Method of control of cluster of satellites in geostationary orbit (versions)
FIELD: control of group of satellites in one and the same orbit or in crossing longitude and latitude ranges of geostationary orbit.
SUBSTANCE: proposed method consists in measurement of parameters of satellite orbits, determination of orbital elements, comparison of them with required ones and performing of correcting maneuvers with the aid of thrusters. Satellite inclination vectors are brought to circular areas of their permissible change which are spaced apart so that angle between line connecting the end of vector with center of its circular area and direction to Sun should exceed right ascension of Sun by 180°. According to first version, vectors of satellite eccentricity are shifted to similar circular areas so that similar line lags behind direction to the Sun by half angular displacement of vector over circumference of its natural drift within circular area. Then, distances between satellites are changed within required limits compensating for quasi-secular increment of inclination vector and correcting eccentricity vector so that at passing the center of interval between point of circumference entry of its natural drift to its circular area and point of exit from this area, line connecting the center of this circumference and center of circular area coincide with direction to the Sun. In case circular area of permissible change of each eccentricity vector is close to circumference of its natural drift (second version), said line for this area is matched with direction to the Sun and no correction is made in this case.
EFFECT: saving of propellant for correction; protracted flight of satellites at safe distance.
3 cl, 13 dwg
The invention relates to the field of space, and in particular to methods of control groups briscoletti artificial Earth satellites, and more precisely, relates to a control group of satellites placed in geostationary orbit (GSO) in the same or intersecting longitudinal and latitudinal ranges.
The need for joint management of satellites in GSO is caused, at least one of the following problems:
the deployment of new satellites in GSO with a limited number of possible orbital positions;
placing satellites one group in the form of a virtual link object, so that all the satellites are moving inside the directional antenna systems ground segment.
The solution of these problems are issues of keeping satellites in orbital position on geographical latitude and longitude, in combination with the requirements of the safe relative motion. In addition, in the case of merging satellites, from the point of view of communication, in a single object the distance between the satellites should not exceed some maximum value.
There are ways to control clusters of satellites in geostationary orbits (U.S. patent No. 5120007, CL 244/158. R; No. 5506780, CL 364/459; H. Watanabe et al., "Results of operational simulation about orbital control model for plural satellites", Proceedings of the First European Conf. on Space Debris,Darmstadt, Germany, 5-7 April 1993, ESA SD-01, pp.657-662; T. Ono, Montenbruck O. et al., "A study of new colocation control system", AIAA Paper, No. 98-1350, 1998, 10 PP.), based on the diversity of the relative vectors moods and eccentricities:
where Ej, Ijthe vectors of the eccentricity and inclination of the j-th satellite; fj, ijthat Ωjand ωjeccentricity, orbital inclination, longitude of the ascending node and the argument of pericenter; ΔEjkand ΔIjk- relative vectors eccentricity and inclination between j and k satellites. Geometrically diversity of these vectors means that the line of nodes of intersection of the planes of the orbits does not coincide with the line of equal altitude of the satellites. Thus, whatever the position of the satellites in these orbits, the relative distance between them will always be no less than some limit value defined by the vectors ΔEjkand ΔIjk. Features of orbital evolution in these methods, usually are not taken into account.
Known way to control satellites in GSO without explode their orbital elements (Chao C.C., Peterson G.E., Menn M.D. "Formationkeeping strategies for GEO halo collocation", AIAA Paper, No. 2000-4136, 2000, 10 PP.), consisting in the organization of flight satellites on the same orbit with the shift of the m their positions along the orbit (mean anomaly).
