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Power supply and control method for spacecraft correction system Invention is related to space engineering and may be used for spacecraft correction by means of electric propulsion plasma engines (EPPE). EPPE are selected for switching on, required time of EPPE operation is selected, used and unused electrodes of the engines are selected and connected to the power supply sources by means of contactors, the power supply sources are switched on and off for EPPE start-up and operation within the required period, high-ohmic resistive network if formed for discharge of electric charge from EPPE electrodes to the spacecraft frame, two modes are formed for commutation of the engine electric circuits, the main and reserve power supply sources with capacitive filters are connected to electrodes of non-operating EPPE, electric circuits of the selected EPPE remain connected to the used power supply sources, electric circuits of the other EPPE are switched off from the used power supply sources and still left connected to the unused power supply sources, the used power supply sources are switched on and off in compliance with the defined algorithm. |
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Method for orientation of artificial earth satellite Invention relates to controlling orientation of artificial earth satellites with solar panels. An artificial earth satellite (3) further includes a self-contained circuit for controlling orientation of the artificial earth satellite relative to the direction towards the sun (2). Upon violation of accuracy of said orientation, orientation of the artificial earth satellite is stopped using an on-board computer simultaneously relative to the direction towards the sun and the earth (1). Said self-contained circuit is turned on, and the solar panels (5) are installed in a fixed position relative to the body of the artificial earth satellite to achieve maximum illumination thereof. Resumption of orientation of the artificial earth satellite using the on-board computer is carried out at a radio command from the earth. The accuracy of orientation of the artificial earth satellite towards the sun can be evaluated from current parameters of the power supply system of the artificial earth satellite. A sign of violation of said orientation can be the beginning of operation of the power supply system in discharge mode of on-board batteries when flying outside shadow portions of the orbit (4). |
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Method of control over program turn of accelerating unit Invention can be used for control over program turn of accelerating unit with the help of fixed constant-thrust engines of orientation. Angular velocity is increased at acceleration and inertial flight and decreased to zero at deceleration and pulsed initiation of orientation engines. Level of fuel component in the tank that brings about the most tangible effect on turn dynamics is measured Angle mismatch and acceleration unit angular velocities are intermittently measured at turn as well as deflection of said fuel level from acceleration unit lengthwise axis Orientation engine are shut down at the ends of acceleration path and switched on at deceleration start path. |
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Invention relates to control of orientation of a space, in particular, a transport cargo ship (TCS) with stationary solar battery panels (SB). The method includes the rotation of TCS around a normal to the working surface of SB facing towards the Sun with an angular speed of at least 1.5 deg/sec. Meanwhile within the time interval of at least one round the components of the angular speed of TCS in a structural coordinate system are measured. Using the measured values, directions of the main central axes of inertia of TCS are determined. Among these axes an axis other than the axis of the minimum moment of inertia and making the minimum angle with the normal to the working surface of SB is found. The angle between the direction towards the Sun and the plane of the TCS orbit is determined. If this angle exceeds a certain value depending on the specified minimum angle and also - on the minimum and maximum SB currents, TCS is turned. Meanwhile the named found inertia axis is combined with the direction, perpendicular to the orbit plane and making a sharp angle with the direction towards the Sun. TCS is rotated around this axis towards the direction opposite to orbital rotation. During the rotation the current from SB is measured. At the achievement by the current of the minimum value TCS is again turned until the alignment of the named found axis of inertia of TCS with the named perpendicular direction and again the named rotation of TCS is performed. |
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Invention relates to aerospace engineering. Proposed method comprises supply spaceship spinning around perpendicular to solar battery working surface directed to the Sun at angular velocity of at least 1.5 degree/s. During said spinning at time interval of duration making at least one circuit supply angular velocity components are measured in structural system of coordinates. Measured magnitudes are used to define the directions of the main central axis of inertia of supply spaceship. Angle between direction to the Sun and spaceship orbit plane is defined. Spaceship orbit height is defined to determine the half-angle of the Earth disc visible from spaceship. In case said angle exceeds said half-angle the spaceship gravity orientation is constructed at aligning the axis of its minimum moment of inertia that makes the minimum angle with perpendicular to solar battery working surface, with direction to the Earth centre. Spaceship gravity orientation is maintained by spinning it around the axis of minimum moment of inertia at angular velocity defined from the condition of stability of the given gravity orientation of spaceship. |
