Elevation doppler system for emergency object positioning

FIELD: space-system engineering.

SUBSTANCE: system can be used on the earth satellite vehicle in orbit except geostationary centers which are stabilised by rotation along the vertical axis, and the ground reception centers. The system contains an emergency object transmitter, spacecraft equipment and ground reception center equipment. Spacecraft equipment includes a horizon sensor, receiving antenna, comparison device, receiver, Doppler frequency meter, blocking-generator, two gating circuits, two valves, pulse generator, pulse counter, plug board chart, and magnetic storage device, transmitter, transferring antenna, modulation code generator, high-frequency generator and power amplifier. The ground reception center equipment includes reception antenna, high-frequency amplifier, two mixers, calibration frequency block, phase doublet, three narrowband filter, scale-of-two phase circuit, phase detector, Doppler frequency meter, computing unit and registration block.

EFFECT: it enlarges functional possibilities of the system.

5 dwg

 

The proposed system applies to space technology and can be used on the spacecraft that are in orbit of an artificial Earth satellite, in addition to geostationary stabilized by rotation along the vertical axis, and on the ground receiving stations.

Known methods and systems of determining the coordinates of the emergency object (RF patents №№2.155.352, 2.158.003, 2.040.860, 2.059.423, 2.174.092, 2.193.990, 2.201.601, 2.206.902, 2.226.479, 2.240.950; U.S. patent No. 4.161.730, 4.646.090, 4.947.177; Scuba R.A. and other Companion at the helm. - Leningrad: Sudostroenie, 1989, s and others).

Known methods and systems closest to the present invention is a system that implements "Elevation time-Doppler method for determining coordinates of the emergency object (patent RF №2.174.092, 64G 1/10, 1999), which is selected as a prototype.

According to the known system searches for such a spatial position of the receiving antenna of the satellite in the event of operation of the transmitter of the emergency object, when the Doppler frequency of the received signal is equal to zero. At this point, measure the angle between the axis of the receiving antenna and the axis of the horizon sensor. The coordinates of the ground point of the track of the spacecraft at the time of measurement are calculated. The measurement is carried out twice. The coordinates of the two ground points and two dimensions of the specified angle determines the location is their emergency object.

The known system provides a clear definition and a more accurate calculation of the coordinates of the emergency object placed on the surface of the Earth, as well as the expansion of the area of the viewing surface and increase the signal-to-noise ratio in the receiving radio.

However, the known system is not fully realizes its potential. It can be used to Refine the orbital elements of the spacecraft as it passes over the ground adoptive point.

An object of the invention is to expand the functionality of the system by clarifying the elements of the orbit of the spacecraft as it passes over the ground adoptive point.

The problem is solved in that the elevation-temporal Doppler system for determining the coordinates of the emergency object that contains, in accordance with the closest analogue, the transmitter of the emergency object, onboard equipment of spacecraft and ground-based equipment of the receiving point, and the axis of rotation of the spacecraft rejected from the local vertical, the spacecraft consists of a casing, a pulsed infrared horizon sensor, placed on the same axis opposite to the reception antenna, the mechanical axis which does not coincide with the axis of rotation of the spacecraft, onboard apparatus is of atur spacecraft consists of a series of the receiving antenna, receiver, a second input connected to the first output of the master oscillator, measuring the Doppler frequency, a second input connected to the second output of the master oscillator, comparator, the blocking oscillator, the first schema matching, the second input of which is connected to the second output of the receiver, the second circuit matches the second input of which is connected to the second output of the blocking oscillator, the first gate, the second input is through a pulse counter connected to the output of the pulse generator and horizon sensor, circuit switching, and magnetic storage device to the second output of the switching circuit serially connected transmitter and the transmitting antenna, to the third the output of the master oscillator connected in series temporary device and a second gate, a second input connected to a second output of the second circuit matches, and the output connected to the second input of the switching circuit differs from the closest analogue of the fact that the on-Board transmitter made in the form of sequentially connected to the second output of the switching circuit of the high-frequency generator, a phase manipulator, the second input is through the shaper modulating code is connected to the output of the magnetic storage device, and a power amplifier connected to a transmitting antenna, ground equipment receiving the unkt made in the form of cascaded receiving antenna, amplifier high frequency, a first mixer, a second input connected to the first output of the reference frequency amplifier intermediate, doubler phase, the first notch filter, divider phase two, the second notch filter, phase detector, a second input connected to the output of the amplifier intermediate frequency computing unit and the registration unit, and to output the second notch filter connected in series to the second mixer, a second input connected to the second output of the reference frequency, the third narrowband filter and measuring the Doppler frequency, the output of which is connected to a second input of the computing unit.

