WO2012092933A1 - Structure spatiale de désorbitation auto-déployable - Google Patents

Structure spatiale de désorbitation auto-déployable Download PDF

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Publication number
WO2012092933A1
WO2012092933A1 PCT/DK2012/050009 DK2012050009W WO2012092933A1 WO 2012092933 A1 WO2012092933 A1 WO 2012092933A1 DK 2012050009 W DK2012050009 W DK 2012050009W WO 2012092933 A1 WO2012092933 A1 WO 2012092933A1
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WO
WIPO (PCT)
Prior art keywords
deorbiting
sdss
configuration
space structure
self
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Application number
PCT/DK2012/050009
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English (en)
Inventor
Anders Schmidt Kristensen
Lars Damkilde
Original Assignee
Aalborg Universitet
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Publication date
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Publication of WO2012092933A1 publication Critical patent/WO2012092933A1/fr

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    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/62—Systems for re-entry into the earth's atmosphere; Retarding or landing devices
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/222—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles for deploying structures between a stowed and deployed state
    • B64G1/2221—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles for deploying structures between a stowed and deployed state characterised by the manner of deployment
    • B64G1/2222—Folding
    • B64G1/2224—Folding about multiple axes
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/10—Artificial satellites; Systems of such satellites; Interplanetary vehicles

Definitions

  • the present invention relates to the field of satellite deorbiting.
  • the invention relates to space structures which facilitate deorbiting of satellites from Low Earth Orbit (LEO), i.e. altitude region between ca. 230-1000 km above ground level.
  • LEO Low Earth Orbit
  • the present invention moreover relates to a method for stowing space structures that can be released in a controlled way to facilitate deorbiting of satellites.
  • Rocket motors may be used to reduce the speed of the satellite provoking its fall and consequently burn up in the atmosphere.
  • the technology takes advantage of the development of missile technology which has now reached commercial level.
  • rocket The main advantage of rocket is the control over the dismantling process by precise angled firing.
  • rocket technology is extremely costly, very heavy, technically very complex, e.g. including keeping rocket fuel under pressure for up to 10 years before its use.
  • a mass of propellant equal to 20 % of the total spacecraft mass is needed, but the exact amount required varies with the mass-to-area ratio and the initial altitude of the spacecraft.
  • a retro-rocket to deorbit a satellite at 600 km altitude must carry 5 kg of fuel for each 100 kg satellite to ensure dismantling.
  • Other leading deorbiting strategies namely electrodynamic tethers and
  • An electrodynamic tether is a conductive tie extending between a deorbiting spacecraft and a deployed end-mass some distance, perhaps several kilometres, below. Tether technology therefore involves deploying and towing a conductive wire from satellite through Earth's magnetic field.
  • the conventional current flowing up the tether cuts the flux lines of Earth's magnetic field as a result of the orbital motion, which in turn produces a Lorentz force against the orbital motion of the spacecraft.
  • the interaction produces a propulsion force that can be used to both raise and lower the satellite altitude orbit depending on the current flow.
  • the deployment unit of the Tether system will remain largely dormant during the active life of the host spacecraft to which it is attached. It can be activated on a regular basis to check the status of the mission. When deorbit is required the system will deploy the end-mass to begin the deorbit procedure. Electrodynamic tethers have potential but require a high degree of active control both at deployment and during the deorbit procedure itself.
  • tether wires length will depend on the satellite's mass and orbit altitude. For a typical LEO satellite tether wires connecting the deployed mass will be more than 5 km long leading to a high danger of entanglement with the End of Life (EOL) spacecraft or collision with other operational satellites.
  • EOL End of Life
  • Aerodynamic drag enhancement relates to an increase of air resistance.
  • Air resistance also referred as "drag”
  • drag depends on the altitude as the atmosphere progressively becomes thinner and thinner with increasing altitude, with no definite boundary between the atmosphere and outer space.
  • a drag system can be effective up to approx. 1000 km altitude.
  • Satellites in LEO therefore experience an aerodynamic drag which is rather limited and the reduction of the satellite speed due to air resistance is very small.
  • a system which is able to enhance the aerodynamic drag, increases air resistance so that the speed of the satellite is reduced at a rate which deorbits the satellite within a desired time.
  • WO 2009/149713 relates to an aerodynamic drag enhancement device comprising a flexible element, wherein the flexible element comprises a plurality of magnetic elements, such as electromagnet so that upon application of electrical energy, the shape of the flexible element is controllable by means of magnetic forces.
  • an improved mechanism for unfolding a deorbiting space structure from a folded condition would also be advantageous, and in particular a simple, lightweight and/or reliable mechanism for unfolding such a deorbiting space structure connected to a spacecraft or a satellite when the spacecraft or satellite is outside the atmosphere, would be advantageous.
  • the invention preferably seeks to mitigate, alleviate or eliminate one or more of the above mentioned disadvantages singly or in any combination by providing a method for deorbiting a spacecraft or satellite in a safe, simple and reliable way.
  • SDSS self deployable deorbiting space structure
  • the deorbiting space structure is therefore self deployable in the sense that its characteristic folding provides the strain energy provoking the deployment without the need of addition of energy to the system.