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Invention relates to aerospace engineering. Proposed method comprises supply spaceship spinning around perpendicular to solar battery working surface directed to the Sun at angular velocity of at least 1.5 degree/s. During said spinning at time interval of duration making at least one circuit supply angular velocity components are measured in structural system of coordinates. Measured magnitudes are used to define the directions of the main central axis of inertia of supply spaceship. Spaceship is turned to alignment of the central axis of inertia making the minimum angle with perpendicular to solar battery working surface with direction to the Sun. Spaceship is spun around spaceship around said axis to measure solar battery current. After current reaches minimum permissible magnitudes spaceship is, again, turned to align said axis of inertia with direction to the Sun. Again, spaceship is spun around said axis. |
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Invention relates to aerospace engineering and can be used in spacecraft engines. Power plant comprises cryogenic tank with shield-vacuum heat insulation and channel with heat exchanger, flow control valve, booster pump, intake with capillary accumulator with heat exchanger and throttle and hydropneumatic system with pipeline. Channel cross-section sizes comply with maximum outer sizes of heat exchanger cross-section. |
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Method of control over spacecraft descent in atmosphere of planets Invention relates to spacecraft control in atmosphere of planet by adjusting its aerodynamics. Spacecraft velocity in atmosphere at initial flight part increases (spacecraft flies toward conditional orbit pericentre). Atmosphere density is low yet to cause notable spacecraft deceleration. As spacecraft reaches atmosphere dense layers its velocity decreases to reach atmosphere enter velocity for angle of roll (γ) γ=π to be changed to γ=0. This manoeuvre allows changing the spacecraft to flight part with maximum aerodynamic performances. In flight with γ=0 continuous skip path is maintained whereat spacecraft velocity decreases monotonously. Maximum skip height reached, angle of attack o spacecraft increases, hence, spacecraft intensive deceleration occurs. |
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Method of control over spacecraft descent in atmosphere of planets Invention relates to spacecraft control in atmosphere of planet by adjusting its aerodynamics. Proposed method consists in selection of conditions for changing the angle of roll to zero at changing the spacecraft from isothermal descent section (IDS) to skip path. With spacecraft in IDS, angle of roll (γ) is, first, increased to decrease aerodynamic performances and to maintain constant temperature at critical area of spacecraft surface. As flight velocity decreases angle (γ) is decreased from its maximum. In IDS, increase in aerodynamics does not cause further temperature increase over its first peak. Therefore selection of the moment of changing to γ=0 allows efficient deceleration of spacecraft at the next step of flight. The best option is the descent of spacecraft of IDS when γ reaches its maximum. Here, angle of attack is set to correspond to maximum aerodynamic performances. This increases the duration of final flight stage and deceleration efficiency. Increase in angle of attach after descent from IDS and completion of climb results in increased in drag, hence, decrease in velocity at initiation of soft landing system. |
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Method of controlling orbiting spacecraft Invention relates to controlling movement of a spacecraft fitted with a heat radiator and a solar panel. The method includes flying the spacecraft on an orbit around a planet and turning the solar panel in a position corresponding to the alignment of the normal to the working surface of the solar panel with the direction towards the Sun; performing orbital orientation of the spacecraft, where the plane of rotation of the solar panel is parallel to the plane of the orbit of the spacecraft and the solar panel is located relative to the plane of the orbit on the side of the Sun; determining the maximum value of the angle between the velocity vector of the spacecraft and the perpendicular to the transverse axis of rotation of the solar panel, passing through the surface of the radiator; determining the orbital altitude of the spacecraft and the angle between the direction towards the Sun and the plane of the orbit of the spacecraft; based on the orbital altitude and the angle, determining the orbit passes where the duration of the illuminated part of the pass exceeds the difference between the orbiting period of the spacecraft and the required duration of the heat release by the radiator on the pass; on the said orbit passes, when the spacecraft passes through the illuminated part of the pass, the solar panel is turned around the transverse axis of rotation until the intersection of the line passing through the region of the surface of the radiator facing the Sun and directed towards the Sun with the solar panel; turning the solar panel around the longitudinal axis of rotation until the angle between the normal to the working surface of the solar panel and the direction towards the Sun assumes a minimum value. The said solar panel rotations are performed within a calculated time interval. |
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Invention relates to space cable systems (SCS) and can be used for the transfer of SCS to a spinning mode in the orbit plane without the application of jet engines. SCS development is executed from its initial compact state in the circular orbit by the repulsion of objects at a low relative speed. SCS end weights are connected by a cable, its length being varied by a cable feed-haul-in device arranged on one of the end objects. The objects are separated by a vector of local peripheral speed, for example, by a pusher. The objects are driven by a start pulse to separate the objects in practically free paths at the free feed of the cable. The cable development is terminated by the SCS transfer to a stable mode of associated pendulum motion at the stretched preset-length cable. At a definite range of angular phases of this mode the SCS objects are stretched by hauling in the cable at a definite constant speed. This results in changing the SCS into the spinning mode at a preset power integral and fixed final end of the cable. |
