RU2779783C2 - Rocket and space system for highly detailed remote sounding of the earth in the visible and (or) infrared observation range - Google Patents

Abstract

FIELD: rocket and space technology.

SUBSTANCE: invention relates to the field of rocket and space technology, more specifically, to remote sounding of the Earth. Rocket and space System (RSS) for highly detailed remote sounding of the Earth in the visible and/or infrared observation range includes a launch vehicle for delivering space vehicles (SV), with a plane of attachment to the LV perpendicular to the longitudinal axis of the SV, and placed in a system for attachment and detachment from the LV, into the insertion orbit. The composition of the SV includes electrooptical surveillance equipment, adjustment propulsion unit, and means of turning the SV relative to the centre of inertia. The attachment and detachment system is made in the form of an adapter. Placed on the adapter are multiple detachment apparatus with SV installed thereon, the longitudinal axes whereof are parallel to the longitudinal axis of the LV. Racks are secured on the adapter between the SVs, connected with the platform whereon the SV detachment apparatus is secured. The SVs have a fuel reserve ensuring transfer of the SV after detachment from the LV from the insertion orbit to the working orbit with a minimum height of H. The maximum transverse dimension of the SV does not exceed 0.6 of the diameter D of the payload area of the LV. The maximum height of the SV does not exceed D, the aperture size of the electrooptical equipment d can be in the range from 0.11 D to 0.25 D.

EFFECT: ensured sounding by means of the RSS.

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Description

The invention relates to the field of space technology and can be used to create multi-satellite systems that provide remote sensing of the earth (ERS) with high resolution and a high frequency of the possibility of observing any object on the surface of the planet in a wide range of latitudes.

A well-known space-rocket system (RS), including a launch vehicle (LV) for delivery into orbit for launching spacecraft (SC) placed in the attachment system and separation from the LV, each of which incorporates a corrective propulsion system (KDU ) and a means of turning the spacecraft around its center of inertia (RF patent No. 2428358). The fastening and separation system is a cone-shaped adapter mounted along the longitudinal axis of the launch vehicle on the launch vehicle adapter and the separation devices (SD) placed on it, on which the spacecraft is installed. In this case, each of the spacecraft is fixed on the CR by a plane parallel to the longitudinal axis of the spacecraft.

The system operates as follows. After the last stage of the launch vehicle enters the launch orbit, the spacecraft is separated from the launch vehicle. In this case, the spacecraft move at an angle to the longitudinal axis of the adapter, which eliminates the collision between the spacecraft during their separation. Then each of the spacecraft is oriented by its means of turning to the required positions and, due to the operation of the CPS, is transferred to the working orbit. In this case, an orbital structure is formed from several spacecraft launched by the launch vehicle. The described RCS makes it possible to dispense with the launch block, which reduces the cost of the RCS. This is justified in the case when the fuel consumption of the CPS during the delivery of the spacecraft to the working orbit is sufficiently small, for example, when deploying an orbital structure from spacecraft located in the same plane of the orbit.

However, the possibilities of using the described RCS are quite limited. For example, it is not possible to use spacecraft for remote sensing of the Earth in the visible and (or) infrared range with high resolution as part of such a RCS for the following reasons:

1. The use of a cone-shaped adapter placed along the axis of the launch vehicle leads to the fact that the spacecraft can be placed in an annular zone around the adapter, limited from the outside by the payload area of the launch vehicle, which limits the transverse dimensions of the spacecraft. At the same time, to obtain high-resolution information, one has to use a telescope with a very large diameter of the primary mirror. Accordingly, all supporting equipment of the spacecraft is assembled around the telescope, which leads to an increase in the transverse dimensions of the spacecraft. Since the telescope has an elongated shape, in combination with the use of the KDU as part of the spacecraft, this leads to an increase in the longitudinal dimensions of the spacecraft. All this limits the number of spacecraft that can be placed in the payload area of the launch vehicle. Of course, you can use a launch vehicle with a larger payload area, but a larger area will have a launch vehicle with a higher carrying capacity and, accordingly, a higher cost, which makes its use impractical.

2. For spacecraft equipped with a telescope of large diameter and considerable length, the most rational layout is in which the plane of installation of the spacecraft on the DO is perpendicular to the longitudinal axis of the telescope, which ensures optimal loading of the spacecraft structure, as well as its components and assemblies during ground operation, for example, during ground transport and launch into orbit. The layout of such a spacecraft with the placement of its installation plane parallel to its longitudinal axis will lead to irrational loading of the spacecraft and its systems and assemblies in ground operation and during launch into orbit, which will lead to an increase in the mass of the spacecraft. This also limits the number of spacecraft that can be delivered into orbit by one launch vehicle. In addition, such a design is less technologically advanced to manufacture, and this will lead to an increase in the cost of manufacturing the spacecraft.