There is a method of management of shared safe flight of two geostationary satellites (Pattinson L. "EUTELSAT satellite colocation", AIAA Paper, No. 96-1187, 1996, 9 pp.), selected for the prototype. The method of controlling the cluster located in a geostationary orbit satellites, placed in the same or intersecting longitudinal and latitudinal ranges, consisting in the measurement of the orbital parameters, defining for them the current values of the orbital elements, comparing them with the required and on the basis of this maneuver with the help of Executive bodies in the form of chemical high-thrust engines for the correction of orbital inclination, eccentricity and orbital period, is also in the diversity of the relative vectors inclination and eccentricity, but this takes into account features of the orbital evolution of the eccentricity vector is a long - term evolution of the locus of the vector eccentricity, describing a circle in the annual motion.
However, this method requires a consistent carrying out maneuvers on all the satellites in the same day, which significantly complicates the management of all satellites. Features of the orbital evolution of the vector inclination in the prototype are not taken into account, which leads to additional fuel consumption.
The basis of the invention is the development of cost-saving top is willow and does not require a concerted conduct maneuvers on the satellites of the way to long-term management cluster satellites in the geostationary orbit, ensuring their flight on the relative distances, varying within certain limits, including joint safe flight.
The technical result of the invention is achieved in that in the method of controlling the cluster located in a geostationary orbit satellites, which consists in the measurement of the orbital parameters of each satellite, determining them to the current values of the orbital elements for each satellite, comparing them with the required and conduct on the basis of this maneuver with the assistance of the Executive authorities for the correction of orbital inclination, eccentricity and orbital period, in contrast to the known according to the first variant the execution of the maneuvers on each of the satellites is performed with the help of Executive bodies in the form of thrusters, translating vector inclination in spaced relation to one another annular region so that the angle between the line connecting the current position of the end of each vector from the center of the annular region, and the Sun was equal increased by 180 degrees the value of the right ascension of the Sun, while conducting maneuvers translating vector eccentricity of the satellites in spaced relation to each annular area so that the line connecting the current position of the end of each vector with the center sootvetstvuyuschemu annular area, lagged from the direction to the Sun at half the angular distance when the motion vector of eccentricity around the circumference of natural drift within the annular area, then the entire flight produce a change in the relative distance between the satellites within the required limits at the expense of compensation quasivector increment vector indicative of each satellite in combination with correction of the eccentricity vector, in which at the moment of passing the vector eccentricity of the middle of the interval between the entry point of the circumference of natural drift in the annular region of permissible variation of the eccentricity vector and the exit point of the line connecting the center of the circle natural drift and the center of the annular region of permissible variation of the eccentricity vector, coincided with the direction to the Sun, thereby leading to the constancy of the relative vectors inclination and eccentricity.
The technical result of the invention is achieved in that in the method of controlling the cluster located in a geostationary orbit satellites, which consists in the measurement of the orbital parameters of each satellite, determining them to the current values of the orbital elements for each satellite, comparing them with the required and on the basis of this maneuvering through the Executive is lanow for correcting the inclination of the orbit, eccentricity and orbital period in contrast to the known according to the second variant the execution of the maneuvers on each of the satellites is performed with the help of Executive bodies in the form of thrusters, translating vector inclination in spaced relation to one another annular region so that the angle between the line connecting the current position of the end of each vector from the center of the annular region, and the Sun was equal increased by 180 degrees, the value of the right ascension of the Sun, while they conduct maneuvers translating vector eccentricity of the satellites in spaced relation to each annular area so that the line connecting the current position of the end of each vector from the center of the annular region, coincides with the direction of the Sun, then the entire flight produce a change in the relative distance between the satellites within the required limits at the expense of compensation quasivector increment vector inclination for each satellite without correction of the eccentricity vector, thereby leading to the constancy of the relative vectors inclination and eccentricity.
The second option unlike the first involves the absence of maneuvers that alter the eccentricity vector. In this case, the size of the annular area say the th vector of eccentricity of each machine should be close to the average diameter natural drift. By choosing the initial conditions, is directly dependent on the position of the Sun, we can ensure that the motion vector of eccentricity will be carried out within the annular area it acceptable changes without conducting further additional maneuvers. The parameters and the position of the annular region of allowable change vectors eccentricity and inclination are selected on the basis of latitudinal and longitudinal ranges and characteristics of the satellites. The feasibility of using thrusters, compared with chemical engines, due to the high specific impulse (14000-25000 m/s), which leads to significant fuel savings and thus to reduce the mass of the satellite.
