# Self-contained onboard control system of "gasad-2a" spacecraft

FIELD: astro-navigation, control of attitude and orbital position of spacecraft.

SUBSTANCE: proposed system includes control computer, star sensor, Earth sensor, storage and timing device, processors for control of attitude, processing angular and orbital data, inertial flywheels and spacecraft orbit correction engine plant. Used as astro-orienters are reference and navigational stars from celestial pole zone. Direction of spacecraft to reference star and direction of central axis of Earth sensor to Earth center are matched with plane formed by central axes of sensors with the aid of onboard units. Shift of direction to reference star relative to central axis of Earth sensor is considered to be latitude change in orbital position of spacecraft. Turn of navigational star around reference star read off sensor base is considered to be inertial longitude change. Point of reading of longitude is point of spring equinox point whose hour angle is synchronized with the board time. This time is zeroed upon completion of Earth revolution. Stochastic measurements by means of static processing are smoothed-out and are converted into geographic latitude and longitude parameters. Smoothed inertial parameters are compared with parameters of preset turn of spacecraft orbit found in storage. Revealed deviations of orbit are eliminated by means of correction engine plant.

EFFECT: enhanced accuracy of determination of spacecraft attitude and orbital position; automatic elimination of deviation from orbit.

44 dwg

The invention relates to navigation and to control the angular orientation and orbital location of the SPACECRAFT (SC) and is intended for use on Autonomous functioning of the AC.

As navigation is widely known and traditionally used radio ground measurement points (NIP). NPCs widely separated in longitude-latitude thus, to cover as much as possible the space. This placement of NPCs due to the fact that the sessions of range measurements to the satellites, azimuth and elevation are implemented only in those areas where conditions range between NIP and KA. And yet, when the orbital period of the SPACECRAFT, for example about one and a half hours from 15-16 per diem of coils, at least 6 are outside the zones range from the territory of the former USSR. Moreover, there is a limitation due to the fact that every NPC takes some time to prepare for further work. On the navigation measurements can be imposed constraints related with the impossibility of simultaneous operation of two different NPCs.

Thus there is a mismatch between the need for global measurements and existing opportunities. Moreover, the accuracy requirements for measuring parameters of the motion of the SPACECRAFT combined with the cost of equipment is URS, Nipah and maintenance cost involved number of measurement points.

Other well-known navigation tool is the cosmic system, which is composed of the navigation satellites and ground-based measurement system. The opportunity to use the measurements on this system depends on the relative position of the navigation satellites and SPACECRAFT. There is uncertainty in the implementation of navigational measurements in the specified or all areas of space, as their ability connected with standby time of the onset of at least one navigation satellite and the time interval for different KA different. In other words, in various subspaces of the scope of the satellite navigation system the probability of navigation definitions. To speak in this case about the globality as the maximum value of the ratio of the accepted system of subspaces to all intended for the navigation of the near-earth space can only conditionally. In the same restrictive sense to speak of the continuity of navigation definitions, as continuity is due to the simultaneous geometric visibility of multiple satellites that are not always consistently available. Also you should not speak in the absolute sense and efficiency.

Meanwhile, as in the case of NPCs, both in aerospace is the first navigation system continuity measurement is the means which provides the necessary accuracy figures navigational attributes.

At the same time accurate, consistent data on the location of the SPACECRAFT in space are the source material, which is used to calculate a consistent angular orientation of the SPACECRAFT, for example, relative to the Earth. Note that among the dimensions considered by means of parameters, there is no definition of the spatial position of the direction of the KA - center of the Earth", which is essentially the angular control of the SPACECRAFT at any target function. In General, the purpose of navigation definitions should be this option. However, the adopted approach, it is not observable directly. Therefore, its definition will provide additional navigational tools, in particular celestial cues. Essentially this means that in addition to artificial navigation system uses a second navigation system - the system of natural reference astroarena. Different systems with different reference systems split a single motion of the SPACECRAFT in two movements - spatial and angular. It must be recognized that such oposredstvovaniya technique designed not initially, and at the end of the navigation process to characterize the motion of the SPACECRAFT as a corner, creates not only additional is sustained fashion the scope of work but the accuracy of the orientation error.

Thus, the task of navigation KA traditionally solved regardless of the task angular SPACECRAFT control and practice algorithms for navigation and control the angular orientation of the SPACECRAFT initially do not depend on each other.

Ultimately, consideration of the level of conventional admittedly, there is the problem of providing real-time and on a global scale the most accurate, continuous and direct measurements of such navigation variables that could directly correspond to control variables and would contribute to the synthesis of control algorithms and navigation in one Autonomous onboard control system of the SPACECRAFT.

A characteristic feature of this problem is that to solve it we need to go beyond the known methods. It should be addressed with those points of view that allow you to consider the navigation and management process as something unified and whole.

The famous "Autonomous on-Board control system of the spacecraft "HASAD" [application 93007754 (patent 2033949), 1993], which does not claim to solve this problem, however, indicates some of the ways of obtaining and processing of astronomical measurements. This system contains: sensor Earth sensor, the Polar star and the navigation star (star sensor), calculate the fir, the device memory, a temporary device and Executive bodies. When you do this:

the outputs of the sensor of the Earth, star sensor, the memory device associated with the inputs of the calculator on the relevant signals; temporary entry transmitter is connected to the output of the temporary device, and the input of the Executive bodies is connected with the output of the transmitter;

device memory contains parameters reference inertial longitude, and these settings are the right ascension of the pole star and the angle relative to the common plane of the sensor sensitivity of the Earth and the star sensor is equal to the angle between the plane containing the center of the Earth - the North star" and "the center of the Earth pole of the world", and the plane containing the center of the Earth - the North star" and "the center of the Earth - navigation star";

the calculator is arranged to determine the latitude of the SPACECRAFT at the corner of "the center of the Earth - KA - polar star";

- the transmitter is configured to generate the control signal for the angular orientation of the SPACECRAFT in the yaw channel based on the corresponding measured angular misalignments for combining the General plane of the sensor sensitivity of the Earth and the star sensor containing the longitudinal axis of the SPACECRAFT, with the plane containing the center of the Earth, the SPACECRAFT and the North star;

- the transmitter is done is n determine inertial longitude of the location of the SPACECRAFT on the azimuthal angle of the navigation star around the direction of the KA - Polar star", which is counted in the appropriate field of view from the base angular position of the plane containing the direction of the KA - navigation star" and "KA - polar star".

Consider whether this system in all of its characteristics to serve as a prototype of the invention. But first let us note the following. Strange, of course, to say that of course. Thus, the range of land measuring point or select navigation satellite as an artificial point of reference, are not equivalent. This choice is usually justified by the uniqueness of a specific case. For example, the specic tasks used orbit etc.

Of similar importance is the choice of one or the other polar star as a natural point of reference.

To ensure the global space navigation definitions of continuity and precision of astronomical measurements of the reference star must be:

1) observed from any point used orbit;

2) relatively immobile (change of right ascension of the star should be as minimal as possible for the lifespan of the).

Note here that the polar zone of stars due to the precession of the earth axis is generally the area of special direct ascents.

In this case, it becomes obvious that the por is the provision of global space and continuity of the navigation measurements of any particular used orbit impose limitations on the freedom of choice of one or the other stars from the zone to the North pole
the world or from the zone of the South pole of the world. So, the motion of the SPACECRAFT in perigynous zone orbits of type ' Zip polar star is obscured by the Earth. At the same time supporting the star of the South pole observed from all points of the same orbit. In addition, the change of right ascension of the pole star for 10 years is 4.3 minutes of that degree in terms 1°4,5' (4,3_{min}×15'/_{min}). To speak in this case of high accuracy performance is not necessary. It should be noted that the polar star by the "instability" of their relative location in space surpasses all the stars as the North and South polar zones.

However, the coordinate changes of the Polar star can be periodically updated in computational algorithms through correction of their current actual coordinates. Given that the change in coordinate time known to implement such a correction is possible due to the significant complexity of the onboard software. Of course, if there are other options to complicate software irrational.

We must recognize that globalization, continuity and accuracy of navigational measurements using the North star in this case is not provided. The system is initially designed to work in conjunction with TPA the investment funds navigation and is designed to solve a limited narrow task. It also follows from the fact that it was not intended positioning of the SPACECRAFT relative to the geographic latitude-longitude grid of the Earth. Moreover, the system is not designed to ensure compliance with the parameters of the real orbit parameters specified and does not contain a means of processing the measured angular data to obtain high-precision values.

Essentially all of the characteristics of the system based on the use of the Polar star as a fundamental navigational landmark. In addition, these signs also contain the following uncertainties and disadvantages:

1) the Stated task is to determine the latitude of the location of the AC at the corner of "the center of the Earth - KA - polar star" is not solved due to the fact that is not taken into account the amendment, which is due to the polar distance Polar star and the value of which depends on the inertial longitude.

2) Not defined zero the instrument base reference inertial longitude corresponding to the external zero-based reference with respect to which, and the time.

3) does Not determine the point of zero flight time.

4) In the General plane of the sensitivity of this system contains the longitudinal axis CA that is not possible to reorient the axis to the specified target point of the Earth without thereby interrupting the session navigat is I.

Thus, the characteristics of this system are not the closest to the features of the invention, and therefore this system cannot be considered the prototype of the invention.

Analogue of the invention may serve as a typical system the attitude control and stabilization of a SPACECRAFT, comprising mainly traditional computer, sensor, Earth, star sensor, inertial flywheels.

This system is designed to solve the problems of the angular orientation and stabilization.

These modes are identified and tracked the stars that fall within the field of view of the sensor, and stabilization that can also be attributed to the General with the invention of the signs.

The basis of the invention is the technical task is to create the onboard control system of the spacecraft, which would allow high precision to provide Autonomous control of the angular orientation, the Autonomous determination of the angle and location of the device relative to the latitudinal-longitudinal grid of the Earth, an Autonomous definition of misalignments of the real orbit from the set and stand-alone fixing them.

The technical result is:

the exception is the use of ground-based navigation infrastructure and SPACECRAFT control;

- ensuring global space navigation measurements made onboard means KA;

- the implementation of the synthesis of control algorithms and navigation;- implementation of direct and continuous measurements of the orbital motion parameters in accordance with each other angles of orientation and navigation;

- definition real-time angle and location of the device as in inertial space, and relative to the latitudinal-longitudinal grid of the Earth;

- to ensure a high accuracy of angular parameters and location of the AC current and projected interval;

- to identify evolutionary changes of the orbital parameters on a given orbit and implementation of their automatic removal.

The technical result, which is reduced to a few results on the functional level of generalization can be represented as autonomy, independence, CA from complexes of terrestrial and space-based, i.e. from an external artificial navigation and control.

In the structural composition of the proposed control system includes, in addition to the host computer, multiple processors, designed in parallel to implement each function, as well as the Executive bodies of the inertial flywheels and engine pulse correction.

All devices of the system, including the sensor of the Earth, the star sensor, the storage device and the temporary installation within the appropriate way and at the appropriate signals are connected by communication lines.

The technical result is achieved due to:

I. AUTONOMOUS ANGULAR ORIENTATION:

(involved: the CPU controls the position sensor of the Earth, star sensor, inertial flywheels and corresponding algorithms)

1). System astroarena is formed from two stars polar zone: reference and navigation and (center) of the Earth.

2). The plane containing the direction of the Central axis of the sensor Ground to the center of the Earth and the direction of the SPACECRAFT on the reference star, combined continuously in the process of orbital motion with a rigid plane formed by the Central axes of the sensors and skoordinirovanno in the coordinate system of the star sensor.

II. The AUTONOMOUS DEFINITION of the INERTIAL LONGITUDE:

(involved: the star sensor, the processor processing the angular data and the corresponding algorithms)

3). The rotation of the navigation star (due to the orbital motion of the SPACECRAFT around the reference stars in the field of view of the star sensor is fixed in the coordinates of the star sensor as inertial longitude changes of the orbital position of the SPACECRAFT.

III. COMMIT ONBOARD REFERENCE POINT INERTIAL LONGITUDE:

(involved: design calculations, the processor processing the angular data and the corresponding algorithms)

4). When design calculations basic inertial longitude of the SPACECRAFT is determined by direct ascent to the CHRA is Noah stars in degree terms at the time of orbital motion, when the base angle between the rigid plane sensors and plane directions with the AC on both stars in the field of view of the star sensor corresponds to the culmination of the reference stars.

5). Side of the zero reference point of the inertial longitude coordinated in the coordinates of the star sensor as the position of the plane directions with the AC on both stars at the time when the project inertial longitude is set to zero, and the range of changes of the basic angle reaches the value of the base inertial longitude.

IV. FIXING an EXTERNAL POINT of REFERENCE INERTIAL LONGITUDE AND TIME:

(involved a temporary device)

6). As an external point of reference inertial longitude becomes (see item 4, 5) the vernal equinox point.

7). Board time KA is synchronized with the hour angle of the vernal equinox relative to the Greenwich Meridian and is reset after the time of a complete rotation of the Earth.

V. AUTONOMOUS geo LONGITUDE:

(involved: the star sensor, the processor angular data, a temporary device and corresponding algorithms)

8). Geographic longitude of the SPACECRAFT is determined by the relevant mathematical dependences, taking into account the Earth's rotation speed, the current time and the measured inertial longitude.

VI. AUTONOMOUS DETERMINATION of LATITUDE IS MESTOPOLOJENIE:

(involved: design calculations, sensor Ground, star sensor, the processor angular data storage device and corresponding algorithms)

9). In design calculations, the angle between the direction from the SPACECRAFT to the reference star and the Earth's axis is determined by the polar distance of the reference star at the time when the underlying inertial longitude KA corresponds to the right ascension of the reference stars in the degree calculation. The value of the specified angle at different values of the inertial longitude are determined by known mathematical relationships, and are represented in the permanent memory (ROM).