3. Limiting the number of spacecraft deployed in working orbits reduces the range of latitudes in which a high frequency of observation of any object on the earth's surface can be ensured, even if the spacecraft can be turned for observation.

Known RCS, which includes a launch vehicle for delivery into orbit of the launch unit and spacecraft having a plane of attachment to the UO perpendicular to the longitudinal axis of the spacecraft and placed in the system of attachment and separation from the launch unit (RF patent No. 2293689).

The fastening and separation system is a cylindrical adapter, on which the UO with spacecraft attached to them are placed in the annular zone. In this case, the control units are set so that the longitudinal axes of the spacecraft are at a certain angle to the longitudinal axis of the launch vehicle. On the adapter, along its axis, coinciding with the longitudinal axis of the launch unit, a conical stand is installed, on which the UO is placed with the spacecraft fixed on it. In this case, the longitudinal axis of this spacecraft coincides with the longitudinal axis of the launch unit. Placement of the spacecraft on the adapter at a certain angle to the longitudinal axis of the launch unit is dictated by the need to avoid collision between the spacecraft and the stand during the separation of the spacecraft from the launch vehicle in orbit.

The RCC functions as follows. After the launch vehicle is delivered from the spacecraft to the launch orbit, the launch unit is separated from the launch vehicle, after which the launch unit sequentially places the spacecraft into working orbits. After delivery of each spacecraft to its working orbit, it is separated from the SO. Upon completion of the delivery of all spacecraft to their working orbits, the launch unit is slowed down and deorbited by an artificial Earth satellite.

This variant of constructing the RCS is inefficient for deploying an orbital structure consisting of a group of remote sensing satellites in the visible and (or) infrared range with high resolution due to the following factors:

1. Placing a group of remote sensing satellites in one plane can provide information about any object on the earth's surface in a certain range of latitudes with a certain periodicity. However, the use of the launch block in the RCS, with a group of spacecraft in the same plane, seems unjustified. The use of the launch block leads to a decrease in the LV payload area intended for spacecraft placement, and therefore limits the number of spacecraft that can be placed in this area. At the same time, in order to maintain the orbital structure of the spacecraft during the period of active existence, it is necessary to have a CPS in its composition and, therefore, by increasing the fuel supply in the CPS, using it, each of the spacecraft, moving from the insertion orbit to the working orbit, can take its place in orbital structure. It should also be noted that the use of the launch unit significantly increases the cost of deploying an orbital constellation, since the cost of the launch unit can reach 30% of the launch vehicle cost.

2. In this RCS, to place the spacecraft in the payload zone of the launch vehicle, a conical stand is used, to place the spacecraft along the longitudinal axis of the launch unit, and an annular zone above the adapter around the stand. At first glance, this allows you to place more spacecraft in the payload area. However, the presence of a stand makes it problematic to place a spacecraft with high-resolution remote sensing equipment in the annular zone due to significant transverse dimensions (similar to the RCS considered above). At the same time, the placement of the spacecraft in the annular zone at an angle to the longitudinal axis of the launch unit limits the length of the spacecraft.

Known RCS for remote sensing of the Earth in the visible and infrared range of observation, which includes a launch vehicle for delivery into orbit of the launch unit and spacecraft having an attachment plane to the UO perpendicular to the longitudinal axis of the spacecraft and placed in the system of attachment and separation from the launch unit, each of which has in in its composition optoelectronic observation equipment, corrective propulsion system and means of turning the spacecraft relative to its center of inertia (website of VNIIEM Corporation JSC www.vniiem.ru).

In this case, we are talking about the two Canopus-V spacecraft that are part of the RCS. The spacecraft is equipped with optical-electronic observation equipment in the visible and near infrared observation ranges. The fastening and detachment system is made in the form of an adapter installed on the launch block, on which the UO with spacecraft placed on them are fixed. In this case, the spacecraft are a passing load to a group of other spacecraft located on the launch block.

The RCC functions as follows. After the launch vehicle is delivered from the spacecraft to the launch orbit, the launch unit is separated from the launch vehicle, after which the launch unit sequentially places the spacecraft into working orbits. After delivery of each spacecraft to its working orbit, it is separated from the SO. In this case, spacecraft are placed in the same plane in a near circular orbit at an angle of 180° relative to each other.

This RCS has the following disadvantages:

1. The use of the payload area to accommodate the launch unit and other passing loads makes it possible to place only 2 remote sensing satellites in it. At the same time, the orbital structure consisting of two remote sensing satellites will make it possible to obtain information about any object on the earth's surface once a day in a narrow range of latitudes. This worsens the competitiveness of the RCS, which leads to a decrease in the number of users of the information received and, consequently, to a decrease in the economic efficiency of the RCS.