  • the deployment of the SDSS spans at least one flexible frame supporting at least one flexible sheet that causes a reduction of the speed of the satellite to which it is attached leading to deorbiting of the satellite, due to aerodynamic drag of the increased area.
  • the speed of deorbiting is a function of the satellite speed, satellite orbit height and ratio between SDSS and satellite dimensions.
  • a self deployable deorbiting space structure comprising at least one flexible frame supporting at least one flexible sheet wherein the at least one flexible frame is twistable into at least two loops and the at least two loops are foldable into an overlapping configuration.
  • the flexible frame supporting a flexible sheet is therefore twisted in at least two loops and the at least two loops are folded into an overlapping configuration to store strain energy into a stressed and compact configuration, also referred to as pre-stressed structure.
  • the word “space” is herein generally used as defining planet's atmosphere, controlled or uncontrolled airspace, outer space or on objects in the outer space.
  • the at least one flexible frame may have any form which has the ability of storing energy in the frame by the method described by the invention and to be folded into a stressed and compact configuration.
  • the SDSS comprising at least one flexible frame supporting at least one flexible sheet is advantageous for deorbiting satellites from an orbit around Earth into the atmosphere.
  • the SDSS comprising at least one flexible frame supporting at least one flexible sheet is advantageous for deorbiting satellites from an orbit around Earth into the atmosphere.
  • the invention may provide a cheap and reliable system useful for small satellites which have no means for intentional re-entry into Earth's atmosphere.
  • the invention is particularly, but not exclusively, advantageous in providing a SDSS which is low cost, simple and independent from the carrier.
  • the SDSS can easily, quickly and effectively be integrated into all types of satellites.
  • SDSS herein disclosed ensures a high unfolded area in relation to the low volume and weight to be transported in space.
  • a method for stowing a self deployable deorbiting space structure comprising at least one flexible frame supporting at least one flexible sheet, the method comprising : deforming by twisting the at least one flexible frame into at least two loops and by folding the at least two loops into an overlapping configuration, thereby providing the self deployable deorbiting space structure into a stressed configuration and storing strain energy;
  • the method allows for storing in the stressed configuration of the energy needed for deployment.
  • This stressed configuration is obtained by twisting and folding the SDSS following the method described by embodiments of the invention. In this way the energy is introduced and stored in gravity conditions and released in non- gravity conditions upon deployment of the SDSS.
  • the invention may also be advantageous for reducing the size of SDSS before being lunched in space.
  • Reducing size of the SDSS is herein defined as reducing the space occupied by it, by reducing its dimensions, high, length, width, volume, cross section or area.
  • the SDSS stressed configuration can be further compressed into a compact configuration, so as to provide a better stowing.
  • the at least one flexible frame is twistable by 360°.
  • the twisting is by 360°.
  • the at least two loops lie on intersecting planes.
  • Intersecting planes are defined as planes that meet in a single line.
  • the at least one flexible frame is twistable by 180° and the at least two loops lie on the same plane.
  • the twisting is by 180° and the at least two loops lies on the same plane.
  • the flexible frame is therefore twisted and folded so that the two loops formed lie on the same plane.
  • the deformation by twisting turns the SDSS into a coil configuration, to which it follows folding of the twisted frame into a closed coil configuration having at least two intermeshing convolutions.
  • folding of the twisted frame into a closed coil configuration leads to three intermeshing convolutions.
  • folding of the at least two loops may lead to three intermeshing convolutions having a closed coil
  • Folding of the twisted frame into a closed coil configuration leading to three intermeshing convolutions may comprise moving two diametrically opposite holding points of the frame, which has been twisted, e.g. mutually rotated by 360°, towards each other along a straight line.
  • Closed is defined as having neither beginning nor end, i.e. endless.
  • the closed coil configuration is induced by the closed configuration of the SDSS.
  • the self deployable deorbiting space structure is connected to a carrier structure. In some embodiments according to the second aspect of the invention the self deployable deorbiting space structure is connected to a carrier structure.
  • the SDSS may be connected to structures which carry the SDSS, i.e. carrier structures.
  • carrier structures are satellites, spacecrafts, and spacestations.
  • spacecraft is meant to denote any a vehicle or device designed for spaceflight.
  • spaceflight is meant to denote an object which has been placed into orbit by human endeavour.
  • a satellite is thus an example of a spacecraft.
  • the self deployable deorbiting space structure further comprises connecting means for connecting the deorbiting structure to a carrier structure.
  • connection between the carrier structure and the SDSS has the characteristic of withstanding a high degree of torsion forces which will be experienced by the SDSS connected to the carrier structure once the SDSS has been deployed. Resistance to torsion stress is therefore an essential characteristic of the connection between the carrier structure and the SDSS.
  • a connection between the SDSS and the carrier structure may be obtained by welding, gluing or bolting means.
  • the means for connecting may be characterized by an element fixed on the carrier structure and a complementary element fixed to the SDSS. These elements because of their structural characteristics may be linked by applying a certain degree of pressure.
  • the flexible frame supporting at least one flexible sheet may be connected to the carrier structure by connecting means.
  • connecting means may be a snap connection.