The geometric layout of the spacecraft 1, pulsed infrared sensor 2 horizon and the receiving antenna 3, is hosted on the same axis opposite to the sensor 2 of the horizon, shown in figure 1. The principle of determining the Doppler frequency shift of the transmitter SPACECRAFT is illustrated in figure 2. The dependence of the Doppler frequency with time is shown in figure 3. The structural scheme of the system is shown in figure 4. Timing diagrams explaining the operation of the system depicted in figure 5.

The system includes a transmitter 20 of the emergency object (emergency position-indicating radio beacon EPIRB), onboard equipment of spacecraft and ground-based equipment receiving p is NCTA.

Onboard equipment KA contains consistently included receiving antenna 3, a receiver 5, a second input connected to the first output of the oscillator 19, 6 meter Doppler frequency, the device 4 comparison inhibited the blocking generator 7, the first schema matching And 8, the second input of which is connected to the second output of the receiver 5, the second scheme matches And 9, the second input of which is connected to the second output of the blocking oscillator 7, the first valve 10, the second input is through the counter 13 pulses connected to the outputs of the sensor 2 of the horizon and the pulse generator 12, the switching diagram 14, the magnetic the storage device 15, the imaging unit 21 of the modulating code, phase arm 23, the second input is through the generator 22 high frequency is connected to the second output of the circuit 14 of the switching amplifier 24 power and transmitting antenna 17. The generator 22 high frequency, phase arm 23 and the amplifier 24 power form the transmitter 16.

To the third output of the oscillator 19 are connected in series temporary device 18 and the second valve 11, a second input connected to a second output of the second circuit matches 9, and the output connected to the second input circuit 14 of the communication.

Ground equipment receiving item 25 contains consistently included receiving antenna 26, an amplifier 27 high frequency, the first is mesial 28, the second input is connected to the first output unit 29 reference frequencies, the amplifier 30 intermediate frequency doubler 31 phases, the first narrow-band filter 32, the divider 33 phase two, the second narrowband filter 34, a phase detector 35, a second input connected to the output of the amplifier 30 of the intermediate frequency computing unit 39 unit 40 and the Desk. The output of the second narrowband filter 34 connected in series to the second mixer 36, a second input connected to the second output unit 29 reference frequencies, the third narrowband filter 37 and the meter 38 Doppler frequency, the output of which is connected to the second input of the computing unit 39.

The principle of the proposed system is to find a spatial position of the receiving antenna 3 KA, stabilized by rotation along the vertical axis, if the fact of operation of the transmitter 20 of the emergency object, when the Doppler frequency of the received signal is equal to zero, the measurement at this point in time, the angle between the mechanical axis of the receiving antenna 3 KA and the axis of the horizon with reference measurements to the onboard temporary device 18. Measurements are recorded in the magnetic storage device 15 and transmitted by radio to a ground receiving item 25. Coordinate of sub-satellite point at the time of the measurements is calculated. The measurements are performed at least twice. On the oordinates two ground points and two measured angles between the mechanical axis of the receiving antenna 3 KA and the axis of the horizon is determined by the location of the emergency object.

The principle of determining the parameters of the orbit SPACECRAFT using Doppler no-request system when on Board the SPACECRAFT is the transmitter, and on the Earth - measuring device illustrated in figure 2 and 3.

The Doppler frequency is determined based on the ratio

where λ - working wavelength,

r is the current distance from the SPACECRAFT to ground receiving item (0).

