SU1764523A3 - Energy converter - Google Patents

Abstract

A method for disposal of radioactive waste in space provides for placing the radioactive waste in a cosmic module, launching the cosmic module with the radioactive waste into a near-earth orbit and then transferring it into a heliocentric orbit. The heliocentric orbit of disposal is conjugated with the orbit of one of the selected planets of the solar system, provided that the full revolution period on the heliocentric orbit, corresponding to the time of probable impact of the cosmic module with the selected planet, is no less than the time of decreasing of the radioactive radiation intensity of the waste up to a desired level. The plane of the disposal orbit is inclined to the ecliptic plane at an angle chosen from the condition of the passage of said orbit in relation to the other planets of the solar system at distances no lesser than the gravitation field sphere radii of those planets. <IMAGE>

Description

one

(21) 5003606/25

(22) 10/4/91

(46) 09/23/92. Bul No. 35

(76) Yu.V. Bryzhinsky, V.V.Ivanik, V.GKinelev, E.V.Oblonsky, A.D.Parashin,

V.F.Suranov, A.M. Khabarov, B.V.Shagov and

V.G. Kirsanov

(56) Patent of the USSR

No. 803874, cl. G 21 F 9/34, 1978. (54) METHOD FOR BURIAL OF RADIOACTIVE WASTE IN SPACE

(57) Use safe disposal of radioactive waste. Summary of the Invention: A method for the disposal of radioactive waste in space includes the disposal of radioactive waste in a space module, the withdrawal of a space module with waste into Earth orbit, and then

its transfer to the heliocentric burial orbit. The heliocentric burial orbit is mated with the orbit of one of the selected planets of the Solar System, and a full rotation period is established in the heliocentric orbit corresponding to the time of a possible meeting of the space module with the selected planet no less than the time for reducing the intensity of radioactive radiation of the waste to a predetermined level. The plane of the burial orbit is inclined to the ecliptic plane with an angle chosen from the condition of the passage of this orbit relative to the orbits of other planets of the Solar system at distances not less than the radii of the spheres of action of the gravitational fields of these planets. 33 il.

cl

The invention relates to protection against radioactive radiation by dumping radioactive waste in space.

With the expansion of the use of atomic energy plants and devices with radioactive isotopes, the amount of radioactive waste is growing, the elimination of which is a big problem.

Protection against radioactive radiation from waste by dumping them in the Earth, the ocean, is widespread. Thus, weakly active liquid and gaseous wastes are discharged to dissolve them in water into open water bodies or sea currents or are released into the atmosphere previously diluted with water or air, respectively. To remove industrial

radioactive wastes of average specific activity are concentrated and then enclosed in special sealed containers, which are installed for long (tens of years) storage in burial grounds - isolated underground rooms. However, this method does not completely exclude the possibility of their impact on the Earth’s biosphere during long-term storage, it requires significant construction and operation of burials, does not allow reliable (for hundreds of thousands of years) burial of radioactive waste (with a high specific activity) emit a large amount of heat, heating the surrounding rocks, which can lead to disruption of the ecology of the Earth.

XJ O N ate

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Methods are also known for the disposal of radioactive waste in space, including heliocentric circular orbit, located between the orbits of the Earth and Venus.

This method (adopted as a prototype) consists in placing radioactive waste in the space module and launching it into orbit, then forming a heliocentric disposal orbit by communicating two velocity pulses to the space module, the first of which is implemented when departing from Earth orbit to go to a migratory heliocentric elliptical orbit, and the second on this heliocentric orbit in its aphelion to go to the final heliocentric circular burial orbit.

However, if the first impulse of speed is not obtained or the second impulse of speed is not available to the space module, there is a risk that the space module will return to the gravitational field of the Earth after some time, which does not exclude the possibility of the module falling on its surface.

The implementation of such a method for the disposal of radioactive waste requires high energy and fuel costs and, accordingly, is associated with a decrease in the mass of the payload delivered by the space module to the disposal orbit.

The task of the invention is to create such a method of disposal of radioactive waste in space, which would ensure the withdrawal and retention of a space module with radioactive waste in a selected orbit for a specified time, ensuring the Earth’s ecological safety and preserving the natural environment on the planets of the Solar System while reducing energy costs.