  • a snap connection may be a snap device comprising engagement elements.
  • a carrier structure may comprise a female engagement element, corresponding to a male engagement element on the SDSS, so that a connection between the carrier structure and the SDSS can be obtained by coupling the female-male engagements, e.g. by means of pressure.
  • the location of the male and female engagement elements on the carrier structure and on the SDSS, as well as the use of other snap's based mechanism are within the person skilled in the art and remain within the scope of the invention.
  • connection between the carrier structure and the SDSS may comprises magnetic means.
  • the SDSS and the carrier structure may comprise at least a magnetic element so that connection between the carrier structure and the SDSS may be controlled by means of magnetic forces.
  • the SDSS may be contained into a container located on the external surface of the carrier structure. Within the container the SDSS may be further connected to the carrier structure by a connecting means, e.g. a rotational joint such as a hinge connecting the carrier structure and the flexible frame.
  • a connecting means e.g. a rotational joint such as a hinge connecting the carrier structure and the flexible frame.
  • the container may act as a physical/mechanical protection for the SDSS, so that accidental collision with space debris or interaction with cosmic radiations will not damage the SDSS before its deployment.
  • the container may also act as
  • the container may be also considered as a holding means as it may be designed to hold the SDSS in the stressed configuration and, upon opening, to allow the deployment of the SDSS.
  • the container may be surrounding the SDSS as a semi-flexible envelope which can rapidly open through an external input that can be generated remotely, such as a pyrotechnic means.
  • an external input such as a pyrotechnic means.
  • the self deployable deorbiting space structure further comprises means for holding the overlapping configuration.
  • the step of maintaining the deorbiting space structure in the stressed configuration comprises the use of means for holding the stressed configuration.
  • Means for holding comprise devices with the function of holding tight and securing the stressed configuration, i.e. the overlapping configuration, e.g. a coil configuration, so as to avoid undesired opening during the transport.
  • These devices may be clamps, clips, couplers, tube locks, magnetic locks which can be opened by either mechanical or electrical intervention, or by the use of pyrotechnic devices, such as small explosion charges, or by electrical resistance that can be burned.
  • Means for holding may also include a pouch, such as the container located on the external surface of the carrier structure, with the function of securing the stressed configuration during transport.
  • a pouch such as the container located on the external surface of the carrier structure, with the function of securing the stressed configuration during transport.
  • the opening of this pouch allows for self unfolding of the SDSS with no need of external intervention, as once removed the holding means, the frame returns into it initial configuration without needing active or external deployment means.
  • the SDSS may comprise at least one electromagnet as holding means.
  • the electromagnet may be controlled remotely though electrical inputs so as to release the SDSS from its stressed configuration.
  • the SDSS may comprise at least two magnets.
  • the mechanism behind the opening and closing of the holding means may be remotely controlled, e.g. by a signal sent from Earth.
  • the opening of the holding means may be triggered without an external intervention.
  • Means for holding are meant to keep the overlapping or stressed configuration of the SDSS until the EOL of the satellite.
  • a trigger mechanism may be required.
  • this trigger mechanism does not produce the deployment of the SDSS, which is caused by the release of the deformation energy previously stored.
  • the trigger mechanism simply releases the device that locks the stressed or overlapping configuration.
  • the opening of these holding means lies within the person skilled in the art.
  • the opening may be triggered by an electrical or a mechanical input.
  • a timer device may be configured to detect the EOL of the satellite and to respond to that by opening of the holding means.
  • the SDSS may be contained internally to the satellite.
  • a hatchway protected by a door may be open so that access to the SDSS is gained.
  • the opening of the door may be remotely triggered.
  • the door may be hinged.
  • the door may have the function of a lid so that once opened is freed in space. In this case, any
  • the deforming of the at least one flexible frame further comprises compressing the at least two loops into a compact configuration.
  • Compressing may comprise pressing, bending, folding or twisting so as to obtain a compact configuration.
  • Compact configuration is herein defined as a configuration where the loops in the overlapping configuration are closely and firmly united, e.g. by pressing the loops together. It is not necessary to compact the twisted and folded frame but it is advantageous as it substantially reduces the encumbrance of the SDSS.
  • the above described object and several other objects are intended to be obtained by a method for deorbiting a carrier structure, the method comprising : stowing a self deployable deorbiting space structure according to the second aspect of the invention; releasing the self deployable deorbiting space structure from the stressed configuration.
  • Carrier structures are therefore not only the structure carrying the SDSS into space but also the structures which will have to be deorbited, such as EOL satellites.
  • deorbiting is herein generally used as defining the variation of the orbit a carrier structure, such as a spacecraft or a satellite for dismantling purposed, such as destruction through burns up in the atmosphere or for decommissioning purposes such returning a spacecraft back to Earth from orbit.
  • the invention in that respect is particularly, but not exclusively, advantageous for providing a rapid deployment of the SDSS without the need of external
  • Another advantage of the method for deorbiting a carrier structure described is that the specific twisted and folded position for stowing avoids the unfolding problems as described in the prior art, e.g. entanglement as no wires are present.
  • the releasing of the self deployable deorbiting space structure from the stressed configuration is triggered by the release of the means for holding the stressed configuration.
