{3 × 1}, g _f∈ R ^{3 × 3}, and have:

$J=\left[\begin{array}{ccc}0& 1& 0\\ \sin \mathrm{\&alpha;}& 0& -\cos \mathrm{\&alpha;}\\ -\cos \mathrm{\&alpha;}& 0& -\sin \mathrm{\&alpha;}\end{array}\right]---\left(18\right)$

$f_f={[\frac{(I_{\mathrm{yy}}-I_{\mathrm{zz}})}{I_{\mathrm{xx}}}\mathrm{qr},\frac{(I_{\mathrm{zz}}-I_{\mathrm{xx}})}{I_{\mathrm{yy}}}\mathrm{pr},\frac{(I_{\mathrm{xx}}-I_{\mathrm{yy}})}{I_{\mathrm{zz}}}\mathrm{pq}]}^T---\left(19\right)$

$g_f=\mathrm{diag}\{\frac{1}{I_{\mathrm{xx}}},\frac{1}{I_{\mathrm{yy}}},\frac{1}{I_{\mathrm{zz}}}\}---\left(20\right)$

Divide attitude mode based on Multiple Time Scales, in view of the dynamic response rate of inner ring is far faster than outer shroud, the design of controller be divided into fast, slow two parts:

4-1) our department is divided into design buffer control unit, guidances command ω for generation of fast loop _c; Choose following sliding-mode surface function: $\mathrm{\&sigma;}={\mathrm{\&gamma;}}_e+{\&Integral;}_0^{\mathrm{\&tau;}}(a_1{\mathrm{\&gamma;}}_e+b_1{\mathrm{\&gamma;}}_e^{q_1/p_1})\mathrm{dt},$ Wherein: γ _e=γ-γ _c, γ _cfor needing guidanceing command of tracking; q ₁, P ₁for positive odd number, and meet q ₁< P ₁< 2q ₁; a ₁, b ₁be positive definite diagonal matrix; Order,

$\stackrel{\&CenterDot;}{\mathrm{\&sigma;}}=-{\stackrel{\&CenterDot;}{\mathrm{\&gamma;}}}_e+a_1{\mathrm{\&gamma;}}_r+b_1{\mathrm{\&gamma;}}_r^{\frac{q_1}{p_1}}=\stackrel{\&CenterDot;}{\mathrm{\&gamma;}}-{\stackrel{\&CenterDot;}{\mathrm{\&gamma;}}}_c+a_1{\mathrm{\&gamma;}}_e+b_1{\mathrm{\&gamma;}}_e^{\frac{q_1}{p_1}}=\mathrm{J\&omega;}+\mathrm{\&Delta;f}-{\stackrel{\&CenterDot;}{\mathrm{\&gamma;}}}_c+a_1{\mathrm{\&gamma;}}_e+b_1{\mathrm{\&gamma;}}_e^{\frac{q_1}{p_1}}$

Get then design following controller:

$w_c=J^{-1}({\stackrel{\&CenterDot;}{\mathrm{\&gamma;}}}_c-a_1{\mathrm{\&gamma;}}_e-b_1{\mathrm{\&gamma;}}_e^{\frac{q_1}{p_1}}-\mathrm{\&zeta;}\mathrm{sgn}(\mathrm{\&sigma;}\left)\right)---\left(21\right)$

4-2) our department is divided into the fast loop control unit of design, for generation of rolling, pitching and driftage control moment M _c:

Choose following sliding-mode surface function:

$s={\mathrm{\&omega;}}_e+{\&Integral;}_0^t(a_2{\mathrm{\&omega;}}_e+b_2{\mathrm{\&omega;}}_e^{q_2/p_2})\mathrm{dt};$

Wherein ω _cfor inner ring needs guidanceing command of tracking;

Q ₂, p ₂for positive odd number, and meet q ₂< p ₂< 2q ₂; a ₂, b ₂be positive definite diagonal matrix,

