The way in which the initial orientation of the spacecraft

 

(57) Abstract:

The invention is intended for use in spacecraft control systems orientation of the satellite. The technical result - the simplification of the laws of management and technical implementation of the system the initial orientation of the spacecraft. How is the initial damping of the oscillations of the spacecraft on the signals of the gyroscope. By calculation determine the correspondence between the positions of stable equilibrium of the gyro stops and the ratio of the magnitudes and signs of the angular velocity of the apparatus. For each stable position of the gyro stops distinguish the dominant angular velocity. Compensate for this speed signals from the angle sensors gyroscope using onboard computers and the Executive bodies of the orientation system, after which the gyroscope is a new stable equilibrium state in accordance with the remaining uncompensated angular velocity, which is similarly offset by signals from the angle sensors gyro to the output of the gyroscope in the working area. 3 Il.

The invention relates to the field of control of spacecraft (SC) and can be ispolzovanie rotates due to perturbations of the push.

Operation initial orientation imposes specific requirements as to the sensors and to the control device. So often it uses its own set of instruments.

There is a method of damping vibrations KA [1] on the basis of the relay system, which includes a free gyro, OS amplifier, relay, solenoid valve and jet engines.

The method consists in the fact that the control moment generated by the jet engines, is formed on the basis of the signals of the free gyroscope due to the on and off of the engine by an electromagnetic valve, which controls the three-position relay.

The main disadvantages of this method are significant fuel consumption, the necessity of law control angle and angular velocity deviation KA to ensure system stability, i.e. high degree of complexity and the large dimensions of the system.

Known methods [2] initial solar orientation of the spacecraft on the basis of the special subsystem initial orientation, which includes: an optical sensor of the Sun, the angular velocity sensors (VCS), the control unit. As grant the ability of operations, but with used equipment. Criteria for comparative evaluation of various schemes of solar orientation can serve as their operational characteristics: the time required for a full cycle of primary solar orientation, the flow of compressed gas during this operation, the complexity of the technical implementation of the system, reliability of operations.

The complexity of the technical implementation of the system can be described by required set of instruments. The requirement for reliability of operation start of the exhibition is very great, because from the success it is often entirely the possibility of the functioning of the AC.

Distinguish the following classes of possible schemes solar orientation:

1) Scheme of parallel casting (both axes), using the angular velocity signals CAx,y,zfrom 3-4 Usov and two guide cosine for the unit vector pointing to the Sun.

2) the successive casts, using the signals of the angular velocityx,y,zfrom 3 Dosov, and signals the presence of the Sun in a limited field of view.

3) Schemes with limited use VCS:

(a) scheme that uses the signals to the provisions of jsousa only angular information.

The known method [2] , adopted as a prototype, the initial damping of the oscillations of the SPACECRAFT, namely, that the power gyrostabilization solves this problem by using jet engines, forming an external braking torques around the axis of the apparatus, for which he served as command-sensitive device is a two - component sensor of angular velocity.

In [2] it is shown that the rotation angle around the axis of the gyroscope Kardanov suspension provide information on the projections of the angular velocity of the apparatus, i.e., confirmed the possibility of using gyroscopic stabilizer as component of the angular rate sensor capable of managing jet engines, creating a braking torques around the axis of the apparatus.

The aim of the invention is the simplification of the laws of management and technical implementation of the system the initial orientation of the spacecraft.

This goal is achieved by defining the correspondence between the positions of stable equilibrium threefold gyro (TG), located on the lugs, and the ratio of the magnitudes and signs of the projections of the angular velocities of the SPACECRAFT on the axis associated with the SPACECRAFT coordinate system in the case when the angular skorostyami, similar to the characteristics of the proposed method. This allows us to conclude that the proposed method has significant differences.

The method is illustrated by drawings, where Fig.1 shows all possible cases of stable equilibrium of the gyro on the fence. In Fig.2 shows associated with housing the camera coordinate system and the angular velocity vectors of the framework of TG. In Fig.3 shows an algorithm for processing information about the signs of the projections of the angular velocities of the SPACECRAFT on the axis of the associated coordinate system.

The method is implemented as follows. The position of the TG stops the stops and, therefore, the signs of the signals from the angle sensors TG is determined by the direction of rotation of the SPACECRAFT with respect to the axes associated with the SPACECRAFT coordinate system and the ratio of modules angular velocity,

As can be seen from Fig.1, the position of the focus TG under the influence of the angular velocities exceeding limit values, a certain regularity.

When various combinations of modules and marks the angular velocities of the SPACECRAFT exceeding limit values, leads to the same state. For example, to the state of the +yandybring the following cases: -x; -x+z; -x-zwhen |isz| |x|.

dimensions can be distinguished in all other States. Based on this way of quieting the object when working TG stops.