  • the releasing of the SDSS from the stressed configuration deployment may occur in a single step.
  • the releasing of the SDSS from the stressed configuration may occur in two or more steps.
  • the SDSS in its stressed configuration may be firstly positioned at a certain distance from the carrier structure, depending on the connection means used for connecting the SDSS to the carrier structure. In this way the SDSS may be firstly released from its compact configuration to its stressed configuration. Secondly, the full deployment can be triggered by, for example, releasing the holding means holding the stressed configuration.
  • the release of the means for holding the stressed configuration is remotely activated.
  • Remotely activated is herein defined as activated at a distance, for example through a signal from Earth.
  • the release of the means for holding the stressed configuration may be a combination between signals sent from Earth and a timer device, present on the satellite, configured to detect the EOL of the carrier structure.
  • the released deorbiting space structure self restores its initial configuration without any external intervention.
  • the method further comprises releasing the SDSS from the stressed configuration. Once released the SDSS self restores its initial
  • the method according to the second aspect of the invention describes a way of deforming the SDSS, by twisting and folding, which allows for storing of the energy necessary for the unfolding so that virtually no external energy, e.g. by external intervention, is needed during the unfolding process.
  • the simple removal of the holding means allows for self restoring of the initial configuration of the SDSS. Once released the SDSS restores its initial configuration because it was folded into an unstable configuration, i.e. the stressed configuration, and the at least one flexible frame and at least one flexible sheet of the SDSS are made with materials with properties which allows for storage of the energy introduced through the folding.
  • External intervention is herein defined as an intervention, e.g. an act performed, to the frame by bodies which are not the flexible frame in itself.
  • an intervention e.g. an act performed
  • a deployment performed by an astronaut on the already released SDSS can be defined as an external intervention.
  • Another example of external intervention is the deployment of the SDSS into its initial configuration after its release by e.g. a boom.
  • the absence of an external intervention is particularly advantageous, as generally restoring of the initial operative configuration of SDSS connected to a satellite, may require deployment means.
  • the unfolding process involves no rotating or moving parts, and deployment of the SDSS occurs automatically when the holding means are released so that the release of the deformation energy stored in the stressed configuration produces the SDSS deployment.
  • the SDSS released from the stressed configuration restores its initial configuration with substantial modification of the carrier structure momentum.
  • Carrier structures connected to the SDSS have a specific momentum which needs to be preserved in order to avoid undesired deviation from their desired orbit until decommissioning is needed.
  • the deployment of the SDSS can be tested on Earth so that the exact deployment movement can be reproduced in space allowing for the desired modification of the carrier structure momentum.
  • the speed of unfolding of the SDSS can also be controlled by constructing the flexible frame of the SDSS using appropriate structural materials. Some structural materials or combination of materials and composites can be devised to exhibit high structural damping. This is advantageous in order to control the speed of unfolding so as to avoid undesired speed of deployment.
  • the at least one flexible sheet comprises elements.
  • the at least one flexible sheet maybe formed by a continuous sheet or by modules, i.e. elements arranged as to facilitate the twisting and folding process of SDSS.
  • modules i.e. elements arranged as to facilitate the twisting and folding process of SDSS.
  • These elements may be textile, metal foils, and composite or polymer films.
  • the flexible sheet consists of a sheet having a thick mesh, e.g. of textile, where voids are not present or are limited to the extent that are not negatively affecting the
  • the flexible sheet may also have properties like a high resistance to laceration.
  • the flexible sheet may deform so as to avoid lacerations or tiny holes may be produced in the sheet avoiding rough and jagged tear.
  • the flexible sheet comprises textile materials.
  • the elements are mounted on a substrate along reinforced folding lines, so as to facilitate the twisting and folding process of SDSS.
  • the SDSS may also have the function of a solar sail for providing driving force to a spacecraft or a satellite.
  • the invention provides a reliable system for unfolding the solar sail attached to a spacecraft, in that the system does not contain complex mechanical unfolding structures.
  • SDSS may be used for any other appropriate use and that it is not meant to be limited to the examples given here.
  • Figure 1 is a schematic diagram of a SDSS according to the invention when attached to a spacecraft.
  • Figure 2 is a perspective drawing of a SDSS according to the invention in its folded condition, when attached to a spacecraft.
  • Figures 3 and 4 are perspective drawings of a SDSS according to the invention in its unfolded condition, when attached to a spacecraft.
  • FIG. 5 is a schematic drawing of the SDSS according to the invention comprising a flexible frame and a flexible sheet.
  • Figure 6 is a drawing of the SDSS according to the invention in the unfolded configuration.
  • Figure 7a is a drawing of the SDSS according to the invention in the twisted configuration.
  • Figure 7b and figure 7c shows two steps of the folding of the SDSS from the twisted configuration.
  • Figure 7d shows the folded configuration of the SDSS in a compact configuration.
  • Figure 8a and 8b show the twisted and compact folded configuration of the SDSS with three loops, respectively.
  • Figure 8c is a drawing of the SDSS according to the invention in the twisted configuration where the two opposite side of the frame have been mutually rotated by 360°.