10) the latitude of the SPACECRAFT is determined by converting the measured angle between the Central axis of the sensor Ground, aimed at the center of the Earth, and the direction from the SPACECRAFT to the reference star (coordinated by the star sensor) and taking into account the angle between the direction from the SPACECRAFT to the reference star and the Earth's axis is represented in memory by a value corresponding to the value of the inertial longitude, which is determined by the latitude.

VII. AUTONOMOUS determine the ACTUAL ORBIT AND the EXACT ANGULAR ORIENTATION:

(involved: the processor angular data, the CPU controls the position storage device, suspension device, inertial flywheels and the corresponding Sal the rhythms)

11). Orbit in inertial space is defined by a set of points generated at each time quantum inertial longitudinal and inertial-pulse measurements.

12). The level of precision of the star sensor (in the light of the incomparability of the error values of both sensors) is treated as the base level and therefore the accuracy of angular measurements related to the error sensor of the Earth.

13). Measurement error is interpreted in the form of the imaginary error nature of orbital motion.

14). The parameters of the real and essentially "smooth" motion of the SPACECRAFT is determined by statistical processing of stochastic inertial longitudinal and inertial-pulse measurements using appropriate software module (neural network) and as represented by entries in the random access memory device (RAM).

15). With this processing eliminates errors in the measurements and multi-clock-step-ahead predicted values of the inertial longitude and inertial latitudinal angles for the subsequent implementation of the adjusted angular orientation of the SPACECRAFT.

VIII. AUTONOMOUS DETERMINE VARIANCES OF THE PARAMETERS OF THE REAL ORBIT PARAMETERS OF A GIVEN ORBIT.

(used: calculation, memory (ROM and RAM), processor orbital data and comply with what their algorithms)

16). On Board the SPACECRAFT via entries in the persistent storage device (ROM) is consistent with the requirements of the objective function KA) given orbital revolution of their traditional settings and in the form of inertial longitudinal and inertial latitudinal angles to each time quantum.

17). On the relevant mathematical dependences and directly reportedly determined period, the semimajor axis, eccentricity, argument of perigee, inclination, and right ascension of the ascending node of the real, is presented in RAM orbital revolution.

18). Comparing the parameters of the real orbit and presents in ROM settings given orbit are determined by the desired deviation.

19). For known dependencies and identified deviations are defined: a sequence of corrective action, the type of corrective pulse, the point of its application, the value of the speed pulse, the value of the energy consumption and operating time of the propulsion system (PS).

IX. AUTONOMOUS IMPLEMENTATION OF THE CORRECTION.

(involved: processor management position, a temporary device, inertial flywheels, motors correction and corresponding algorithms)

20). Point and the direction of application of the pulse defined and implemented the angular spreads of the SPACECRAFT relative to modeled on the ORT coordinate system, providing the necessary direction vector of the corrective pulse, and pulse width - work time control.

X. AUTOMATIC control of TECHNOLOGICAL PROCESS.

(involved: the host PC and peripherals)

21) In terms of multitasking host computer cyclically sampled and assessed the current state of the peripheral device, the data is updated internal database and taking the necessary commands impacts for proper process.

It should be noted that the combination of the above signs refers to a group of inventions so linked together by a single idea - a single synthesized as part of the navigation and control of Autonomous on-Board system of the SPACECRAFT.

Figure 1 shows the hierarchical structural diagram of the Autonomous onboard control system; figure 2 - schematic representation of astronomical spatial and temporal correlations as the basis of navigation definitions; figure 3 - spacecraft in earth orbit with the scheme of the measured angles, stable in respect to the center of the Earth and directions to the reference star; figure 4 - table of the stars around the North celestial pole; figure 5 is a table of the stars around the South celestial pole; figure 6 - space is a mini device with an extended field of view of the star sensor and the ability to Orient the longitudinal axis to any point on the Earth; figure 7 - diagram of shading the Earth star benchmarks on different orbits; on Fig spherical triangle, is intended to determine the angle between the plane containing the directions on the anchor and navigational stars, and the plane containing the directions on the reference star and the pole of the world; figure 9 - diagram definition on-Board reference inertial longitude; figure 10 is a scheme for determining the angle between the direction from the SPACECRAFT to the reference star and the Earth's axis; figure 11 - orbit generated by points, which are coordinated by the inertial longitude and inertial latitudinal angles; Fig diagram illustrating the relative position values of longitude and inertial longitude; Fig - scheme of the geocentric vertical with 4 scanning sensors; Fig diagram of the geometry of measurements geocentric vertical; Fig - monoblock of astrogation KA, its "hard" plane and the coordinate axes of the sensors in the monoblock; Fig diagram of differences between systematic and random errors; Fig and 18 presents diagrams illustrating the angular error of the spatial position of the SPACECRAFT in the plane of the longitudinal channel and in the latitudinal plane of the channel, respectively; Fig diagram illustrating the real and imaginary orbit in the moments of the measurement pulse angle; Fig diagram, illustrious is I'm real and imaginary position of the SPACECRAFT in the moments of the measurement of the latitudinal angle; on Fig - block diagram of the statistical processing of the measured data; Fig - time dependent stochastic latitudinal and longitudinal data for training the neural network is a software module processing; Fig - table of results of processing the measurement data on the orbital area from 500 to 559 sec; Fig - graphical dependency on time actual measured and predicted by the neural network inertial longitude on the orbital area from 500 to 559 sec; Fig - graphical dependency on time actual measured and predicted by the neural network latitudinal equivalent to the orbital area from 500 to 559 sec; on Fig - table of results of processing the measurement data on the orbital area with 2999 on 3058 s; Fig - graphical dependency on time actual measured and predicted by the neural network inertial longitude on the orbital area with 2999 on 3058 s; Fig - graphical dependency on time actual measured and predicted by the neural network latitudinal equivalent to the orbital area with 2999 on 3058 s; Fig - graphic time dependence of the real and predicted by the neural network inertial longitude on the full interval loop 0 through 5402 s; Fig - graphic time dependence of the real and prognosis the bath neural network latitudinal equivalent to the interval of a full round with 0 5402 s; on Fig - comparative table of parameters processed by different methods - linear trend and a neural network on the orbital area with 3049 on 3058 s; Fig - graph comparisons longitudinal parameters, the predicted linear trend and a neural network on the orbital area with 3049 on 3058 s; Fig - graph comparisons pulse parameters, the predicted linear trend and a neural network on the orbital area with 3049 on 3058 s; Fig - defined orbit and the actual orbit and the elements that characterize each of the orbits; not Fig - table three orbits explaining the definition of apogee and perigee; Fig - time graph of the speeds of the two orbits zones and points of apogee and perigee; Fig - orbit and the corresponding axes of change of inclination, argument of perigee and right ascension of the ascending node under the action of lateral corrective force; Fig - orbit and lateral impulse, which in General changes all three elements that define the orientation of the orbit; Fig diagram explaining a function of the angle of spread to the orbit plane of the magnitude of lateral momentum and the angle describing the direction of the pulse; Fig diagram of a direct reversal of the orbital plane to change its inclination; Fig diagram explaining the change vnutripolostnyh elements of the orbit under the action of tangents is a pressing force; on Fig diagram illustrating the correction of the position of perigee (argument of perigee) under the action of tangential momentum; Fig - orbit and the points at which changes in semimajor axis, eccentricity and argument of perigee under tangential forces have extreme values; Fig - table, where the equations increments orbital elements under the action of tangential or lateral forces.

Shown in the figures denote: 1 - GND, 2 - the celestial sphere, 3 - direction on the celestial pole, 4 - direction on the reference star, 5 - direction of the navigation star, 6 - equator, 7 - celestial equator, 8 - Meridian of Greenwich, 9 - celestial Meridian absolute reference point 0,γ10 to the plane of the Greenwich Meridian, 11 - Meridian reference stars 12 - the first plane 13 to the second plane 14 - spacecraft, 15 star gauge, 16 - gauge Ground 17 - plane sensitivity, 18 - geocentric vertical, 19 - orbit, 20 - solution of the field of view of the star sensor, 21 - monoblock astrobingo, 22 - two-stage actuator monoblock astrobingo, 23 - longitudinal axis KA, 24 - motor installation, 25 - the center of the Earth, 26 - point touch horizon line of sight of the sensor of the Earth, 27 - horizon, 28 - instantaneous field of view 29 - scan area of the instantaneous field of view, 30 - side reference system of measurements on which tcheka Land,
31 - layer CO_{2}a 32 - shadow area of the orbit, 33 - first tangent at the vertex of angle λ, 34 - second tangent at the vertex of angle λ, 35 - scale inertial longitude.

The presented figures also indicated: γ - point of the vernal equinox, "0", the instrument base reference 0 - absolute direction of the external starting point at t=0, formed by the intersection of the plane of the celestial equator plane of the Greenwich Meridian, α - the right ascension of the reference stars, μ - the polar distance of the reference stars, t is the hour angle of the vernal equinox relative to the Greenwich Meridian, t_{m}- timeline full Earth rotation, scale Board time, σ - angle triangle, the apex of which is the pole of the world, ρ - angle triangle, the apex of which is the navigation star, β - the polar distance of the navigation star ε - the angular distance between stars, x, y, z - coordinate axis sensor of the Earth, x_{0}, y_{0}, z_{0}- coordinate axis of the star sensor, ω - the angle between the Central axis of the sensor Ground and the direction of the KA - reference star, ω_{0}the angular relationship between the coordinate systems of the star sensor and the sensor of the Earth, λ - the angle between the first and second planes, ψ - angle, locking instrument base frame relative to the hard pleskot the sensors,
ϕ - the angle between the first plane and the absolute reference direction, the angle between the rigid plane sensors and absolute reference direction, the angle between the rigid plane of the sensor and the plane of the Greenwich Meridian, which is at t=0 coincides with the direct ascent of the reference stars, And the apogee of the orbit, P is the perigee, the ascending node, N is the descending node, Ω - the right ascension of the ascending node, i is the orbital inclination, ∈ coincides with the direct ascent of the reference stars, And the apogee of the orbit, P - perigee, In the ascending node, N is the descending node, Ω - the right ascension of the ascending node, i is the orbital inclination, ∈ argument of perigee, u - argument KA ϑ is the true anomaly of the SPACECRAFT, 2α - the distance between the perigee and apogee of the orbit, e is the eccentricity, F is the focus of the orbit, D and D_{1}- the point of intersection of the orbit with minor radius, ι - inertial longitude, ∂ - geographic longitude, η latitude ζ longitude error, ξ - latitudinal error, Z is measured by the sensor, the angle between the side frame of reference and direction to the visible horizon of the Earth, ∑ - measured by the angle sensor to the directions on the horizon in the opposite direction, ΔV is the pulse velocity, V_{η}- transversal speed, Δχ - turn angle of the orbital plane, j is the angle describing the direction of momentum ΔV.

Vnutrennego funds CA, which is schematically depicted in figure 1, is structurally and functionally divided into three levels:

at the level of local control by the star sensor and sensor Ground measured parameters of the external part of the orbital motion, which takes the form of angular change, and by the inertia of the flywheel and engine correction are of influence on this process in order to change its settings;

- at the level of the control angle and the location of the SPACECRAFT by the processor, control the position of balance laws control the angular position of the SPACECRAFT, its orbital location of the matching results (processor orbital data) parameters of the real orbit with parameters given (storage device) and the results of statistical processing of the measured data (processor angular data, which is also the transformation of the inertial parameters in geographic);

at the level of strategic management in real time and multitasking series-parallel mode control computer cycles (using grid frequency temporary device) queries and evaluates the current state of the peripheral device, updates the data inner the database and takes the necessary commands impacts for proper flow of the overall process.

Consider the astronomical basis of the navigation and control, and to do this, refer to figure 2, which presents a schematic representation of astronomical spatial and temporal correlations.

Pay attention to the following circumstances:

1) the point of the vernal equinox and the pole world in the sky nothing is checked, this abstract point;

2) corners α, μt affected in one way or another to change.

Reason for change α is that the point of the vernal equinox (the point of intersection of the celestial equator with the Ecliptic) is shifted relative to the fixed stars due to precession.

Reason for change μ is that the pole of the world (the axis of rotation of the Earth) moves due precessions around the pole of the Ecliptic, continuously changing Equatorial coordinates of the stars.

Note that the effect of the precession effect is much stronger for stars close to the celestial pole.

Reason for change t is that the reference point of time of the corners of the adopted point (the point of intersection of the celestial equator with the celestial Meridian), not involved in the daily rotation of the celestial sphere.

The situation, seized certain angular correlations in the time of the spring equinox, in the next moment "thawed".

Suppose, however, to link the situation (systems the guidelines) new angular relationship so to these abstract points (point γ and the pole of the world) would have indicated already in the planes, i.e. clarify the situation.

In fact, if we are on any system look in isolation, it does not mean that it exists in isolation and the point here is to determine the communication system. Pay attention to the fact that the direction 3 direction 4 are contained in the plane 12, in which the reference star is the 0^{h}Greenwich universal time at its upper end, and areas 4 and 5, see the second plane 13, which is constant in inertial space. These planes intersect at an angle λ.

We emphasize the following fact - the second plane 13 by angle λ records in the space of the first plane 12, which contains the direction of the center of the Earth - the North pole of the world". When the pole is fixed in space, not only in this plane, and the polar distance μ.

In turn, recorded thus in inertial space, the first plane 12 through an angle from her and equal in degree terms of the right ascension of the reference star 4, is fixed in space, the position of the Meridian of the vernal equinox γthat, too, nothing is marked in the sky.

Further defined the e external relations frame of reference associated with the manifestation of the angular relationship of the first plane 12 to the plane of the Greenwich Meridian.

Greenwich sidereal time is the hour angle of the vernal equinox relative to the Greenwich Meridian, that is, t in the degree measurements. The position of the vernal equinox γ with respect to the first plane 12 defined above direct ascent of the reference stars, i.e. the angle α. It is obvious that the first plane 12 and the plane of the Greenwich Meridian 10 intersect at an angle equal to α-t. At t=0 the plane of the Greenwich Meridian coincides with the direction to the vernal equinox γi.e. at this point the plane 12 and the plane of the Greenwich Meridian 10 intersect at an angle equal to the right ascension α reference stars.