The task is solved by the fact that in the method of disposal of radioactive waste in space, which consists in placing radioactive waste in a space module and placing it in a near-earth orbit with subsequent transfer of the space module to a heliocentric burial orbit, the latter is matched by at least one of the selected planets The solar system, on which a full rotation period is established, corresponding to the time of a possible meeting of the space module with the selected planet, is not less than the time of reduction and the radiation intensity of the radioactive waste to a predetermined level, and a plane which is inclined to the ecliptic plane at an angle selected from the conditions of the orbit passing other orbits relative to the planets of the solar system at a distance not less than the sphere radius of action of gravitational fields of these planets.

Conjugation of the heliocentric disposal orbit with the selected planet orbit improves the safety of waste disposal by reducing the number of speed pulses transmitted by the space module to perform transition maneuvers from near-earth orbit to a heliocentric disposal orbit, which significantly reduces energy costs for inter-orbital flights and increases the likelihood The success of the flight of the space module to the burial orbit. In addition, this makes it possible to reduce the size of the region of the space around the solar space used for the disposal of radioactive waste.

The established period of a complete rotation of the space module in the selected orbit, corresponding to a decrease

5 intensity of radioactive radiation of waste to a predetermined level, guaranteed to exclude the possibility of a premature meeting with the planet and thereby contributes to improving the safety of the disposal of radioactive waste.

The inclination of the plane of the selected burial orbit to the plane of the ecliptic ensures the safety of the planets of the solar system.

5 Before launching a space module into a heliocentric orbit, it is usually brought into an intermediate elliptical heliocentric orbit conjugated to the Earth’s orbit, then at the conjugate point

0 of these orbits translate the space module into a heliocentric burial orbit, which is combined with the Earth's orbit, and the distance of the space module to the Earth is not less than the radius

5 spheres of action of the gravitational floor of the Earth.

This contributes to the environmental safety of the landfill by using ballistic-resistant

0 orbit of the Earth, as the most studied space object of the Solar system, as well as reducing energy costs for the flight to a burial orbit,

In one of the options before the withdrawal

5 space module in a heliocentric orbital burial its output into an intermediate elliptical heliocentric orbit, conjugated with the Earth's orbit, which is implemented relative to the Earth in the form of a halo-orbit with a center located in the Earth's orbit, and any point of the halo-orbit is spaced no less than the radius of the sphere of action of the gravitational field of the Earth. This provides, along with environmental safety, the possibility of observing the disposal of radioactive waste in a disposal orbit for a period of time, and thus the possibility of monitoring and carrying out corrections in accordance with the results of the monitoring, which increases the reliability of disposal and safety for the Earth.

In another embodiment, before launching a space module into a heliocentric burial orbit, it is put into an intermediate elliptical orbit mating with the Earth's orbit and the orbit of another selected planet of the Solar System, through which the space module is directed into the gravitational field of the selected planet, and then translated to the heliocentric burial orbit by performing a perturbation maneuver in the gravitational field of the selected planet with a simultaneous change in the perihelion radius, eccentricity and loneni to the plane of the ecliptic. A heliocentric burial orbit is implemented as an ellipse with perihelion that is at a distance from the Sun at which solar radiation affecting radioactive waste converts them to a plasma state, or heliocentric burial orbit is realized with an eccentricity of at least one,

Such an option allows to reduce energy costs when delivering a space module with radioactive waste to a disposal orbit, where the waste is completely destroyed, transformed into a plasma, or along which it is sent outside the solar system.

Thus, the proposed method for the disposal of radioactive waste in space solves the problem of disposal with higher environmental safety for both the Earth and other planets of the Solar System while reducing energy costs.

Figure 1 schematically shows a general view of a spacecraft with a space module docked to it, carrying radioactive waste;

figure 2 is a schematic general view of a launch vehicle with a spacecraft installed on it;

on fig.Z - schematically the trajectory of the launch vehicle during the withdrawal of the space module in earth orbit; on

4 is a schematic of the initial portion of the flight path of a space module in a heliocentric orbit;

Fig. 5 schematically shows the orbits of the planets and the burial orbit external to the Earth's orbit (plan view);

figure 6 is a view along arrow A in figure 5;

7 - the same, the spatial position of the orbits;

FIG. 3 shows schematically the conjugation of the burial orbit and the orbits of two selected planets;

Fig. 9 schematically shows the orbits of the planets and the burial orbit internal to the Earth's orbit;

