RU2487823C1 - Method of adaptive control over displacement of centre of gravity of spacecraft - Google Patents

Abstract

FIELD: transport.

SUBSTANCE: invention relates to control over flight of the group of spacecraft and may be used for tracking one spacecraft by another spacecraft at preset distance. Proposed method comprises measuring trajectories and making corrections, minimizing orbit eccentricity and defining position of subject spacecraft in inertial space. Note here that control and navigation system is provided with set of transceiving radio hardware and optical transducer of angles "Pole Star - subject spacecraft - object spacecraft". Distance to object spacecraft is measured to define is deviation from mean magnitude at measurement steps. At termination of every cycle of measurement steps, dynamics of variation of said mean magnitude is revealed to define increment of oscillation period of subject spacecraft relative to similar period of object spacecraft. In one orbital period, angle between planes of orbits of object spacecraft and subject spacecraft are defined as well as time of crossing of said planes by readings of angle transducer. In case said increment exceeds preset threshold, parameters of correction of said period are computed. At estimated time of orbit crossing, correction engines are initiated giving preference to engine of orbit inclination correction engine.

EFFECT: higher accuracy of tracking, lower poser consumption of spacecraft correction system.

1 dwg

Description

The invention relates to the field of space technology and can be used to control the movement of the center of mass of a spacecraft (SC) when the spacecraft is accompanied by another spacecraft at a given distance.

1. The well-known "Method for simultaneous correction of the retention of the inclination vector of the orbit and the period of revolution of a three-axis stabilized spacecraft" (RU 2381965, IPC B64G 1/24), including the correction, characterized in that they determine the required angle of deviation of the engine thrust vector from normal to orbit in the plane yaw based on ensuring the specified accuracy of the correction of the period of revolution of the spacecraft and the required changes for the correction of the transversal and orthogonal components of the space velocity vector of the device, set the deviation direction of the thrust vector engines, engine run time is calculated according to the formulas

$${\tau }_{\mathrm{one}}=\frac{J_n\cdot sin\left|{\mathrm{<figure-callout\; id="2"\; label="\theta "\; filenames="00000017.png"\; state="\{\{state\}\}">\theta </figure-callout>}}_2\right|+J_{\tau }\cdot cos{\theta }_2}{F_{\mathrm{one}}\cdot sin(|{\theta }_{\mathrm{one}}|+|{\mathrm{<figure-callout\; id="2"\; label="\theta "\; filenames="00000017.png"\; state="\{\{state\}\}">\theta </figure-callout>}}_2|)}$$

$${\mathrm{<figure-callout\; id="2"\; label="\tau "\; filenames="00000017.png"\; state="\{\{state\}\}">\tau </figure-callout>}}_2=\frac{J_n\cdot sin\left|{\theta }_{\mathrm{one}}\right|-J_{\tau }\cdot cos{\theta }_{\mathrm{one}}}{{\mathrm{<figure-callout\; id="2"\; label="F"\; filenames="00000017.png"\; state="\{\{state\}\}">F</figure-callout>}}_2\cdot sin(|{\theta }_{\mathrm{one}}|+|{\mathrm{<figure-callout\; id="2"\; label="\theta "\; filenames="00000017.png"\; state="\{\{state\}\}">\theta </figure-callout>}}_2|)}$$

where τ ₁ , τ ₂ - the duration of the engines, s;

J _n , J _τ - thrust impulses required for corrections, respectively, of the inclination vector of the orbit and the spacecraft orbital period, N · s;

F ₁ , F ₂ - engine thrust, N;

θ ₁ , θ ₂ are the angles of deviation of the engine thrust vectors from the normal to the orbit plane in the yaw plane along the smallest arc,

and carry out the correction by a pair of engines installed on different sides from the normal to the orbit, for which they are turned on sequentially for the estimated duration of work. The following sequence of operations is performed (all angular values are expressed in radians).

1. The required angle (θ) of the deviation of the engine thrust vector from the normal to the orbit is determined based on ensuring the specified accuracy of the correction of the spacecraft rotation period and the required changes for the correction of the transversal and orthogonal components of the spacecraft velocity vector.

