Device to control passive spacecraft orientation
SUBSTANCE: 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.
EFFECT: decreased loads at docking assemblies.
The invention relates to space technology and can be used when running in space operations convergence, flight, freezing, docking with the docking of SPACECRAFT (SC), as well as for positioning actuators when performing construction and Assembly operations and other operations with robotic means.
As taken similar laser visual planting proposed in the U.S. Navy for landing deck aircraft (LA) on an aircraft carrier . There are two laser systems designed to facilitate the landing deck of the aircraft.
Laser pointer planting rate (Laser Localizer Centerline - LCL) allows to enter the landing pattern at a distance of more than 10 nautical miles.
Lasers that emit red, green and yellow, help the crew of LA to keep landing course and glide path when landing on the deck of a moving ship. To denote deviations of the left and right are used lasers with different wavelengths, continuous and modulated radiation. Deviation to the right is green, left is red. With increasing deviation from the course green and red lights begin to flash. Flash rate increases with the deviation. In strict keeping boarding rate is only visible flame yellow. The total power of light sostavljaet CD.
The laser beams are visible in the sector of 21.9°. At the distance of 0.5 mile of the setting on the USS LA goes out of range of the LCL and focuses on optical landing system on the Fresnel lenses.
Laser glide slope indicator (Laser Glideslope Indicator - LGD) allows the crew to LA to get on the glide path with a distance of 10 miles. Also use different wavelengths and modulation of radiation with a significant deviation from the glide path. At a distance of less than 0.5 mile LGD ceases to vary, and the orientation is performed on the optical landing system on the Fresnel lenses.
With the right approach, the pilot LA sees on the aircraft carrier two yellow lights on the left (LCL and right LGD. The accuracy rate for LCL is in the range of 0.25°, line LGD is determined with accuracy of 0.1°.
Maximum detection range of LGD is 12 miles. Tests have shown that these devices make it easy to detect the slightest deviation in the landing approach and can be used in carrier-based aircraft.
The disadvantage of analog is a large number of radiation sources, which increases the weight and size of equipment consumed her power, as well as direct observation by the operator emitting laser apertures, which requires careful calculation of laser safety.
As a prototype adopted is designed to control the rendezvous and docking of SPACECRAFT is mehanicheskij target convergence .
Mechanical target includes a base (round or asymmetrical) and external cross black with applied white cross and labels. Coverage targets are produced according to special technologies, as they need a long time to maintain optical performance during space flight. As the primary means of visual inspection with the approach used cameras or optical aiming device.
On the target (or target complex) visually determined by the position of the line of sight of the active SPACECRAFT during rendezvous respect to passive SPACECRAFT pitch and yaw, as well as the position of the axes of the passive SPACECRAFT relative to the active pitch, yaw and roll.
The corners of the active SPACECRAFT are determined by the offset of the front cross target relative to the crosshairs cameras or optical viewfinder, and the corners passive KA - offset front cross about cross Foundation.
Most of the existing targets require the blowout at run time mode mooring. On the night side of the orbit of the target is exposed headlight installed on the active CA.
On the sunlit side of the orbit is imposed on the position of the Sun relative to the target axis, excluding the shading body both active and passive KA and specifies the minimum angle saswad the base of the target.
The disadvantages of the prototype are the complexity and the lack of accuracy of determining the relative angular position of the SPACECRAFT at the site of the mooring due to the fact that for small angles of deflection corresponds to a small shift of the cross of the target relative to its Foundation, the definition of which can be difficult, and serious restrictions on lighting conditions at the time of closing.
The objective of the invention is to reduce the total angle between the longitudinal axes of the active and passive SPACECRAFT by increasing the accuracy of estimating the angles of deflection of the passive SPACECRAFT from the line of sight mode orientation, in addition, requirements are reduced lighting conditions during approach, but it is not required organization blowout at run time mode mooring.
The problem is solved using the monitoring device orientation passive SPACECRAFT, comprising a housing, a radiation source and two planar reflective diffraction grating, and plane and In a plane diffraction gratings perpendicular to the planes C and D, respectively, intersecting at right angles on the axis "
Thus, each flat diffraction grating allows to separately determine the angles in the plane of the pitch or yaw.
Figure 1-3 shows the construction of the device in three dimensions, where:
1 - body;
2 - radiation source;
3 - flat diffraction grating;
4 - output apertures.