10 - the same, with the intersection of the orbit of the burial of the orbits of the planets;

in fig. 11 shows schematically the orbital sections of the burial and the Earth as they approach each other; Fig. 12 shows the functional dependence of the distance relative to the earth;

FIG. 13 shows the functional dependence of the parameter characterizing the level of radioactivity of the waste on time;

Fig. 14 shows the functional dependences of the number of turns of keeping the waste in a disposal orbit and the total impulse of speed against the period of a complete revolution in this orbit;

in fig. 15 is a schematic of the Earth's orbit and burial (in plan) when the module is placed in the Earth's orbit and the external location of the intermediate orbit of the module

FIG. 16 is the same in the orbital coordinate system relative to the Earth;

on Fig - the same as on Fig, with the internal location of the intermediate orbit of the module;

on Fig the same as on Fig, in the orbital coordinate system relative to the Earth;

in fig. 19 is a schematic of the orbits of the burial and the Earth (in plan) during deployment

module in halo orbit relative to the Earth;

on Fig - the same, in the orbital coordinate system relative to the Earth; on Fig - the same as on Fig, when

the rotated plane of the halo-orbit;

Fig, 22 is a view along arrow B in Fig. 21; Fig. 23 is a schematic of the module trajectory relative to the Earth in the implementation of a halo-orbit through one orbit of the module in an intermediate orbit;

Fig.24 is a view along arrow B in Fig.23; Figure 25 shows schematically the flight paths of the module relative to the Earth when implementing a halo-orbit around the Earth (in plan);

7

FIG. 26 is a view along arrow D in FIG. 25;

, fig.27 - schematically the planets' orbits and burial (plan view) during the flight of the module to the Sun; FIG. 28 - schematically the module perturbation maneuver trajectory in the gravitational field of the planet when implementing the flight scheme of FIG. 27;

on Fig - schematically the mutual position of the velocity vectors of the module and the planet when performing the maneuver on Fig;

on fig.ZO - the same as on fig.28, the spatial position of the trajectory;

FIG. 31 is the same as FIG. 29, the prostatic position of the velocity vectors, where C46, CAI is the normal to the orbits;

Fig. 32 shows schematically the orbits of the planets and their burial (in plan) when the module is flying beyond the limits of the Solar system;

FIG. 33 schematically shows the trajectory of the perturbation maneuver of the module in the gravitational field of the planet when implementing the flight scheme of FIG. 32;

.fig.34 - schematically the mutual position of the velocity vectors of the module and the planet when performing the maneuver of fig.32.

The proposed method for the disposal of radioactive waste into space is as follows.

Radioactive waste 1 (figure 1) to be buried in space is loaded into transport container 2 and delivered to the cosmodrome, where they are placed in space module 3, docked with transporting spacecraft 4 of known construction, which has accelerating stages 5.

Module 3 with a spacecraft 4 is installed in a freight transport container 6 (FIG. 2), docked with a launch vehicle 7 of known construction, and then taken from earth 8 (FIG. 3) to the near-earth orbit 9. Then, in section 10, acceleration 4) inform module 3 of the required increment of the flight speed at which module 3, moving along trajectory 11, goes beyond the limits 12 of the action of the gravitational field of the Earth 8. The radius r0 of the area 12 of action for the Earth and takes from 0.95 to 2.5 million km depending on the formulation of the problem.

The velocity vector V0 of flight of module 3 with the spacecraft 4 relative to the Earth at point 13 on the sphere 12 of the Earth’s gravitational field, adding to the velocity vector w0 of the Earth’s orbital motion around the Sun 14, determines the vector of departure velocity V0 module 3 s braid

eight

The medical device 4, which, together with the vector RO of the position of module 3 relative to the Sun, sets the parameters of the heliocentric burial orbit 15.

When the departure speed V0 satisfies the condition Vo2 Vo Wo (where Voa is the release speed in Earth’s orbit, at which the module leaves the solar system), orbit 15 (FIG. 5) has the shape of an ellipse external to Earth’s orbit 16. Orbit 15 mates with orbit 16 of the selected planet of the solar system (in this example, the Earth) by type of intersection or touch (when these orbits are separated from each other by the minimum distance, not more than r0), or by type of opposition (when the minimum the distance between these orbits is no more than 10 r0, but not less than r0).