Recall that in the astronomical yearbooks defined Greenwich hour angles of the vernal equinox γand thereafter in intervals.

This spatial situation, including such geometric elements, as the directions of 0, 3, 4, 5 and planes 12, 13 and 10, characterized by lower level of abstraction and are defined angular relationships.

Of particular importance here is the point 0, which coincides in the present moment with a point γ. Special significance of this point is due to the fact that it is intended to be a zero base timing and longitude.

At t=0 the point 0 is zero base reference time is military action hour angle.

As for the zero reference longitude, the case is a bit tricky.

The question in the following. Point γ moves along the Ecliptic to the West with the speed 50,3 in the year, so it should be correlated to the time to the equinox which are the coordinates of the stars.

Of course, the coordinates of the stars in our case should be referred to a certain age.

It should be noted that only the Equatorial stars annual increment their direct ascent is equal to the annual shift of the vernal equinox γ. Annual increment direct ascents polar stars are not subject to such compliance.

Polar stars zone is an area of particular direct ascents, where the annual increment direct ascents can be positive, negative, and even very close to zero.

Since sidereal day is the time interval between two consecutive same culmination point γ, it is obvious that the displacement of a point γ 0",133=0^{s},0084 per day leads to a mismatch between the duration of the sidereal day and the period of complete rotation of the Earth and, thus, starry nights become shorter.

If the coordinates of stars assigned to a certain age (frozen) and annual (daily) increment direct vos is ojdani meet annual (daily) offset point γ then the problem is not here and starry night will be equal to the period of complete rotation of the Earth.

In this regard, preferred for use as benchmarks are the polar star, the annual increment direct ascents which is equal (or close) the annual displacement of a point γ.

At t=0 point 0 after each full rotation of the Earth is fixed in space the plane of the Greenwich Meridian and, thus, the following situation will repeat the previous one.

In this case, the values of the geographical longitudes will coincide with the values of direct ascents stars (e.g., Equatorial).

Thus, the point 0, which is formed by the intersection of the plane of the celestial equator plane of the Greenwich Meridian, becomes the absolute zero point of reference not only time, but also the longitude.

In General precession alters not only the direct ascent of the stars, but their polar distance.

For engineering applications, it is preferable to such a spatial frame of reference in which the coordinate changes are minor.

Figure 3 schematically shows the spacecraft in earth orbit with the scheme of the measured angles, stable in respect to the center of the Earth and directions to the reference star.

Arose the et question: how identified astrometric framework could be useful for our purpose?

It is known that the usefulness means the ratio of the subject under consideration as part of, a certain relationship of objects considered as a whole.

Whole, in our case, it is the relationship of on-Board devices KA with the spatial structure of the environment.

A common characteristic of each astrometric basis, i.e. each particular subject, be it a frame of reference, absolute reference, inertial longitude or the relationship of longitude is the angular measure, which is calculated depending on the need and temporary measure.

Therefore, use can be detected angular relationship.

As material points of reference using the reference star 4, the navigation star 5 and the center of the Earth.

Note that the anchor and navigational stars appear in the star sensor with appropriate coordination relative to the instrument axis.

The Earth centre is implemented by the sensor of the Earth as the estimated location of a point relative to the on-Board reference. The direction of this point is treated as a geocentric vertical.

The plane that contains the Central axis of the star sensor and the sensor Ground, and displaying the reference star and the center of the Earth, we will call the plane of sensitivity 17.

To use the coordinate autosimulations basis, "freeze" the location of the SPACECRAFT in the orbital point belonging to the first plane 12, so that the plane of sensitivity 17 is aligned with the first plane 12, in which the reference star is at its upper culmination, and geocentric vertical coincided with the direction of the KA - center of the Earth".

Obviously, in this case, the plane of sensitivity of 17 will be the first plane 12, and the plane formed by directions "KA - reference star" and "KA - navigation star" - the second plane 13 and these planes intersect at an angle λ.

At this point, the SPACECRAFT is similar to the reference star 4 culminates in the first plane 12 and its right ascension is equal to the right ascension of the reference stars or, in the case of the calculation, is equal to the angle ϕ, i.e. equal to the inertial longitude, which at zero hour angle point γ relative to the Greenwich Meridian is the geographic longitude.

Knowing the coordinate plane sensitivity 17 in axes star of the instrument and the angle ϕ, it is not difficult to calculate the axis of the star sensor zero position of the instrument reference inertial longitude corresponding to the absolute reference point 0.

The value of the same time point angle γ relative to the Greenwich Meridian caliroots side sidereal time, nullable through which the each 24 hours.

In addition to corners ϕ and λto coordinate the Earth in space is also used polar distance of the reference stars μ. This angle is used to coordinate Land in the first plane. Obviously, the angle of the center of the Earth - SPACECRAFT - supporting star"marked ωserves a similar purpose - it coordinates the position of the SPACECRAFT in the plane of sensitivity 17.

Note that the first plane 12, which contains the center of the Earth, the pole and the reference star, considered a fixed point, combined with the plane of sensitivity 17 formed by the Central axis of the earth sensor and the Central axis of the star sensor and in which the coordinated direction to the center of the Earth and the reference star.

If "unfreeze" the situation and to keep the orbital motion of the plane position sensitivity combined with the plane containing the direction of the KA - reference star" and "KA - the center of the Earth", you'll get the effect of the rotation of the plane of sensitivity of CA around the direction of the reference star 4.

This effect is due to the orbital motion of the SPACECRAFT and is supported by a respective angular control of the SPACECRAFT in space.

As bringing the SPACECRAFT to a specified angular position and maintain this position in time is impossible without the implementation of a managed motion of the SPACECRAFT.

Consider anybodies more details.

Managed is movement, which under the action of the control power AC goes from some initial state to a specified final state.

The nature of the controlled motion is caused thus the characteristics of the initial and final States.

In the General case, the angular position of the SPACECRAFT can be specified relative position of the two coordinate systems, axes one of which establishes the desired and the other is rigidly associated with the apparatus of the actual orientation of the device.

The first of these coordinate systems is called the basic reference system, the second - bound coordinate system.

Since in the General case, the axis of the associated and underlying systems do not match, the task orientation of the SPACECRAFT can be defined as the problem of combining the axes of the associated coordinate system with the axes of the base.

It is clear that to accomplish this, the motion of the vehicle around its center of mass should be managed. This control allows you to attach CA in any desired position in space and to maintain this position by the action of the apparatus of the various disturbing moments. Thus it is possible to distinguish two characteristic control mode. The first corresponds to the process of bringing the axes of the associated coordinate system to the axes of the base frame. If rassmatrivat is this mode since the separation of the SPACECRAFT from the booster, it includes:

- inhibition (damping) the initial angular velocity;

search physical landmarks used for technical play of the underlying reference system;

- turning maneuver with the aim of combining the axes of the base and related coordinate systems.

The second mode, called mode stabilization, is designed to maintain a desired orientation of the SPACECRAFT with precision.

The position of the coordinate axes of the coupled system in the reference coordinate system is determined by the angles of roll, yaw and pitch.

In our case, the basic coordinate system adopted a system of two directions "KA - the center of the Earth" 28 and KA - reference star 4, the mutual angular position of which varies depending on the location of the SPACECRAFT in space. As the third direction is connecting the image in the star sensor reference and navigation stars, we have used the projection plane 13, which is formed by directions "KA - reference star 4 and KA - navigation star 5. The angular position of the given direction relative to the base plane formed by the first two directions, also varies depending on the location of the SPACECRAFT in space.

For operations orientation to ensure discussed combining first, you need to repay the ü angular velocity of the SPACECRAFT.

When the separation from the launch vehicle SPACECRAFT acquires considerable momentum arbitrarily located in space.

Mode damping the initial angular velocity is implemented algorithm control signals measuring the angular velocity of SPACECRAFT and using the corresponding control point of the Executive bodies.

Next, the orientation of the SPACECRAFT in the Sun. This mode is required to provide service systems KA energy and retrieve a specific spatial orientation of the associated axes of the SPACECRAFT. Capture the Sun, bringing and keeping it in sight of the solar sensor is the traditional method of using solar signals of the sensor measuring angular velocity and control aspects of the Executive bodies of the channels of the roll, yaw and pitch.

By known algorithms are also implemented search Land 1 use directions "KA - Sun", the subsequent capture of the landmark earth sensor 16 and the construction of the geocentric vertical.

Appropriate turning maneuvers of satellites with change in coordination of the Sun in the solar sensor is the orientation of the field of view of the star sensor 15 at the area of the celestial sphere, which is the reference star 4.

It should be noted that the orientation of the field of view of the star sensor at pornozvezdy can be made to search and seizure of Land by the turn of the SPACECRAFT around the direction of the KA - The sun".

At the end of this mode, the television camera of the star sensor registers on the CCD-matrix images of stars that fall within its field of view. As a result of processing the received video data is determined by the position of the stars in the instrument coordinate system. In memory of the star sensor (on-Board star catalog) stores information about the position of stars in the star coordinate system.

The mapping instrument and catalogue of coordinates of the same stars makes it possible to determine the mutual orientation of the instrument and the star coordinate system. Recognition registered on the CCD matrix of the stars is carried out according to the criterion of sameness of the angular distances between them in both the above-mentioned coordinate systems.

Thus recognized the reference star 4, which is then combined with the plane formed by the Central axis of the star sensor and the sensor of the Earth.

When the motion of SPACECRAFT in orbit 19, the CPU controls the position of the continuous indications of the star sensor 15 generates setpoint intended to resolve the angular misalignment between the specified plane and the direction of the reference star. To resolve the error involved the Executive bodies and the control channel on the roll.

Managing channels of pitch and yaw by using dates is the IR of the Earth, intended to implement the geocentric vertical.

Dynamic processes in the implementation of these modes is theoretically well developed and in practice do not cause insurmountable problems.

Recall that the state of CA as a dynamical system in General is defined as a set of the state variables at time t_{0}on which it is possible to predict the behavior of a dynamic system at any time t>t_{0}.

The set of such variables is represented as a state vector.

It is known that not only the traditional slant range, azimuth, elevation, and their derivatives, but also the equivalent parameters can generate the state vector, moreover, the missing parameters can make mathematical dependencies, allowing to get the values immeasurable parameters.

Consider the state vector of the AC generated by equivalent parameters.

As already mentioned, during the operation of the spacecraft continuously oriented so that the plane of sensitivity persisted sensitive elements of the star sensor, has always been two directions: direction to the center of the Earth the Central axis of the sensor Ground and the direction of the KA - reference star.

Note that the coordinate fixation on all the time quantum of the provisions in the star sensor plane, containing directions for the stars, is relative to the instrument base count "0", which is discussed in the following material. Here we can only say that relative to the instrument base reference point "0" on-Board processor is calculated inertial longitude of the SPACECRAFT.

Position as "KA - reference star relative to the Central axis of the sensor, the Earth is calculated on-Board processor, and when the amendment is dependent on the value of the inertial longitude is determined by the inertial latitude of the SPACECRAFT.

The results of these measurements are determined not only passed the latitude and longitude angular distance, and the appropriate speed, which can be decomposed along the axes of the SPACECRAFT control, and when the vector addition - orbital velocity.

Thus become known angle and the location of the SPACECRAFT in inertial space, and converting taking into account the Earth's rotation speed and current Board time, the projection of the location of the SPACECRAFT relative to the latitudinal-longitudinal grid of the Earth, i.e. the track of the SPACECRAFT.

In General, for the implementation of spatial and angular motion of the SPACECRAFT from the state at time t_{0}in state at time t>t_{0}this set of state variables that characterized not only the initial angle and the location of the SPACECRAFT relative activities is but the latitudinal-longitudinal grid of the Earth,
but also the subsequent corner and the location of the SPACECRAFT.

When this state at the time t>t_{0}must conform to the required condition with reliable accuracy.

In this regard, we should note the following circumstance: measurement inertial latitude and inertial longitude angles are processed on Board the SPACECRAFT special software module of the statistical processing.

The purpose of statistical data processing is to identify the trend component in the stochastic input data and obtain the predicted values of the inertial-latitudinal and inertial longitude of the corners on some number of clock steps after the input values.

In the result of implementation of this procedure, the angles that define the position of the axes of the SPACECRAFT relative to the direction to the center of the Earth and relative to the latitudinal-longitudinal grid Land, acquire values increased accuracy.

Moreover, high-precision values of these angles are estimated from recent measurements at a few steps forward, allowing you to control the angle and the location of the SPACECRAFT in space with high precision.

It should be noted that the determination of the angular position of the SPACECRAFT and determining the location of the SPACECRAFT in space are not two separate modes.

High-precision mode determining the angular Polo is possible axes of the SPACECRAFT relative to the direction to the center of the Earth and relative to the latitudinal plane of the Earth, containing sub-satellite point at the same time is a high-precision positioning mode the SPACECRAFT in space.

From the above it follows that the set of state variables KA at time t_{0}sufficient to predict the behavior of a dynamic system at time t>t_{0}.

It should also be noted that the parameters required for the formation of the state vector KA is determined on Board the SPACECRAFT autonomously. If necessary, traditionally used parameters can be calculated on Board the SPACECRAFT on the relevant mathematical relations, which will be discussed in subsequent materials.

In the above materials noted that it is particularly important that the correct choice of reference stars.

Figure 4 and 5 shows a table of stars from around the North and South poles of the world, respectively. Previously it was noted that the polar stars zone is an area of particular direct ascents, where the annual increment direct ascents can be positive, negative, and even very close to zero.

Include coordinates of stars in our case, to the era of 2000, and to compare the coordinate increments imagine these epoch 1950.

For these epochs of data, including, in addition to coordinate, also symbols of stars and stellar magnitude, bring in two tables: tablice imagine coordinates of the stars around the North pole, in table 5 coordinates of stars around the South pole.