Figure 4 shows to explain the design of the ISO.
Enclosure 1 provides protection of structural elements from external effects of space flight factors, as well as shielded from the observer to the source (having a greater brightness compared to dragirovaniya beam radiation) and has outlet openings 4, servants output apertures for the diffracted radiation.
The radiation source 2 is used to receive optical radiation with the required parameters.
Flat diffraction grating 3 decompose the radiation source 2 in range and are indicators of deviation from the line of sight.
Outlet openings 4 are designed to monitor the diffracted radiation in a predetermined range of angles.
Line of sight is a flat diffraction grating 3 coincides with the direction inwhich is maximum for wavelengths from 0.50 to 0.56 μm (green). The deviation from the line of sight corresponds to the shift of the luminescence in a more short - or long-wavelength region (blue or red). The principle of orientation is to maintain the direction corresponding to the green glow of both planar diffraction gratings 3. On the area of the green light (from 0.50 to 0.56 μm) also has a maximum sensitivity of the human eye (0,38-0,76 µm).
In the variant represented in figure 1-4, the deviation to the right corresponds to the red glow of the first planar diffraction grating 3, and the deviation to the left is blue. Deviation up corresponds to the blue glow of the second planar diffraction grating 3, and down - red.
With little care to the left of the green glow of the outlet 4 is replaced by blue, with the increase of angle - passing in blue, and at large angles becomes purple. Care to the right is indicated by changing the green glow on the yellow, passing into alternating orange and red. Each corner has its own wavelength.
Unlike mechanical targets, where for determining small changes passive corners it is necessary to detect the deviation of the front cross about cross Foundation, the proposed device, changing the angle corresponds to the color change, which facilitates the determination of the angular misalignment of the longitudinal axes of the active and passive the CSOs CA.
As the radiation source 2 can be used incandescent lamp or led lamp with a continuous spectrum in the visible region.
When installing the unit on a passive object to bind the lines of sight planar diffraction gratings 3 building axes. For precise alignment device may be provided with adjustable fastenings, are installed in the housing is a flat diffraction grating 3.
The zone of action of the device is determined by the near-simultaneous observations of the two outlet openings 4 without spontaneous changes color based on the base length is the distance between the planar diffraction gratings 3.
Each of the two output holes 4 has its line of sight, which are parallel to each other, so while their observation, a small parallax, increasing in the near zone. When approaching to a certain distance determined by the length of the base, the parallax increases, which causes a noticeable color change, complicating the orientation.
The maximum distance is limited resolution of optical or television system that allows separately to observe both outlets 4, and also depends on the length of the base.
The solution to this problem is the use of multiple is komplektov flat diffraction grating 3 with different length of the base for orientation in the far and the near zone, or joint application of the described device with traditional mechanical target. The use of planar diffraction gratings will not require an increase in power (they use only a small part of the light source, and in principle, their number is limited only by the proper detection and identification) and considerable design complexity.
The application of the invention can reduce the load on the connecting nodes and structures of active and passive SPACECRAFT during docking by reducing the total angular misalignment of the longitudinal axes of the active and passive SPACECRAFT pitch and yaw, which is achieved by increasing the accuracy of determining the angles of deflection of the passive SPACECRAFT from the line of sight.
The device described in the present invention can also be used for visual control of orientation of the interacting products in industry, construction and engineering fields that require precise control of the relative position of the interacting machines and mechanisms. In addition, the device can be used in aircraft for a visual approach aircraft on landing in low visibility conditions, refueling LA in flight using the pin-cone and other similar tasks.
1. Nordwall, Bruce D. Navy Tests Lasers To Help Carrir Pilots. Aviation Week & Space Technology, Nov. 19, 1990, pg.46.
2. Device control the orientation of the observed object. Patent No. 2093432.
The control device orientation passive spacecraft, comprising a housing, a radiation source, a flat diffraction grating and output apertures, characterized in that plane and In a plane diffraction gratings perpendicular to the planes C and D, respectively, intersecting at right angles on the axis
FIELD: instrument making.
SUBSTANCE: invention is designed for shaping of information field of laser teleorientation and navigation systems, optical connection, and can be used at control, landing and docking of airborne vehicles, escort of ships through narrow zones or bridge sections, remote control of robotic devices in zones that are dangerous for human health, etc. The proposed method is based on scanning by means of acoustooptical deflectors of the laser emission with a pencil-beam directional pattern; at that, laser beam movement trajectory provides formation both of information frames used for measurement of the controlled object coordinates, and command frames used for transfer of additional commands to the controlled object. The peculiar feature of the method is simultaneous formation of two lines of the information raster, which are displaced relative to each other by N/4 lines, by alternating formation of single cycles in the first line and then in the second line, where N is number of lines in a raster.