  • Figure 8d shows an intermediate state of the folding of the twisted configuration where the two opposite side of the frame have been mutually rotated by 360° leading to three intermeshing convolutions
  • Figure 8e shows the folding of the twisted configuration where the two opposite side of the frame have been mutually rotated by 360° leading to three
  • Figure 8f shows a way to reduce the diameter of two convolutions, thereby increasing the diameter of the third convolution so as to allow an easier further twisting and folding of the frame, according to the invention.
  • Figure 9 and 10 show examples of two types of flexible frames of a SDSS supporting a flexible sheet according to embodiments of the present invention.
  • the SDSS of the invention is a deorbiting system which can increase the effect of aerodynamic drag by increasing the surface area of the satellite to which it is attached, accelerating its removal from space after its end-of mission.
  • the SDSS provides a solution to spacecraft deorbiting using the already present aerodynamic drag. Instead of using a complex and heavy fuel- or reactant-based propulsion system for deorbiting the satellite, the SDSS may be used, since by increasing the area of the satellite system colliding with air molecules in the residual atmosphere of Earth, the orbit time is decreased.
  • a 1 kg pico satellite will stay 10- 100 years in low Earth orbit, e.g. at 600 km, if it has no deorbiting mechanism.
  • the cross sectional area of the satellite is increases at least by a factor of 10, thus bringing the satellite down in a matter of months.
  • FIG 1 is a schematic diagram of a SDSS 18 according to the invention when attached to a spacecraft, such as a satellite 20.
  • the SDSS 18 is advantageous for deorbiting the satellite 20 from an orbit around Earth into the atmosphere.
  • the SDSS 18 is generally stowed to be launch into space attached through a connection 19 to a satellite 20 in its folded position see also figure 2.
  • the SDSS 18 comprises a flexible frame 1 and a flexible sheet 2 as shown in figure 5.
  • the flexible frame 1 of the SDSS 18 support a flexible sheet 2 of a material such as a woven fabric or any other material with the properties of a flexible foil, e.g. a polyimide film which can remain stable in a wide range of temperatures, for example between -273° to +400 °C, i.e. Dupont Kapton foil.
  • the flexible sheet 2 may also be made of composite materials or reinforced composite materials with aramid fibres, such as Kevlar.
  • Metal foils may also be used as flexible sheet 2, e.g. aluminium foils.
  • the flexible frame 1 of the SDSS 18 is connected to a satellite 20 by connecting means 19.
  • the connecting means 19 may be arranged so as to keep the SDSS 18 at a distance from the satellite 20 when the SDSS is in its operative condition, e.g. unfolded.
  • the connecting means 19 may be at least one rotational joint located on the carrier and connected to the SDSS, e.g. to the flexible frame. Through the rotation of the SDSS along the rotational joint, the SDSS may be positioned at a distance from the carrier structure so as to allow for deployment. Once the EOL of the carrier structure is reached, the SDSS may swing out by 90° from the carrier structure along the hinge, such as a pivot hinge. The SDSS now located at a certain distance from the carrier structure, but still connected to it by the rotational joint, may be deployed by releasing the holding means.
  • the SDSS may be contained into a container located on the external surface of the carrier structure. Within the container the SDSS may be further connected to the carrier structure by a connecting means, e.g. a rotational joint such as a hinge connecting the flexible frame to the container or directly to the carrier structure.
  • a connecting means e.g. a rotational joint such as a hinge connecting the flexible frame to the container or directly to the carrier structure.
  • the rotation along the rotational joint and the release of the holding means may occur in a single step.
  • a hatchway protected by a door may be open and the SDSS may be hinged to the door.
  • the connecting means may be flexible wires between the carrier structure and the SDSS, providing a tie extending out of the carrier structure that allows deployment of the SDSS at a certain distance from the carrier structure.
  • This embodiment has the advantage of allowing for a deployed flexible sheet surface higher than the case when the SDSS is rigidly linked at a short distance from the carrier structure.
  • the SDSS 18 Once the SDSS 18 has been unfolded, the residual atmosphere will keep the fabric of the flexible sheet 2 stretched out giving the highest possible area exposed to the aerodynamic drag. Once unfolded, the SDSS 18 will cause the spacecraft to dissipate mechanical energy due to the drag forces and in effect decrease its orbital altitude towards Earth and into the dense atmosphere thereof, where the air friction will heat up the spacecraft 20 and the SDSS 18 until they incinerate.
  • This incineration of a spacecraft 20 is a known way of disposing spacecraft;
  • the present SDSS provides a simple and reliable mechanism and reduces or eliminates the need of pyro-active or heavy device has to be brought onboard the spacecraft for deploying the space structure.
  • Figure 2 is a perspective drawing of a SDSS according to the invention in its folded condition, when attached to a spacecraft 15.
  • Figure 2 shows one element 16, such as a flexible frame supporting a flexible sheet, in its folded configuration (for perspective reasons only one element of the SDSS is shown in figure 2).
  • FIGS 3 and 4 are perspective drawings of a SDSS according to the invention in its unfolded condition, when attached to a spacecraft 15.
  • the SDSS comprises two elements 16 which are shown unfolded.