The choice of stars as a reference or navigation due to a valid value of its coordinate changes, and detecting the ability of the star sensor and the size of its field of view.

Magnitude should correspond to the detecting ability of the sensor. Mastered by industry sensors able to detect stars down to 7-8 magnitude. Presented in tables stars satisfy this condition.

As for the issue of coordinate changes of the stars, this issue should be considered in detail.

If we proceed from that point γ moves with a speed of 50",24 a year for 50 years, its offset will be 2512" or 168^{s}(2^{m}48^{s}), then for those same 50 years, the right ascension of the fixed stars should increase by the same 2^{m}48^{s}. However, as already mentioned, the polar zone is a special direct ascents stars and their increments for 50 years, as evidenced by data presented in tables coordinates vary in a fairly wide range.

When the increment of right ascension of the stars in the 2^{m}48^{s}equal to the offset point γit is fair to assume that the star is stationary, while the increment in 00^{m}00^{s}star for 50 years conditionally offset by 2^{m/sup>
48s.}

For our case of the North polar zone, we choose as the reference star 1642, and navigation - 454, from the southern polar zone as a reference star, 3983, as well as navigation - 459. Increment direct ascents of these stars are close to the specified value in the 2^{m}48^{s}.

This choice is motivated by the following consideration.

A sidereal day is the interval of time between the first culmination point γ and then the same culmination of this same point, which is shifted during this period. Therefore, the sidereal day is shorter (amount of displacement) of the period of rotation of the Earth. To the period of time between the first culmination of the star and follow the same culmination of the same star was equal to a full period of the Earth's rotation, the selected star unlike point γ must be stationary, i.e. the increment its right ascension should be at least close to the value in the 2^{m}48^{s}.

Here it is pertinent to note that the increment of the declinations of the stars of the Northern zone (-16 37.56 and -16 40.01) should be close to each other, as the increment of the stars of the southern zone (+42.22 16 +16 at 38.67).

It is also advisable to polar distances of the reference stars (1642 and 3983) were possible, and the angular distance between the reference and navigation is wesdome fit at least half the size of the field of view of the star sensor field of view of modern sensors are in the range of 8^{
°}×8^{°}to the 13^{°}×13^{°}and astrocytoma recognition constellation development MBB and SYRAH has a field of view of the three optical heads, 30^{°}×40^{°}).

In some specific cases, the choice of the stars can be produced at a more lenient criterion relative to the coordinate increments.

Also, in some specific cases, there may be different requirements for the size of the field of view of the star sensor.

It is obvious that (except stationary orbit) the net latitudinal range of the angular solution of the field of view of the star sensor should be increased depending on the values of the latitudinal locations of the SPACECRAFT in orbit.

Mastered by industry star sensors do not have sufficient for our purposes, a field of view in the latitudinal plane.

The necessary angular pulse solution of the field of view of the star sensor can be provided with a corresponding set of multiple optical heads, and/or use of the actuator that implements two fixed positional status.

When using a two-position actuator for its control will require a certain logic switching fixed positions of the optical head. For example, the command to switch from the first position to the second may be formed upon contact image TNA is Noah stars in some marginal coordinate area of the field of view of the sensor.

The option of increasing the field of vision through a set of optical heads can be implemented in a single monoblock together with the sensor of the Earth.

Compared with on-off option option monoblock is more preferred due to the absence of the positioning logic and exceptions are possible ljuftovyj errors fixing the position of the optical head.

One or another variant of the expansion of the field of view of the star sensor, a particular size of its latitudinal solution depends on the inclination used orbit (which is expected to provide latitudinal dimension in full) and is determined by the requirements of each specific project.

However, it is possible and common option for maximum latitudinal size of the field of vision through a set of optical sensor heads, as well as the common variant of one optical sensor head, which "holds" the reference star in the center of your field of view, tracking via the external drive variable angular position of the SPACECRAFT relative to the direction on the reference star. These options satisfy any inclination used orbit particular case, however, with unequal accuracy characteristics.

Note that the choice of the values of the latitudinal angle of the solution may affect the implementation of the target function KA, which may in some cases paragraph is to require the angular reorientation of target equipment.

With the active functioning of the AC often to execute the target function there is a need to move the Central axis of the target equipment with directions to the center of the Earth and refocusing it on one or another area of the earth's surface.

In the absence of the target apparatus corresponding drive the implementation of this orientation leads to inconsistency persisted plane sensitivity astrogation with the direction of the KA - center of the Earth" and to relocate in onboard algorithms reference inertial longitude and latitude location of the SPACECRAFT.

Using for this purpose a special Mat. ensure it would not be feasible.

The required angular orientation of the SPACECRAFT is rigidly mounted on it a target apparatus can be provided, if you use a two-stage actuator monoblock astrobingo.

Figure 6 presents the spacecraft with the advanced field of view of the star sensor and the ability to Orient the longitudinal axis to any point on the Earth through a two-stage actuator 22.

Monoblock 21 with extended field of view 20 in this case consists of two star sensors 15 and sensor Ground 16.

When time tracking of the same area of the earth surface is the best option.

However, the predictable phases of flight when consider is Ino minimum time perenatselivanie you can do without a special drive.

If the target hardware has the ability differently oriented relative to the longitudinal axis KA its Central axis, the drive is also not necessary.

Obviously, in some situations, the selected stars can be shaded by the Earth and therefore may not be available to astrogation KA in some areas of the orbital motion.

Figure 7 presents: left - option, which involved the stars around the North celestial pole, right - stars of the South pole.

To the left are low Equatorial circular orbit 19, middle circular orbit with some "optimal" inclination and orbit with "suboptimal" inclination.

It is clear that, if the inclination of the circular low or average circular orbit does not exceed some threshold value, then all points of the orbit have the visibility of the involved stars 4 and 5 of the North polar zone. The point of the orbit is not visible stars involved when they are shaded by the Earth 1 as on the shady part 32 of the orbit with "suboptimal" inclination.

In General it should be noted that the length of the shadow area (and its availability) is determined by the diameter of the Earth, the orbital inclination and altitude.

The geostationary orbit in this case, the special situation - her point is observed involved as North, t the to the South stars.

The image on the right is involved, the stars around the South celestial pole and presents: low circular orbit, the average circular orbit and orbit type "Lightning", the inclination of which about 65°.

What was said for option involved with the North stars, is true for the variant involved with the southern stars.

Especially it is necessary to emphasize the following: in HEO orbit type "Lightning" perihelia plot is not observable Northern stars, apogee same area, which is usually the target area functioning SPACECRAFT, located in the Northern hemisphere at a considerable height, and therefore the observed and the Northern and southern stars.

Circular polar orbit unlike examined have the following features: due to the inadmissibility of the so-called "folding" of the axes of the control they have in the polar zones of the sites implemented passive mode KA - continuous solar orientation (PSO).

As already noted, the first plane 12 and the second plane 13 astronomical bases are combined, respectively, at some point, the orbital motion of the SPACECRAFT with its plane of sensitivity 17 and the plane defined by directions "KA - reference star" and "KA - navigation star.

Angle λ between these planes define Elaida in advance during the design of the triangle, the nodes on the celestial sphere are: the pole, the reference star and the navigation star.

On Fig this triangle with the target angle is depicted on the sphere. In this regard, it is shown tangent 22 and 23 to the sides at the top of the desired angle. But as the radius of the celestial sphere of infinite size, to calculate the required angle using the following formula planar trigonometry that connect to the sides and angles of a triangle:

1) cos ε=cos β cos μ+sin β sin μ cos σ;

2) cos β=cos μ cos ε+sin μ sin ε cos λ.

The determined angle λ corresponds to the moment when the SPACECRAFT is in the plane of the climax of the reference stars. At other times the plane sensitivity, containing the Central axis of the sensors, due to the effect of rotation omelet for orbital revolution of the heavenly sphere, and the navigation star 5 (display) moves in the star sensor 15 by a closed curve.

Refer to Fig.9, which shows two characteristic positions of SPACECRAFT in orbit: right, coinciding with the plane in which culminates reference star, and left, coinciding with the absolute reference direction of 0.

The position on the orbit space 19 of the apparatus 14 shown in the figure to the right, is characterized by the angle λin which inertial longitude KA equal direct is shoedini reference stars.

Obviously, at some point in the orbital motion of the inertial longitude of the SPACECRAFT will be equal to 0°. As the plane 13, containing directions on the stars, constant in space, then any angular rotation of the SPACECRAFT orbital motion would entail a change in the angle between this plane and the plane of sensitivity 12. This occurs continuously during the motion of the SPACECRAFT in its orbit when the Earth centre is 25 and the reference star 4 (display) are held in the plane of sensitivity.

When switching the AC from the right orbital position in the left plane of sensitivity combined with an absolute reference direction. The angle between the plane of sensitivity and display in the star sensor plane 13 (second plane astronomical basis) takes the value that captures the instrument base reference inertial longitude and which corresponds to the zero value of the inertial longitude.

We emphasize that the angle ψfixing in the star sensor instrument base counting "0" relative to the plane of sensitivity, is determined in advance by calculation when zeroing inertial longitude KA equal to the right ascension of the reference stars. The angle at any location of the reference stars in the plane of sensitivity has the same value, which is taken into account then, boron is the new algorithms for the calculation of inertial longitude.

The inertial latitude of the location KA is defined as the difference between the measured angle of the center of the Earth - SPACECRAFT - supporting star and 90°however, it is necessary to consider the amendment, which is due to the polar distance of the reference stars μ. Cm. figure 10.

The light of this amendment, provides essentially the angle from the center of the Earth - KA - pole of the world" in any orbital points, which, of course, she has a certain value.

Note that the spatial position of the SPACECRAFT is characterized by the fact that the plane sensitivity coincides with the direction of the center of the Earth pole of the world" in moments of alignment with the plane of the angle μ, that is, in the moments passing plane in which the reference star is in the upper or lower culmination.

Obviously, in these moments, the measured angle of the center of the Earth - SPACECRAFT - supporting star" is necessary when determining the latitude to reduce the magnitude of the polar distance μ at the top of the climax of the stars and to increase by the same amount at lower culmination.

In orbital positions in which the plane of sensitivity of the SPACECRAFT perpendicular to the same plane sensitivity KA placed in position 1, the correction due to the polar distance of the reference stars, accepts null values.

In General, the current adjustment value depends on the current values of the inertial to the Goths.

Each value of the inertial longitude strictly complies with a specific correction value, which can be pre-calculated and are presented in the on-Board storage device.

The current amendment is determined by the product of the polar distances of the reference stars on the cosine of the angle whose value is equal to the difference of the current values of the inertial longitude and angle between the first plane and the absolute reference direction.

To determine the latitude of the location of the SPACECRAFT, thus, you should use the ratio:

η=(ω-90°)-μ cos(ι-ϕ),

where η latitude ω - the angle of the center of the Earth - SPACECRAFT - supporting star, μ - the polar distance of the reference stars, ι - inertial longitude, ϕ - the angle between the first plane 12 and the absolute reference direction.

The polar distance of the reference stars also causes fluctuations in the AC roll.

Retention areas on the reference star and the center of the Earth in the plane of sensitivity airborne sensors leads when the orbital motion of the SPACECRAFT to fluctuations in the plane of the feed roll with the amplitude equal to the polar distance of the reference stars.

The angular position of the SPACECRAFT relative to a North-South axis depends on the inertial longitude of the SPACECRAFT.

Obviously, in the orbital positions of the SPACECRAFT, when ι=ϕ and when ι=ϕ+180°plane-the th sensitivity KA coincides with the plane of the angle μ and with the North-South axis.

In orbital terms KA, when ι=ϕ+90° and when ι=ϕ+270°, plane sensitivity deviates from a North-South axis by an angle equal to the polar distance of the reference stars.

Each value of the inertial longitude strictly corresponds to a certain angular position of the SPACECRAFT relative to a North-South axis, which can be pre-calculated and are presented in the on-Board storage device.

The current angular position of the SPACECRAFT is determined by the product of the polar distances of the reference stars on the sine of the angle whose value is equal to the difference of the current values of the inertial longitude and angle between the first plane 12 and the absolute reference direction.

The scheme of measurements of the inertial latitude and longitude orbital location CA presented on 11 showing: 1 - Ground, 14 - KA, 19 - orbit, 25 - the center of the Earth, ι - inertial longitude, η- breadth.

Inertial longitude is the angle of absolute space in the range from 0 to 360°measured in the direction of rotation of the Earth in the plane of the celestial equator from the absolute reference point 0 and coincides with the geographical longitude through each complete rotation of the Earth.

Here it should be clear that the right ascension coordinates of the fixed stars with respect to the point γand in realna longitude coordinates moving against the background of fixed stars, the spacecraft relative to the absolute reference point 0.

To transfer inertial longitude in geographic uses the processor angular data.

The translation is performed as follows. The mutual ratio of longitude and inertial longitude at different points in a 24-hour cycle (the period of complete rotation of the Earth) is represented in the variants a) and b) on Fig (view from the North celestial pole).

In this figure, the following notation: 0 - absolute reference direction, 1 - Ground, 2 - the celestial sphere, 8 - Meridian of Greenwich, ι - inertial longitude, ∂ - geographic longitude, t is the time at Greenwich. In option (a) illustrates the relative position of the values of longitude and inertial longitude at 0^{h}on the Meridian of Greenwich. At the moment the values of longitude coincide with the values of the inertial longitude. Thus, the value of the inertial longitude 30° (designated in the figure by an arrow) corresponds to the geographic longitude in the 30°. Option b) illustrates the relative position of longitude and inertial longitude at time t=10^{h}. At the moment, as well as other non-zero points in time, the values of longitude does not coincide with the values of the inertial longitude.

Consider two possible ways in which variously defined otnosheniya.jdu longitudes.

1) Determine at this time the importance of geographical longitude, which should correspond to, for example, 240° inertial longitude (denoted in the figure by an arrow).