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Method of orientation of space vehicle and device for its implementation Group of the inventions relates to control of angular motion of space vehicle (SV). The method includes additional generation of signals for assessment of orientation angle and angular velocity of rotation of space vehicle. Also the reference signals of the orientation angle, angular velocity and control assessment signal are generated. For the named orientation angle and angular velocity their differences with their assessed signals, and also the difference with their reference values are determined. The difference of control signal and its assessed value and, at last, the signal of correction of the signal of assignment of mathematical model and the signal of assessment of external noises using the respective formulas are determined. On this base the signals of assessment of orientation angle and angular velocity of space vehicle are determined, which are used for space vehicle control. The device in addition contains the reference model of the basic circuit of orientation of space vehicle and other necessary devices and connections. |
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Navigation satellite orientation system Invention relates to controlling orientation of an artificial earth satellite with solar panels. The disclosed method includes performing necessary turning of the artificial earth satellite along with solar panels and, separately, the solar panels about a first and a second axis. The antenna of the artificial earth satellite is directed towards the earth and the normal to the solar panels is directed towards the sun. Independent programmed turns about the first and second axes of the artificial earth satellite are performed in intervals of uncertainty of orientation of the artificial earth satellite on shadow orbits. In different versions of said turns, after the first turn, the artificial earth satellite is held in an intermediate position and normal orientation of the artificial earth satellite is then restored. This improves the accuracy of predicting movement of the artificial earth satellite on shadow orbits and accuracy of measuring the range to the artificial earth satellite. |
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Control over orbital spacecraft Invention relates to in-flight control over spacecraft equipped with heat radiator and solar battery. Proposed process comprises spacecraft flight in orbit around the planet with solar battery turn to position corresponding to normal to solar battery working surface directed to the Sun. Spacecraft orbital orientation is constructed whereat solar battery spinning plate is parallel with spacecraft orbit plane while solar battery is located on the Sun side relative to orbit plane. Spacecraft orbit altitude and angle between direction to the Sun and spacecraft orbit plane are defined. Magnitude of said angle (β*) is defined whereat duration of turn shadow section equals the necessary time of radiator heat release in said turn. Orbit turns are defined wherein current magnitude of said angle is larger than β*. In said turns, solar battery is turned around crosswise and lengthwise rotation axes unless shadowing of solar battery radiator. Note here that minimum departure of orientation of solar battery working surface to the Sun. Spacecraft orbital flight is conducted in near-circle orbit at altitude not exceeding a definite design value. |
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Method of holding spacecraft in geosynchronous 24-hour orbit Invention relates to control over spacecraft, particularly, to holding of geosynchronous spacecraft in preset are of stay and collocation with the other geostationary spacecraft. Proposed method comprises determination and correction of initial inclinations and longitude of injection orbit ascending node with allowance for epoch of spacecraft placing in orbit and term of its active existence. Note here that the time of beginning of operation in geostationary orbit when spacecraft orbit inclination reaches maximum permissible value iper. area. The latter corresponds to permissible reach in latitude at the boundary of nominal spacecraft stay area in altitude. Stable and minimum eccentricity magnitudes are defined. Eccentricity vector is corrected so that it equals the nominal value for spacecraft collocation and spacecraft orbit apse line is aligned with that of nodes. Spacecraft active collocation is executed at changing the inclination from 0 to iper without interaction with adjacent spacecraft control centres. At inclination larger than iper , eccentricity is increased to minimum with setting of Laplace vector in direction from the Sun. Note here that eccentricity vector is not corrected unless the end of spacecraft active existence term termination. At inclinations larger than iper, eccentricity vector equals modulo and is spaced apart relative to eccentricity vectors of the other spacecraft. |
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Space vehicle correction engine test method Invention relates to control of movement of space vehicle (SV). According to the proposed method, thrusts of correction engines (CE) (control accelerations) are determined as per total changes in a period of SV revolution from correction to correction. The latter are performed with one and the same CE and a thrust level of those CE is estimated. For reliable understanding of thrusts of a pair of mutually opposite located CE, series control activations of the same pair are performed from time to time with equal pulses. Discrepancy as to total pulse of thrust is entered in equal parts with an opposite sign to the implemented pulses. As a result, reliable thrust levels of CE in operation are obtained. |
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Method for generation of control actions on spacecraft Invention relates to spacecraft (SC) motion control using solar radiation pressure forces distributed over SC working zones. The latter are formed as flat parallel optically transparent droplet flows. Distance between droplets of R radius in each flow in its lengthwise direction (Sx) and frontal-lateral direction (Sy) is divisible by 2 R . Number of flows is n = ( S x / 2 R ) − 1 . By mutual bias of flows in direction of their motion for 2 R distance droplet mist flows are generated in number of m = ( S y / 2 R ) − 1 . Each of the mentioned flows is biased relative to previous flow for 2 R distance in frontal-lateral direction. Thus opacity in frontal-lateral direction and transparency in direction of plane perpendicular to a flow is created. Unit distributed light pressure force is regulated by changing radius and number of droplets coming to point of it application in unit time. Total action value is regulated by changing number of droplet jets. |