The angle θ is calculated by the formula

$$\theta =\{\begin{array}{c}arcctg(\gamma ),PR\mathrm{and}\gamma \le \frac{\Delta V_n}{\Delta V_{\tau }};\\ arcctg\left(\frac{\Delta V_n}{\Delta V_{\tau }}\right),PR\mathrm{and}\gamma >\frac{\Delta V_n}{\Delta V_{\tau }},\end{array}$$

where θ is the acute angle of deviation of the engines on different sides from the normal to the orbit in the yaw plane;

due to the fact that $$\delta V_{\tau }=\frac{sin(\theta +\delta \theta )}{sin(\theta )}-\mathrm{one}$$

should

$$\gamma =\frac{\mathrm{one}+\delta V_{\tau }-cos(\delta \theta )}{sin\delta \theta }$$

δθ is the error with which the position of the spacecraft in the yaw plane is maintained relative to the center of mass;

δθ _τ is the specified maximum relative error in the implementation of the correcting pulse to change the spacecraft revolution period;

ΔV _τ is the required maximum change for the correction of the transversal component of the spacecraft velocity vector during the period of its active existence, m / s;

(составляющие вектора наклонения i_x=sin(i)·cos(Ω); i_y=sin(i)·sin(Ω), Ω-долгота восходящего узла орбиты), м/с.ΔV _n is the required change for the correction of the orthogonal component of the spacecraft velocity vector, corresponding to the calculated maximum change in the inclination vector $$\overline{\Delta i}$$

(components of the inclination vector i _x = sin (i) · cos (Ω); i _y = sin (i) · sin (Ω), Ω is the longitude of the ascending node of the orbit), m / s.

It should be noted that the method does not require angular turns of the spacecraft. Engines are installed structurally at pre-calculated angles θ ₁ and θ ₂ .

2. Define the directions of the thrust vectors of the engines.

Two engines are installed relative to both axis of the normal to the orbit. The directions of the engine thrust vectors are now set by deviations from the normal to the orbit in the yaw plane at the angles “+ θ” and “-θ”. In the general case, the moduli of these angles may not be equal.

3. Adjust the direction of the thrust vectors of the engines.

When installing engines on a spacecraft at angles “+ θ” and “-θ”, the geometric axis of the engine is taken as the direction of the engine thrust vector. However, due to the error in the installation of the engine and the deviation of the actual direction of the engine thrust vector from its geometric axis, the actual angles of direction of the thrust vectors differ from the calculated ones. Therefore, an adjustment is made in which the actual angles of deviation of the engine thrust vectors from the normal are determined. To carry out the adjustment, the engines are switched on in turn and, after each switching on, trajectory measurements are carried out. By changing the parameters of the orbit and determine the actual angles θ ₁ and θ ₂ respectively for the first and second engines of each of the semi-axes of the normal to the orbit. For example, for a geostationary orbit, the angles θ ₁ and θ ₂ can be determined by the formula

$$\theta =arcsin(\frac{\mathrm{<figure-callout\; id="6"\; label="\mu "\; filenames="00000017.png"\; state="\{\{state\}\}">\mu </figure-callout>}}{6\mathrm{<figure-callout\; id="2"\; label="\pi \; R"\; filenames="00000017.png"\; state="\{\{state\}\}">\pi \; </figure-callout>}{R}^{}{}^2}\cdot \frac{\Delta T}{a\cdot \tau })$$

where µ is the gravitational parameter of the Earth, km ³ / s ² ;

ΔТ - change in the spacecraft rotation period due to engine operation (determined by the results of trajectory measurements), s;

R is the radius of the nominal geostationary orbit of the spacecraft, km;

а - acceleration created by the engine, km / s ² ;

τ is the duration of the engine, s.

4. Calculate the duration of the engines.

The sum of the projections of the thrust impulses of the first and second engines on the normal to the orbit should be equal to the required impulse for the correction of the inclination vector of the orbit, i.e.

