Determination of coordinates of emergency object by elevation angle and time doppler method

FIELD: space engineering; operation of spacecraft flying in orbit of artificial earth satellite, but for geostationary orbit, which are stabilized by rotation along vertical axis, as well as ground reception points.

SUBSTANCE: system used for realization of this method includes emergency object transmitter, onboard equipment of spacecraft and ground equipment of reception point. Onboard equipment of spacecraft includes horizon sensor, receiving antenna, comparison unit, receiver, Doppler frequency meter, blocking oscillator, two AND gates, two rectifiers, pulse generator, pulse counter, switching circuit, magnetic memory, transmitter, transmitting antenna, modulating code shaper, RF generator and power amplifier. Ground equipment of reception point includes receiving antenna, RF amplifier, two mixers, standard frequency unit, phase doubler, three narrow-band filters, phase scale-of-two circuit, phase detector, Doppler frequency meter, computer and recording unit. Proposed method consists in search of such space position of space object by receiving antenna when Doppler frequency of received signal is equal to zero. Measurement at this moment of angle between mechanical axle of receiving antenna and horizon axis is carried out referring to onboard receiving unit.

EFFECT: extended functional capabilities; enhanced accuracy of determination of spacecraft orbit elements; reduction of time required for search of emergency object.

5 dwg

 

The proposed method belongs to the 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 (patents of the Russian Federation№2155352, 2158003, 2040860, 2059423, 2174092, 2193990, 2201601, 2206902, 2226479, 2240950; U.S. patent No. 4161730, 4646090, 4947177; Scuba R.A. and other Companion at the helm. - Leningrad: Sudostroenie, 1989, s and others).

Known methods and systems closest to the proposed is "Elevation time-Doppler method for determining coordinates of the emergency object (patent RF №2174092, 64G 1/10, 1999), which is selected as a prototype.

According to a known method 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 determine the location of the emergency object.

Known JV the property provides an unambiguous definition and improving the accuracy of calculating the coordinates of the emergency object located on the Earth's surface, 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 method does 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 enhance the functionality of the method by clarifying the elements of the orbit of the spacecraft as it passes over the ground adoptive point.

The problem is solved by the fact that according to elevation and temporal Doppler method for determining coordinates of the emergency object placed on the surface of the Earth, with the spacecraft stabilized by rotation along the vertical axis, namely, that when the signal transmitter of the emergency object is displayed with a spacecraft strip on the surface of the Earth measured Doppler frequency no-request method, find the spatial location of the spacecraft at the time when the Doppler frequency of the received signal is equal to zero, measure at this point in time, the angle between the mechanical axis of the receiving antenna of the spacecraft and the axis of the horizon sensor with reference measurements to b is RevEmu time, compute the coordinates of the sub-satellite point at the time specified dimension, and the measurement is carried out twice and the coordinates of the two ground points and two measurements of the angle between the mechanical axis of the receiving antenna of the spacecraft and the axis of the horizon sensor to determine the location of the emergency object on the Earth's surface, on Board the spacecraft form a modulation code that contains information about the angle between the mechanical axis of the receiving antenna of the spacecraft and the axis of the horizon sensor, manipulate them phase-frequency vibrations, forming a complex signal with phase shift keying strengthen his power, radiate in the air, take to the ground receiving point when passing over it spacecraft, convert the received complex signal with phase shift keying frequency using a voltage of the first reference frequency, produce a voltage intermediate frequency, sequentially multiply and divide its phase two, emit harmonic oscillation of the intermediate frequency, is used for synchronous detection of the received complex signal with phase shift keying intermediate frequency, emit low-frequency voltage that is proportional to the modulating code, and use it to determine the location of the emergency object n is the surface of the Earth, at the same time the harmonic oscillation of the intermediate frequency transform on the frequency using the voltage of the second reference frequency, highlight the fluctuation of the Doppler frequency, determine the magnitude, sign, and zero Doppler frequency and use them to Refine the orbital elements of the spacecraft.

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. Block diagram of a system that implements the proposed method is presented 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 item.

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, schema matching And 8, the second input of which is connected with is that the output of the receiver 5, schema matching And 9, the second input of which is connected to the second output of the blocking oscillator 7, the 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, a magnetic 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.

Ground equipment receiving item 25 contains consistently included receiving antenna 26, an amplifier 27 high frequency, the first mixer 28, a second input 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 catalogobject to the second input of the computing unit 39.

The essence of the method consists in finding such 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. The coordinates of the 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 KAmay be on the of rawlin 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 the report of 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 condition 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 frequency stability 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 velocitydoes not 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 to what rotectio.

The proposed method is 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. Pulse transmitter 2 to the horizon at the moment of crossing the tracks KA produces a pulse which resets the counter to zero 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 y 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)

where Ucfcthat ϕc- the 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,)

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=kτ i.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

where f1=fc±Fd,

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

.