At time t=10^{h}The earth will turn through an angle equal to 150°. This angle is less than the inertial longitude 240°, their difference will be 90°. In this case, the inertial longitude 240° corresponds to the geographic longitude in the 90°.

2) Determine the same point in time t=10^{h}the importance of geographical longitude, which should correspond to, for example, inertial longitude 90°.

The angle of rotation of the Earth 150° not less, but greater than the value of the inertial longitude and the difference between them is 60°. In this case, geographic longitude will be equal to the difference between 360° and 60°, that is, 300°.

Thus, the geographical longitude and inertial longitude associated with the following dependencies:

1) λ_{g}=λ_{and}-(ωat. t) at ωat. t<λ_{and};

2) λ_{g}=360°-[(ωat. t) -λ_{and}] ωat. t>λ_{and},

where the speed of rotation of the Earth ω=15°/h; λ_{g}- geographic longitude; λ_{and}- inertial longitude; t - time.

As for the orientation axis of the SPACECRAFT relative to the center of the Earth, then this orientation is provided by a sensor of the Earth.

The scheme of the geocentric vertical with four scanning sensors presented on Fig.

The principle of determining the geocentric vertical 18 is based on the measurement of the angle between the direction of the curvature of the Earth 27, defined as the boundary of space-to-Earth, and the axes of the SPACECRAFT at different points of the horizon.

The geocentric direction of the vertical axis is the bisector of the angle between two diametrically opposite points visiwave horizon 27.

The principle is based on the use of temperature contrast that exists between the earth's surface and outer space. Used for this purpose, the sensors must be sensitive to radiation in the range of 11 to 20 microns.

The field of view of sensors are arranged in pairs in two mutually perpendicular planes, and the angle of view of each sensor by scanning the instantaneous field of view 28 covers the area from the boundary of the Earth-to-space.

A receiver of radiant energy fixes the angular position of the horizon of the Earth as a sharp drop of the energy brightness (optical contrast) between the edge of the disk of the Earth and the surrounding space.

Sensor Earth measures the position of the horizon in the onboard system of the SPACECRAFT coordinates.

There are two types of measurement errors:

1)instrumental errors;

2)errors caused not what oceanstor phenomena of the visible horizon (error "phenomenon).

Instrumental errors of the sensor is the electronic noise, quantization errors, and others.

Instrumental errors are represented as a sum of Gaussian random process and errors of the displacement type. Random errors are not correlated (white noise), constant measurement error (offset) is acceptable to assume errors with infinite correlation time.

Error "phenomenon" caused by the uncertainty of the layer of CO_{2}in the Earth's atmosphere. Due to temporal and spatial fluctuations of the atmosphere, the definition of horizon on the position of the layer in CO_{2}made with greater error than in the case of the horizon directly on the edge of the earth's disk (the earth's limb).

Error "phenomenon" is most pronounced at low altitude flight SPACECRAFT and decrease rapidly with increasing height.

At high altitudes is dominated by instrumental errors. In our case, using the sensor, consisting of four scanning devices: two of them produce the dimension in the plane of sensitivity and in two orthogonal planes.

On Fig presents the scheme of the geometry of measurements geocentric vertical, which is "orthogonal" to the plane of sensitivity.

After measuring sensor 16 of the angle Z between the side frame 30 and the direction of the visible horizon of the Earth and angle ∑ to directions on the horizon in the opposite direction is defined geocentric vertical 18, which should be in the plane of sensitivity 17 and strictly focus on the center of the Earth 25.

Note that the error ζcaused by uncertainty horizon, electronic noise and offset, and error ξ when measured in the plane of sensitivity, generally speaking, makes a strict focus on the center of the Earth.

The measured angles are defined by the following relations:

Z=Z_{Mr.}-ζ;

∑=∑_{Mr.}+ζ,

where Z_{Mr.}that ∑_{Mr.}- nominal angle values.

Similarly defined angles in the plane of sensitivity. The angle between the side frame 30 and the direction of the center of the Earth is defined as half the sum of the measured angles.

Note that the accuracy of the sensors used incommensurable values with each other. High precision of the star sensor in 2-3 arcsec becomes extremely important due to the fact that the angle between the Central axis of the sensor Ground and the direction of the KA - reference star can be measured with almost equivalent accuracy, if initially on the Ground to implement and pasportizirovat the relationship between the coordinate systems of the star sensor and the sensor of the Earth.

In particular, this angular relationship between the Central axis x of the sensor of the Land 16 and the sighting axis y_{0}star sensor 15 is necessary in order to suscestvitel and to align in the same block and thus,
to the plane of the two axes x_{0}, y_{0}star sensor and the plane of the two axes x, y of the sensor of the Earth formed a single plane sensitivity 17. Cm. Fig.

The value measured at this angle ω_{0}between the Central axis of the sensor Ground and the sighting axis of the star sensor should be provided in the respective on-Board algorithms.

Obviously, in this case, the angle in the field of view of the star sensor between the sighting axis and the direction of the KA - reference star must be summed with the angle ω_{0}representing the above link.

This enables high-precision measurement of the position of the Central axis x of the sensor of the Earth relative to the direction of the KA - reference star. Whether in this case the Central axis of the sensor with the direction to the center of the Earth, is another question. Here it is important only very high-precision knowledge of this situation.

This is a high precision knowledge refers to the latitudinal angular relation to the plane of sensitivity, in which the level of precision of the star sensor is essentially a basic level.

Longitudinal measurements realized in the plane perpendicular to the plane of sensitivity. But since the instrument base reference inertial longitude is fixed relative to the plane of sensitivity with high calculation accuracy, caviano, the actual measurement of the longitudinal angle can also be carried out with accuracy equivalent to the accuracy of the star sensor.

Thus, when high-precision angular relationship between the coordinate systems of sensors accuracy level of the star sensor is the base level as in the latitudinal dimension and the longitudinal dimension of the angle.

On Fig presents the scheme of the differences between systematic and random errors of the sensor of the Earth.

The center of each target represents the true value of the measurand, and each dot is measurement. The amount of measurements is characterized by the offset and dispersion. In the figure the first and third top presents biased results. The standard deviation or the variance of the individual measurements is a measure of the error. Sensor with good repeatability (or a small random error) is, obviously, a good random error, but not necessarily give the correct output value, since the shift can significantly distort the result, i.e. the accuracy of the sensor is small. The results of the measurements in the figure the second and fourth have a small error, but only the result, shown in the fourth figure is accurate. The diagrams to the right of the true value is represented by a straight line, on which is superimposed the results of the measurements.

large offset + large scatter = low accuracy;

- small offset + large scatter = low accuracy;

large offset + small spread = low accuracy;

- small offset + small spread = high precision.

In terms of the orbital motion of the coordinate system of the sensors is converted in the navigation coordinate system associated with the longitudinal and latitudinal location of the SPACECRAFT. There are a number of other transformations onboard coordinate systems that generate errors. Thus, to instrumental errors and sensor errors caused by the variability of the phenomena of the visible horizon of the Earth, added error transformations. When building a geocentric vertical measurements are made through two channels: pulse, i.e. in the plane of sensitivity, and in the longitudinal orthogonal plane.

According to the results of measurements and subsequent transformations of coordinate systems the Central axis of the sensor by means of a corresponding rotation of the SPACECRAFT is oriented to the center of the Earth with the specified error.

All errors regardless of the nature of their origin, let's agree to consider in a longitudinal channel in one longitudinal error, and all errors in the latitudinal channel as one latitudinal error.

Angular spatial position of the SPACECRAFT in the plane, orthogonal the Noah plane sensitivity, that is, in the plane of the longitudinal channel provided on Fig.

Note that the designation under figure 18 is geocentric vertical, while the direction from the sensor 16 to the center of the Earth 25 rejected on the value of the longitudinal error ζ. Under figure 18 it would have been better to present the Central axis of the sensor of the Earth.

Obviously, if the Central axis of the sensor was strictly focused on the center of the Earth 25, inertial longitude ι would have a different value, namely the value for the situation depicted in this figure, it would be less by the amount longitudinal error ζ.

This reasoning is also valid for the angular position of the SPACECRAFT, depicted on Fig in the latitudinal plane of the channel.

This figure under figure 18 is similar to the notation in the previous figure is bogus geocentric vertical, differ from the true vertical due to the presence of errors ξ. It is obvious that the elimination of latitudinal error ξ angular position of the SPACECRAFT, depicted on the figure in the space of the Northern and southern hemispheres of the Earth, will be the true latitude of your location.

Latitudinal and longitudinal errors lead the Central axis of the sensor in the direction from the direction to pull the center of the Earth. However, whether these errors in the next dimension, neizvesten is. But regardless of this uncertainty, it is safe to say that KA in each moment of measurement of latitude and longitude is the real point of the orbit.

Permissible, however, to imagine that the Central axis of the sensor is always strictly focused on the center of the Earth. However, this assumption SC in each time of measurement must be in real orbital point, and in some imaginary point.

Contact Fig, which presents two types a) and B), illustrating the real and imaginary position of the SPACECRAFT in the moments of the measurement of the latitudinal angle, respectively.

At this illustration: 1 - Earth, 14 - KA, 4 - direction on the reference star, 6 - equator, 18 - geocentric vertical, 19 - orbit, 19_{f}- imaginary orbit, 25 - attractive center of the Earth, η latitude ξ - latitudinal error.

The picture on the left (type a) spacecraft 14 is valid orbital point. He is also depicted and other real point where the SPACECRAFT is at the points of measurements of the latitudinal angle.

Latitudinal angle η measured with error ξ and therefore the Central axis of the sensor is oriented in the direction from the center of the Earth.

The figure on the right (view B) shows the imaginary position where the Central axis of the sensor with the same error ξ in the latitudinal angles η coincides with the geocentric vertical is Yu 18,
but KA is imaginary points, forming imaginary orbit 19_{f}.

It should be noted that all related with the latitudinal dimension to the imaginary points is true for the imaginary points in the longitudinal dimension.

Such an imaginary way, you can represent the error of the sensor of the Earth in the form of an error of the nature of the motion of the SPACECRAFT.

Obviously, latitudinal and longitudinal angles, if they did not contain errors that characterized the motion of the SPACECRAFT, its orbit with high precision. You can say in another way - the nature of the motion of the SPACECRAFT, i.e. its location in space, at different points in time leads to different values of the latitudinal and longitudinal angles at the same time.

Thus, the values of the latitudinal and longitudinal angles depend, on the one hand, from the location of the SPACECRAFT in space, on the other hand, from the measurement errors.

Contact Fig, which illustrates this dependency.

This figure illustrates: 14_{about}HERE in the real point of the orbit at zero time 14_{about} ^{m}HERE in an imaginary point at the same time, 14_{4}ME in real point in time in 4 minutes, 14_{4} ^{m}HERE in an imaginary point in the fourth minute, 3 - direction on the pole, 18 - geocentric vertical, 19 - orbit, 25 - the center of the Earth, η_{0}a spread angle of zero, the PTO is t time,
η_{4}a spread angle at the time in four minutes, ξ_{0}- latitudinal error on the zero moment of time.

Imaginary points located on the figure along the plane of the orbit, forming a cloud of measured pulse data.

When a sufficient number of such points in the cloud data statistical processing.

Obviously, in our case, the statistical processing latitudinal and longitudinal data can be used to identify the nature of the motion of the SPACECRAFT, i.e. to accurately identify the line along which movement occurs KA.

It is also clear that the results of such processing is the Central axis of the sensor should be oriented towards the center of the Earth with high precision.

From a statistical point of view, the measured data should be attributed to the time series, which can be processed in one way or another repayment nonsystematic components.

Comparison of solutions to the problem of approximation of the curve, which moves the SPACECRAFT revealed the advantage of the neural network, in particular RBF-network with radial basis functions).

Thus, for on-Board aggregation is better to use such software module. On Fig presents a flowchart of the processing.

Note that due to perturbations of the orbital parameters of the SPACECRAFT is moving on the intent of the wool line not located essentially in one plane and not being closed.

In this regard, the training of the network (its programming) strongly depends on the requirements of her in the learning process of the training set data. Training data must be representative of the tasks, the solution of which the network is trained. Such programming network can be carried out before launch, but you can train the network to produce and in real conditions.

To confirm the validity of using neural networks as on-Board software module processing a simulation of the motion of the SPACECRAFT.

The simulation was performed on a personal computer using a neural network system STATISTICA.

This choice is dictated by the task of time series, in which the relationships between the data in our case is close to linear in a limited area of the orbit and obviously nonlinear relatively long segments, for example at full turn.

In the simulation as a reference parameter for assessing the estimated inertial latitude and longitude, as measured angles - inertial longitudinal and inertial latitudinal dimension of the stochastic nature and quality of the filtered parameters - parameters that are processed by the neural network.

Assessment at the appropriate time interval were used value is assogbavi between the calculated inertial, the so-called real options and processed, filtered parameters.

The input of the neural network for training were submitted latitudinal and longitudinal data sets of stochastic nature. It should be emphasized that the training data and the measured data (inertial longitude and latitude) are different fluctuations, as evidenced by the graphical dependences presented on Fig. Processing information intended to identify the trend component in conventionally measured data.

On Fig table summarizes the results of processing by the neural network is conventionally measured information on the orbital area from 500 to 559 sec.

As input of the neural network was used to measure inertial latitude (column 2), the measurement inertial longitude (column 6) and the corresponding time (column 1).

The resulting data processing presents the corresponding values for each measure time in columns 3 and 7 of the table.

So, in column 3 presents processed, that is predicted by the neural network, the values of the inertial latitude, and column 7 - the predicted values of the inertial longitude.

In column 4 and 8 shows the reference value of the real inertial latitude and real inertial longitude, respectively, which are, as agreed, to compare them predicted values and determine the effectiveness of the processing module.

Analyzing the errors presented in angular seconds in columns 5 and 9, it is possible to recognize that processed by the neural network data as inertial latitude and inertial longitude possess good convergence with the reference values.