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Method of clearing space debri from orbit Invention can be used to move space debris from working orbits to recycling orbits. The method includes taking towing spacecraft and a self-contained docking module into the region of an orbit from which space debris is to be cleared. The sequence of removing space debris is selected by comparing a criterion, for example the probability of the space debris colliding with other space objects, for each space debris. Compensation for accumulated errors of motion parameters of the towing spacecraft during previous manoeuvres, as well as the pointing system is distributed between correcting pulses of the towing spacecraft at the long-range guidance step and of the self-contained docking module at the self-guidance portion. |
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Antifailure protection of rocket multiplex control system Proposed method comprises generation of data signal in every channel corresponding to a definite combustion chamber as the difference between command signal and feedback signal, generation of control system channel cut-off signal as the signal of setting actuator rod on said channel to mid position. Channel cut-off signal is generated in case the data signal modulus integral calculated at preset-duration time interval exceeds the preset threshold. Note here that command signals of other channels are generated as sums of or difference in pitch, yaw and bank control signals and cut-off channel feedback signal with coefficients depending on cut-off channel number so that required summed pitch, yaw and bank control moments are actuated. |
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Method to control spacecraft placing into orbit of planet artificial satellite Invention relates to spacecraft (SC) motion control during placing it into orbit of planet artificial satellite using aeroassisted maneuver. In the phase of airbraking, following parameters are forecasted for the time of SC exiting planetary atmosphere: SC speed, angle to inclination to local horizon and altitude of transfer orbit apofocus. In this process in each of successive forecast points, SC motion is considered on remaining atmospheric flight segments for roll angles γ = 0 rad and γ = π. The mentioned above forecasted manoeuvring parameters are found for each of these angles. Their values are used during control of SC incidence change (close to its value corresponding to maximum quality) and issuing pulse of SC speed in transfer orbit apofocus. |
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Method of spaceship orienting and device to this end Proposed method comprises generation of signals for estimation of: spacecraft orientation angle, spacecraft sing angle and control. Signal difference of said parameters and their estimates are defined. Several formulas are used to calculate the correction signals of setting and estimation of external interference. Said corrections are allowed for at correction of orienting angle estimate signals and angular velocity signal. The latter are applied in spacecraft orientation control circuit. Proposed device comprises the following extra units: memory, adders, amplifiers, integrators interconnected and connected with other elements via system of switches. Proposed device incorporates models of the spacecraft orientation main circuit and flywheel engine. |
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Method of spaceship orienting and device to this end Invention relates to aerospace engineering, particularly, to spacecraft orientation. Proposed method comprises the steps that follow. Generation of angle estimate and spacecraft spin rate estimate signal are generated. Determination of difference in angle signal and angle estimate signal. Determination of difference in spin rate and spin rate estimate signal. Corrected angle estimate signal and corrected spin rate estimate signal are defined. Control signal is generated using said corrected angle estimate signal and corrected .spin rate estimate signal Proposed device comprises engine-flywheel model, four integrators, four adders, four normally closed switches and two normally-open switches. Output of 2nd adder is connected via engine-flywheel model, 1st integrator, 4th adder, 2nd integrator, 5th adder, 6th adder, 1st switch and 3rd integrator, all being connected in series, with 5th adder 2nd input. Output of the latter is connected via 1st switch with 2nd input of 1st adder. Output of 4th adder is connected via 7th adder, 2nd switch and 4th integrator, all being connected in series, with 4th adder 2nd input. Output of the latter is connected via 2nd switch with 3rd input of 2nd adder. Output of spin rate transducer is connected with 7th adder 2nd input and, via 3rd switch, with 2nd amplifier input. |
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Method of spaceship orienting and device to this end Invention relates to aerospace engineering, particularly, to spacecraft orientation. Proposed method comprises the steps that follow. Generation of angle estimate signal and signal of spacecraft spin rate. Generation of control estimate signal. Determination of difference in angle signal and angle estimate signal. Determination of difference in spin rate and spin rate estimate signal. Determination of difference in control signal and control estimate signal and determination of corrected angle estimate signal, corrected spin rate estimate signal and outer interference estimate signal. Then, generated are control signal with corrected angle estimate signal, corrected spin rate estimate signal and external interference estimate signal. Proposed device to this end comprises five normally-closed switches, two normally-open switches, seven adders, model of engine-flywheel, two amplifiers and five integrators. |