The motion vector KAcan be directed at any angle to the line radio. The relationship of the radial component of Vrmodule V is when you set a specific law of motion of the SPACECRAFT, which determines the form of the function r=r(t).

Let the observed trajectory of the SPACECRAFT S1-S2does not pass through the surface receiving the item Regarding, against which is counting distances. The shortest distance between the transmitter and receiver when the last point of S0is r0(figure 2). This so-called point of the beam. The time is counted from the moment t=0, corresponding to the passage of the SPACECRAFT through the point S1. The distance between the S1and S0denote by l0the moment passed the point of S0- t0.

The dependence of the Doppler frequency from time to time has the following form:

where the plus sign corresponds to the is the condition of 0≤ t≤t0(convergence), and the sign "minus" condition t0<t≤∞ (delete).

This expression shows that the Doppler frequency depends on both V and λand t, r0and l0. Moreover, the time dependent non-linear (figure 3).

On a linear plot near the inflection point

and then

Differentiating this expression with time, you can find the expression for the derivative of the Doppler frequency:

It is seen that the value ofdoes not depend on the beginning of observations (l0).

From the last expression it follows that knowing the speed V and the wavelength λand measuring the derivativeyou can find the shortest distance

The values of V and r0calculate the orbital elements of the SPACECRAFT.

Feature no-request method of measuring radial velocity is the need of the use of frequency standards. Assuming that the measurement error of the radial velocity should not exceed a tenth of a meter per second, valid relative instability of frequency standards at all times during operation of the system should not exceed 10-10. Such high requirements for stability frequent which you satisfy the quantum frequency standards.

In the no-request receiver system for measuring radial velocity is twice the frequency conversion. It is necessary because the relative value of Doppler shiftequal to the ratio of velocitynot exceed 10-4. Under these conditions, the selection of the Doppler shift at a single frequency conversion requires the use of circuits with a very high, almost unattainable quality.

The proposed system works as follows.

The translational motion of the spacecraft, the axis of rotation of which is rejected from the local vertical, moves the scan line pattern of the receiving antenna 3 and consistent view of the strip on the surface of the Earth along the orbit of the spacecraft. The rotation frequency of the AC is selected from the viewing conditions of the Earth's surface without a badge. To disambiguate the mechanical axis of the receiving antenna 3 KA is shifted relative to the axis of rotation at an angle βequal to the width of the directional receiving antenna.

In the initial state before getting signal from the transmitter 20 of the emergency object in the directivity pattern of the receiving antenna 3 at the output of receiver 5 no signal. The output schema matching And 8, 9 is zero. Impulse the th sensor 2 horizon at the moment of intersection of the route the AC generates a pulse, which clears the counter 13 pulses. The valves 10 and 11 are closed.

When the signal from the transmitter 20 of the emergency object in the viewed band on the Earth's surface measuring 6 starts measuring the Doppler frequency no-request method. When reaching the Doppler frequency value of zero, the mechanical axis of the receiving antenna 3 is located at the point of beam. At this point, the measured value of the angle between the axis of the sensor 2 of the horizon and the position of the mechanical axis of the receiving antenna 3 (angle α). Measurements are linked to the onboard temporary device 18.

Upon reaching the values of the Doppler frequency at the output of the meter 6, is equal to zero, opens the device 4 comparison and starts inhibited the blocking generator 7, the outputs of schema matching And 9, you receive the unit. Open the valves 10 and 11. Information about the angle α (the number of pulses stored in the counter 13 pulses) and the measurement time is recorded through the switching diagram 14 on the magnetic storage device 15 and to the input of the shaper 21 modulating code, where code is generated M(t) (figure 5, b), which is supplied to the first input of the phase manipulator 23. To the second input of the latter is fed a high-frequency oscillation from the output of the generator 22 high frequency (figure 5, a)

uc(t)=Uccos(2πfct+ϕc), 0≤t≤Tc ,

where Ucfcthat ϕc, Tc- amplitude, carrier frequency, initial phase, and the duration of high-frequency vibrations.