  • the spacecraft 15, such as a LEO satellite following the deployment of the SDSS experience aerodynamic drag which causes a reduction of its speed.
  • the SDSS comprises two symmetrical elements 16.
  • the SDSS may comprise more than two elements which are not symmetrical.
  • the optimisation of the location and geometry of the elements of the SDSS as a function of the level of aerodynamic drag desired lies with the person skilled in the art.
  • the spacecraft 15 in figure 4 is moving in the direction of arrow 22 and therefore will experience aerodynamic resistance 17 due to residual atmosphere. Once the elements 16 are deployed, as shown in figure 4, the aerodynamic resistance experienced by satellite will increase as shown by arrows 23.
  • the invention also relates to an apparatus for returning a spacecraft, such as a satellite, from orbit, the apparatus comprising a SDSS according to the invention as well as connecting means for connecting the SDSS to the spacecraft.
  • the apparatus 21 of figure 1 together with the connecting means 19 may be seen as such an apparatus.
  • FIG. 1 shows a schematic drawing of the flexible frame 1 on which a flexible sheet 2 is supported.
  • the specific outline of the flexible frame 1 is linked to its function.
  • the flexible frame 1 may have any form which has the ability of storing energy in the frame by the method described by the invention and to be folded into a compact configuration. While in this embodiment the flexible frame 1 is shown circular in other embodiments it may assume different forms, e.g.
  • Figure 6 is a drawing of the flexible frame 1 of the SDSS in the unfolded configuration, the flexible sheet 2 is not shown to simplify the drawing.
  • the flexible frame may have the characteristics of being light in weight, thin, and highly elastic. It may have a very high yield strength point so that the frame deforms elastically and returns to its original shape, i.e. no plastical deformation is present, even when the high level of mechanical stress is applied.
  • the material choice for the flexible frame is within strong, flexible materials with damping properties which can retain strain or deformation energy in the folded position for at least 10 years after the satellite launch.
  • Example of materials that can be used to produce the flexible frame 1 are metal or metal composite such as aluminium or aluminium based alloys, titanium, zinc alloys or magnesium alloys, iron alloys such as steel.
  • Polymers, polymer composites or metal/fiber reinforced materials may be also used as flexible frame materials.
  • Substantial resistance to cosmic rays is also an important characteristic of the flexible frame 1.
  • Polymers or polymer composites maybe not be used as frame materials unless presenting sufficient resistance to cosmic rays degradation.
  • the material used, having diverse properties of strength and elasticity, causes different degree of
  • deformability of the frame i.e. twisting and folding
  • deformation of the frame i.e. twisting and folding
  • the SDSS 18 has to be exposed to aerodynamic resistance it will have to have a high resistance to torsion.
  • Figure 7a is a drawing of the flexible frame in the twisted configuration. In this figure and in the following the flexible sheet is not shown to simplify the drawing.
  • the flexible frame 1 may be hold at its diametrically opposed sides 3 and 4 and then twisted along the axis 5 formed between the two holding positions 3 and 4 as indicated by the arrow 6.
  • the frame is twisted by 180° so that the frame assumes a figure-of-eight shape or two loops 7 and 8 which lie in the same plane. This is particularly advantageous for obtaining a configuration with reduced hindrance, so as to allow the folding of the two loops into a compact and closed coil
  • Figure 7b and figure 7c show the folding of the flexible frame from the twisted configuration when the two convolutions lie in the same plane.
  • One convolution is folded following the arrow 9 in figure 7b so as to produce a coil configuration where the two intermeshing convolutions are overlapping as shown in figure 7d.
  • the coil configuration may be maintained by the presence of means for holding (not shown).
  • the coil configuration may not require a strong holding means.
  • a single or multiple strap thin wire or a burnable resistance or pyrotechnic devices are examples of holding means.
  • Other examples of holding means are pin lock type, magnetic lock system or electrically activated hook devices. Releasing of the holding means allows for restoration of the initial configuration of the flexible frame. The release will be initially slow but will increase its speed of opening proportionally to the elasticity of the frame. The releasing of the stressed configuration by removal of the holding means would be advantageously carried out virtually without introduction of energy in the structure.
  • the degree of twisting may be different, so that the two loops formed may lie in separate planes.
  • the at least two loops may lie on intersecting planes, e.g., perpendicular planes.
  • the twisting may continue so that formation of more convolutions is possible.
  • figure 8a shows the twisted configuration where three loops or convolutions (10, 11 and 12) lie on the same plane.
  • Figure 8b shows the folded compact configuration where the three convolutions formed are folded to create a closed coil, e.g. by interposing the third convolution between the first and the second convolution.
  • more convolutions can be formed with the advantage of reducing the coil diameter and therefore increasing the stowing capacity.
  • folding of the twisted frame into a closed coil configuration leads to three intermeshing convolutions.
  • Figure 8c is a drawing of the flexible frame in the twisted configuration.
  • the flexible frame 1 may be hold at its diametrically opposed sides or holding points 3 and 4. These sides are mutually rotated by 360°.
  • holding point 4 is twisted by 180° around axis 5 as indicated by the arrow 6 in a counter clockwise (CCW) direction
  • holding point 3 is twisted by 180° around axis 5 as indicated by the arrow 24 in an opposite clockwise (CW) direction.