On the other hand, the difference between the projections of the pulses of the thrusts of the first and second engines on the transversal should be equal to the required pulse for the correction of the spacecraft rotation period, i.e.

$$F_{\mathrm{one}}{\tau }_{\mathrm{one}}\cdot \sin \left|{\theta }_{\mathrm{one}}\right|-{\mathrm{<figure-callout\; id="2"\; label="F"\; filenames="00000017.png"\; state="\{\{state\}\}">F</figure-callout>}}_2{\mathrm{<figure-callout\; id="2"\; label="\tau "\; filenames="00000017.png"\; state="\{\{state\}\}">\tau </figure-callout>}}_2\cdot \sin \left|{\mathrm{<figure-callout\; id="2"\; label="\theta "\; filenames="00000017.png"\; state="\{\{state\}\}">\theta </figure-callout>}}_2\right|=J_{\tau }.(2)$$

Solving equations (1) and (2) together with respect to τ ₁ and τ ₂ , we obtain

$${\tau }_{\mathrm{one}}=\frac{J_n\cdot \sin \left|{\mathrm{<figure-callout\; id="2"\; label="\theta "\; filenames="00000017.png"\; state="\{\{state\}\}">\theta </figure-callout>}}_2\right|+J_{\tau }\cdot \cos {\theta }_2}{F_{\mathrm{one}}\cdot \sin (|{\theta }_{\mathrm{one}}|+|{\mathrm{<figure-callout\; id="2"\; label="\theta "\; filenames="00000017.png"\; state="\{\{state\}\}">\theta </figure-callout>}}_2|)}$$

$${\tau }_2=\frac{J_n\cdot \sin \left|{\theta }_{\mathrm{one}}\right|-J_{\tau }\cdot \cos {\theta }_{\mathrm{one}}}{{\mathrm{<figure-callout\; id="2"\; label="F"\; filenames="00000017.png"\; state="\{\{state\}\}">F</figure-callout>}}_2\cdot \sin (|{\theta }_{\mathrm{one}}|+|{\mathrm{<figure-callout\; id="2"\; label="\theta "\; filenames="00000017.png"\; state="\{\{state\}\}">\theta </figure-callout>}}_2|)}.$$

5. Carry out a correction with a pair of engines.

Corrections are made by sequentially turning on the first engine for τ ₁ seconds and the second engine for τ ₂ seconds.

The pulses J _n , J _{τ are} determined by the strategy of real spacecraft retention according to well-known formulas, for example, P.E. Elyasberg, “Introduction to AES flight theory”, M., Nauka, 1965

, $$J_{\tau }=|m\cdot \Delta V|=\frac{m\mu \cdot \left|\Delta T\right|}{6\pi \cdot {\mathrm{<figure-callout\; id="2"\; label="R"\; filenames="00000017.png"\; state="\{\{state\}\}">R</figure-callout>}}^2}$$

,

where J _τ is the required impulse to carry out the correction of the spacecraft orbital period, kg · km / s;

m is the mass of the spacecraft, kg;

ΔV is the increment of speed, km / s;

µ is the gravitational parameter of the Earth, km ³ / s ² ;

ΔТ is the required change in the period for applying for correction, s;

R is the radius of the nominal geostationary orbit, km,

as well as G.M. Cherniavsky, V. A. Bartenev, V. A. Malyshev, “Control of the orbit of a stationary satellite”, M., Mechanical Engineering, 1984, p. 129, 138. Moments of engine start-up are determined from the condition that the middle of the interval of the engines corresponded to the point of optimal application of pulses. With continuous correction by two engines in geostationary orbit, you can use the following working formulas (all angular values are expressed in radians):

; $$t_{\mathrm{at}\mathrm{to}l}=t_0+({\alpha }_{\mathrm{at}\mathrm{to}l}-{\alpha }_0^n-S_0-{\lambda }_{\mathrm{from}t})/n$$

;

, $$J_n=m\cdot a\cdot (2/n)\cdot arcsin(n\cdot V_{\mathrm{from}R}\cdot |\overline{\Delta i}|/2a)$$

,

where t ₀ - some initial time, seconds from the reference era;

α _on = arctan [Δi _y · sign ( a _z ) / Δi _x · sign ( a _z )] - right ascension of the middle of the active section;

в координатах:Δi _y , Δi _x - the required components of the change in the inclination vector $$\overline{\Delta i}$$

in coordinates:

i _x = sin (i) cos (Ω);

i _y = sin (i) sin (Ω);

Ω is the longitude of the ascending node of the spacecraft orbit;

a _z - orthogonal acceleration, km / s ² ;

- отклонение от точки «стояния» в момент t₀; $${\alpha }_0^n$$

- deviation from the point of "standing" at time t ₀ ;

S ₀ - average Greenwich stellar time at time t ₀ ;

λ _st - the longitude of the "standing" of the spacecraft - the center of the orbital position;

n is the average spacecraft motion, s ^-1 ;

V _cp is the average orbital velocity, km / s.