The invention is further explained in the description of specific variants of its implementation and the accompanying drawings, and graphs, on which:
figure 1 depicts the change of the vector inclination at the annual interval during uncontrolled movement;
figure 2 - change vector inclination at the annual interval during controlled movement;
figure 3 - change vectors inclinations of two satellites;
figure 4 - change vectors of the eccentricities of the two satellites;
5 is a diagram of the correction vectors eccentricities;
6 is a diagram of a synchronous change in the relative vectors of the eccentricity and inclinations;
Fig.7 - the selection of the initial vector direction of eccentricity;
Fig - the evolution of the eccentricity vectors for the two satellites at the annual interval;
figure 9 - relative vectors eccentricities at the annual interval;
figure 10 - evolution of vectors indicative for the two satellites at the annual interval;
11 - relative vectors indicative annual interval;
Fig change the relative distances between the satellites at the annual interval;
Fig - phase portrait of the motion of two satellites (the drift velocity in longitude from the values of the average geographical longitude).
While figure 1-13 the following notation:
e* is the radius of the circumference of the free evolution of the vector of eccentricity;
E, E1E2the eccentricity vector and its components;
- the direction vector to the Sun;
I, I1, I2- vector inclination and its components;
ΔE, ΔI - differential vectors eccentricity and inclination;
S is the radius of the free evolution of the vector of eccentricity;
t - time;
ϕ - the angular displacement of the Sun for half of the period of correction of eccentricity;
Θ - the relative angle between the vectors ΔE and ΔI.
The problem is solved by carrying out the maneuvers performed by the thrusters and causing such movement of the centers of mass of the satellites, h is about the evolution of vectors inclination and eccentricity on hodographs are synchronized, and mutual orientation relative vectors eccentricity and inclination remains almost constant. To secure the joint flight of the cluster satellites in the geostationary orbit significant is not the preservation of the absolute values of the vectors of the eccentricities and inclinations (i.e. maintaining the shape and position of the orbit of each satellite), while maintaining constant relative vectors inclination and eccentricity. This can be achieved by keeping the absolute values of the orbital elements of the satellites within narrow limits and, as a consequence of their relative vectors. However valid, if the orbital elements of the vector of eccentricity and inclination of all the satellites have such simultaneous evolution in a long-term scale, in which the constancy of the relative vectors is preserved.
The nature of the perturbed motion of geostationary satellites and management methods single satellites are well known. The most significant perturbation of the orbital elements is kazibekova change of mood, which is ˜0.8-1.0 deg/year (figure 1). About 90% of the cost of fuel to maintain orbital position is connected with the necessity of correction of this perturbation. The magnitude of the required characteristic rate per year is ˜45-55 m/s Correction, etc) the RA inclination is provided by conducting maneuvers on one or two intervals on the circuit. The middle of these intervals are located on the line of nodes of the target and source planes of the orbits of the satellite. The direction of the total working pulse correction orthogonal to the orbit plane and the antipodal sites round (when two intervals maneuvering) the opposite.
There are three types of corrections hold vector inclination in a given latitudinal limits (listed in order of increasing cost (fuel):
correction only kazibekova component of the displacement vector inclination, i.e. average I (accounting averaged perturbation with a period of up to one year (inclusive);
partial correction compensated cyclic perturbations along the direction of the secular displacement vector inclination, i.e. the correction projection quasiaveragyes I in the direction of medium I (annual and semi-annual disturbances);
the maximum correction, i.e. correction of the true I.
The most effective method of correcting the inclination is correction, compensating only the component of the vector inclination parallel to the direction of secular care. While fuel economy will be ˜10-25%. In this case, the vector inclination will be cyclic with semi-annual period and amplitude ˜0.025 deg. (figure 2).