An example of the implementation of the method in which the heliocentric disposal orbit 15 is conjugated to the Earth’s orbit 16 by type of touch is shown in FIGS. 5 to 7. In this case, the conjugation with the Earth’s orbit is chosen based on the minimum energy cost.

The position of the burial orbit 15 in space is determined by the angle i of inclination (Fig. 7) of its plane to the ecliptic plane (the plane of Earth's 16 orbit). This angle is chosen in advance when planning a flight to a burial orbit. This takes into account the inclination of the orbit 17 (Fig. 6) of another planet of the Solar System, which intersects the plane of the orbit 15, as well as the angular position of the line of nodes 18 (Fig. 5) of the orbit 17 relative to some constant direction 19 and the position of the Earth in its orbit 16 departure time module 3 (i.e. start date). The orbital plane 15 intersects the orbital plane 17 of another planet (for example, Mars) along line 20, and the orbit 17 itself intersects the orbit plane 15 at point 21 (in FIG. 5, the orbital sections 15 and 17, which are above the ecliptic plane, are shown as solid lines, and the corresponding orbital planes are shaded). The choice of the inclination angle i of the orbit 15 is made taking into account the evolution of the orbits of the planets of the Solar System and the orbit’s own evolution during the time the module 3 was delivered to the disposal orbit 15 of the burial until the intensity of the waste radioactive radiation decreases to an acceptable level. When choosing the angle of inclination i of orbit 15 on it and on orbit 17 of another planet, points 22 (23) and 24 (25) are determined, respectively, at which the distances between the orbits are minimal. These distances must be no more than the radius n of the sphere.

the action of the gravitational field of another planet, so that the perturbations of the orbit from this floor are insignificant. At the same time, the minimum distance from orbit 15 to orbit 26 of another planet (for example, Jupiter) is controlled, which is reached at point 27 and must be greater than the radius ha, the sphere of action of this planet for the same reason.

During the flight of module 3 in orbit 15, at the initial stage, it is assumed that orbit parameters are corrected at given points 28 to displace the perihelle of orbit 15 from orbit 16, to increase the accuracy of achieving the required disposal parameters and ensure the orbit perihelion phasing 15.

From the point of view of the Earth’s ecological safety during the evolution of its orbit 16 during thousands of years of a waste flight in orbit 15, it is advisable to choose the starting date so that departure from the Earth’s orbit takes place in its aphelion (do) in the case of an external position of the orbit 15. Phasing of the orbit 15, with the help of corrections, is carried out in the interests of the Earth’s ecological safety in order to exclude premature entry of module 3 into the scope of the Earth 8.

An example of a conjugation of a burial orbit with the orbits of two planets selected for this conjugation with the Earth orbit 16 (as opposed to) and with the orbit 26 of Jupiter (as intersection) is given in FIG. 8. To implement the method of disposal of radioactive waste, the same actions are performed as in the previous example, including carrying out a correction to raising the height of the perihelion of the orbit 15, which in Fig. 8 corresponds to position 15 before performing the correction and position 29 after the correction.

Note that the shading in Fig. 8 shows those half-planes of the burial orbit 15 and Jupiter orbits 26, which are located above the ecliptic plane.

When the departure speed satisfies the condition Wo V0 0, orbit 15 has the shape of an ellipse, internal to Earth orbit 16 (Fig.9. 10). And in these cases, the burial orbit 15 can be interfaced with the Earth orbit 16 by touch type (Fig. 9), by intersection (or opposition) (Fig. 10), including with orbit 26 by one of the selected inner planets of the Sun. system, In this case, similarly to the above variants of the method for the same environmental safety considerations, it is advisable to choose the start date in such a way that the departure from orbit

Earth occurred at its perihelion (Lo). Correction of the orbit 15 is carried out with similar objectives as in the previous cases. The methodological foundations of the selection of the parameters of the orbit 15 are illustrated in Figures 11-14 for the case of the external location of this orbit with respect to the Earth's orbit 16 upon conjugation of these orbits by the type of touch. Obviously, with such a conjugation

0 orbits if the corresponding actions are not taken in advance, module 3 can enter prematurely with waste (or waste 1, if module 3 is destroyed due to space flight factors) to the Earth’s scope, resulting in waste 1 or surface Earth, or move to a new heliocentric orbit with unpredictable

0 the consequences.