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Device to control spacecraft position in space with help of orbital gyrocompass Device for control over spacecraft position in space with the help of orbital gyrocompass. Control device comprises local vertical plotter, adders, amplifier-converter units, integrators, compensation units and gyro meter of angular velocity, spacecraft position setting device, cosine angle transducers, sine angle transducers, spacecraft veering control unit and spacecraft bearing setting unit. Connection between device elements are configured to allow spacecraft turn the course through arbitrary angle without loss in orientation relative to orbital system of coordinates. Note here that contour of correction from local vertical plotter is in operating mode. Spacecraft can either spin along the course or stay in definite position relative to orbital system of coordinates with loss in precision of orientation. |
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Method of descending space rocket stage separation part and device to this end Invention relates to aerospace engineering and may be used for descending space rocket stage separation parts (SRSSP) from orbits of payloads. SRSSP comprises propellant compartment and power compartment with bottoms. Upper bottom accommodates rotary chambers of rocket gas engine while lower bottom accommodates mid-fight engine (MFE) with elongated charge electrically connected via switchboard with power supply. SRSSP is oriented and stabilised by energy of liquid propellant gasified residues at application of velocity pulse defined by radii of SP MFE descending path apogee and perigee. |
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Stabilisation of unstable fragments of space garbage Proposed method comprises application of force to the fragment at its design points. Said force is created by air effects applied by gas flare to said fragment produced by satellite located nearby said fragment. Said gas flare can be generated by device, for example, various jet engines. Note here that space garbage fragment orbit can be changed simultaneously. |
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Method of spaceship orienting and device to this end Proposed device comprises eleven adders, five amplifiers, two normally-open switches, five normally-closed switches, four integrators, two multipliers, spacecraft, flywheel engine, flywheel engine simulator, transducers of angular velocity and orientation angle, constant setter and memory. Signals of angular orientation and velocity are measured to generated spacecraft control signals, angular orientation estimation signals, those of angular velocity so that difference between appropriate signals and estimation signals is defined as well as those of spacecraft inertia and external interference estimation and for spacecraft orientation signal is corrected and generated. |
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Invention relates to space engineering, primarily to space cable systems. Proposed method comprises undocking of two objects connected by cable, imparting to the lander of initial velocity of departure, unobstructed release of cable on departure of the lander, fixation of cable length at the end of reverse section, paired pendulum motion and cutting of the cable at the moment whereat lander passes by the line of orbital station local vertical. Separation of the lander is performed against orbital velocity vector with control over cable tension force control at departure of the lander. Free cable is hauled in at trajectory reverse section. |
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Invention relates to control over spacecraft angular flight, particularly, to spacecraft orientation gyro systems equipped with ground object observation system in near-earth orbit. Operation of such spacecraft required lateral shift of ground object image, for example, in observation hardware focus plane caused by daily rotation. For this, spacecraft is turned to track route angle by cosine law at circular frequency equal to orbital angular velocity. In compliance with this method, along with spacecraft course turn, it is turned through track angle of bank by sine angle of the same frequency. This makes instrumental track plane formed by bank and course angle turned relative to line of orbit nodes through constant angle equal to track angle amplitude. Therefore in steady state mode power is consumed only for angular stabilisation to counteract external disturbance effects. Turn of observation hardware bank observing line relative to orbital system of coordinates toward observed ground objects is performed with allowance for current program bank turn angle. Proposed device comprises appropriate instrumentation for implementation of above described method. |
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Method of spacecraft orbital flight correction Invention relates to space engineering and can be used for spacecraft orbital flight correction. Test and correction effects are applied by switching on of engines to make path changes, to define parameters of motion of spacecraft center of gravity, to compute the correction, to generate command-program data with initial conditions of the flight, correction plan and control accelerations, and to transmit said data files to spacecraft. |
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Method of spacecraft orbital flight correction Invention relates to space engineering and is intended for retaining the craft in preset geostationary orbital position. Spacecraft orbital flight parameters are corrected and sent thereto. At the time of tracking the correction engine operation at every step thereof, inertial mass free motion start and end aboard said craft in closed spherical vessel is registered.Control acceleration is defined from equation of uniformly accelerated motion without initial velocity in known path in said vessel by means of high-precision accelerometer of linear accelerating effect at zero-gravity effects. Magneto susceptible ball makes said inertial mass. |
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Method of forming information space-time field Method of forming an information space-time field involves synchronising spaced-apart universal time system generators based on astronomical events, wherein universal time system generators are brought to libration points L4, L5 as spacecraft target equipment. Synchronisation of operation thereof is carried out based on astronomical events in form of signals of pulsar events in the X-ray range. The obtained synchronised signals of the universal time system generators are converted to a digital code bearing information on reference time and frequency, as well as spatial coordinates of the spacecraft, which is emitted in the standard radio range in near-earth space. The universal time system generators are synchronised using signals of pulsar events, coordinates of which are located on the straight perpendicular plane of the orbit of the moon. Spatial coordinates of the spacecraft are determined using signals of at least three pulsar events, coordinates of each of which are located on one of the three mutually perpendicular planes. Universal time system generators are launched from the signal of a unified pulsar event. |