  • CCW counter clockwise
  • CW clockwise
  • Other combination of degrees of twisting may be possible so that of the holding points 3 and 4 mutually rotate across axis 5 by 360°.
  • Figure 8d and 8e show the folding step of the flexible frame from the twisted configuration described above. Following the mutual rotation across axis 5 of the holding points 3 and 4 the flexible frame 1 is folded into a coil having three intermeshing convolutions. These three intermeshing convolutions have a diameter of about one third of the untwisted flexible frame 1. In this way three convolutions 26, 27 and 28 are formed, where, as shown in figure 8d the third convolution formed 27 is interposed between the first convolution 26 and second convolution 28, creating a closed coil.
  • Folding of the twisted frame into the closed coil configuration having the three intermeshing convolutions 26, 27 and 28 may be achieved by moving two diametrically opposite holding points 3 and 4 of the frame, which has been mutually rotated by 360°, towards each other along a straight line. In this way convolution 26 leads into convolutions 28, and convolution 28 leads into convolution 27 that leads back into convolution 26.
  • the three convolutions of the closed coil are then compressed in a compact configuration as shown in figure 8b.
  • the very same compact configuration of figure 8b may be achieved through the twisting as shown in figure 8a.
  • the compact configuration may be maintained by the presence of means for holding (not shown).
  • Increased stowing capacity may be achieved by further overlapping more convolutions reducing the closed coil diameter.
  • the compact configuration comprising the three convolutions as shown in figure 8b may be further compacted by subsequently: reducing the diameter of two convolutions, thereby increasing the diameter of the third convolution; deformation the third convolution by carrying out the twisting and folding as previously described, thereby achieving a compact configuration consisting of five intermeshing convolutions. These steps may be repeated so as to achieve an even more compact configuration, for example consisting of seven or nine intermeshing convolutions.
  • a single twisting by 180°, as shown in figure 7a may also be possible at this stage producing a compact configuration consisting of four intermeshing convolutions. By repeating these steps more compact configurations may be achieved
  • the diameter of two convolutions may be reduced by pulling the two opposite sides 29 and 30 of the third convolution towards the same direction, as shown by arrows 31 and 32, thereby increasing the diameter of the third convolution.
  • Figure 9 shows an example of a flexible frame 1 supporting a flexible sheet 2 which is produced in a continuous sheet.
  • Flexible sheets can be provided in materials which are light in weight, not brittle, elastic and can be folded without substantial loss of their mechanical, physical and optical properties.
  • the flexible sheet is formed by separate textile elements.
  • Figure 10 shows an embodiment where the flexible frame supports a flexible sheet which is formed by textile elements 13.
  • the substrate 14 and textile elements 13 may be made of different materials. This configuration using textile elements may be advantageous as more robust to the mechanical stress which occurs during the twisting and folding as the textile elements 13 may be mounted on a substrate 14 along reinforced folding lines.
  • the SDSS may be connected to a satellite, such as a Cubesat.
  • CubeSat is a type of miniaturized satellite for space research, usually has a volume of exactly one litre and weighs no more than 1.33 kilogram. Typically uses commercial off-the-shelf electronics components.
  • Tests to optimize the selection of materials for flexible frame and flexible sheet may be made on prototype models of reduced diameter, e.g. between 100-250 mm. In these tests the diameter is maintained in an appropriate aspect ratio and freezing and sand blasting system may be used to emulate the conditions in space. Tests condition may include specified extreme temperature variations, e.g. between +50° to -130° degrees Celsius, and mechanical stress, such as shocks and vibrations tests, to ensure that the SDSS may preserve its properties even after prolonged exposure to mechanical stress, such as from a rocket system. Materials for the flexible frame and the flexible sheet may be identified also though test in relation to strength, flexibility, spring action and use in space through the exposure to artificial solar radiation and to particles collision. For example, the SDSS may be exposed to sandblasting at different particle velocity and size both in the folded and unfolded state, to ensure that the SDSS can deploy and have the necessary life time.
  • Material for the flexible frame and the flexible sheet may be identified also through burning tests as after serving as SDSS they will have to burn in atmosphere during the deorbiting process.
  • Materials used for the SDSS may have to withstand hard space conditions for the length of the operational life of the satellite to which it is connected.
  • the materials used to produce the SDSS may be degradable in space, e.g. sensitive to UV radiation, temperature differences, cosmic radiation, as the SDSS will have to survive space condition only for the limited time required for the deorbiting process.
  • unfolding test may reveal degradation of the flexible frame or of the flexible sheet upon unfolding, e.g. burst of the flexible sheet due to shrinkage in the folded configuration.
  • the choice of the desired holding means may be base on the desired degree of control of the speed of deployment of the SDSS.
  • Tests for optimising the resistance to air for continuous flexible sheet or fro flexible sheet formed by several elements may be also performed on Earth.