It is possible to eliminate such a situation, if in advance for each orbit of the 15 burial orbit with number n to analyze the relative position of module 3 and the Earth in

5 moments of their closest approaches and realize such a period P of a total rotation of module 3 in orbit 15 around the Sun 14 and such an initial phase position of module 3 in orbit 15 (including using corrections on the first orbits of the orbit), that a meeting with the Earth is possible, those. the module 3 (or waste 1) will enter the Earth’s sphere of action at the orbit of coil 15 with number n No, where No is determined by the requirements for

5 times Tplah keeping the waste in orbit 15 to a safe level of their radiation. Accepted into the Earth’s initial position (position 30) at the time of departure of module 3 from the Earth’s orbit 16 for the initial position (Fig. 11),

0 further, the current position 31 of the Earth as it moves around the Sun 14 in orbit 16 with a radius r can be characterized by the magnitude of the central angle F. The angular positions of the Earth at points 32, 33, 34 are the closest to the module 3 corresponding to the orbits of FIG. 11 the burials with the numbers n 2, N0 - 2, N0 - 1. In these positions, the distance D (n) from module 3 to the Earth must be at least g0.

0 Graphically, the dependences of the current distance L from the Earth to module 3 in the vicinity of their approach are shown as functions of the angle F in Fig. 12, where the dotted line and curved curves 1 (F) for orbit turns

5 directly precedes the entry of module 3 into the sphere of action of the Earth.

To select the period P of the total rotation of module 3 in orbit 15, it is necessary to use the dependences of the radiation characteristics of the waste on time t. As

such a characteristic can be, for example, the radiation activity of a waste is chosen, understood as the number of acts of spontaneous nuclear transformations in a given isotope of waste per unit of time, or the exposure dose rate, and the like. Typical dependences of the characteristic I of the waste radiation are shown in Fig. 13, where curves 35 and 36 correspond to different isotopes (for example, curium-245, americium-243).

FIG. 13 it can be seen that the maximum permissible level of residual radiation of 1 ct corresponds to the time Tmax of the decrease in the controlled radiation characteristic for wastes consisting of two isotopes, the masses of which are in a certain ratio.

Similarly, a certain intermediate level of the radiation parameter t and the corresponding time Tmax Tt of holding the waste in orbit 15 can be established, at which the possible local deviation of the radiation background will not lead to environmental disturbance in the event of the fall of radioactive waste fragments on the Earth.

Based on the data on the allowable time for holding waste outside the Earth, it is possible to determine the allowable number of complete revolutions of the No module in a disposal orbit, depending on the period P of one complete revolution in this orbit (Fig. 14).

Having determined using the appropriate calculations, the typical results of which are shown in Fig. 12, the dependence on the period P of the number of turns N in the burial orbit 15, through which module 3 falls into the sphere of the Earth’s trampling, we can distinguish the ranges of acceptable values of the periods P of the total burial in the orbit 15 (on Fig not shaded). For the periods found, the required energy expenditure (characteristic speed Ws) for reaching the disposal orbit by the method shown in FIG. 5 is determined from a typical curve 37.

The use of the proposed method of burial provides increased environmental safety of the Earth and the likelihood of waste being delivered to a burial orbit, reducing energy costs, which is achieved by reducing the number of interorbital transitions (active areas of operation of the engines of the transporting spacecraft that implement velocity pulses); the use of burial orbits closer to the Earth’s orbit, while keeping waste in orbit for a period of time after

which waste by its radioactivity becomes safe and can fall to Earth or another planet; providing the possibility of creating additional energy resources on the means of delivering radioactive waste to space, which allows in more cases of a failure of the cruise engine on these means to push module 3 with waste into orbit

0 burial of emergency, abnormal program.

It should be noted that in the vicinity of P values close to the period of the complete revolution of the Earth around the Sun, the implementation of the method of the planet requires additional operations that can be performed using one of the following options.

In the variant of combining the heliocentric orbit 15 of the burial with the heliocentric orbit 16 of the Earth (Fig. 15-18), before the module 3 is withdrawn with radioactive waste into the orbit 15, it is transferred to an intermediate elliptical orbit 38, conjugated to the orbit 16 of the Earth (according to

5 touch) and located relative to the orbit 16 of the Earth outside (Fig. 15. 16) or inside it (Fig. 17, 18). At least one revolution of module 3 in orbit 38 around the sun 14 translates module 3 into

0 orbit 15, which is combined with the earth orbit 16, and the distance L0 (Fig. 16, 18) of the module from the Earth establishes no less than the radius of the sphere of action of the gravitational field of the Earth. After entering orbit 15 on

At the initial stage of keeping the waste on it, orbits 15 are corrected to improve the accuracy of realizations of the indicated orbital parameters.