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Set of invention relates to rocketry and serves to generate control commands to fuel components second consumption adjustment means in acceleration unit flight. Proposed method consists in maintaining preset ratio of fuel components second fuel components flow rates on the basis of continuous data of fuel components second fuel flow rate transducers. Note here that discrete fuel store time derivative is defined by measurements of two adjacent measurement points of fuel tank fuel level transducers. Proposed system comprises fuel second consumption transducers arranged in feed lines to transmit signals to computation device and, therefrom, to actuators. Additionally, this system incorporates fuel level discrete transducers mounted in fuel tanks. It includes also amplifying converter of signals from discrete transducers to generate logical 1 signal at their outputs at fuel passage through sensor of appropriate transducer and amplifying converter of oxidiser and fuel second flow rate transducer signals into digital code corresponding to measured volume second flows of fuel components. Fuel component temperature transducers are connected to signal amplifying converter to generate temperature correction to second flow rate transducer readings. Computer processes data from amplifying converters and used the control unit to generate instruction for throttle that controls ratios of fuel component second flow rates. |
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Invention relates to aerospace engineering, particularly, to means for placing spaceships in orbit. Carrier rocket comprises one sustainer engine at gimbal suspension and jettisonable first stage. Said first stage comprises aerodynamic roll system with hydraulic servo drives at maximum ram pressure in orbiting of carrier rocket. Said hydraulic servo drives are equipped with fluid-driven locks for automatic locking in preset position with shutdown sustainer and no pressure in said hydraulic servo drives of aerodynamic roll system arranged at rocket carrier jettisonable first stage nose. |
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Method of adaptive control over displacement of centre of gravity of spacecraft Invention relates to control over flight of the group of spacecraft and may be used for tracking one spacecraft by another spacecraft at preset distance. Proposed method comprises measuring trajectories and making corrections, minimizing orbit eccentricity and defining position of subject spacecraft in inertial space. Note here that control and navigation system is provided with set of transceiving radio hardware and optical transducer of angles "Pole Star - subject spacecraft - object spacecraft". Distance to object spacecraft is measured to define is deviation from mean magnitude at measurement steps. At termination of every cycle of measurement steps, dynamics of variation of said mean magnitude is revealed to define increment of oscillation period of subject spacecraft relative to similar period of object spacecraft. In one orbital period, angle between planes of orbits of object spacecraft and subject spacecraft are defined as well as time of crossing of said planes by readings of angle transducer. In case said increment exceeds preset threshold, parameters of correction of said period are computed. At estimated time of orbit crossing, correction engines are initiated giving preference to engine of orbit inclination correction engine. |
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Device to control passive spacecraft orientation Invention relates to space engineering and may be used in approach, buzzing, hovering, docking jobs etc using robotic systems. Device comprises casing, radiation source, flat diffraction gratings and outlets. Four planes of flat diffraction gratings are perpendicular in pairs, two of them intersect at right angle to axis extending through common radiation source and parallel with passive spacecraft construction axis while remaining two make the angle of 0 to 90 degrees with the axis. |
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Method of retaining geostationary spacecraft in preset orbital position Invention relates to space engineering and is intended for retaining in preset geostationary orbital position. After spacecraft gravity center control term extended, no ground means of navigation parameters measurement are used to compute the schedule of spacecraft gravity center motion displacement by one correction engine to fix, at every correction step, the start and finish of free displacement of inertial mass aboard spacecraft inside spherical closed vessel. Magneto susceptible ball makes said inertial mass. Control acceleration is defined from equation of uniformly accelerated motion without initial velocity in known path in said vessel by means of high-precision accelerometer of linear accelerating effect in no-gravity effects. Correction engine operation duration is defined to define correction engine switch-off time. |
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Method of controlling spacecraft center of inertia in docking Invention relates to space engineering. For control over spacecraft center of inertia in docking displacement angle relative to boresight with lag and angular velocity of boresight with lag are measured. In case displacement angle exceeds preset operation threshold or angular velocity exceeds preset operation threshold or displacement angle is smaller than preset operation threshold control action is applied to center of inertia. Duration of control action varies with modulus of displacement angle and modulus of boresight angular velocity with due allowance for distance while the sign is opposite the displacement angle and that of boresight angular velocity. Mean velocity is defined by drift angle. Drift angle is defined by boresight angular velocity at time interval as the sum of lags in determination of displacement angle and boresight angular velocity. With mean velocity equal to or exceeding half the operation threshold in boresight angular velocity, control action is applied to center of inertia. Duration of control action varies with modulus of mean velocity with due allowance for distance while sign is opposite that of mean angular velocity. Accumulated drift angle and time interval are zeroed to start defining the drift angle and mean velocity. |