The “Method for the simultaneous correction of the retention of the inclination vector of the orbit and the period of revolution of a three-axis stabilized spacecraft” is today the main version of the strategy of holding the spacecraft at a given orbital position. It is successfully used to solve the collocation problem, i.e. coexistence without mutual interference of two or more spacecraft in a narrow area of confinement, but with its help it is impossible to follow the spacecraft by the spacecraft subject at a well-defined predetermined distance.

2. There is a known method of planning corrections, described in the working documentation of the enterprise, for example, in 757A.IE87. The Luch-5A spacecraft, Ballistic Support Instructions [1], as part of the general technological ground-based cyclogram for solving ballistic problems (the cyclogram is given in the appendix, where the GCA is a geostationary spacecraft), which is taken as a prototype. In the prototype method, the following sequence of operations is performed (non-essential details are omitted, the differences between the ground cyclogram and the cyclogram implemented onboard the spacecraft are highlighted in italics).

1. A correction plan is being developed on-board navigation and traffic control system.

Corrections, according to the plan, are carried out by 2 correction engines on a 2-day interval (1 engine - 1 day), or 2 engines on a daily interval.

2. Conduct trajectory measurements.

Trajectory measurements are a regular cycle of measurements of current navigation parameters (ITNP), the number of measurement sessions and the number of intervals between sessions is for the daily interval and the presence of two ground measurement points from 4 to 6. In the presence of autonomous (airborne) radio navigation, trajectory measurements are carried out in continuous mode .

3. Run the program for determining the motion parameters of the center of mass of the spacecraft.

The operation consists in determining the position of the spacecraft in inertial space.

4. Clarify control accelerations by changing orbital parameters.

The refinement does not allow one to determine the control accelerations, more precisely, the range of acceleration values specified by the manufacturer. It guarantees tracking the abnormal operation of correction engines, and in the case of a protracted, and possibly constant situation, when (so far) the correction engine failure is not fixed onboard the spacecraft, still calculate the correction plan. When clarifying, the heuristic method is used: there are initial conditions (NU) of movement according to the previous ITNP, there are current NUs according to items 1-3, there is a previous correction plan that includes up to three conditional numbers of correction engines, the problem of coming to current NUs is solved without large errors in the controlled movement parameters.

5. Execute the program for calculating the parameters of the spacecraft corrections in the vicinity of the orbital position in the interval from the date of calculation to the beginning of the next regular ITN cycle. If there are airborne trajectory measurements, a correction plan is drawn up for a daily time interval.

When calculating the parameters of corrections, they adhere to the chosen strategy for controlling the motion of the center of mass of the spacecraft, which is reflected in the selection or prolongation of the "aiming" points. For simple retention in a given region in terms of longitude of the spacecraft and inclination of the orbit of the spacecraft, the “aiming” points are the center of the orbital position in longitude and the optimum point on the phase plane [i; Ω]. The chosen strategy does not imply direct targeting and data of the point, but launches the long-term planning algorithm with the goal of not leaving the spacecraft from the region in longitude and from the region of a given radius ε relative to the selected point on the plane [i; Ω]. For the needs of collocation (retention of several spacecraft in a given narrow region), the point of "aiming" in addition to the listed ones is also a point on the phase plane [e - eccentricity of the spacecraft orbit; π is the longitude of the pericenter of the spacecraft’s orbit]. Usually a simple minimization of eccentricity is required.

6. Execute the program for generating arrays of command-program information (KPI) containing NU (vector of kinematic motion parameters), a plan of corrections, projections of accelerations from correction engines on the axis of the coordinate system associated with the spacecraft.

7. Sending a generalized form of the CRPD onboard the spacecraft.

Further, items 1-7 are repeated throughout the entire spacecraft operation for its intended purpose. This scheme [1] of ballistic support for the spacecraft flight is universal.