On Fig and 25 presents the results of processing in graphical form. Presents a graphical time-dependent real, measured and predicted information on the area from 500 to 559 seconds (0 to 60 sec). From the graph it is seen that the sawtooth line (measurement) converted by the processing in the smooth dotted line (forecast), which coincides with the solid line, the real line of motion of the SPACECRAFT.

On Fig, 27 and 28 shows the table and the graphical dependence of the results of processing by the neural network is conventionally measured information on other orbital plot 2999 on 3058 s.

Interpretation of the data and findings of this processing is similar to the previous.

On Fig and 30 presents a graphical depending on the time of processing results is conventionally measured information on full turn. The graphs indicate good convergence of the processed data with the reference values (dashed line).

On Fig presents a comparative table of the parameters processed by a linear trend and a neural network on the orbital area with 3049 on 3058 s.

In this the table in columns 2 and 5 shows the reference value of the real inertial latitude and real inertial longitude, intended for comparison with them processed the predicted values.

Processed by a linear trend and neural network values of the inertial latitude are presented in columns 3 and 4, respectively.

Processed by a linear trend and neural network values of the inertial longitude are presented in columns 6 and 7 respectively.

Comparative analysis, therefore, subject to the parameters columns 3 and 4 and column options 6 and 7.

The result of this comparison is the degree of their approximation to the parameters presented in columns 2 and 5.

Obviously, the preferred method of processing the measurement data is the method based on the use of neural networks.

On Fig and 33 presents graphs comparing treated by different methods on the plot with 3049 on 3058 sec. It is obvious that most of the convergence criteria have parameters that are processed by the neural network.

Thus, the simulation showed that the most preferred method of processing the measurement data is the method based on the use of neural networks.

For a given target function KA is always the choice of an appropriate orbit.

Its spatial-temporal characteristics must meet all of the criteria for decisions facing the KA task, whether it's about the zorny observation of a given area of the earth surface, whether it's shooting a specific ground-based object sensing in the interests of economic activity or connectivity between the areas.

Specific optimum for the target orbit is the basic environment in which complex interactions of various spacecraft systems are in space corresponding to a specific time, and in time corresponding to a particular space.

That is why it is so important to keeping the parameters of the estimated orbit within acceptable ranges.

Meet the requirements of the orbital stage, given their traditional settings and in the form of inertial longitudinal and inertial latitude values with a clock time step equal to the quantum of onboard measurements, seems to be on Board the SPACECRAFT through its records in a persistent storage device.

In orbit as statistical processing of the measurement data of the next real revolution is written in the random access memory for subsequent comparison with the set.

Any orbit is uniquely determined by six independent parameters Keplerova movement, four of which characterize the motion of the SPACECRAFT inside the orbit plane and two - plane orientation of the orbit in space. The orbital elements are parameters that determine the abused its position in space, the size and shape. Contact Fig, which includes: a given orbit, the real orbit and all the elements that characterize each of the orbits.

This figure shows: 1 - Ground, 6 - plane of the equator, 19_{1}, 19_{2}- real orbit and given, respectively, 0, γ - the point of the vernal equinox, B_{1}(B_{2}and H_{1}(H_{2}) - ascending and descending nodes real (given) orbit Ω_{1}( Ω_{2}) - right ascension of the ascending node of the real (given) orbit, P_{1}(P_{2}and A_{1}(A_{2}- the perigee and apogee real (given) orbit, i_{1}(i_{2}- the inclination of the real (given) orbit ∈_{1}(∈_{2}- the argument of perigee real (given) orbit 2α_{1}(2α_{2}- the distance between perigee and apogee real (given) orbit ϑ is the true anomaly of the SPACECRAFT (the argument KA u=ϑ+∈).

The parameters and characteristic orbital points of the actual orbits are determined on Board the SPACECRAFT as follows.

The period of revolution T_{1}is measured as the time interval between successive passages of the KA of the two points where the value of the latitude is the same (for example, points of the ascending or descending node of the orbit).

When the inclination of 0°, T_{1}is the time interval between successive p is kodename KA two points,
in which the value of the inertial longitude is the same.

The semimajor axis α_{1}is determined by the formulawhere k is the gravitational parameter, equal 3,986·10^{5}km^{3}/s^{2}.

The angular velocity at the perigee and apogee of the real orbit is known from the results of measurements.

The point of maximum angular velocity becomes the point of perigee and coordinated by the angles of latitude and inertial longitude. Likewise determined and coordinated and the point of climax.

The ratio of the speeds of the SPACECRAFT's perigee and apogee is inversely proportional to the ratio of the radius-vectors of these points of the orbit, i.e., V_{p}/V_{a}=r_{and}/r_{p}where r_{and}, r_{p}- the distance from the center of the Earth to apogee and perigee, respectively. Due to the fact that 2α=(r_{and}+r_{p}), the velocity ratio can be represented in the following form V_{p}/V_{a}=(2αr_{p})/r_{p}where is determined by the radius-vector of the perigee. And then from the formula 2α=(r_{and}+r_{p}) after the substitution it r_{p}is defined by r_{and}.

The eccentricity of the orbit is known from the following formula e_{1}=(r_{A1}-r_{P1})/(r_{a1}+r_{P1}).

Upward B_{1}(and downward H_{1}node of the orbit is determined directly as a point with a certain value of the inertial longitude IP zero latitude.

Knowing the coordinates of the point of perigee P_{1}and the ascending node In_{1}is defined in the orbit plane angle corresponding to the argument of perigee ∈_{1}.

Similarly calculates the argument KA u or true anomaly KA ϑ.

The inclination of i_{1}the real orbit is defined as the maximum value of the latitudinal angle.

Thus become known following the parameters of the real orbit: α_{1}e_{1}that ∈_{1}that Ω_{1}, i_{1}and u (or ϑ).

Comparing the parameters of the real orbit and presented in the MD parameters given orbit are determined by the desired deviation.

So, by comparing the values of the inertial longitude at the time of passage of the SPACECRAFT ascending nodes of each of the orbits (i.e. at the moment when both values of latitude is equal to zero), we get the error Δ Ω as the difference of these longitudes.

Similarly, it is enough to compare the values of the inclinations i_{1}and i_{2}and their difference will reveal the value of their disagreements Δi.

As a result of implementation of the correction of the real orbit to eliminate mismatches Δ Ω and Δi the real plane of the orbit will be aligned with the plane of a given orbit.

Here it should be noted that when the same value Ω and when the same value of i, the size and shape of the two orbits that lie in the same PLO the bone, can not match.

Different orientation of perigee two orbits shows that the values of their arguments are different, and therefore Δ∈ quantifies the specified orientation and the corresponding Δα and Δe quantitatively characterize the differences in size and shape.

It should also be noted and the difference in the form of measurements of the orbital velocity: the speed of the real orbit is presented in deg/sec, and the velocity of a given orbit - km/sec (in deg/sec).

The following formula allows us to determine the velocity in km/sec real orbit at perigee and apogee:

V^{2} _{p}=(k/α)[(1+e)/(1-e)] and V^{2} _{a}=(k/α)[(1-s)/(1+e)].

The comparison of these velocities with angular velocity allows to ensure their mutual convertibility at any orbital area.

This ensures that the determination of the deviations of the real orbit from the set: Δα, Δe, Δ∈, Δi, Δ Ωand the convertibility of velocity.

The measured on-orbit orbit data, as shown above, smoothed by the neural network at each time step with second errors. Such processed data is written to the on-Board memory for use in determining the parameters of the orbit. The time quantum for a particular orbit for the optimum. In the profile is discussed below from three types of orbits (see Fig) say that in our case, when determining the perigee and apogee of the time step is set to 10 seconds. On Fig presents a summary table of the design parameters of the three orbits (three subtables). High elliptical orbit with period values, altitude perigee and apogee and eccentricity of the set and with the time step 1000 sec, the measured angle in degrees and a speed of V/sec at each time after the expiration of each step. With the appropriate options set to "steady-state" and a low elliptical orbit. Note that the data provided to the memory detected angular distance traversed KA for every 10 seconds. Determine for each of the three orbits of the angular distance traversed per time steps, and will postavim the values obtained in the empty third columns of each subdatasheet (empty fourth columns of each of the subdatasheets are used to commit the zone of location of the point of perigee and apogee). It is easy to verify from the subdatasheets that the lowest values of angular distances, such as 1,86°, 8,16° and of 10.05° correspond to the zones of finding points of the peak, and maximum values of the angular distances of the zones of location of the point of perigee. In the fourth columns of each subdatasheet these zones are marked by letters A (apogee) and P (perigee). The range of such areas is restricted to be equal (or close) values of the angular distances for the 10-second step(V/c× 10 (C), which significantly differ in their values from intraband values when the available precision. Ranges perigynous zone (PZ) and apogee zone (AZ) (table) orbits will be such:

- high PZ: ˜1530",4→˜2991",6←˜1640",16 (1530"); AZ: ˜80",028→˜66",96←˜79",92;

- fixed PZ: -150",84→˜154",08←˜150",84; AZ: ˜150",12→˜146",88←˜150",12;

- low PZ: ˜2309",4→˜2325",9←˜2301",8; AZ: ˜1814",4→˜1804",3←˜1819".

Note that for quasi-stationary orbit requires a step of more than 10 seconds. The median angular range values zones define the point of perigee and apogee. It should be borne in mind that the points a and d spaced apart an angular distance of 180°. On Fig illustrated above, and presents the locus of angular velocities circular and HEO orbits. Vectors OR equal to here the angular velocity in successive moments of time.

When you commit the ascending node of the orbit, perigee on the smoothed data, the argument of perigee is defined as the angle between the longitudes of the ascending node and perigee (in the direction of motion of the SPACECRAFT).

Thus defined apogee, perigee and its argument, and of the following formulas - all necessary parameters: T=2π√α^{3}/k, where α; 2α=(r_{a}+r_{p}but, because V_{p}/V_{
=ra/rpthen Vp/Va=(2αrp)/rpwhere rp;}

r_{a}=(2αr_{p});

e=(r_{a}-r_{p})/(r_{a}+r_{p});

r=α(1-s^{2})/1-e cosϕwhere ϕ here the angle from the axis of the ellipse connecting apogee to perigee.

Correction of parameters of the real orbit to ensure their compliance with the parameters specified orbit depends on the magnitude, point of application and direction of the correction pulse.

The direction of action of the known lateral and tangential impulses, under the influence of which is the elimination of the deviations of the real orbit from the set.

It should be noted that in our case the corrective maneuvers should be implemented under the action of pulsed thrust.

Consider what consequences arise under the application of lateral momentum.

Contact Fig showing: 1 - Ground, 10 - plane of the equator, 30 - orbit line of nodes NR - axis changes the orbital inclination i, the axis of the world F_{1}P - axis change of right ascension of the ascending node Ωaxis F_{1}Q - axis change of argument of perigee ∈ (this axis is perpendicular to the orbit plane and passes through the center of the Earth), γ - the direction to the vernal equinox And the point of climax, t - perigee.

It is obvious that the side force changes i Ω, ∈,ie the orientation of the orbital plane in space, but it doesn't change the shape of the orbit, nor its size.

The point of application of the lateral force in the maneuver to align the planes of the orbits is one of the two points on the line of intersection of the planes of the orbit (the point closest to the peak).

This point is determined when comparing the real and the set of orbits as the point with the General coordinates, i.e. as a point that belongs to both orbits.

Consider Fig, which presents the real orbit 30_{1}with the ascending node of the B_{1}with an inclination of i_{1}and the argument of perigee ∈_{1}. These parameters of the real orbit does not comply with the same parameters In_{2}, i_{2}that ∈_{2}given orbit 30_{2}. After application of the corresponding pixel side of the corresponding impulse ΔV_{in}the parameters of the real orbit will acquire parameter values given orbit.

Under the action of lateral momentum will be a reversal of the velocity vector with respect to an axis passing through the point of application of lateral forces (point M, Fig) and the center of the Earth (point F_{1}).

When flying SPACECRAFT in an elliptical orbit, the velocity vector at any point, except perigee and apogee, does not coincide with the line of the local horizon. The projection of velocity vector on the line of the local horizon is called the transversal velocity V_{η}=VcosΘwhere the angle Θ=arctan[(1+ecosϑ)/esinϑ]. Here e, ϑ and V is known.

When the lateral maneuver at the point M of the pivot axis of the orbit is the axis M M passing through the point of application of the lateral force and the center of the Earth. On this axis and directed vector of pivot orbits Δχ, i.e. the pivot axis of the orbit does not coincide with any of the axes changes i Ω and ∈. In this case, the increment of all three elements (Δi_{in}that Δ Ω_{in}that Δ∈_{in}are determined by the projections of the vector of spread to the orbit plane Δχ on the corresponding axes of their changes.

Change the orbital inclination, right ascension of the ascending node and argument of perigee under the action of lateral momentum depend on the argument KA u at the point of the maneuver (u=ϑ+∈).

For change of right ascension of the ascending node, the impulse to lateral forces must be applied at the points with argument equal to 90° or 270°. However, when the orbital inclinations of i not equal to 90°the application of lateral momentum in the pixels of the original orbit with u_{0}=90° or u_{0}=270° will cause a change in argument of perigee. Since u=∈+ϑ, the change of argument of perigee ∈ will lead to the change of the argument KA u at the point of the maneuver. Therefore, the argument SPACECRAFT into orbit after the maneuver will not be equal to 90° or 270°. In is the result of lateral momentum will change not only the right ascension of the ascending node,
but the inclination of the orbit.

Contact Fig showing: 30_{1}- the source plane of the orbit (the plane of the real orbit), 30_{2}- the plane of the orbit after the maneuver (the given plane of the orbit), Δχ - turn angle of the orbital plane, ΔV_{in}- lateral impulse, V_{η}- transversal velocity, j is the angle describing the direction of the vector ΔV_{in}.