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Spacecraft angular motion stabilisation system Invention relates to aerospace engineering and may be used in spacecraft onboard control systems. Proposed system comprises three control channels. Every said control channel comprises control signal setting device, angular velocity transducer, unit of intermittent switching, switch, unit of jet engines, angle transducer, comparator, adding amplifier and relay element connected in series. Comparator is connected with control signal setting device, adding amplifier is connected with angular velocity transducer, and intermittent switching unit is connected with switch. |
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Method of retaining geostationary spacecraft in preset orbital position Invention relates to aerospace engineering and may be used for retaining spacecraft in preset range of altitudes and longitudes at working position in orbit. Error in controlling displacement of spacecraft center of gravity is eliminated by using the coefficient of converting voltage and current in plasma engines into engine thrust and displacing spacecraft period check plane into center of orbit active section center. Nominal line (paradigm) of retaining in pane is selected while stable centripetal effect of spacecraft evolution at orbital position is triggered and maintained by correction for long time interval. |
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Method of adjusting and stabilising pressure in silphon-type working units Invention relates to space technology and may be used for stabilisation of preset engine thrust by correction of spaceship motion. Tank with working medium (WMT) has three chambers. All supercharge gas (SG) is kept in extra permanent-volume tank (EPVT) adjoining WMT wall opposite the bellows. In case current and preset fuel pressures differ, defined are valid current SG temperature and pressure between bellows and EPVN, fuel mass residue, current SG volume, SG portion of EPVT required to reach operating pressure proceeding from current pressure in EPVT and interchamber channel cross-section, as well as duration of transfer of this portion into central chamber. Interchamber valves are opened and closed at preset time. |
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Invention relates to termination control of rocket acceleration unit uncontrolled-thrust sustainer engines. Proposed method comprises forecasting sustainer engine shutdown in changing over to termination control before spaceship separation when functional reaches preset power. For this, during said transition determined is conditional time of rocket propellant combustion (propellant consumption) and difference between said conditional time and preset time specified in flight mission. Said difference describes sustainer engine operation available time to be compared with calculated on the basis of flight mission. Is calculated time exceeds preset time then maneuver termination moment (engine shutdown) is defined as the sum of the interval of changing to termination control available time of engine operation. Said summed time is memorised to forecast acceleration unit motion at every interval of termination control with constant integration step. Said step equals relation of period to termination of maneuver to preset number of integration steps in acceleration unit motion model. Acceleration fuel propellant and final maneuver may terminate at shortage of power functional. In this case, knowledge of the moment of maneuver termination allows setting the spaceship orbit required inclination. |
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Method of plotting manned spaceship orbital attitude Invention relates to manned spaceship orbital attitude control in on-orbit navigation. Spaceship is equipped with planet surface scanner. Proposed method comprises plotting orbital attitude in local vertical. Thereafter, scanner screen grid is turned to align its lines with direction of check point motion. Grid turn angle is defined to set angular speed of spaceship rotation about center of inertia relative to local vertical. Said rotation is completed after grid turn angle reaches definite magnitude. Then, screen grid is moved back into initial position to control alignment of underlying surface references displacement with grid lines. |
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Invention relates to space engineering, particularly, to strapdown integral orientation system angular velocity meters, namely, to methods of their correction. Proposed method consists in executing three sequential correction plane revolutions of spaceship about axes of roll, yaw and pitch through preset angle. Before first correction turn, spaceship is stabilised in preset position to define angular position increment and angular position as analytical solution of kinematic equations and average magnitudes of projections of spaceship speed of rotation on angular velocity meter measurement axes. Discrepancies at program turns and calculated constant drifts are used to calculate errors in scale factors and errors in single-axis meter measurement axis setting. Correction device comprises angular velocity meters, stellar-measurement system, correction parameters registration unit connected to angular velocity meter, correction parameter calculation unit, unit of integration of angular velocity projections on meter sensitivity axis, and to angular motion parameters calculation unit. Angular motion parameter calculation unit is connected to stellar-measurement system. Discrepancies calculation unit is connected to correction parameter calculation unit and memory unit. The latter is connected to integration unit, angular motion parameter calculation unit and programmer unit. Zero signal calculation unit is connected to correction parameter calculation unit, discrepancy calculation unit and motion programmer unit. The latter is connected to discrepancy calculation unit and correction parameter calculation unit. |