The disadvantage of the prototype [1], as well as the analogue method, is the inability to synchronously accompany one spacecraft with another spacecraft at a certain distance with a given accuracy. It should be noted that synchronous tracking of the apparatus by the apparatus through typical trajectory measurements (ground or airborne), due to typical work of the space complex (ground-based optical observations are possible with respect to the spacecraft), does not bring due result due to the delay in the reactions of the control of the spacecraft unknown strategy for managing a spacecraft. The quality of the spacecraft’s own retention is deteriorating and the energy consumption of the correction system is unreasonably increasing to ensure the spacecraft operates for its intended purpose (justification of the fuel budget is not included in the main objective of the invention, therefore, only mentioning the increase in energy consumption, as this is obvious, can be considered sufficient).

The aim of the invention is the creation of a method of adaptive control of the motion of the center of mass of the spacecraft that meets all the accuracy requirements for tracking the spacecraft-object and reducing the energy consumption of the spacecraft correction system.

This goal is achieved by the fact that in the method of holding the geostationary spacecraft at a given orbital position, including trajectory measurements, determining the position of the spacecraft in the inertial space, calculating the correction parameters depending on the deviation of the adjusted parameter from the nominal value, applying the corrective action to the spacecraft body by turning on the engine, new operations have been introduced, consisting in the fact that on board the spacecraft are installed and attached to the navigation and motion control system Plect transceiver radio equipment and optical angle sensor "Polar Star-Subject-Object"; the antenna and the sensor are installed so that their geometric axes coincide with the transverse direction; carry out continuous measurements of distances to the spacecraft-object over time, while the largest projections of the solar panel on a plane orthogonal to the line of sight "Subject-Object" are greater than the wavelength of the antenna radiation; at each measurement step, the distances to the object are averaged in relation to the selected moments of the steps; at the end of each cycle, including the measurement steps, identify the dynamics of changes in these quantities and calculate the osculating period (major axis of the orbit) of the CA object; determine the increment of the osculating period of circulation of the KA-subject to the corresponding osculating period of circulation of the KA-object; determine per revolution the angle between the planes of the orbits of the object and the subject and the time of intersection of the planes of the orbits of the spacecraft according to the readings of the angle sensor; if the increment of the osculating circulation period exceeds a predetermined value, the tuning adjustment parameters of the circulation period are quickly calculated, and at the estimated time of intersection of the orbit planes of the spacecraft, the correction engine is turned on, giving priority to the inclination correction engine, according to the vectors of relative increments of the orthogonal and transverse speeds, for the correction.

The implementation of the proposed method involves the following sequence of operations.

1. The navigation and motion control system is attached with tracking equipment (AS): a set of radio transmitting and receiving equipment and an optical angle sensor "Polar Star-Subject-Object" (PSO).

The idea of the proposed method of tracking one spacecraft with another is based on the fact that the relative proximity of the spacecraft allows one to obtain mutual distances, even if there are no active signal transponders; A complete set of optical sensors for determining the location of a spacecraft (solar and stellar sensors) is not required - to determine the angle between the planes of the orbits of the spacecraft and the time of intersection of the planes of the orbits of the spacecraft, it is necessary and sufficient to have one optical sensor for the angles of the SAR. The angle sensor is no different, in fact, from, say, the standard angle sensor "Polar Star-Object-Earth", which is part of the on-board orientation and stabilization system. AC macrocycle time - period of revolution at a given flight altitude.

2. The antenna and the sensor are installed so that their geometric axes coincide with the transverse direction.

This direction, unlike the tangential one, does not depend on the parameters of the spacecraft orbit, which means that in principle it does not require rotary devices of the antenna, but in near-circular orbits, with the eccentricity value less than 0.0003 and the maximum distance to the object no more than 74 km, even at the highest - the geostationary orbit, is an ideal reference direction, allowing to receive high-quality radio signals reflected from the design elements of the KA-object, and in the presence of an optical PSO sensor operating on the object, as in a star - and deviation angles KA-object in the orthogonal direction. The fixed direction of aiming of the antenna and the angle sensor minimizes the errors of pointing at the object (in the orientation of the direction of aiming in space, only the errors of stabilization of the KA-subject are not taken into account). The relative proximity of the spacecraft and the large absolute speeds of their motion allow us to assume that the entire spectrum of passive forces in orbit (the Sun, the Moon, the off-centerness of the Earth’s gravitational field) acts on both devices in exactly the same way, and any deviations beyond the accuracy of determining the spatial position of the spacecraft are due to spacecraft maneuvers -object.