For the vector eccentricity of the following types of adjustments:
the retention module of the vector e is centricitee within the specified range;
the transition at a given sighting vector eccentricity;
forced drift vector eccentricity of a circle of given radius (i.e. in the annular region).
Average fuel consumption on these correction is equivalent to ˜2 m/s is the characteristic velocity for the year. To control the position of the satellite longitude uses the traditional method of formation of the phase curve (see Cherniavsky G.M., Bartenev, VA Management orbit geostationary Earth", M - mechanical, 1978) and the longitude correction is performed in conjunction with the correction of eccentricity (A.Kamel and .Wagner, "On the Orbital Eccentricity Control of Synchronous Satellites", Journal of the Astronautical Sciences, Vol.XXX, 1982, No. 1, pp.61-73).
Currently, for the management of shared safe flight of two or more satellites in the vicinity of one operating point geostationary orbit using the diversity vectors inclinations and eccentricities adjacent satellites and maintaining these differences within certain limits by the choice of the scheme carrying out maneuvers. Physically, this means that near the line of nodes of intersection of the orbits of the satellites carry them in height, and near the line of equal altitude carry them in the side with respect to the planes of the orbits of the direction that eliminates the dangerous proximity of the satellites in relative motion. In addition, the satellites must be held in a slave is whose point defined in the tactical-technical task accuracy in latitude and longitude.
The traditional way of implementing control the joint motion of two satellites in the vicinity of one point on the geostationary orbit is a periodic correction of the true vector inclination or partial correction of this vector with compensation of periodic perturbations along the direction of the secular displacement vector inclination. Correction of the eccentricity vector is held so that its locus was represented by a circle with a one-year period traffic. On satellites with relatively large thrust maneuvers to change the inclination (which about 90% of all fuel costs on hold at the working point of geostationary orbit) are held at intervals of 2 to 6 weeks of the characteristic speed of the manoeuvre ˜2-10 m/s the vector inclination receives a significant displacement and therefore this procedure should be carried out consistently over time for all satellites to reduce the relative rotation vector inclination. As mentioned above, the most economical way of correction vector inclination is to eliminate the secular component, which leads to residual evolution vector inclination (see figure 2), mainly close to the circle of radius ˜0.025(with semi-annual period. The size of this circle is comparable to the size of the field holding the satellite in rabotaet longitude and latitude (usually it ±0.05°0.1...°). Therefore, such a correction (i.e. the elimination of the only secular perturbations) in the joint management of the satellites, as a rule, do not use, because inconsistent with the provisions of satellites in part semi-disturbing members vectors indicative of their difference can be ˜0.05°. However, the special choice of the initial positions of the vectors inclination satellites and at low frequency for correction of inclination (i.e. satellites with relatively low thrust) can be provided by correcting only the secular part (see figure 3) phase evolution of vectors indicative in circles of radius ˜0.025 so that their vector difference will save close to a constant direction. When conducting maneuvers correction inclination because of their small size relative vector inclination will not be rotated. In addition, from a practical point of view, very important is the lack of necessary coordination of the torque correction on the satellites.
Scheme selection and correction of the eccentricity vectors shown in figure 4. Due to the offset of the centers of the circles of forced motion (radius e*) in the desired direction relative to the vector difference of inclinations (ΔI) and by maneuvering the thrusters with panago movement in these circles, line of equal elevation will be almost stationary. If the radius of e* is equal to the radius S of the circle corresponding to the free evolution of the eccentricity vector, caused by the pressure of sunlight, to stabilize the direction of this line correction is not required, since the common-mode motion on any identical trajectories ensures the stability of the direction of the initial exhibition. If e*<S, which is possible due to the limitations of the magnitude of the eccentricity, you will need special correction of the eccentricity vector and the corresponding costs of the working fluid.