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)

where

To1the gain of the mixer;

- intermediate frequency;

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

where

To2- 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 Tcsignalwhile the spectral width 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 Δfwithinput

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)

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

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

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

where

To3- gain of the phase detector is 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 according to the coordinates of the two ground points and two measured angles α1and 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

where

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

The output of the second mixer 36 is formed fluctuation

where

Fp=±Fd+Fabout,

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

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 determines the orbital elements of the SPACECRAFT.

The method enables to determine the coordinates, to reduce the search time of the 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 method 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 method is expanded.

Elevation time-Doppler method for determining coordinates of the emergency object placed on the surface of the Earth, with the spacecraft stabilized by rotation along the vertical axis, namely, that when the signal transmitter of the emergency object is displayed with a spacecraft strip on the surface of the Earth measured Doppler frequency no-request method, find the club is to promote the position of the spacecraft at the moment when the Doppler frequency of the received signal is equal to zero, measure at this point in time, the angle between the mechanical axis of the receiving antenna of the spacecraft and the axis of the horizon sensor with reference measurements to the on-Board time, compute the coordinates of the sub-satellite point at the time specified dimension, and the measurement is carried out twice and the coordinates of the two ground points and two measurements of the angle between the mechanical axis of the receiving antenna of the spacecraft and the axis of the horizon sensor to determine the location of the emergency object on the Earth's surface, characterized in that on Board the spacecraft form a modulation code that contains information about the angle between the mechanical axis of the receiving antenna of the spacecraft and the axis of the horizon sensor, manipulate them phase-frequency vibrations, forming a complex signal with phase shift keying strengthen his power, radiate in the air, take to the ground receiving point during the passage over it of the spacecraft, convert the received complex signal with phase shift keying frequency using a voltage of the first reference frequency, produce a voltage intermediate frequency, sequentially multiply and divide its phase two, emit harmonic oscillation of the intermediate frequency, use it to sync the CSO detection of the received complex signal with phase shift keying intermediate frequency, emit low-frequency voltage that is proportional to the modulating code, and use it to determine the location of the emergency object on the Earth's surface, while the harmonic oscillation of the intermediate frequency transform on the frequency using the voltage of the second reference frequency, highlight the fluctuation of the Doppler frequency, determine the magnitude, sign, and zero Doppler frequency and use them to Refine the orbital elements of the spacecraft.



 

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EFFECT: enhanced efficiency; increased number of mini-satellites carried on adapter.

7 dwg

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

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

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

2 dwg

FIELD: control of group of satellites in one and the same orbit or in crossing longitude and latitude ranges of geostationary orbit.

SUBSTANCE: proposed method consists in measurement of parameters of satellite orbits, determination of orbital elements, comparison of them with required ones and performing of correcting maneuvers with the aid of thrusters. Satellite inclination vectors are brought to circular areas of their permissible change which are spaced apart so that angle between line connecting the end of vector with center of its circular area and direction to Sun should exceed right ascension of Sun by 180°. According to first version, vectors of satellite eccentricity are shifted to similar circular areas so that similar line lags behind direction to the Sun by half angular displacement of vector over circumference of its natural drift within circular area. Then, distances between satellites are changed within required limits compensating for quasi-secular increment of inclination vector and correcting eccentricity vector so that at passing the center of interval between point of circumference entry of its natural drift to its circular area and point of exit from this area, line connecting the center of this circumference and center of circular area coincide with direction to the Sun. In case circular area of permissible change of each eccentricity vector is close to circumference of its natural drift (second version), said line for this area is matched with direction to the Sun and no correction is made in this case.

EFFECT: saving of propellant for correction; protracted flight of satellites at safe distance.

3 cl, 13 dwg

FIELD: global satellite information systems.

SUBSTANCE: proposed system and method of organization of communication with the aid of this system includes injection of satellites into inclined elliptical orbits ensuring simplified tracking of satellites by means of ground tracking stations. Satellite orbits form pair of repeated routes (130, 140) embracing the earth's globe in projection on ground surface. Satellites are activated on each of these routes only on active arcs located considerably higher or lower relative to equator, thus emulating some essential characteristics of geostationary satellites. Parameters of satellite orbits are so set that final points of active arcs of two routes coincide; point at which active arc terminates in one route coincides with point where active arc starts on other route. Satellites placed on such active arcs are accepted by ground station located in satellite servicing zone as satellites slowly moving in one direction at rather large elevation angle. Their trajectory in celestial sphere has shape of closed teardrop line.

EFFECT: increased capacity of global satellite communication system with no interference in operation of geostationary satellites; simplifies procedure of tracking satellites.

39 cl, 15 dwg, 1 tbl

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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