3. Conduct trajectory measurements.

The operation is similar to the operation of claim 2 of the prototype.

4. Minimize the eccentricity of the orbit.

The orbit of the KA subject should always be almost circular, for which inclination correction is planned to be carried out by engines that give thrust projections of the direction necessary to minimize the eccentricity, and periodically correct the eccentricity of the orbit.

The operation is similar to paragraph 5.1 prototype.

5. Continuous measurements of the distances to the spacecraft-object are carried out over time, while the largest projections of the solar panel on the plane orthogonal to the line of sight "Subject-Object" are greater than the radiation wavelength of the antenna.

This condition is necessary, it follows from the physics of processing reflected signals.

6. At each measurement step, the distances to the object are averaged in relation to the selected moments of the steps.

7. At the end of each cycle, including the measurement steps, identify the dynamics of the average change in the steps of measuring the distance to the spacecraft-object.

8. The osculating period (major axis of the orbit) of the CA object is calculated.

With a possible eccentricity of the orbit of the KA-object 0.0005 and a distance of 74 km to it, the angle of deviation of the KA-object in the field of view of the OSS sensor will not exceed 16 °, which is quite satisfactory for the purpose of tracking the object using a device that is stationary relative to the structures of the KA.

To determine the entire spectrum of motion parameters of a spacecraft-object from the values of the distances to the spacecraft-object and their first derivatives, it is necessary and sufficient that the measurement cycle is a quarter of a revolution around the earth in time. They solve the typical problem of determining the motion parameters of the center of mass of the spacecraft, considering the orbit of the spacecraft-subject known.

The operation is similar to paragraph 3 of the prototype.

9. The increment of the osculating period of circulation of the KA-subject to the corresponding osculating period of circulation of the KA-object is determined.

According to the speed of the spacecraft divergence, we obtain an increment of the osculating period of circulation. The oscillating period of the circulation of a KA-subject is always known at any moment of time - this is a typical task of ballistic tracking of “one’s” KA.

10. The angle between the planes of the orbits of the object and the subject and the time of intersection of the planes of the orbits of the spacecraft according to the readings of the PSO sensor are determined per revolution.

Obtaining during the macrocycle (revolution) the extreme values of the deviations of the spacecraft from the transverse direction with reference to time, average values of time and the angle of intersection of the planes of the orbits of the spacecraft are obtained.

11. Calculate the parameters of the tuning correction period of circulation.

As noted in paragraph 5 of the prototype, when calculating the correction parameters, the selected strategy for controlling the motion of the center of mass of the spacecraft is followed, which is reflected in the selection or prolongation of the “aiming” points. For simple retention in a given region in terms of longitude of the spacecraft and inclination of the orbit of the spacecraft, the “aiming” points are the center of the orbital position in longitude and the optimum point on the phase plane [i; Ω]. Since there is no need to deal with high-quality collocation (this process is mutual for devices), the need for a “aiming” point [e; π] is dropped. The difference between this operation and the operation of clause 5 of the prototype lies in the unconditional priority of the tracking procedure for the KA-object over the procedure (strategy) of its own retention of the KA-subject at a given orbital position. The tracking KA should recognize and follow the strategy of controlling the KA-object, which means, if necessary, to approximate (change) the “aiming” points and carry out corrections more often and more quickly than if it had no task of tracking and tracking.

Correction parameters (time, duration and conditional number of the correction engine) depend on the parameters being corrected [i is the angle of intersection of the orbital planes of the spacecraft (inclination analogue), ΔТ is the required change in the osculating period] and are calculated by known methods (for reference, you can return to the analogue method ) To correct the period of circulation, inclination correction engines are more often used. Longitude correction engines are used when the orbit planes of both spacecraft practically coincide.

If the increment of the osculating period of circulation exceeds the two-day rate of increment due to the corrective action, then go to step 12, otherwise - plan the correction of inclination with equal impulses by the correction engines of the inclination of the opposite direction of the thrust projection onto the transversal. Fulfillment of this condition allows one to respond to the evolution of a spacecraft with an acceptable delay, which saves fuel consumption on board. It is possible that at the next round, the periods of circulation of the vehicles “by themselves” will come closer, since the strategy for controlling the spacecraft is unknown.