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Method for active-passive damping, orientation and stabilisation of spacecraft Invention relates to spacecraft engineering. Method for active-passive damping, orientation and stabilisation of spacecraft on orbits with height up to 400 km includes combined moment action from gravity stabiliser and jet thrust. Roll and pitch orientation of space craft is executed in response to a signal from control unit at the moment of maximum deviation of gravity stabiliser longitudinal axis from its dynamic balance position along local vertical by single or multiple switching on the jet engines in the plane of deviation and towards deviation. Jet engines are located at remote relative to spacecraft mass centre end of gravity stabiliser so that thrust vector of each jet engine could be perpendicular to longitudinal axis of gravity stabiliser. Damping and active roll and pitch stabilisation of spacecraft is executed in response to a signal from control unit by single or multiple switching on the jet engines at the moment when gravity stabiliser end crosses local vertical, in the plane of deviation and opposite to direction of deviation. Yaw damping, orientation and stabilisation of spacecraft is executed in response to a signal from control unit by deploying additionally introduced aerodynamic shield with ballistic coefficient exceeds ballistic coefficient of the spacecraft. Connection of spacecraft and the shield is performed using flexible link, swivel and at least three slings where total length of the mentioned connection is not less than ½ middle of the shield. |
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Invention relates to control of orientation of space vehicle (SV) with immovable relative to SV body panels of solar batteries (SB). Control method includes SV gravitational orientation and its spin around longitudinal axis (minimum momentum of inertia). When the Sun is near orbit plane this plane is aligned with SB plane by the time of passing morning terminator. Angle between normal to SB active surface and direction to the Sun is measured and monitored. At the moment of passing morning terminator, SV spin is performed in direction corresponding to decrease of the mentioned angle, where spin angular velocity is selected from the range of 360°/T - 720°/T, where T is SA orbiting period. |
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Method of controlling space rockets Invention relates to rocketry and may be used in designing optimum programs for control over space rocket first stage descent proceeding from softened restrictions related with fall of ejection stages onto existing land disposition zones. Proposed method comprises forecasting location of ejection stage fall, comparing forecast coordinates of fall point with coordinates of allowed fall zone, and estimating fuel residues in tanks. After ejection-stage separation, pulse is applied to center of gravity of said stage, its magnitude and direction being defined subject to possible variation in inclination of trajectory of space rocket descent orbit. To allow increase in velocity of ejection stage center of gravity for descent in preset area, onboard self-contained drift system is used using liquid-propellant residues in tanks. |
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Invention relates to space engineering. Proposed method comprises forecasting parameters of accelerating unit at cruising engine shutdown to define deviation of radius and radial speed from their magnitudes in preset orbit. Sensitivity of said parameters to orientation program parameters variation is defined. Control correction signals are generated for pitch angle and pitch angular speed to sum up corrections with their programmed magnitudes. Pitch angular acceleration is loaded into said program as extra parameter to define sensitivity of radius and radial speed sensitivity to angular acceleration change. Forecast deviations of radius and radial speed from their magnitudes in preset orbit are used to define corrections to pitch angular speed and pitch angular acceleration. Pitch angular speed correction signal is limited to preset level to vary pitch angular acceleration correction signal with pitch angular speed correction signal and inversely with generated pitch angular speed correction signal. Said corrections are summed up with programmed pitch angular speed and angular acceleration to define new parameters of orientation program. |
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Invention relates to operation of rockets with the first stage multi-engine pack. Said pack may consist of central engine and two lateral engines arranged in common plane, e.g. in rocket yaw plane. In case one of said lateral engines fails in launching, less the explosion, remaining engines or their part are switched over into thrust augmentation. Rocket is stabilised in vertical position to preset lift altitude. Then, rocket is turned down bank unless engine location plane is aligned with that of rocket safe zone plane. Note here that faulty engine is directed toward ram airflow. Angular parameters of rocket deviation in direction of launch path end are kept within tolerances to allow rocket further stabilisation. To ensure shock-less release of chambers from launch table recesses, lateral engines are mounted with equal angular inclination from rocket lengthwise axis. Chambers are registered in zero position relative to said axis while registration is cancelled at rocket motion start. Note that lateral axes of lateral chambers are aligned, after their release from recesses, with those of lateral engines. |
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Aerodynamic surface of satellite aerobraking Invention relates to means of satellite aerodynamic braking used to deorbit satellites after expiration of their life. Aerodynamic surface 2, 3 of satellite 1 serves to increase satellite drag without special stabilisation in orbit. Proposed surface comprises one or several elements that, when unfolded, make 3D structure made up of at least two panels 2a, 2b, 3a, 3b that form a dihedron. Each pair of panels unfolds from one mast of unfolding arranged along edge 6, 7 of appropriate dihedron. Panels 2, 3 are made up of flat flexible membranes tightened on both sides of retaining elements perpendicular to said mast. |
Another patent 2551154.