For carrying out correction of the eccentricity can be accepted by the scheme shown in figure 5. According to it every maneuver puts the vector eccentricity in the position from which the natural drift (due to the pressure of sunlight) the locus of the vector will touch the circumference of his forced changes approximately in the middle of the planned time interval passive (eccentricity) of the flight. Thus, the locus of the vector eccentricity will be inside the ring area.
Scheme of synchronous changes in relative vectors eccentricity and inclination on the example of a cluster of three satellites is shown in Fig.6. It is seen that in the case of a movement locus of vectors in circles the relative angles between vectors Ɗ E and ΔI remain almost constant and approximately equal to a preset angle Θ. To ensure that such simultaneous evolution of vectors eccentricity and inclination at the start point, they should be formed in such a way that the line connecting the center of the circle of eccentricity vector with initial pointwas aimed at the Sun, and the line of relative vector inclination formed by the vector of eccentricity of the specified angle.
As an example we choose the following values: coordinates of the centers of the circumferences of the evolution of the locus vectors eccentricities 1 satellite (0,2·10-4), 2 satellites (0, -2·10-4), e*=1.5·10-4the coordinates of the centers of the circumferences of the evolution of the locus of vectors indicative 1 satellite (0, 0.045°), 2 satellites (0, 0.045°). Accepted θ=0, and the frequency correction of the eccentricity in accordance with the practice adopted is 25 days, which corresponds to the offset position of the Sun ˜25°. Thus ϕ˜12.5°. On Fig-13 presents the results of modelling the long-term mission of two geostationary satellites in the same longitudinal and latitudinal ranges ±0.1 deg. at the annual interval: change absolute vectors eccentricities (Fig) relative to the vectors of eccentricities (Fig.9), the absolute age of the Directors inclination (figure 10), relative vector eccentricity inclination (11). The change in the relative distance between the satellites is shown in Fig, and their average geographic longitude - Fig. The relative distance between the satellites changed within 18-110 km Mutual orientation relative vectors eccentricity and inclination is maintained with an accuracy of ±8-12°i.e. no worse than the value of the angle ϕ. Flight simulation satellites was carried out in the exact setting, taking into account decentralist the Earth's gravitational field, the gravitational attraction of the Sun and moon, the pressure of sunlight. Also take into account errors in the orbit determination and execution of maneuvers. The simulation results confirm the technical feasibility of the proposed method of control by the cluster satellites.
1. The method of controlling the cluster located in a geostationary orbit satellites, which consists in the measurement of the orbital parameters of each satellite, determining them to the current values of the orbital elements for each satellite, comparing them with the required and conducting with the assistance of the Executive bodies maneuvers correction of the orbital period, inclination and eccentricity of the orbit, characterized in that the maneuvers on each of the satellites is carried out with the assistance of the Executive bodies in the form of thrusters, translating vectors inclination satellites in R is znesennya relative to each other, the annular area of their allowable changes so to the angle between the line connecting the current position of the end of each vector from the center of the annular region, and the Sun was equal increased to 180° the value of the right ascension of the Sun, while conducting maneuvers, which moves the eccentricity vectors of the satellites in spaced relation to each other, the annular area of their allowable changes so that the line connecting the current position of the end of each vector from the center of the circular area behind the direction to the Sun at half the angular distance when the motion vector of eccentricity around the circumference of natural drift within the annular area, then the entire flight produce the change in the relative distance between the satellites within the required limits at the expense of compensation quasivector increment vector indicative of each satellite in combination with correction of the eccentricity vector, in which at the moment of passing the vector eccentricity of the middle of the interval between the entry point of the circumference of natural drift in the annular region of permissible variation of the eccentricity vector and the exit point of the line connecting the center of the circle natural drift and the center of the annular region of admissible vector eccentric is the Humanities, coincides with the direction of the Sun, thereby leading to the constancy of the relative vectors inclination and eccentricity between satellites.