The output of the phase manipulator 23 is formed a complex signal with phase shift keying (QPSK) (figure 5,)

u1(t)=Uccos(2πfct+ϕk(t)+ϕc), 0≤t≤Tc,

where ϕk(t)={0, π} - manipulated component phases, reflecting the law of phase manipulation in accordance with the modulating code M(t) (figure 5, b), and ϕk(t)=const kτe<t<(k+1)τeand may change abruptly at t=kτei.e. at the boundaries between elementary parcels (K=1, 2,..., N-1);

τeN - the length and number of basic assumptions which form the signal duration Tc(Tc=Nτe);

which after amplification in the amplifier 23 power through the antenna 17 is radiated to the air. With the passage of the SPACECRAFT above the ground adoptive paragraph 25 above signal is captured by the receiving antenna 26. At the output of the receiver 27 in this scenario, you receive the signal

u2(t)=U2cos(2πf1t+ϕk(t)+ϕc), 0≤t≤Tc,

where f1=fc±F,

F- Doppler shift frequency due to the motion of the SPACECRAFT relative to the surface receiving the item, which item is admitted to the first input of the first mixer 28, to the second input of which is applied the voltage of the first reference frequency from the first output unit 29 reference frequencies

uE1(t)=UE1cos(2πfE1t+ϕE1).

At the output of mixer 28 are formed voltage Raman frequencies. The amplifier 30 is allocated to the intermediate voltage (differential) frequency (figure 5, g)

uCR(t)=UCRcos(2πfCRt+ϕk(t)+ϕCR), 0≤t≤Tc,

where;

To1the gain of the mixer;

fCR=f1-fE1=fc±F-fE1- intermediate frequency;

ϕCRcE1,

which is supplied to the information input of phase detector 35 and to the input of the doubler 31 phase. The output of the last formed a harmonic oscillation (figure 5, d)

u3(t)=U3cos(4πfCRt+2ϕCR), 0≤t≤Tc,

where

K2- transfer coefficient of the multiplier.

It should be noted that the doubler 31 phase represents a multiplier, two inlet of which a QPSK-signal intermediate frequency uCR(t).

As 2ϕk(t)={0,2π}in the specified oscillation manipulation phase is already absent. Spectrum width Δf2the second harmonic is determined by the duration of Tcis ignal whereas the width of the spectrum ΔfcFMN signal is determined by the duration of τehis elementary parceli.e. the spectral width Δf2the second harmonic signal is N times smaller than the width of the spectrum Δfcinput.

Therefore, by doubling the phase of the QPSK signal, its spectrum will minimize N times.

Harmonic oscillation u3(t) appears as a narrow-band filter 32 and is fed to the input of the divider 33 phase two, the output of which is formed of a harmonic oscillation (figure 5, e)

u4(t)=U4cos(2πfCRt+ϕCR), 0≤t≤Tc,

given narrowband filter 34 and is fed to the reference input of phase detector 35.

Therefore, the reference voltage required for synchronous detection of QPSK signals and operation of the phase detector, is allocated directly from the received QPSK signal.

The output of phase detector 35 is formed of a low-frequency voltage (figure 5, g)

un(t)=Uncosϕto(t)

where,

To3- gain of the phase detector,

proportional to the modulating code M(t) (figure 5, b), which is supplied to the first input of the computing unit 39. In the computing unit 39 by coordinates on the wow sub-satellite points and two measured angles α 1and α2between the mechanical axis of the receiving antenna of the spacecraft and the axis of the horizon is determined by the location of the emergency object.

At the same time harmonic oscillation u4(t) (figure 5, e) from the output of the narrowband filter is fed to the first input of the second mixer 36, to the second input of which is applied the voltage of the second reference frequency

uE2(t)=UE2cos(2πfE2t+ϕE2),

where fE2=fc-fE1-F0,

F0- frequency stand, which is introduced to determine the sign of the Doppler shift F.

The output of the second mixer 36 is formed fluctuation

up(t)=Upcos(2πFpt+ϕp),

where;

ϕpCRE2;

Fp=±F+F0,

given a narrow-band filter 37 and is fed to the input of the meter 38 Doppler frequency.