Order,

$\begin{array}{c}\stackrel{\&CenterDot;}{s}=-{\stackrel{\&CenterDot;}{\mathrm{\&omega;}}}_e+a_2{\mathrm{\&omega;}}_e+b_2{\mathrm{\&gamma;}}_e^{\frac{q_2}{p_2}}=\stackrel{\&CenterDot;}{\mathrm{\&omega;}}-{\stackrel{\&CenterDot;}{\mathrm{\&omega;}}}_c+a_2{\mathrm{\&omega;}}_e+b_2{\mathrm{\&omega;}}_e^{\frac{q_2}{p_2}}=f_f+g_{f}M+\mathrm{\&Delta;d}-{\stackrel{\&CenterDot;}{\mathrm{\&omega;}}}_e+\\ a_2{\mathrm{\&omega;}}_e+b_2{\mathrm{\&omega;}}_e^{\frac{q_2}{p_2}}\end{array}---\left(22\right)$

Get then design following controller:

The invention provides a kind of Spacecraft Attitude Control; the method and access of this technical scheme of specific implementation is a lot; the above is only the preferred embodiment of the present invention; should be understood that; for those skilled in the art under the premise without departing from the principles of the invention; can also make some improvements and modifications, these improvements and modifications also should be considered as protection scope of the present invention.The all available prior art of each ingredient not clear and definite in the present embodiment is realized.

Claims (4)

1. a Spacecraft Attitude Control, is characterized in that, comprises the following steps:

1) the translation motion equation of aircraft barycenter is set up;

2) consider that earth rotation is on the impact of gesture stability, corresponding rotation around center of mass equation is obtained according to the translation motion equation of aircraft barycenter, rotation around center of mass equation determines angle and its angular speed of aircraft rotation around center of mass, is used for realizing the gesture stability of aircraft;

3) Aerodynamic Model of reference is provided;

4) adaptive controller based on Terminal sliding formwork designs, comprising buffer control unit design and fast loop Controller gain variations.

2. a kind of Spacecraft Attitude Control according to claim 1, is characterized in that, step 1) the translation motion equation of aircraft barycenter be obtain according to following condition:

Aircraft is considered as controlled particle, considers that Spherical Earth rotation is on the impact reentering motion, following Three Degree Of Freedom can be obtained and be loaded into motion model:

Wherein: state of flight r _e, θ, v, x, γ represent the earth's core distance, longitude, dimension, flying speed, course angle and flight-path angle respectively; M represents vehicle mass; G=g ₀r _e ²represent gravitational acceleration, g ₀represent terrestrial gravitation constant, Ω represents rotational-angular velocity of the earth; L, D, Y represent that aircraft reenters lift, resistance and the side force received in process respectively.

3. a kind of Spacecraft Attitude Control according to claim 2, it is characterized in that, step 2) in rotation around center of mass equation consider earth rotation on the impact of aircraft manufacturing technology, the Three Degree Of Freedom attitude motion model obtained under body axis system is:

Wherein: state p, q, r, α, β, u represent roll angle speed, pitch rate, yawrate, the angle of attack, yaw angle and pitch angle respectively; M _x, M _y, M _zbe respectively the control moment of rolling, pitching and jaw channel; I _ij(i=x, y, z, j=x, y, z) represents the moment of inertia of aircraft.

4. a kind of Spacecraft Attitude Control according to claim 3, it is characterized in that, step 3) in reference gas dynamic model be the aerodynamic data of hypersonic aircraft X-33 adopted, the lift L that in loading process, aircraft is subject to, resistance D and side force Y are respectively:

L＝q _dSC _L(M _a，α(13)

D＝q _dSC _D(M _a，α)(14)

Y＝q _dSC _Y(M _a，α(15)

Wherein flight vehicle aerodynamic area of reference S=2690ft ², dynamic pressure q _d=0.5 ρ (r) v ², lift coefficient C _l(M _a, α), resistance coefficient C _d(M _a, α) and lateral force coefficient C _y(M _a, α) and be expressed as angle of attack Q and Mach number M _a, M _abe defined as the function of the ratio of flying speed and the velocity of sound.