RU2180729C2 - Method of initial orientation for spacecraft - Google Patents

Abstract

FIELD: space engineering; spacecraft control systems. SUBSTANCE: proposed method consists in damping initial oscillations of spacecraft by gyroscope signals. Correspondence of positions of stable equilibrium of gyroscope on stops and magnitude and signs of spacecraft angular velocities is found through calculations. Predominant angular velocity is selected for each stable position of gyroscope on stops. This velocity is compensated for by signals received from gyroscope attitude sensors by means of on-board computer and actuating members of attitude control system, after which gyroscope takes new position of stable equilibrium in accordance with residual non-compensated angular velocity which is compensated for similarly by signals from gyroscope attitude sensors till getting into working zone. EFFECT: simplified of control laws; facilitated procedure. 3 dwg

Description

 The invention relates to the field of spacecraft (SC) control and can be used in Earth satellite orientation systems.

 At the first moment after the separation of the spacecraft from the carrier, the spacecraft rotates randomly due to perturbation-shock.

 The initial orientation operation presents specific requirements for both the sensors and the control device. Therefore, often it uses its own set of devices.

 A known method of damping oscillations of a spacecraft [1] on the basis of a relay system that includes a free gyroscope, an OS amplifier, a relay, an electromagnetic valve and jet engines.

 The method consists in the fact that the control moment created by jet engines is formed on the basis of signals from a free gyroscope by turning the engines on and off using an electromagnetic valve, the operation of which is controlled by a three-position relay.

 The main disadvantages of this method are significant fuel consumption, the need to organize a control law for the angle and angular velocity of the spacecraft deviations to ensure the stability of the system, i.e. the degree of complexity is large and the dimensions of the system are significant.

 Known methods [2] of the initial solar orientation of spacecraft based on a special subsystem of initial orientation, which includes: an optical sensor of the Sun, angular velocity sensors (DLS), a control device. As executive bodies, gas-jet nozzles are used. The methods differ from each other not only in the sequence of operations, but also in the equipment used for this. The criteria for a comparative assessment of various solar orientation schemes can be their operational characteristics: the time spent on the full cycle of the initial solar orientation, the flow of compressed gas during this operation, the complexity of the technical implementation of the system, and the reliability of the operations.

 The complexity of the technical implementation of the system can be characterized by the required set of devices. The requirement for the reliability of the operation of the initial exhibition is very high, since the possibility of the spacecraft functioning entirely depends on its success.

The following classes of possible solar orientation schemes are distinguished:
1) Parallel reduction circuits (simultaneously along two axes) using spacecraft angular velocity signals ω _x , ω _y , ω _z from 3-4 DOSs and two cosine guides for a unit vector directed at the Sun.

2) Sequential reduction schemes using signals of angular velocity ω _x , ω _y , ω _z from 3 DOSs, and signals of the presence of the Sun in a limited field of view.

3) Schemes with limited use of CRS:
a) schemes using signals about the position of the Sun and a signal about the angular velocity of rotation of the spacecraft only around the "solar" axis ω _x ;
b) schemes using only angular information.

 The known method [2], adopted as a prototype, damping the initial oscillations of the spacecraft, which consists in the fact that the power gyrostabilizer solves this problem using jet engines that form external braking moments around the axes of the apparatus, for which it serves as a command sensitive device - a two-component angular sensor speed.

 In [2], it was shown that the rotation angles of the gyroscope around the axes of the cardan suspension give information about the projections of the angular velocity of the apparatus, the possibility of using a gyroscopic stabilizer as a two-component angular velocity sensor that can control jet engines that create braking moments around the axes of the apparatus is confirmed.

 The aim of the invention is to simplify the laws of control and technical implementation of the initial orientation system of spacecraft.

 This goal is achieved by determining the correspondence between the positions of stable equilibrium of a three-stage gyroscope (TG) located on the stops and the ratio of the values and signs of the projections of the angular velocity of the spacecraft on the axis of the coordinate system associated with the spacecraft in the case when the angular velocity of the spacecraft exceeds the measurement range of the TG.

 None of the known technical solutions revealed signs similar to those of the proposed method. This allows us to conclude that the proposed method has significant differences.

 The method is illustrated by drawings, where figure 1 shows all possible cases of stable equilibrium of the gyro on the stop. Figure 2 shows the coordinate system associated with the spacecraft body and the angular velocity vectors of the TG frames. Figure 3 shows the algorithm for processing information about the signs of the projections of the angular velocity of the spacecraft on the axis of the associated coordinate system.