EFFECT: improving informativity of laser teleorientation system owing to increasing the repetition frequency of information and command rasters in information field of laser teleorientation system by reducing the duration of time delays between cycles, and owing to reducing light losses.
SUBSTANCE: scanning laser beacon has a housing, a laser light source mounted in a scanning unit, a base and an axle. The device includes an anamorphic optical system mounted in the scanning unit on the same optical axis as the laser light source. The axis around which the scanning unit rotates lies at an angle of 120° to said optical axis, and the anamorphic optical system is a wide-angle lens in a section perpendicular to the scanning direction, said lens having a 90° field of view. A rotating drive, which is in mechanical connection with the scanning unit, rotates in the scanning plane.
EFFECT: possibility of detecting a passive spacecraft in half the solid angle at distances of up to 160 km when pointing an active spacecraft on said passive spacecraft.
SUBSTANCE: scanning laser beacon has a housing and a laser light source mounted in a scanning unit in a gimbal suspension. The device includes an anamorphic optical system mounted in the scanning unit on the same optical axis as the laser light source. The axis of the gimbal suspension is perpendicular to said optical axis, and the anamorphic optical system is a fisheye lens in a section perpendicular to the scanning direction. A swinging drive, which is in mechanical connection with the scanning unit, swings in the scanning plane.
EFFECT: possibility of detecting a passive spacecraft in half the solid angle at distances of up to 160 km when pointing an active spacecraft on said passive spacecraft.
FIELD: physics, navigation.
SUBSTANCE: invention relates to instrument making and is intended for generation of data field of laser teleorientation systems (DF LTS) and navigation, optical communication, and can be used in control, landing and docking of aircraft, etc. continuous length-adjusted laser radiation band is generated as well as delay between three scanning cycles originating in object banking is generated by a certain law, the object accommodating control field generation system.
EFFECT: control over object with no zones wherein object laser control does not exist, scissors-like laser radiation directional pattern.
FIELD: control of moving objects with tele-orientation in the laser beam.
SUBSTANCE: the system has a laser, optoelectronic scanning system, output optical system and a control unit of deflectors. The control unit of deflectors has a formation unit of sync signals and raster parameters, driver of raster codes, driver of shift codes, adder and a double-channel frequency synthesizer. Raster codes Zs and Yt from the outputs of the raster code driver and shift code Kφ from the output of the shift code driver are fed the inputs of the adder connected to the inputs of the double-channel frequency synthesizer, codes Zs=Zt, Ys=Yt+Kφ or Zs=Zt+Kφ, Ys=Yt or Zt+Kφ, Ys=Yt+Kφ are formed. The control inputs of the shift code driver are connected to the control outputs of the formation unit of sync signals and raster parameters and the driver of raster-codes. The laser system of tele-orientation is made for input of the "DESCENT" command to the input of the formation unit of sync signals and raster parameters.
EFFECT: enhanced noise immunity of the system and enhanced methods of control of objects.
2 cl, 5 dwg
SUBSTANCE: 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.
EFFECT: higher accuracy, lower power consumption, longer life.
1 dwg, 2 tbl
SUBSTANCE: 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.
EFFECT: higher accuracy of boresight angular velocity adjustment.
SUBSTANCE: 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.
EFFECT: lower power consumption, decreased weight, minimised quantity of jet engines.
SUBSTANCE: 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.
EFFECT: higher accuracy, ruled out longitude period correction, reduced power consumption.
SUBSTANCE: 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.
EFFECT: increased and stable thrust, accurate computation of correction parameters.
SUBSTANCE: 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.
EFFECT: higher accuracy.
SUBSTANCE: 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.
EFFECT: higher accuracy of attitude plotting.
SUBSTANCE: 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.
EFFECT: comprehensive and accurate determination of angular velocity of rotation and orientation.
2 cl, 2 dwg
SUBSTANCE: 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.
EFFECT: lower fuel consumption.
SUBSTANCE: 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.
EFFECT: providing enough SB energy receipt on orbits with maximum eclipse period.
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.