Angle Δχwhose plane is perpendicular to the line of intersection of the planes are real and the set of orbits, is defined as the maximum difference of the coordinate values of the orbital points lying in the plane of the angle.

If known Δχ required quantity of impulse ΔV_{in}can be determined from the relation: ΔV_{in}=2V_{η}sin0,5Δχ. In this case, the direction of the vector ΔV_{in}relative to the plane of the initial orbit will be characterized by the angle j=90°+0,5Δχ. For circular orbits, as is known, the velocity vector at any point of the orbit is perpendicular to the radius-vector. Therefore, for a circular orbit V_{η}=V_{kr}.

For elliptical orbits of transversal speed is the magnitude of the variable.

The maximum value it has in the perigee, and the minimum in height. Therefore, the magnitude of the impulse ΔV_{in}when turning the plane of this orbit at perigee is going to be the maximum,
and apogee minimum.

If lateral force is applied in one of the nodes of the orbit, the pivot axis is coincident with the axis changes the orbital inclination. Therefore, the lateral force applied at the nodes of the orbit, will cause a change in only the orbital inclination.

The scheme is a direct reversal of the source plane of the orbit to change its inclination is shown in Fig. In one of the nodes of the orbit 30_{1}applied pulse speed ΔV_{in}, which is a reversal of the orbital plane at an angle Δχ. As a result, the inclination of the orbit 30_{2}will be different on the corner Δi=Δχ.

Consider what consequences arise for the application of a tangential impulse.

The direction of the tangential force or coincides with the direction of the velocity vector flight, or opposition to it.

Therefore, the tangential force can only change the value of the speed of flight. The direction of velocity in the tangential force does not change. The tangential force does not change the orbital inclination and right ascension of the ascending node. To determine what changes in the orbit will cause the tangential force, consider the equation V=√k[(2/r)-(1/α)]. Because k is a constant, and r at the point of the maneuver is not changed, then the maneuver increase (decrease) value / min net and will cause a change α .

Consider Fig. It shows the elliptical orbit 30_{1}with foci F_{1}and F_{2}to maneuver.

Let the execution time of the pulse maneuver the SPACECRAFT was located at the point M with the true anomaly ϑ. The distance from this point to the first focus is r, and the distance to the second focus - r'.

As you know, the ellipse is called the locus of points for which true equality r+r_{1}=2a. If the velocity of the point M to zoom in on ΔV_{t}this will cause an increase in semimajor axis on Δα_{t}.

Obviously that will increase the distance to the second focus and he from point F_{2}will move to point F_{2}'. The magnitude of this move Δr'=2Δα_{t}. It is also known that the tangent at any point of the ellipse is the bisector of the external angle formed by the radii-vectors that point. The velocity vector is directed along the tangent to the trajectory (ellipse at the point M). Therefore, the angle τ=τ'. Since after the maneuver direction of the velocity vector and the radius vector of the SPACECRAFT does not change, do not change the angles τ and τ'. This means that the direction of the second focus will remain and it will move along the radius-vector r', size 2Δα_{t}in point F_{2}'. Moving the second focus point F_{2}in point F_{2}' leads to a change misfocusing is astonia on size 2Δ
with_{t}and this means that changes the eccentricity of the orbit e (e=s/α). Moving the second focus also leads to a rotation of the line of apses on the corner Δ∈_{t}i.e. the tangential force, changing the semi-major axis of the orbit changes and the argument of perigee.

Contact Fig. If the correction of the position of the perigee to implement due to the tangential impulse is applied at one of the points of intersection of the orbit with minor radius (point D), the result will change and the value of the semimajor axis.

The latter also causes the change of the orbital period. As a result, when such optical correction will need to hold the correction value of the semimajor axis. Therefore, in those cases, when the correction of the position of perigee change in semimajor axis is unacceptable, this correction should be carried out using two tangential impulse is applied at the points D and D' (Fig). The magnitude of these pulses should be the same, but the signs are different. When double-pulse maneuver under the action of the positive pulse ΔV_{T1}applied at point D, the argument of perigee will change to Δ∈_{T1}and the semimajor axis increases by Δα_{T1}. Under the action of the negative pulse ΔV_{T2}applied at point D' intermediate orbit, the semi-major axis will be reduced by Δα_{T2}and argument perigee ∈
will change as and when the first pulse.

Note that the energy cost of performing the maneuver is not only determined by the selected orientation of the control pulse (e.g., tangential or lateral), but the point of its application in orbit.

In orbit, there are not only the point at which a pulse of force does not cause changes of individual elements of the orbit, but the point at which this force provides the biggest change for the specified element of the orbit.

Contact Fig.

In this figure indicate the points at which changes αe and ∈ under the applied tangential force have extreme values. In equation Δα_{t}=(2Δα^{2}V/k)ΔV_{t}the semimajor axis α - a constant for any point on this orbit. Variable in this equation is only flying speed, the maximum value of which is at perigee, and the minimum - in the climax. Therefore, the same in magnitude and direction the speed pulse ΔV_{t}at perigee, will cause a maximum change of semimajor axis, and apogee is minimal.

The maximum change of the eccentricity of the orbit under the action of a tangential force is at perigee and apogee of the orbit. However, it should be borne in mind that to increase the eccentricity of the orbit is tangentially pulse at perigee must be positive, and peak - negative.

The tangential force applied at the point of perigee and apogee, does not cause rotation of the line of apses. In these cases, the directions r and r' coincide with the line of the apses and the second focus will move along this line. For correcting the position of the perigee (apogee) of the orbit, when the value of the semimajor axis must be preserved, it is advisable to apply two tangential impulse is applied at the points (D, D') of the intersection of the orbit with minor radius. The results of both pulses should be the same, but their signs are different.

The necessary changes the orientation of the orbital plane is performed only under the action of lateral forces. To combine the two orbital planes of the side pulses applied to the two points on the line of intersection of the planes, which is closer to climax. When the change of right ascension of the ascending node the application of lateral forces at the points with argument equal to 90° 270°, leads to changes in the inclination of the orbit.

To change the orbital inclination of the control pulse is applied to one of the nodes of the orbit.

When comparing the parameters of the real orbit with parameters given orbit defined by the following error: Δ Ω, Δi, Δ∈, Δα and Δthat is, To address each of these mismatches in the above is the materials identified type of corrective pulse, its direction and point of application. It is obvious that first of all the real plane of the orbit must be aligned with the plane of a given orbit. Given that the correction of the longitude of the ascending node of the source orbit leads to a change in inclination i, argument of perigee ∈, the first corrective action should be to address the mismatch Δ Ω. So the true misalignment Δ∈ may be detected only after elimination Δ Ω. A second pulse must be aimed at eliminating Δi, i.e. on the final alignment of the planes of the orbits. Two of these pulse are lateral. Troubleshoot Δα and Δe is the application of tangential impulses. Correction of the position of perigee, i.e. correction ∈in that case, when the value of α must be preserved, are two tangential impulses applied at the point of intersection of the orbit with minor radius.

Equation increments of orbital elements at a particular corrective pulse presented on Fig.

Included in equations such elements of the real orbit, as α, e, i, u, and V_{η}as shown above, are determined. Value of each of the increments were observed when comparing the desired and actual orbits. The value of the velocity increment ΔV is required for the OS is enforced each type of maneuver,
is determined from the presented formulas.

The cost rate associated with the cost of fuel the engine of Tsiolkovsky formula ΔV=ω_{e}In(m_{0}/m_{0}-m_{t}), where ω_{e}effective flow rate of gases from the nozzle of the engine, m_{0}is the mass of the SPACECRAFT to maneuver, m_{t}is the mass of fuel consumed for the maneuver. Resolving it relative to the weight of fuel received

All the actions are deterministic, and therefore when the corresponding on-Board software can be implemented in stand-alone mode.

Autonomous on-Board control system of the SPACECRAFT (SC)containing connected by communication lines to the sensor Ground, at least one star sensor, a temporary device, the device memory, Executive bodies and the CPU controls the position determining angular misalignment between the axes of the AC and the external axes coordinate system and generates control signals for these Executive bodies to maintain the relative position of these coordinate systems, characterized in that the processor and Executive bodies in the form of inertial flywheels provide orbital motion of the SPACECRAFT alignment direction on the reference star, selected from the pole of the world, visa is nuclear biological chemical (NBC plane, containing the Central axis of the star sensor and the Central axis of the sensor Earth - with a focus on the center of the Earth, and in the plane of sensitivity defined by these axes defined by the specified processor offset guide star describes the change in the latitudinal angle of the orbital position of the SPACECRAFT, and the rotation in the star sensor of the navigation star around an anchor star - inertial longitude changes of the orbital position of the SPACECRAFT and inertial longitude is measured from the side of the zero reference point, which is coordinated by the angle between the plane defined by the directions with the AC on both the stars and the plane of sensitivity at the time of resetting measured on Board the SPACECRAFT right ascension of the reference stars, dependent inertial longitude is the angle between the direction from the SPACECRAFT to the reference star and the Earth's axis is recorded in the storage device, and Board time synchronized with the AC fragile time of the vernal equinox and is reset to a temporary device at the time of a complete rotation of the Earth, with the system processor angular data statistical processing of stochastic inertial longitudinal and inertial-pulse measurements, which is designed with the ability to determine the trend components for each cycle of measurement, broadcast the x to write to the storage device, forecast angular data over multiple clock steps forward for the subsequent implementation of the adjusted angular orientation of SPACECRAFT and security, taking into account the speed of rotation of the Earth and the current time, convert sophisticated inertial longitude in the longitude and converting the adjusted latitudinal angle in latitude with regard to the angle between the direction from the SPACECRAFT to the reference star and the Earth's axis, the system also introduced the processor orbital data that defines a period, semi-major axis, eccentricity, argument of perigee, inclination, and right ascension of the ascending node real orbital revolution based on the results of the statistical smoothing recorded in the storage device, and detecting by comparing these parameters with the parameters recorded in the storage device, the deviation of the actual orbit parameters from the set and on the revealed deviations determines the kind, the point of application, the pulse value and consistency of the implementation of corrective action by the propulsion system of the SPACECRAFT, the system also introduced the host PC.

**Same patents:**

FIELD: methods of control of spacecraft angular motion.

SUBSTANCE: proposed method includes measurement of spacecraft orientation parameters at flight time interval, forming control moment by means of reactive flywheels, measuring the present vectors of main and additional reactive flywheels and determination of vector of total moment of momentum of spacecraft. Belonging of to area S of its magnitudes is checked as well. Area S consists of sub-areas of main (S_{0}) and additional (S_{d}) of reactive flywheels. Otherwise "n" versions of change of signs of angular velocities of main and additional reactive flywheels are checked, re-distributing the moment of momentum among reactive flywheels and orientation parameters are checked for presence within permissible limits. If this condition is observed, control of spacecraft is continued and if this condition is not observed, respective versions are noted re-distributing moment of momentum among main and additional reactive flywheels. Relief of main flywheels, setting of directions and modes (acceleration or deceleration) of rotation of main and additional reactive flywheels and conditions of connecting them are carried out on basis of check of indicated versions of change of reactive flywheel angular velocities and belonging of moment of momentum of main and additional reactive flywheels to respective sub-areas S_{0} and S_{d}. Upon completion of flight interval, moment of momentum accumulated in additional reactive flywheels is zeroed by re-distributing it on main reactive flywheels; their moments of momentum shall belong to sub-area S_{0}. If they do not belong to this area, main reactive flywheels are subjected to relief. Such control is repeated for subsequent maintenance of spacecraft orientation at next time interval.

EFFECT: enhanced accuracy of maintenance of preset spacecraft orientation by exclusion of relay change of control moments from reactive flywheels.

8 dwg, 1 tbl

FIELD: methods and systems for control of angular motion of spacecraft.

SUBSTANCE: proposed method includes measurement of parameters of spacecraft orientation at flight interval with the aid of control moment reactive flywheels, measurement of present vectors of reactive flywheels, determination of vector of total moment of momentum of spacecraft. If belongs to area S of its magnitudes, control of spacecraft is carried out with relief of reactive flywheel, otherwise control is carried out without relief of reactive flywheel. At control of spacecraft, "n" versions of change of signs of angular velocities of reactive flywheel are checked distributing the moment of momentum among reactive flywheels and keeping the parameters of orientation within permissible ranges. If this condition is observed, control of spacecraft is continued and is these conditions are not observed, respective "k" versions of change of angular velocities are noted. Upon completion of all "n" versions of change of signs of flywheel angular velocities, change of vector is predicted at subsequent flight interval separating vector from vector of moment of momentum of reactive flywheel system. According to prediction of , versions of change of signs of flywheel angular velocity change are determined at indicated interval and are compared with previous "k" versions. According to comparison, initial conditions are found for vector for observing definite conditions for change of signs of reactive flywheel angular velocities, belonging of magnitude to area S and allowance of spacecraft orientation parameters. Duration of flight interval may be corrected. For realization of this method, control system with reactive flywheel relief unit and other units with couplings is proposed.

EFFECT: avoidance of jump-type change of control moments of reactive flywheel; enhanced accuracy of maintenance of spacecraft orientation.

3 cl, 8 dwg, 1 tbl

FIELD: control of spacecraft angular motion.

SUBSTANCE: proposed method includes measurement of total moment of momentum G in power gyroscope system and prediction of present magnitudes for performing each dynamic mode of spacecraft. At prediction, condition of belonging of vector G to definite fields of its magnitude for selecting required fields and composition of power gyroscope is checked. Many configurations of location of main and redundant power gyroscopes are fixed relative to axes of body basis. Fields of magnitudes of G are determined for each configuration. Configurations of power gyroscopes depend on spacecraft structural peculiarities and its dynamic modes. In the course of control of power gyroscope system, rational redistribution of service life of working power gyroscopes takes place at saving of service life of other systems of spacecraft. Proposed method provides for increase of fields of G magnitudes, thus making it possible to control this vector without switching-on power gyroscope jet engines.