If on Fig, 17 shows the movement

0 module 3 in absolute coordinates relative to the Sun 14, then in Fig.16, 18 - in relative coordinates yx relative to the Earth, where the y axis is directed along the current radius-vector R of the Earth from the Sun, the axis

5 is perpendicular to it and directed in the direction of the motion of the Earth (in terms of speed W0). The origin of xy coincides with the center of the earth, so in these coordinates the earth’s orbit 16 is represented by a line 39 on the x axis

0 coordinates, the burial orbit 40 in the ideal case also coincides with the axes (module 3 is displayed as a point at a distance L0 from the origin). The motion path of module 3 along an intermediate elliptical

5 orbit 38 (fig. 15, 17) in relative coordinates xy has the form of cycloid 41 (figs 16, 18).

It should be noted that the transition from the intermediate orbit 38 to the disposal orbit 15 is carried out through one or

a greater number of full revolutions in orbit 38 by issuing a Vi speed speed brake (with an external tangency of orbits 16 and 38) or a V2 acceleration speed pulse (with an internal tangency of orbits 16 and 38), which are realized by spacecraft engines 4. In ideal cases, after the output of the speed pulses Vi or V2 module 3 must be at a distance LO along the x axis and have zero speed relative to the Earth.

Possible inaccuracies in the implementation of these pulses are compensated by successive corrections at points 28 at the initial stage of the module 3 flight along a burial orbit, which ensures the secular movement (drift) of module 3 relative to the Earth with the entry of the earth’s holding time to the safe radiation level.

Using this option will allow

reduce energy costs for delivering waste to a disposal orbit;

to improve the conditions for the disposal control using remote and contact technical means;

to dispose of waste in the long term (on demand), when the degree of their radiation hazard is reduced or when appropriate methods for their safe processing and use are found, and this disposal is possible with a decrease in the required energy consumption;

reduce the size of the space where the waste is located and, accordingly, increase the safety of spacecraft inter-orbit flights;

improve the ecological safety of the Earth.

In another variant of burial of waste in space, a burial orbit with a period P of a complete revolution around the Sun, close to the period of a complete rotation of the Earth in its orbit, is realized relative to the Earth in the form of a halo orbit.

In this case, before launching module 3 into the burial orbit 15, it is transferred to an intermediate elliptical orbit 38 (Fig. 19), conjugated with Earth orbit 16 (by type, for example, touch). Then module 3 is put into a heliocentric orbital 15 burial, which is implemented in the coordinate system relative to the Earth in the form of a halo orbit (Fig. 20-26) with a center located in the orbit 16 of the Earth (displayed on the x-axis of the coordinates as line 39 ),

moreover, any point of the halo-orbit 42 is separated from the Earth at a distance not less than the radius r0 of the sphere of action of its gravitational field. Orbit 15 is conjugated in this case to orbit 16 by the type of intersection. The transition from intermediate orbit 38 to orbit 15 can occur according to different flight patterns. Fig. 19 is a diagram in which, through a half-turn of the flight of module 3, in orbit 38, a speed pulse Uz is output to enter orbit 15, i.e. into halo-orbit 42 (Fig. 20), which is in the ecliptic plane and is separated from the sphere of the Earth's gravitational field at the minimum distance

.

In this embodiment, in the xy coordinate system relative to the Earth, the orbit 38 is displayed in the form of a cycloid 43, and the Earth's orbit 16 in the form of a line 39 on the x axis, as in Figs. 16, 18.

21, 22 is given a similar scheme, different from the previous one, in that the impulse speed pulse Uz for the transition from the intermediate orbit in the form of a cycloid 43 to the halo-orbit 42 of the burial simultaneously rotates the plane of the halo-orbit 42 at a given angle relative to the ecliptic (impulse vector The speeds at this transition have two components: Uzh and Uzg), as a result of which the halo-orbit 42 has a minimum distance from the surface of the Earth's sphere of action at point E, equal to Amin 0. Component z in the relative coordinate system yz in FIGS. 22, 24 , 26 corresponds to vector points in a systematic way kinematic Earth orbit y component along the radius vector of the current position of the Earth.