2. The use of the launch unit leads to an increase in the cost of deploying the spacecraft's orbital structure.

The task to be solved by the invention is to minimize the deployment time of the orbital constellation of remote sensing satellites of high resolution, providing a high frequency of observation of any object on the surface of the planet in a wide range of latitudes.

The problem is solved by the fact that the attachment and separation system is made in the form of an adapter, on which several separation devices are placed, with spacecraft installed on them, the longitudinal axes of which are parallel to the longitudinal axis of the launch vehicle or close to such a position, the separation devices are made with the possibility of separating the spacecraft at an angle not less than 15° to the longitudinal axis of the launch vehicle, and on the adapter between the spacecraft there are racks connected to the platform on which the separation device of the spacecraft is fixed, which is installed so that its longitudinal axis is parallel to the longitudinal axis of the launch vehicle or close to this position, the separation device is made with the possibility of separating the spacecraft installed on it in the direction of the longitudinal axis of the launch vehicle, the corrective propulsion systems of the spacecraft are equipped with a fuel reserve, which ensures the transfer of the spacecraft after separation from the launch vehicle from the insertion orbit to the working orbit with a minimum height H, the maximum transverse size of the spacecraft does not exceed 0.6 of the diameter D of the zone payload launch vehicle, the maximum height of the spacecraft does not exceed D , the size of the aperture of optoelectronic equipment d can be in the range from 0.11 D to 0.25 D, and the means of turning the spacecraft relative to its center of inertia provide the possibility of turning it in a plane perpendicular to the plane of the orbit by an angle α, the value of which is determined by the expression:

α≥arc sin R/(R+H)

where R is the radius of the Earth.

The essence of the invention is illustrated by graphic materials, where:

- in Fig. 1 shows the RCS in its original state;

- in Fig. 2 shows view A on the RCS from FIG. one;

- in Fig. 3 shows the RCS at the time of separation of the spacecraft from the MA;

- in Fig. 4 shows the calculation scheme for determining the angle α.

In the initial position, the RCS is located on the earth's surface and includes a launch vehicle 1, an attachment and separation system installed on the launch vehicle 1 under the fairing, consisting of an adapter 2, on which separation devices are located, each of which includes a hinged 3 on the adapter 2 frame 4, fixed relative to adapter 2 by locks 5, equipped with locks 6 of spacecraft 7, pushers and turn device (pushers and turn device are not shown conventionally). The fastening and separation system also includes racks 8 fixed on the adapter 2 and a platform 9 installed on them, on which the separation device is fixed, including a frame 10 fixed on the platform 9, equipped with fixation locks 6 of the spacecraft 7 and pushers (the pushers are conditionally not shown). The longitudinal axes of the spacecraft 7 placed on the adapter 2 are parallel to the longitudinal axis of the launch vehicle 1, and the longitudinal axis of the spacecraft 7 mounted on the platform 9 coincides with the longitudinal axis of the launch vehicle 1. Each of the spacecraft is equipped with optoelectronic equipment 11, a corrective propulsion system 12 and means for turning the spacecraft 7 in flight around its center of inertia (these tools are made in the form of power gyroscopes placed inside the spacecraft 7 and not shown conventionally). The parameter D denotes the diameter of the payload zone of the launch vehicle 1, the parameter T is the maximum transverse size of the spacecraft 7, and the parameter L is its maximum height. The parameter d denotes the diameter of the main mirror of the optoelectronic equipment 11, and the parameter f denotes the gap between the diameter d and the diameter of the inlet of the optoelectronic equipment 11.

The RCC functions as follows. After the launch of the launch vehicle, its last stage, together with the attachment and separation system, as well as with the spacecraft 7 placed in it, is delivered to the launch orbit. Then from LV 1 a command is issued to the locks 5, as a result of which the frames 4 of the separation devices are released, after which the turn devices turn the frames 4 together with the spacecraft 7 fixed on them to the required angle and fix in this position. After that, a command is issued to the locks 6, with which the KA 7 is attached to the frames 4 and frame 10, as a result of which the KA 7 is released relative to the frames 4 and frame 10, as a result of which, under the action of the pushers that are part of the UO, the KA 7 is separated from the frames 4 and frame 10 at the required speed.