12. Apply corrective action.

The operation is similar to claim 1 of the prototype.

Justification of the idea of passive-reflected radio measurements

For satellite radar purposes, centimeter wavelengths are used: S-band (4-2 GHz), wavelength (λ) 7.5-15 cm and C-band (8-4 GHz), wavelength 3.8-7.5 cm. The choice of range affects the resolution and type of reflection (specular, for L >>λ; diffuse, for L≤λ, L> λ). L - target size (width-height). Choose an S-band. With a minimum active reflective area of solar panels of 2.7 m ² we will have a mirror image. Under the S-band - the range of satellite communications is easier to "assemble" equipment, i.e. some elements of the system will not have to be developed again. Another - the frequency is tied to the size of the antenna, with the radiation pattern: the smaller the wavelength, the smaller the size of the antenna. For the geostationary orbit, attenuation is not taken into account, there is no atmosphere. The transmitter power of about 32 watts will be super sufficient. From a distance of 74 km a signal with a power of (-165) dB · W, no less, will return. The figures for output and input power correspond to domestic autonomous radio navigation. In reality, a complete reflection of the radio signal from the target will not work, we can lose about 20%, which will not be critical. Antenna type - PAR (phased array) or with synthesized aperture (SA). We choose an antenna with a SA, this will reduce the overall dimensions of the airborne radar system (radar). The approximate size of the antenna array for the S-band is 25X25 cm, more precisely, the size can be specified only after calculation, taking into account the frequency of the emitted signal and other parameters).

The main difference between the headlamp and the SA antenna in principle. The principle of operation of the antenna with CA is based on the use of moving the onboard antenna of the radar system for the sequential formation of the antenna array. The on-board antenna, as a rule, is small in size and has a fairly wide radiation pattern. A wide daylight rate will improve reception quality.

Figure 1 table shows the characteristics of domestic radar (SCHAR-slotted antenna array). We are interested in the table only the mass characteristics of the radar. From figure 1 it follows that the transceiver equipment is guaranteed not to exceed 100 kg, which is quite acceptable for the desired target task of the KA-object.

The proposed method of adaptive control of the motion of the center of mass of the spacecraft fully meets the requirement of tracking the target and has high accuracy of the task. In addition, obviously, the proposed method with respect to the prototype is more rational in terms of the cost of the working fluid of the correction system.

The invention is proposed to be introduced into the working documentation of the enterprise in 2011, to refine the above retention method in 2012 and to fully use it on the corresponding geostationary spacecraft from 2013.

Claims (1)

A method of adaptive control of the motion of the center of mass of a spacecraft (SC), including trajectory measurements, minimizing the eccentricity of the orbit, determining the position of the KA subject in inertial space, calculating correction parameters depending on the deviation of the adjusted parameter from the nominal value, applying a corrective action to the spacecraft body by turning on the engine, characterized in that on board the spacecraft they install and attach a set of transceiver to the navigation and motion control system radio equipment and the optical angle sensor "Polar Star - KA-subject - KA-object", and the antenna and optical sensor are set so that their geometric axes coincide with the transverse direction, conduct continuous measurements of the distance to the KA-object over time, until the greatest the projection of its solar panel panel onto a plane orthogonal to the line of sight “KA-subject - KA-object” is greater than the antenna radiation wavelength, at each measurement step, the distances to the KA-object are averaged in relation to the selected moments steps, at the end of each cycle, including measurement steps, identify the dynamics of changes in these values and calculate the osculating period of circulation (the major axis of the orbit) of the CA object, determine the increment of the osculating period of circulation of the CA subject with respect to the corresponding osculating period of circulation of the CA object, determine the angle between the orbital planes of the KA-object and KA-subject orbits and the time of intersection of the orbital planes of these KA according to the readings of the optical angle sensor, in this case, if the increment exceeds of the period of predetermined value, the parameters of the adjustment correction of the period of revolution are quickly calculated and, at the estimated time of intersection of the orbital planes of the indicated spacecraft, the correction engine is turned on, giving priority to the inclination correction engine, according to the vectors of the relative increments of the object’s orthogonal and transversal velocities for the correction.

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