2. The method of controlling the cluster located in a geostationary orbit satellites, which consists in the measurement of the orbital parameters of each satellite, determining them to the current values of the orbital elements for each satellite, comparing them with the required and conducting with the assistance of the Executive bodies maneuvers correction of the orbital period, inclination and eccentricity of the orbit, characterized in that the maneuvers on each of the satellites is carried out with the assistance of the Executive bodies in the form of thrusters, translating vectors inclination satellites in spaced relation to each other, the annular area of their allowable changes so that the angle between the line connecting the current position of the end of each vector from the center of the circular region and the Sun was equal increased to 180° the value of the right ascension of the Sun, while they conduct maneuvers, which moves the eccentricity vectors of the satellites in spaced relation to each other, the annular area of their allowable changes so that the line connecting the current position of the end of each vector from the center of the annular region, coincides with the direction of the Sun, then the and the entire flight produce a change in the relative distance between the satellites within the required limits at the expense of compensation quasivector increment vector indicative of each satellite without correction vector eccentricity, thereby leading to the constancy of the relative vectors inclination and eccentricity between satellites.
FIELD: spacecraft power supply systems on base of solar batteries.
SUBSTANCE: proposed spacecraft has form of right-angle prism with cross-section in form of equilateral tetragon (rhomb). Mounted on side faces of prism are solar battery panels. Spacecraft is provided with passive or combined system of gravitational stabilization in orbit. Acute angle of tetragon ranges from 50 to 90° to ensure required power supply for spacecraft equipment. Main central axes of symmetry of spacecraft in transversal plane are parallel to tetragon diagonal. Lesser axis is parallel to larger diagonal, thus enhancing stable gravitational orientation of spacecraft by larger diagonal perpendicularly to orbit axis.
EFFECT: enhanced efficiency.
FIELD: spacecraft for interplanetary flights, research and development of celestial bodies.
SUBSTANCE: proposed spacecraft is equipped with solar sail, central fixed module and movable module which is coaxial relative to first module and is provided with bio-energy complex. Laid spirally on surface of movable module are growth tubes with plant conveyers which ensure turn of movable module around central axis. Connected with modules are generator and electric power accumulator. Fixed module is provided with cylindrical separable ice melting modules. Each module is provided with parachute for descent on planet, its own bio-energy complex and ice melting chamber for forming shaft in ice cover of planet. Ring of reactors located around central axis of module are combined with toruses. On side of central axis reactors are coated with warmth-keeping jacket and are provided with heaters and units for filling the reactors with water in lower part and with oil in upper part. These units ensure operation of hydraulic generator generating vapor for melting ice and supplying distilled water to bio-energy complex. Modules are provided with envelope pressurizing units, deploying their parachutes and supplying sea water from shafts to envelope surfaces for forming ice domes. When domes are combined, stations may be formed for research of planet followed by its populating. Modules are equipped with descent bathyspheres for research of under-ice ocean and robots for performing jobs on planet surface. Spacecraft may include manned separable raiders and bathyscaphs for research of ocean depth. Both of them may be provided with their own bio-energy complexes.
EFFECT: extended boundaries of research and development of far celestial bodies, mainly planets and satellites with thick ice cover.
3 cl, 12 dwg
FIELD: spacecraft equipment deployed in orbit from transportation position to working position.
SUBSTANCE: proposed complex contains components (1.1-1.n) rigidly connected with side (3) of soft inflatable mat (4). In transportation position of components, mat (4) is in deflated state and is folded in such way that components of complex are located on both sides of fold (5.1-5.n-1) of mat in pairs.
EFFECT: simplified construction of complex; enhanced reliability of deploying the complex in working position.
10 cl, 5 dwg
FIELD: spacecraft engine systems; construction of solar sails.