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

Each stable position TG stops corresponds dominant over the rest of the angular velocity of the SPACECRAFT. For the first quadrant, corresponding to the positive signals from the angle sensors TG (+y, +y), a dominant negative angular velocity isxrelative to the X-axis KA, for the second quadrant (+y, -y), positive angular velocity +zalong the Z-axis KA, for the third quadrant (-y, -y) - positive angular velocity +xrelative to the X-axis KA, for the fourth quadrant (-y, +y) negative angular velocity iszalong the Z-axis of the SPACECRAFT.

Originally compensation is the dominant angular velocity of the SPACECRAFT. After payment of the dominant angular velocity TG takes the new position of stable equilibrium, which is determined by the remainder of the measured data TG angular velocity. Payment of the remaining angular velocity continues to return TG with stops in the working area.

Theoretical background the possibility of creating algorithm Denia on the set TG (in this case, the basis of the TG is KA) with angular velocities exceeding the measurement range of the TG, there is a "knockout" of TG. We show that in this case the reading TG can be used for damping KA.

The equations of motion of TG in gimbals are [3]:

< / BR>
< / BR>
where A, C, A1B1C1, A2- moments of inertia, respectively, of the rotor of the gyroscope, the inner and outer frame Kardanov suspension;

- angular velocity of rotation of the rotor;

the absolute angular velocity of rotation of the inner frame;

the absolute angular velocity of rotation of the inner frame;

L- the moment of external forces about the axis of rotation of the outer ring;

LN- the moment of external forces about the axis of rotation N of the inner ring.

For gyro type HPA, the following notation:

Ie- Equatorial moment of inertia of the rotor TG;

N - kinetic torque TG;

ythe ultimate USBC deviations internal frame, limited focus;

y- limit the deflection angle of the outer frame, the limited focus;

the vector of absolute angular velocity of the object,

vector brand is istula axis TG.

Given the legend, ignoring the moments of inertia frames and considering the angles of deflection of the gyro small, the equations of motion of a gyroscope can be written in the form:

< / BR>
< / BR>
Mutual arrangement of the axes AC and TG HPA shown in Fig.2, where we have introduced the following notation:

XcYcZc- axis associated with KA;

XNRTHENRZNR- axis associated with the inner frame;

XnYnZn- axis associated with the outer frame.

It is evident from Fig. 2 the resulting expression is the absolute angular velocity of the gyroscope with a closed backward links:

< / BR>
< / BR>
Thus, the equation of motion TG type HPA on the movable KA when closed feedbacks in accordance with Fig.2 have the form:

< / BR>
< / BR>
For the casex= const,y= const,z= const equation have the form:

< / BR>
< / BR>
Neglecting the influence of the angular velocity ydirected along the axis of the angular momentum of the gyroscope for small angles, and given that

Mz= -k,

Mx= -k,

where k is the transmission coefficient in the correction circuits gyro, get

< / BR>
< / BR>
In normal operation of the gyroscope, i.e., | | <yand || <ywe will limit ourselves to consideration of the precessional uravnenii and SPACECRAFT from the launch vehicle angular velocity of the SPACECRAFT can reach significant values, which exceed the measuring range of the TG. It is well known that there is a "knockout" TG, i.e., TG refers to first one and then another stop, after which occupies the position of stable equilibrium.

We define the ratio between the angular velocities of the object and conditions of stable equilibrium of the gyroscope.

Suppose thatz> 0,x>0 and |z| > |x|. At the moment of contact lugs t=t1;

The initial conditions for the second leg movements define at t=t1:

< / BR>
< / BR>
Integrating the equations of motion of a gyroscope using these initial conditions, we will have:

< / BR>
For time t1you can write:

< / BR>
where

< / BR>
< / BR>
where

< / BR>
Thus, when you touch the wall stopsythere is uniform motion TG around the axis X, and > 0 and touch focusy.

At time t= t2TG will occupy a stable position of equilibrium.

Then the equations of motion of a gyroscope will take the form:

Hx= -ky+Mzp,

Hz= -ky+Mxp,

i.e., gyroscopic moments at theyandywill be counterbalanced by moments of supports and reactions TG will conduct the first effect of the present invention is that the use of the proposed method eliminates the need for additional equipment for the initial orientation of the SPACECRAFT.

Literature

1. Alekseev, K. B. , byabenin,, office of space flight vehicles. M., engineering, 1974

2. Rauschenbach B. C., Turner E. N. Attitude control of a spacecraft. Publishing house "Nauka, main editorial Board for physical and mathematical literature, M., 1974

3. Nicolai E. L. Gyroscope in gimbals. Ed. 2-e, M, Science, 1964

The way in which the initial orientation of the spacecraft, consisting of the initial damping of the oscillations of the spacecraft on the gyro signals, characterized in that the calculation determines the correspondence between the positions of stable equilibrium threefold gyro stops and the ratio of the magnitudes and signs of the angular velocities of the spacecraft and for each stable position threefold gyro stops distinguish the dominant angular velocity, compensate for this speed signals of the angle sensors threefold gyroscope using onboard computers and the Executive bodies of the orientation system, after which tregs vannoy angular velocity, which similarly compensate for the signals of the angle sensors threefold gyro to the release of the latest in the working area.

 

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