EFFECT: enhanced controllability of spacecraft; minimization of consumption of propulsive mass of relief jet engines.

6 dwg

FIELD: engineering of devices for orientation and navigation of objects, moving in some environments, in particular, pipe-internal inspection bodies of pipeline mains.

SUBSTANCE: in proposed method, two-axial gyroscopic and additionally inserted one-component angular speed indicators are used, and also a block of three acceleration meters, which are mounted on the body of module. Firstly, module axes are combined with horizon axes and direction to geographical north, determining and recording averaged zero signals of angular speed indicators and acceleration meters block. Then for the module limited turns are set for angles of course, bank and pitch, measuring and recording signals of aforementioned indicators in rotated positions. Further, module in serially mounted in six fixed positions. As a result, angular drifting speed of angular speed indicators is determined as well as angles of imprecise setting of their measuring axes, scale coefficients of these indicators and acceleration meters block, and also zero shifts and angle shifts of imprecise setting of acceleration meters measuring axes. Also determined is non-parallelism of three measuring axes of indicators to appropriate three axes of acceleration meters block.

EFFECT: expanded set of calibrated parameters of inertial measuring module.

6 dwg

FIELD: spacecraft attitude control and mass motion systems.

SUBSTANCE: proposed method includes maintenance of spacecraft attitude by means of powered gyroscopes in the course of correction of orbit by means of jet attitude-control jet engines. Prediction is made for estimation of position of vector of spacecraft momentum of momentum in region of its magnitudes for present and final moments of correction time. Deviation of measured magnitude of this vector from predicted value is maintained within limits of preset region. In case vector does not belong to these regions center of spacecraft mass is shifted depending on configuration of spacecraft inertia tensor and control moment created by attitude-control jet engines. Upon completion of correction, spacecraft center of mass is restored. Proposed method makes it possible to form preset vector of spacecraft moment of momentum at the moment of completion of correction and to minimize number of starts of attitude-control engines in performing the flight program which is necessary for relief of load of powered gyroscopes.

EFFECT: maximum rate of correction at minimum errors of control and minimum consumption of propulsive mass of jet engines.

2 dwg

FIELD: rocket-space equipment, mainly means and methods for water supply to low-orbital spacecraft.

SUBSTANCE: the offered method provides for use of the energy of formation of the raw material, in particular, of water from the fuel components for increasing the efficiency of the means of its injection into orbit. The offered rocket power plant has a chemical reactor, in which the given product is formed, as well as a heat-exchange unit, in which the heat of the chemical reaction is transferred to the fuel components. The latter results in the growth of the power plant specific impulse. The reaction product is cooled, and a condensate (water) is obtained which is accumulated in the storage tank. The offered rocket may use one of the cleared fuel tanks for accumulation of condensate. The offered transportation system includes the offered rocket, orbital station equipped with a system of water processing to fuel components, and means of delivery of the space vehicle to the station together with the non-filled boosting unit. The offered transportation-fueling station includes also an orbital fueling complex. Space vehicles injected into high-altitude orbits, in particular, into a geostationary orbit, as well as space vehicles returning on the Earth, may be refueled there. At injection of the space vehicle into a geostationary orbit the dependence of the efficiency of injection on the latitude of the cosmodrome is essentially reduced (by 2-3 times).

EFFECT: reduced cost of supply of the orbital stations and cost of injection of the space vehicle into a geostationary orbit, as well as into other trajectories, reduced dependence of the cost of injection of the space vehicle into a geostationary orbit on the latitude of the cosmodrome.

19 cl, 3 dwg

FIELD: space engineering; transportation of payloads in extravehicular space and in atmosphere.

SUBSTANCE: proposed flying vehicle has case, piston, shock absorber with control unit, jet engine and safety stops. Jet engine is rigidly connected with piston and safety stops are rigidly connected with case which is rigidly connected in its turn with shock absorber. Shock absorber is hydraulically connected with output of its control unit. Piston has projection located behind shock absorber and two flat projections having sections of lesser thickness in the middle. Case has two recesses receiving the said flat projections; safety stops are mounted between walls of sections of lesser thickness and walls of said recesses. At acceleration of vehicle, piston and jet engine rigidly connected with it perform reciprocating motion relative to case. Reciprocating motion is ensured by shock absorber control unit which ignites the gases contained in shock absorber; these gases repel piston from case. Return motion is performed due to operation of jet engine.

EFFECT: reduced overall dimensions; enhanced wear resistance of device.

1 dwg

FIELD: rocketry and space engineering; multi-module launch vehicles of cluster arrangement at transfer of propellant among modules.

SUBSTANCE: proposed launch vehicle has central and side modules; it may be made in three versions at various modifications. According to first version, propellant for transfer is so distributed that each side two-tank module has excessive amount of only one propellant component. Each of these modules is provided with one inter-module main for transfer of propellant components. Some modifications of first version make use of displacement system for delivery of propellant component from upper tanks of side modules to lower tank of central module. According to second and third versions, use is made of one-tank side modules filled with different propellant components. In some modifications of these versions, modules are combined in clusters and are interconnected by means of simplified inter-module propellant lines. Propellant system of launch vehicle is simplified due to specialization of modules excluding dangerous processes. Efficiency is enhanced at use of kerosene-liquid oxygen propellant and engines of "closed" system employed in Russia.

EFFECT: low cost of manufacture and operation of launch vehicle; enhanced reliability.

11 cl, 18 dwg

FIELD: spacecraft inter-planetary flights with the aid of cruise jet engines, mainly electrical rocket engines.

SUBSTANCE: proposed method includes injection of spacecraft into heliocentric trajectory at distance of spacecraft from Sun followed by its approach to Sun. Active motion of spacecraft is realized in part of this trajectory behind Earth's orbit during operation of jet engines. Then, spacecraft returns to Earth at velocity increment and increases its heliocentric velocity in the course of gravitational maneuver near Earth. After spacecraft has crossed Earth's orbit in section of its approach to Sun and before entry into Earth's gravisphere, spacecraft is accelerated by repeated switching-on of cruise jet engines. At the moment of fly-by over Earth when gravitational maneuver is performed, angular motion of Earth and spacecraft relative to Sun are equalized. During fly-by over Earth, spacecraft is subjected to its gravitational field changing the vector of spacecraft heliocentric velocity, thus ensuring further acceleration of spacecraft and forming inter-planetary trajectory of flight to target.

EFFECT: reduction of time required for organization of gravitational maneuver in Earth's gravity field at injection of spacecraft into required inter-planetary trajectory.

2 dwg

FIELD: flying vehicles transported by other flying vehicles.

SUBSTANCE: rocket has first stage with first thrust forming unit and device for control of this first unit during motion over trajectory; rocket has also second stage with thrust forming unit, device for securing the first stage to aircraft and for separating it from aircraft during flight and spacecraft. Thrust forming unit of first stage makes it possible to control thrust vector at limitation of velocity head. Thrust forming unit of second stage ensures terminal control of thrust pulse for minimum deviation from preset trajectory at moment of engine shut-down. Thrust forming unit of first stage is made in form of solid propellant rocket engine provided with nozzle and drive for deflecting the nozzle for control of thrust vector. Aircraft is designed for launching the rocket with spacecraft in flight. Proposed aerospace complex is used for launching the rocket in payload trajectory; this complex has rocket, aircraft and device for securing the first stage of rocket to aircraft and for separating of rocket from aircraft in flight. Proposed method consists in launching the rocket with spacecraft in ascending module in trajectory at control of thrust vector.

EFFECT: reduction of effect of losses caused by aerodynamic drag and reduced losses of energy caused by gravitational forces.

26 cl, 8 dwg

FIELD: terminal control of motion trajectory of cryogenic stages injecting spacecraft into preset orbits by means of cruise engines.

SUBSTANCE: swivel combustion chamber of cruise engine is used for angular orientation and stabilization of cryogenic stage of spacecraft. Proposed method includes predicting parameters of motion of cryogenic stage at moment of cut-off of cruise engine; deviation of radius and radial velocity from preset magnitudes are determined; angle of pitch and rate of pitch are corrected and program of orientation of thrust vector for subsequent interval of terminal control is determined. By projections of measured phantom accelerations, angle of actual orientation of cruise engine thrust vector and misalignment between actual and programmed thrust orientation angles are determined. This misalignment is subjected to non-linear filtration, non-linear conversion and integration. Program of orientation of cryogenic stage is determined as difference between programmed thrust orientation angle and signal received after integration. Proposed method provides for compensation for action of deviation of cruise engine thrust vector relative to longitudinal axis of cryogenic stage on motion trajectory.

EFFECT: enhanced accuracy of forming preset orbit.

5 dwg, 1 tbl

FIELD: aero-space engineering; flights in atmosphere and in space.

SUBSTANCE: proposed flying vehicle has casing, gas shock absorber of combustible fuel with control unit, cylinder and piston, two exhaust pipes and two resilient safety stops for piston which is engageable with shock absorber and performs reciprocating motion inside cylinder. Gas shock absorber is communicated with control unit for metered delivery of fuel to shock absorber. Mounted inside cylinder is cylindrical strut rigidly connected with rear wall of cylinder and with two said safety stops. Piston is provided with projection inside strut and additional shock absorber of fuel being ignited is provided inside strut behind piston projection. Control unit may be used for control of two shock absorbers due to alternating metered delivery of fuel. Provision is also made for two bent exhaust pipes rigidly connected with cylindrical strut, exhaust nozzle behind gas shock absorber which is rigidly connected with rear wall of cylinder and two booster jet engines rigidly connected with casing. Additional gas shock absorber and exhaust nozzle perform functions of jet engine (which was earlier mounted on piston) and this engine may be excluded.

EFFECT: enhanced reliability and durability; reduced mass of flying vehicle.

1 swg

FIELD: aero-space engineering; flights in atmosphere and space.

SUBSTANCE: proposed flying vehicle has casing, gas shock absorber of combustible fuel with control unit, cylinder and piston, two exhaust pipes and two resilient protective stops for piston. Piston is engageable with shock absorber and performs reciprocating motion inside cylinder. Gas shock absorber is connected with control unit for metered delivery of fuel. Cylindrical strut mounted inside cylinder is rigidly connected with its rear wall and with two protective stops. Piston is provided with projection located inside cylindrical strut. Bent exhaust pipes are also located in this strut. Provision is made for mechanical shock absorber which is connected with casing and is located inside cylinder in front of piston and exhaust nozzle mounted behind gas shock absorber and rigidly connected with rear wall of cylinder. Gas shock absorber and exhaust nozzle perform function of jet engine (which was mounted on piston in previous construction), therefore there is no need in jet engine.

EFFECT: simplified construction; reduced mass of flying vehicle.

1 dwg

FIELD: spacecraft engine plants, mainly geostationary communication satellites; orientation of satellite and correction of its orbit.

SUBSTANCE: proposed module of spacecraft engine plant includes low-thrust engine (or engine package) mounted in biaxial suspension on rod and components of engine plant pneumatic system. Engine is mounted on one end of rod and pneumatic system components are mounted on other end of rod. Articulation securing the rod to spacecraft case is located between engine (or engine package) and pneumatic system components. Biaxial suspension makes it possible to perform independent rotation of engine (or engine package) relative to rod at two perpendicular axes through 90° around each axis. Rod securing articulation makes it possible to rotate the rod through 360° relative to spacecraft case. Provision is made for shaping correcting pulse during operation of low-thrust engine (or engine package) in any direction for maintenance of point of sight of spacecraft on orbit and creating required control moment around center of mass of spacecraft for its orientation.

EFFECT: enhanced efficiency.

4 cl, 2 dwg

FIELD: cosmonautics; approach and tethering in the course of docking.

SUBSTANCE: proposed system includes angle and angular velocity output units, first and second adders, relay amplifier with dead zone and actuating members for control of motion of center of mass and around center of mass of spacecraft. Output of angle read-out unit is connected to first input of first adder whose second input is connected to output of angular velocity read-out unit and output of first adder is connected to relay amplifier. System contains also units for starting the engines of motion of center of mass, test starting of actuating organ of control of motion of center of mass, units for calculation of increment of angular velocity and duration of advance connection of actuating organ of control of motion around center of mass, as well as first and second switching units. Output of test connection unit is connected with actuating member of control of motion of center of mass and first input of first switching unit. Output of unit of connection of engines of motion of center of mass is connected with actuating member of control of motion of center of mass and first input of second switching unit. Output of angular velocity read-out unit is connected with first input of angular velocity increment calculation unit whose output is connected with second input of second switching unit through advance connection duration calculation unit. Second input of angular velocity increment calculation unit is connected to output of test connection unit. Output of second switching unit is connected with second input of second adder whose output is connected with second input of first switching unit. Output of first switching unit is connected with actuating member of control of motion around center of mass and output of relay amplifier is connected with input of second input of second adder.

EFFECT: enhanced accuracy of stabilization of spacecraft; improved quality of transients at presence of external disturbing moment.

1 dwg

FIELD: rocketry, applicable at an air start, mainly of ballistic missiles with liquid-propellant rocket engines.

SUBSTANCE: the method consists in separation of the missile with a payload from the carrier aeroplane and its transition to the state with initial angular parameters of motion in the vertical plane. After separation the missile is turned with the aid of its cruise engine, preliminarily using the parachute system for missile stabilization. The parachute system makes it possible to reduce the duration of the launching leg and the losses in the motion parameters (and the energy) in this leg. To reduce the missile angular bank declination, the strand of the parachute system fastened in the area of the missile nose cone is rehooked. To reduce the time of missile turning towards the vertical before the launcher, the cruise engine controls are preliminarily deflected to the preset angles and rigidly fixed. By the beginning of missile control in the trajectory of injection this fixation is removed. In the other modification the missile turning is accomplished by an additional jet engine installation. It is started depending on the current angular parameters of missile motion so that by the beginning of controlled motion in the trajectory of injection the missile would have the preset initial angular parameters of motion.

EFFECT: enhanced mass of payload injected to the orbit.

4 cl