Depending, Fp>F0or Fp<F0determine the sign of the Doppler shift, and hence the direction of the radial velocity.

Knowing the velocity V and the wavelength λand measuring the derivativein the computing unit 39, determine the orbital elements of the SPACECRAFT.

The system allows to unambiguously determine the coordinates, to reduce the time of the CIP is the AC emergency object to increase the area of the viewing surface of the Earth due to the scanning of the receiving beam, to increase the ratio signal/noise radio through the use of the receiving antenna with a narrow beam pattern.

Thus, the proposed system is compared with the prototype, you can Refine the orbital elements of the spacecraft as it passes over the ground point. Thus the functionality of the system expanded.

Elevation-temporal Doppler system for determining the coordinates of the emergency object containing the transmitter of the emergency object, onboard equipment of spacecraft and ground-based equipment of the receiving point, and the axis of rotation of the spacecraft rejected from the local vertical, the spacecraft consists of a casing, a pulsed infrared horizon sensor, placed on the same axis opposite to the reception antenna, the mechanical axis which does not coincide with the axis of rotation of the spacecraft, onboard equipment of the spacecraft consists of cascaded receiving antenna, a receiver, a second input connected to the first output of the master oscillator, measuring the Doppler frequency, a second input connected to the second the output of the master oscillator, comparator, the blocking-generator is a, the first scheme matches the second input of which is connected to the second output of the receiver, the second circuit matches the second input of which is connected to the second output of the blocking oscillator, the first gate, the second input is through a pulse counter connected to the output of the pulse generator and horizon sensor, circuit switching, and magnetic storage device to the second output of the switching circuit serially connected transmitter and the transmitting antenna, to the third output of the master oscillator connected in series temporary device and a second gate, a second input connected to a second output of the second circuit matches, and the output connected to the second input of the switching circuit, characterized in that the on-Board transmitter made in the form of sequentially connected to the second output of the switching circuit of the high-frequency generator, a phase manipulator, the second input is through the shaper modulating code is connected to the output of the magnetic storage device, and a power amplifier connected to a transmitting antenna, ground equipment of the receiving item is made in the form of cascaded receiving antenna, amplifier high frequency, a first mixer, a second input connected to the first output of the reference frequency amplifier intermediate frequency doubler phase, per the CSO narrowband filter, divider phase two, the second notch filter, phase detector, a second input connected to the output of the amplifier intermediate frequency computing unit and the registration unit, and to output the second notch filter connected in series to the second mixer, a second input connected to the second output of the reference frequency, the third narrowband filter and measuring the Doppler frequency, the output of which is connected to a second input of the computing unit.



 

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FIELD: space engineering; spacecraft flying in earth artificial satellite orbit, but for geostationary orbit stabilized by rotation along vertical axis.

SUBSTANCE: system used for realization of this method includes spacecraft case, infra-red horizon pulse sensor, receiving antenna, comparison unit, receiver, Doppler frequency meter, biased blocking oscillator, two AND gates, two rectifiers, pulse generator, pulse counter, switching circuit, magnetic storage, transmitter, transmitting antenna, onboard timing device, onboard master oscillator and emergency object transmitter. Doppler frequency meter includes 90-deg phase shifter, two mixers, two difference frequency amplifiers, 180-deg phase inverter, two AND gates and reversible counter. Frequency of received oscillations is preliminarily reduced in two processing channels.

EFFECT: enhanced accuracy of determination of coordinates due to accurate measurement of minor magnitudes of Doppler frequency and recording its zero magnitude.

3 dwg

FIELD: power supply systems for high-orbit and geostationary orbit communication satellites whose orbits are corrected by means of electric jet engines.

SUBSTANCE: proposed method consists in determination of power requirements for each items of onboard using equipment for all spacecraft of cluster, electric jet engines inclusive. In case some items of onboard using equipment are not provided with electric energy at interval of dynamic mode with the aid of electric jet engines, this mode is changed-over for another permissible interval. In case of absence of permissible intervals, duplicate spacecraft of equivalent payload are selected. Items of equivalent using equipment of main spacecraft which are not provided with electric power are changed-over to duplicate spacecraft which are provided with required electric power. Spacecraft equipment items are changed-over till restoration of power supply on board main spacecraft (upon completion of dynamic modes of these spacecraft). Then control of power supply is performed for spacecraft of orbital cluster of later performance of dynamic modes. Main spacecraft are used as duplicate spacecraft.