The method is implemented as follows. The position of the TG on the stops and, therefore, the signs of the signals from the TG angle sensors are determined by the direction of rotation of the spacecraft relative to the axes of the coordinate system associated with the spacecraft and the ratio of the angular velocity moduli,
As can be seen from figure 1, the position of the stop TG under the influence of angular velocities exceeding the limiting values, has a certain pattern.

With various combinations of modules and signs of the angular velocity of the spacecraft exceeding the limiting values, it leads to the same state. For example, the following cases lead to the state + α _y and β _y : - ω _x ; - ω _x + ω _z ; - ω _x -ω _z for | -ω _z | ≫ | ω _x |.
It is obvious that in all these cases there is an angular velocity of -ω _x .
Similar patterns can be distinguished in all other states. This is the basis for the method of calming the object during operation of the TG at stops.

 The method is illustrated by the algorithm shown in figure 3.

Each stable position of the TG on the stops corresponds to the angular velocity of the spacecraft that dominates over the rest. For the first quadrant, corresponding to the positive signals from the TG angle sensors (+ α _y , + β _y ), the negative angular velocity -ω _x relative to the X axis of the spacecraft is dominant, for the second quadrant (+ α _y , -β _y ), the positive angular speed + ω _z relative to the spacecraft axis Z, for the third quadrant (-α _y , -β _y ) - positive angular velocity + ω _x relative to the spacecraft X axis, for the fourth quadrant (-α _y , + β _y ) negative angular velocity -ω _z relative to the Z axis of the spacecraft.

 Initially, the dominant angular velocity of the spacecraft is compensated. After compensating for the dominant angular velocity, the TG takes a new position of stable equilibrium, which is determined by the angular velocity remaining from the measured TG data. Compensation of the remaining angular velocity continues until the TG returns from the stops to the working area.

Theoretical prerequisites for the possibility of creating an algorithm for damping object vibrations from TG signals at stops
It is known that during rotation of the base onto which the TG is mounted (in this case, the base of the TG is a spacecraft) with angular velocities exceeding the range of TG measurement, the TG is knocked out. We show that in this case the TG reading can be used to damp the vibrations of the spacecraft.

где A, C, A₁, B₁, C₁, A₂ - моменты инерции соответственно ротора гироскопа, внутренней и наружной рамок карданова подвеса;
ω - угловая скорость вращения ротора;

абсолютная угловая скорость вращения внутренней рамки;

абсолютная угловая скорость вращения внутренней рамки;
L_ζ- момент внешних сил относительно оси вращения ζ наружного кольца;
L_N - момент внешних сил относительно оси вращения N внутреннего кольца.The equations of motion of the TG in the gimbal have the form [3]:

where A, C, A ₁ , B ₁ , C ₁ , A ₂ - the moments of inertia, respectively, of the gyro rotor, the inner and outer frames of the cardan suspension;
ω is the angular velocity of rotation of the rotor;

absolute angular speed of rotation of the inner frame;

absolute angular speed of rotation of the inner frame;
L _ζ is the moment of external forces relative to the axis of rotation ζ of the outer ring;
L _N is the moment of external forces relative to the axis of rotation N of the inner ring.

вектор абсолютной угловой скорости объекта,
Ω - вектор абсолютной угловой скорости гироскопа относительно инерциального пространства;
М_х, М_z - моменты, действующие по осям ТГ.The following designations are accepted for a GPA gyroscope:
I _e - equatorial moment of inertia of the TG rotor;
N is the kinetic moment of TG;
β _y - utop utop deviation of the inner frame, limited focus;
α _y - the maximum deviation angle of the outer frame, limited by the emphasis;

vector of the absolute angular velocity of the object,
Ω is the vector of the absolute angular velocity of the gyroscope relative to inertial space;
M _x , M _z - moments acting on the axes of the TG.

Взаимное расположение осей КА и ТГ ГПА показано на фиг.2, где введены следующие обозначения:
X_cY_cZ_c - оси, связанные с КА;
Х_внУ_внZ_вн - оси, связанные с внутренней рамкой;
X_нY_нZ_н - оси, связанные с наружной рамкой.Given the accepted notation, neglecting the moments of inertia of the frames and considering the angles of deviation of the gyroscope small, the equations of motion of the gyroscope can be written in the form:

The relative position of the axes of the spacecraft and TG GPA is shown in figure 2, where the following notation is introduced:
X _c Y _c Z _c - axis associated with the spacecraft;
X _{VN VN} Z _VN - axis associated with the inner frame;
X _n Y _n Z _n - axis associated with the outer frame.