Fig. 23, 24 shows a flight diagram in which an ultrasound speed impulse is reported to module 3 after one complete revolution in orbit 38. In the diagrams of Figs 25, 26, an ultrasonic speed impulse is issued through one incomplete rotation, resulting in a halo orbit 42 in the relative coordinate system, the radius r0 of the sphere of action of the gravitational field of the Earth.

Note that on Fig-26 shows the ideal cases of the implementation of the method. In the real case, the position of the halo-orbit 42 is chosen, in which, taking into account corrections in the course of the secular motion relative to the Earth, its possible contact with the sphere of the Earth's action takes place not earlier than the established time for keeping the waste in a burial orbit. It is advisable to deflect the halo-orbit 42 from the ecliptic plane at some angle, which depends on the parameters of the halo-orbit, the radius of the sphere of action of the planet and which provides a gap (minimum distance) Lpsh 0.

The use of this option will improve the conditions for controlling the disposal, dispose of waste (on demand) with a decrease in energy consumption; reduce the size of the landfill area; to increase the environmental safety of the Earth and the safety of interorbital flights of spacecraft.

The proposed method also includes more complex options for its implementation, shown in FIG. 27 through 34, which are realized when the intermediate orbit 38 is interfaced with the orbit 26 of another selected planet. At the same time, after module 3 is brought into intermediate orbit 38 (figs. 27, 32), it is directed along it into the sphere of action of another planet (for example, Jupiter) with the help of several corrections (corrective velocity impulses). Module 3 enters the gravitational field of this planet at the moment when the planet is at the point

44, and leaves it when the planet reaches its point during its orbit

45. Within the radius gz of the gravitational field of this planet, a perturbation maneuver is performed, at which module 3, intersects the sphere of action of this planet at points 46, 47 (Fig. 28) and moves in orbit 48, is transferred to a heliocentric burial orbit 49 with by simultaneously changing the perihelion radius, eccentricity e, and inclination angle i to the ecliptic plane relative to intermediate orbit 38. During the maneuver, module 3 can be imparted additional velocity impulses. The required parameters of the orbit 49 are formed in the sphere of action of the planet, mainly by selecting the position of the entry point 46 (Fig. 28, 30, 33) and the target range d (d is the distance from the vector line Vi of the relative speed of entry of module 3 to the center of the planets ). Geometrically, the perturbation maneuver is shown in Figs.28-31 and Fig.33, 34, where the y-axis is directed along the radius-vector RP of the position of the planet relative to the Sun 14, the x-axis is in the plane of the orbit of the planet in the direction of its orbital velocity Vp perpendicular to the radius vector Rp, z axis perpendicular to the orbit plane of the selected planet in the direction of the vector of the kinetic module of the moment of the orbit. As a result of the rotation angle cp, relative to the vector Vi, the relative velocity V2, when the gravitational field of this planet leaves the sphere of action, the absolute velocity vector of module 3 changes: the velocity vector at entry point 46 changes by the velocity vector V47 at exit point 47. At the same time, the radius-vector changes, respectively, with the positions of points 44 and 45, changing the position vector of module 3 relative to the Sun. The magnitude of the rotation angle tp of the velocity vector is determined by the sighting range d and the magnitude of the velocity is determined by the sighting range d and the magnitude of the velocity Vne: when d gz is the angle ip 0, when d is 0 the angle ij). In the case when in the intermediate orbit 38 the velocity pulses are not applied to module 3,

the values of Vi and V2 are equal.

Thus, the implementation of the corresponding parameters of the orbit 38 (R46 and / 4b vectors), the targeting range d and the guidance module 3 to the corresponding entry point 46 in

the sphere of action of the selected planet, the module can be transferred from the intermediate orbit 38 or to the elliptical heliocentric orbit of the burial 49 (Fig. 27) with a perihelion spaced 14 from the Sun

the distance at which exposure to solar radiation converts radioactive waste to a plasma state (at point 50), or to a parabolic (when e 1) or hyperbolic (when e 1) heliocentric orbit 51 (fig. 32) burial with an exit in flight module 3 with waste outside the solar system into interstellar space.