SUBSTANCE: proposed craft has hull, main and additional circular reflecting surfaces, units for forming such surfaces provided with twisting devices and control units for orientation of these surfaces. Orientation control units are made on base of gimbal mounts brought-out beyond craft hull. Each twisting device is made in form of hoop mounted on outer frame of gimbal mount for free rotation; it is engageable with electric motor. Units for forming reflecting surfaces are made in form of pneumatic systems with concentric pneumatic chambers and radial struts. Said struts are provided with flexible tubes with valves mounted at equal distances. Valves have holes. Built on said tubes are pneumatic cells in form of torus or spheres. Each pneumatic system is mounted on respective hoop and is communicated with compressed gas source through concentric hermetic groove found in hoop and in outer frame of gimbal mount.
EFFECT: reduced responsiveness of solar sail control; possibility of deploying solar sail according to preset program; reduction of mass per unit of surface.
4 cl, 11 dwg
FIELD: space engineering; devices and methods of maneuvering of spacecraft with the aid of solar sail.
SUBSTANCE: proposed method includes forming light-sensitive surface of solar sail and orientation of this surface in solar radiant flux. This surface is formed as cloud of finely-dispersed particles charged by solar photoelectrolizing. Stable shape close to sloping surface is imparted to cloud by means of electrostatic system of spacecraft. This system has at least one central and one concentric charge carriers of opposite signs. Control of shape and sizes of cloud may be performed by screening central charge or moving it relative to circular charge.
EFFECT: facilitated procedure; low expenses for deployment and control of spacecraft.
9 cl, 5 dwg
FIELD: space engineering; spacecraft power supply systems.
SUBSTANCE: proposed solar battery includes panels foldable by "bellows" pattern and frame with drive mechanism. Panels are interconnected together and are connected with spacecraft through frame by means of drive springs and cable run with pulleys. Provision is made for articulated rods of adjustable length and locking units for locking the solar battery in folded and open positions and several contact components (pins, seats, pushers) for interaction of solar battery panels. Locking units are made in form of stops with elliptical holes and spring-loaded retainers. According to first version, solar battery includes fixed pulleys on spacecraft, intermediate pulleys on first panel of solar battery, brace movable relative to spacecraft, strut movably connected with frame and with brace (through spring) and stop engageable with brace. According to other version, drive mechanism is provided with engine and pulley connected with intermediate pulley by means of cable run. Engine and pulley are secured on spacecraft by means of bracket with elliptical locking holes. Movable unit of engine is fastened with frame and frame is provided with spring-loaded pins locked in locking holes of bracket in opening the solar battery.
EFFECT: enhanced reliability.
3 cl, 43 dwg
FIELD: spacecraft power systems using solar batteries and electric jet plasma engines, mainly stationary engines.
SUBSTANCE: proposed method includes stabilization and change of power of power plant through regulation of consumption of engine working medium. When power of solar battery drops to level of maximum permissible power consumed by engine, consumption of working medium is changed in such way that power of solar battery might change in saw-tooth pattern and vertices of saw might be in contact with line of maximum probable power of solar battery. Device proposed for realization of this method includes matching voltage converter whose outputs are connected with engine electrodes and inputs are connected with solar battery busbars, current and voltage sensors showing solar battery voltage and power sensor connected with current and voltage sensors. Comparator connected with power sensor is also connected with controllable power setter and initial power setter. Outputs of controllable power setter are connected with comparator and comparison circuit whose input is connected with power sensor output. Output of comparison circuit is connected with amplifier-regulator of consumption of working medium.
EFFECT: enhanced reliability; simplified construction and facilitated procedure of regulation of power.
3 cl, 2 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: cosmonautics, applicable in space activity - space exploration, exploration of the solar system, observation of the Earth from the space, at which it is necessary to determine the space co-ordinates of the space vehicles and the components of their flight velocity vectors.