Then each of the spacecraft 7, due to the work of the KDU, goes to its working orbit, using the fuel reserve intended for this. In this case, all spacecraft 7 are placed in the same plane in near circular orbits with a minimum height H, with equal angles between them. After that, spacecraft 7 begin to function, carrying out remote sensing of the Earth. To ensure the maximum performance of the SC 7, as well as a high frequency of observation in a wide range of latitudes of any objects on the surface of the planet by a group of SC 7 placed in orbit, each of the SC 7 has the ability, using its own means of turn, to turn around its center of inertia in a plane perpendicular to the plane of the orbit , to direct the longitudinal axis of the spacecraft 7 (the optical axis of optoelectronic equipment) to the required areas. In this case, the turn angle α, at which it is possible to observe the maximum number of regions, is determined by the expression α=arc sin R/(R+Н), and its value is determined from the right-angled triangle OAB (Fig. 4), where AB is directed tangentially to the surface Earth.

During the operation of the spacecraft 7 with the help of the CDU 12 is maintaining the required altitude of the orbit and maintaining the specified angles between the spacecraft 7 in the plane of the orbit.

Carrying out the design of the spacecraft in compliance with the geometric relationships of the design of the spacecraft T≤0.6 D; L≤D; 0.11D≤d≤0.25D makes it possible to place up to 4 remote sensing satellites in the visible and (or) infrared range of observation with high resolution in the payload area of a medium-class launch vehicle of the Soyuz type. The provision of separation devices placed on the adapter with the possibility of separating the spacecraft at an angle of at least 15° to the longitudinal axis of the launch vehicle makes it possible to avoid collision between the spacecraft, as well as between the spacecraft and elements of the attachment and separation system during the separation process. Due to this, it is possible to install spacecraft placed on the adapter with the orientation of their longitudinal axes parallel to the longitudinal axis of the launch vehicle. This makes it possible to give the spacecraft large longitudinal and transverse dimensions. At the same time, up to three spacecraft can be placed on the adapter.

At the same time, the placement of the SD with the spacecraft installed on it on a platform connected to the racks fixed on the adapter makes it possible to increase the number of spacecraft in the payload area of the launch vehicle. At the same time, the placement of the racks between the spacecraft ensures the shock-free separation of the spacecraft placed on the adapter.

The formation of the orbital structure of the spacecraft through the use of the capabilities of the CPS allows you to reduce the cost of deploying the orbital structure, due to the rejection of the use of the launch unit. At the same time, the absence of the launch block increases the volume of the payload area of the launch vehicle, which contributes to the placement of a larger number of spacecraft.

Deployment of an orbital structure from a large number of spacecraft with one launch of the launch vehicle reduces the costs required for the launch of one spacecraft. In addition, such a structure immediately makes it possible to obtain information about any object in a wide range of latitudes once a day, which makes it attractive to consumers.

The ability of the spacecraft turn means around its center of inertia to provide a turn angle in a plane perpendicular to the orbit plane to a position when its optical axis is directed tangentially to the horizon also increases the possibility of observing a large number of objects, and also expands the range of latitudes in which information is obtained about each time per day. It also helps to increase the economic efficiency of the RCS.

The development and manufacture of the proposed RCS at the current level of industrial production does not represent a fundamental difficulty, which will not create technical problems in the implementation of this solution.

Claims (3)

Rocket and space system (RSS) for highly detailed remote sensing of the Earth in the visible and / or infrared range of observation, including a launch vehicle for delivering spacecraft (SC) into orbit with an attachment plane to the launch vehicle perpendicular to the longitudinal axis of the SC, and placed in the attachment and separation system from the launch vehicle, each of which incorporates optical-electronic surveillance equipment, a corrective propulsion system and means of turning the spacecraft relative to its center of inertia, characterized in that the attachment and separation system is made in the form of an adapter, on which several separation devices with spacecraft installed on them, the longitudinal axes of which are parallel to the longitudinal axis of the launch vehicle or close to such a position, the separation devices are made with the possibility of separating the spacecraft at an angle of at least 15° to the longitudinal axis of the launch vehicle, and on the adapter between the spacecraft there are racks connected to the platform , on which the spacecraft separation device is fixed, which installed so that its longitudinal axis is parallel to the longitudinal axis of the launch vehicle or close to this position, the separation device is designed to separate the spacecraft mounted on it in the direction of the longitudinal axis of the launch vehicle, the corrective propulsion systems of the spacecraft are equipped with a fuel supply that ensures the transfer of the spacecraft after separation from the launch vehicle from orbit launching into a working orbit with a minimum altitude H, the maximum transverse size of the spacecraft does not exceed 0.6 of the diameter D of the payload zone of the launch vehicle, the maximum spacecraft height does not exceed D, the size of the aperture of optoelectronic equipment d can be in the range from 0.11 D to 0 ,25 D, and the means of turning the spacecraft relative to its center of inertia provide the possibility of turning it in a plane perpendicular to the plane of the orbit by an angle α, the value of which is determined by the expression α≥arc sin R/(R+H), where R is the radius of the Earth.

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