EFFECT: reduced power requirements of cluster spacecraft; possibility of supplying power for additional items of using equipment.

1 dwg, 1 tbl

FIELD: space engineering.

SUBSTANCE: proposed method is based on continuous measurement of parameters of orbital space station motion. In addition to parameters of space station motion, its motion relative to center of mass is measured. Object of known shape and mass is separated from station in section of orbit when engines are disconnected and orientation of station is maintained with the aid of gyrodynes. Simultaneously parameters of motion of object separated from station in orbit and parameters of its rotary motion are measured. Mid-section of separated object is determined by measured parameters of rotary motion. Density of atmosphere is determined by this mid-section and measured parameters of motion of separated object. Then, mid-section area of orbital station is determined by measured parameters of its rotary motion and attitude of its movable parts. Then, mass of orbital station is determined from the respective mathematical expression by the mid-section area and atmosphere density and by measured parameters of motion. Proposed method does not require supply of calibrating and measuring pulses by power plant of cargo spacecraft coupled with station as distinguished from known methods.

EFFECT: saving of working medium; enhanced operational safety of cargo spacecraft.

FIELD: information satellite systems; forming global radio-navigational field for sea-going ships, ground, air and space vehicles.

SUBSTANCE: proposed system includes many low-orbit spacecraft whose number depends on conditions of global covering of access areas of users. Each spacecraft contains communication unit in addition to navigational equipment for communication of this spacecraft with two other spacecraft in its orbital plane and two spacecraft from adjacent orbital planes. Communication is performed in millimetric wave band absorbed by Earth atmosphere. At least one spacecraft is provided with high-accuracy synch generator. Thus spacecraft group is formed which is provided with noise immunity system of relaying and measuring radio lines connecting all spacecraft groups and navigational radio line covering upper hemisphere.

EFFECT: enhanced reliability and accuracy; enhanced noise immunity of data fed to users of satellite system.

1 dwg

FIELD: highly accurate ground and space systems.

SUBSTANCE: invention relates to frame-type stable-size bearing structures made of laminated polymeric composite material. Proposed integral framed construction made of laminated polymeric composite material consists of right angle ribs and their connecting units forming, together with ribs, a monolithic load-bearing skeleton made of layers of fibrous material impregnated with polymeric binder lying in plane of frame. Each rib and each unit has at least one layer of fibrous material fibers in which are orientated along longitudinal axis of rib, and layers of fibrous material fibers in which are orientated in directions corresponding to direction of longitudinal axes of other ribs.

EFFECT: increased stability and accuracy of positioning of frame units, reduced variations of thermomechical properties along length of ribs, provision of high accuracy of dimensions of articles.

2 cl, 2 dwg

FIELD: rocketry and space engineering; scientific and commercial fields.

SUBSTANCE: proposed method includes placing payloads on injection facility, launching the launch vehicle, separation of injection facility from launch vehicle and injection of injection facility into geocentric orbit where said payloads are separated from injection facility. Main payload is placed on injection facility directly of body of accompanying payload; this body combines its functions with functions of main load-bearing member of adapter system for placing the main payload. After separation of injection facility from launch vehicle, additional acceleration of injection facility is performed and injection facility is injected into reference orbit and then it is shifted to geocentric orbit where main and accompanying payloads are separated. Accompanying payload is separated from injection facility after main payload is at safe distance without waiting for complete turn of main payload. Spacecraft in facility injecting the artificial satellites into geocentric orbit are placed in succession on injection facility beginning with lower one. Main payload in form of one or several spacecraft is placed on body of lower spacecraft through separation device. Body of lower spacecraft combines its functions with functions of adapter load-bearing member for placing the main payload.

EFFECT: increased mass ratio of launch vehicle and injection facility; extended functional capabilities.

3 cl, 2 dwg

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