SUBSTANCE: the method consists in the fact that in the intermediate orbit simultaneously with determination of the co-ordinates of the space vehicle (SV) at initial time moment t0 by signals of the Global Satellite Navigation Systems the determination and detection of radiations at least of three pulsars is carried out, and then in the process of further motion of the space vehicle determination of the increment of full phase ΔФp=Δϕp+2·π·Np of periodic radiation of each pulsar is effected, the measurement of the signal phase of pulsar Δϕp is determined relative to the phase of the high-stability frequency standard of the space vehicle, and the resolution of phase ambiguity Np is effected by count of sudden changes by 2·π of the measured phase during flight of the space vehicle - Δt=t-t0; according to the performed measurements determined are the distances covered by the space vehicle during time Δt in the direction to each pulsar and the position of the space vehicle in the Cartesian coordinate system for the case when the number of pulsars equals three is determined from expression where Dp - the distance that is covered by the space vehicle in the direction to the p-th pulsar; Δt - the value of the difference of the phases between the signal of the p-th pulsar and the frequency standard of the space vehicle, measured at moment Tp - quantity of full periods of variation of the signal phase of the p-th pulsar during time Δϕp; Np - column vector of the position of the space vehicle at moment Δt; - column vector of the space vehicle position at initial moment t0; -column vector of estimates of space vehicle motions in the direction cosines determining the angular position of three pulsars.
EFFECT: provided high-accuracy determination of the space vehicle position practically at any distance from the Earth.
FIELD: space engineering; on-board terminal control facilities of cryogenic stages with non-controllable cruise engines.
SUBSTANCE: parameters of motion of cryogenic stage at moment of cruise engine cutoff are predicted and radius of deviation of radius and radial velocity of cryogenic stage from their preset magnitudes are determined. Signals for correction of pitch angle and rate of pitch are shaped for compensation of said deviation. Pitch angle correction signal is limited at preset level and its excess above this level is determined. When signal is shaped for limitation, addition to correction pitch rate signal is formed. This correction is equal to product of said excess by ratio of functions of sensitivity of radial velocity of cryogenic stage to pitch angle and rate of pitch. Resultant pitch rate correction signal is formed as sum of this signal determined without taking into account pitch angle correction signal limitation and addition. Thus, priority follow-up of velocity error is ensured at limited pitch angle correction.
EFFECT: enhanced accuracy of forming preset orbit due to reduction of disturbance level on angular stabilization loop.
9 dwg, 1 tbl
FIELD: space engineering; designing spacecraft motion control systems.
SUBSTANCE: proposed method is performed by information of orientation unit to Sun by introducing the orbit parameters into on-board computer followed by calculating the Sun position in observation field of orientation unit for each point of orbit for orientation of axes in orbital coordinate system; search angular velocity is set for spacecraft to ensure capture of Sun by observation field of orientation unit, after which angular velocity is decreased to zero ensuring position of Sun in observation field of orientation unit. Then spacecraft is turned in such way that Sun should move to required initial point; turning the spacecraft to preset points is continued for each orbital point.
EFFECT: reduced mass; simplified construction of spacecraft due to reduced number of instruments and units; extended field of application.
FIELD: global satellite information systems.
SUBSTANCE: proposed system and method of organization of communication with the aid of this system includes injection of satellites into inclined elliptical orbits ensuring simplified tracking of satellites by means of ground tracking stations. Satellite orbits form pair of repeated routes (130, 140) embracing the earth's globe in projection on ground surface. Satellites are activated on each of these routes only on active arcs located considerably higher or lower relative to equator, thus emulating some essential characteristics of geostationary satellites. Parameters of satellite orbits are so set that final points of active arcs of two routes coincide; point at which active arc terminates in one route coincides with point where active arc starts on other route. Satellites placed on such active arcs are accepted by ground station located in satellite servicing zone as satellites slowly moving in one direction at rather large elevation angle. Their trajectory in celestial sphere has shape of closed teardrop line.
EFFECT: increased capacity of global satellite communication system with no interference in operation of geostationary satellites; simplifies procedure of tracking satellites.
39 cl, 15 dwg, 1 tbl