This option is characterized by the fact that

when it is carried out, the waste is destroyed completely (when flying to the Sun) or irreversibly removed from the Solar System, ensuring their natural deactivation in flight, which will last in interstellar space for many thousands of crews. When flying to the Sun, the position of the perihelion of the orbit 49 is implemented as such in order to ensure the destruction of the module 3 structure (melting and evaporation) and the conversion of radioactive waste into plasma without falling into the lower layers of the heliosphere (solar atmosphere) and on the sun surface. The specific value of the conditional orbit perihelle 49 depends on

the material of construction of the module, its layout, the material of the container where the waste is located, itself from the composition and form of the waste fragments, and also the filler - retarder surrounding the waste in the container.

The considered variant makes it possible to ensure high safety of burial, since in the event of a shortage of the required departure speed Vo or the impossibility of performing a module 28 correction 28 to target another selected planet, it is possible to execute a backup flight program in an intermediate burial orbit 38, which is formed by auxiliary engines spacecraft 4 and unspent fuel reserves (by simplified control logic).

Compared with the method of disposal of radioactive waste using the direct pattern of removal of module 3 on the Sun, which requires an increment of the speed Ws of the module when launching from Earth orbit by 24 kg / s, for the proposed option it is sufficient to dial 8 km / s, i.e. the energetically proposed option is significantly more profitable. A comparison of the direct scheme of introducing module 3 to the flight path outside the solar system and the proposed option also shows the possibility of gain in energy costs (the required total increment of speed Ws module 3 can be reduced by 2 km / s).

Claims (4)

1. A method for the disposal of radioactive waste in space, which consists in placing radioactive waste in a space module and putting it into a near-earth orbit, then transferring the space module to a heliocentric burial orbit, characterized in that the heliocentric burial orbit is orbited with an orbit. at least one of the selected planets of the Solar System, on which a full rotation period is established, corresponding to the time of a possible meeting of the space module with the selected planet no less than the time with izheni a radiation intensity of waste to a predetermined level, and a plane which is inclined to the ecliptic plane at an angle selected from the conditions of the orbit passing other orbits relative to the planets of the solar system at the distance x is not less than the sphere radius of action of gravitational fields of these planets. 2. A method according to claim 1, characterized in that before a space module is put into a heliocentric orbit, it is put into an intermediate elliptical heliocentric orbit, conjugated with the Earth’s orbit, then at the conjugation point of these orbits, the space module is transferred to a heliocentric burial orbit, which is combined with the Earth's orbit, and the distance of the space module from the Earth is not less than the radius spheres of action of the gravitational floor of the Earth. 3. The method according to claim 1, characterized in that before launching the space module into the heliocentric burial orbit, it is put into an intermediate elliptical heliocentric orbit, conjugated with the Earth's orbit, then the space module is put into the heliocentric burial orbit, which relative to the Earth, they are realized in the form of a halo-orbit with the center located on the Earth’s orbit, and any point of the halo-orbit is separated from the Earth at a distance not less than the radius of the sphere of action of the gravitational field Of the earth. 4. The method according to claim 1, characterized in that before launching the space module into a heliocentric orbit, its burial is placed at an intermediate elliptical orbit mating with the orbit of the Earth and the orbit of another selected planet of the solar system, in which the space module is directed into the sphere of action of the gravitational field of the selected the planets and then the space module are transferred to the heliocentric burial orbit by performing a perturbation maneuver in the gravitational field of the selected planet with a simultaneous change in the perihelion radius, eccentricity and inclination to the ecliptic plane. 5, the method according to claim 4, characterized in that the heliocentric burial orbit is implemented as an ellipse with perihelion spaced from the sun at a distance at which solar radiation affects the radioactive waste, translates them into a plasma state. 6, The method according to claim 4, wherein the heliocentric burial orbit is realized with an eccentricity of at least one. ////////////////. 5 4J eight Fig.Z ft 5 x h H /four FIG. four BUT 26 sixteen I fa) Fig.9 1764523 Fig.Yu Fi'11  L . h p - №, n0 Ttttvqx CD and 2/3 D 7 fy Nfar0) F Fi ft Vottf Wj f 7 B 14  M / FIG. Vyuj - Sch. FIG. 23  6t / de 42 Fig.gv 3 4 Z6 FIG. 17 38 FIG. 28 FIG JO Y Y "Y 9 4i 32 ; $ 7 # Schi 31 38 Fig.ZZ Fi.34