Таким образом, уравнение движении ТГ типа ГПА на подвижном КА при замкнутых обратных связях в соответствии с фиг.2 имеют вид:

Для случая ω_x= const, ω_y= const, ω_z= const уравнения имеют вид:

Пренебрегая влиянием угловой скорости ω_y, направленной по оси кинетического момента гироскопа для малых углов α и β и учитывая, что
M_z= -kα,
M_x= -kβ,
где k - коэффициент передачи в цепях коррекции гироскопа, получим

В нормальном режиме работы гироскопа, т.е. |α| < α_y и |β| < β_y ограничимся рассмотрением прецессионных уравнений

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или при переходе на картинную плоскость

При отделении КА от носителя угловые скорости КА могут достигать значительных величин, которые превышают диапазон измерения ТГ. Известно, что при этом происходит "выбивание" ТГ, т.е. ТГ касается сначала одного, а затем другого упора, после чего занимает положение устойчивого равновесия.From FIG. 2 we obtain the expressions of the absolute angular velocity of the gyroscope with closed feedbacks:

Thus, the equation of motion of a TG of the GPA type on a moving spacecraft with closed feedbacks in accordance with figure 2 have the form:

For the case of ω _x = const, ω _y = const, ω _z = const, the equations have the form:

Neglecting the influence of the angular velocity ω _y directed along the axis of the kinetic moment of the gyroscope for small angles α and β and taking into account that
M _z = -kα,
M _x = -kβ,
where k is the transmission coefficient in the gyro correction circuits, we obtain

In the normal mode of operation of the gyroscope, i.e. | α | <α _y and | β | <β _y restrict ourselves to the consideration of precession equations

where from

or when moving to the picture plane

When separating the spacecraft from the carrier, the angular velocity of the spacecraft can reach significant values that exceed the measurement range of the TG. It is known that in this case the TG is "knocked out", i.e. TG touches first one, and then another emphasis, after which it occupies a position of stable equilibrium.

 We determine the relationship between the angular velocities of the object and the positions of the stable equilibrium of the gyroscope.

Начальные условия для второго участка движения определим при t=t₁:

Интегрируя уравнения движения гироскопа с учетом полученных начальных условий, будем иметь:

Для момента времени t₁ можно записать:

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Таким образом, при касании стенки упора β_y наблюдается равномерное движение ТГ вокруг оси X, причем α > 0 до касания упора α_y.
В момент времени t= t₂

ТГ займет устойчивое положение равновесия.Assume that ω _z > 0, ω _x > 0 and | ω _z | > | ω _x |. At the moment of contact of the stops t = t ₁ ;

The initial conditions for the second section of motion are defined at t = t ₁ :

Integrating the equations of motion of the gyroscope taking into account the obtained initial conditions, we will have:

For time t ₁ you can write:

where from

where from

Thus, when the wall of the stop β _y touches, the TG moves uniformly around the X axis, and α> 0 until the stop touches α _y .
At time t = t ₂

TG will take a stable equilibrium position.

In this case, the equations of motion of the gyroscope will take the form:
Ω _x = -kα _y + M $\genfrac{}{}{0ex}{}{\text{z}}{\text{p}}$ ,
Ω _z = -kβ _y + M $\genfrac{}{}{0ex}{}{\text{x}}{\text{p}}$ ,
those. the gyroscopic moments in the position α _y and β _{y will be} balanced by the moments of reactions of the supports and the TG will perform a forced rotation in inertial space with speeds + ω _x and + ω _z .
The technical effect of the invention is that the use of the proposed method eliminates the need for additional equipment for the initial orientation mode of the spacecraft.

Literature
1. Alekseev K. B., Bebenin G. G. Spacecraft control. M., Mechanical Engineering, 1974

 2. Raushenbakh B.V., Tokar E.N. Spacecraft orientation control. Publishing house "Science", the main edition of the physical and mathematical literature, M., 1974

 3. Nikolai E.L. Gyroscope in a gimbal. Ed. 2nd, M., Science, 1964

Claims (1)

 The method of initial orientation of the spacecraft, which consists in damping the initial oscillations of the spacecraft according to the gyroscope signals, characterized in that the calculation determines the correspondence between the positions of the stable equilibrium of the three-stage gyroscope on the stops and the ratio of the values and signs of the angular velocities of the spacecraft and for each stable position of the three-degree gyroscope on stops emphasize the dominant angular velocity, compensate for this speed according to the signals of angle sensors three degrees of the gyro using the onboard computer and the executive bodies of the attitude control system, then the threefold gyroscope takes the new position of stable equilibrium in accordance with the remaining uncompensated angular velocity, which is similar to compensate for signals threefold gyro angle sensor before the latter in the work area.

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