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  • Engineering & Computer Science (AREA)
  • Remote Sensing (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Details Of Aerials (AREA)

Abstract

La présente invention se rapporte à un procédé et à un système permettant de faciliter la désorbitation de satellites d'une orbite terrestre basse (LEO). La structure spatiale de désorbitation auto-déployable (SDSS) (18) de l'invention fixée à un satellite (20) se déforme en une configuration tendue de manière à stocker l'énergie de déformation sur Terre. L'énergie de déformation est relâchée dans l'espace produisant un déploiement rapide de la SDSS qui augmente la superficie du satellite et augmente à son tour la traînée aérodynamique entrant en collision avec les molécules d'air de l'atmosphère résiduelle de la Terre, entraînant des variations de l'orbite de satellite.
PCT/DK2012/050009 2011-01-07 2012-01-06 Structure spatiale de désorbitation auto-déployable WO2012092933A1 (fr)

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Cited By (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN103887611A (zh) * 2014-02-28 2014-06-25 西安空间无线电技术研究所 一种大型网状天线网面卸载方法
CN105799956A (zh) * 2016-03-18 2016-07-27 南京理工大学 立方体卫星制动帆离轨装置
CN107539500A (zh) * 2017-09-01 2018-01-05 南京理工大学 一种立方体卫星制动帆离轨装置
CN110304274A (zh) * 2019-06-14 2019-10-08 苏州展驭长空空间技术有限公司 一种用于被动离轨的充气增阻球
CN110304271A (zh) * 2019-06-14 2019-10-08 苏州展驭长空空间技术有限公司 一种被动离轨立方体卫星
CN111619828A (zh) * 2020-04-20 2020-09-04 中国卫通集团股份有限公司 一种同步轨道卫星离轨的方法及装置
EP3770352A1 (fr) 2019-07-24 2021-01-27 Instytut Podstawowych Problemów Techniki Polskiej Akademii Nauk Concept de la structure sdt (tensegrite auto deployable) pour le levage rapide et precis des aerostats a helium, en particulier dans la stratosphere
EP3752423A4 (fr) * 2018-02-15 2021-11-17 L'garde Inc. Système de mise en prise et de désorbitation de débris spatiaux
CN117022685A (zh) * 2023-08-04 2023-11-10 南京理工大学 一种空间碎片无源驱动增阻离轨装置

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0666612A2 (fr) * 1994-02-04 1995-08-09 Orbital Sciences Corporation Structure hélicoidale à déploiment autonome
US5996940A (en) * 1997-07-07 1999-12-07 Hughes Electronics Corporation Apparatus and method for combined redundant deployment and launch locking of deployable satellite appendages
US20050183377A1 (en) * 2004-02-05 2005-08-25 Johnson Samuel A. Deployable and retractable space frame
EP1593597A1 (fr) * 2004-05-06 2005-11-09 EADS Astrium GmbH Structure déployable ou extensible avec longerons flexibles et éléments de support
DE102007018564A1 (de) * 2007-04-18 2008-10-30 Deutsches Zentrum für Luft- und Raumfahrt e.V. Verfahren und Vorrichtung zum Entfalten einer Membran im Weltraum
WO2009149713A1 (fr) 2008-06-10 2009-12-17 Gomspace Aps Commande de la forme d'un élément souple

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0666612A2 (fr) * 1994-02-04 1995-08-09 Orbital Sciences Corporation Structure hélicoidale à déploiment autonome
US5996940A (en) * 1997-07-07 1999-12-07 Hughes Electronics Corporation Apparatus and method for combined redundant deployment and launch locking of deployable satellite appendages
US20050183377A1 (en) * 2004-02-05 2005-08-25 Johnson Samuel A. Deployable and retractable space frame
EP1593597A1 (fr) * 2004-05-06 2005-11-09 EADS Astrium GmbH Structure déployable ou extensible avec longerons flexibles et éléments de support
DE102007018564A1 (de) * 2007-04-18 2008-10-30 Deutsches Zentrum für Luft- und Raumfahrt e.V. Verfahren und Vorrichtung zum Entfalten einer Membran im Weltraum
WO2009149713A1 (fr) 2008-06-10 2009-12-17 Gomspace Aps Commande de la forme d'un élément souple

Cited By (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN103887611A (zh) * 2014-02-28 2014-06-25 西安空间无线电技术研究所 一种大型网状天线网面卸载方法
CN103887611B (zh) * 2014-02-28 2016-05-04 西安空间无线电技术研究所 一种大型网状天线网面卸载方法
CN105799956A (zh) * 2016-03-18 2016-07-27 南京理工大学 立方体卫星制动帆离轨装置
CN105799956B (zh) * 2016-03-18 2017-12-12 南京理工大学 立方体卫星制动帆离轨装置
CN107539500A (zh) * 2017-09-01 2018-01-05 南京理工大学 一种立方体卫星制动帆离轨装置
CN107539500B (zh) * 2017-09-01 2024-05-03 南京理工大学 一种立方体卫星制动帆离轨装置
EP3752423A4 (fr) * 2018-02-15 2021-11-17 L'garde Inc. Système de mise en prise et de désorbitation de débris spatiaux
US11713141B2 (en) 2018-02-15 2023-08-01 L'garde, Inc. Space debris engagement and deorbit system
CN110304271A (zh) * 2019-06-14 2019-10-08 苏州展驭长空空间技术有限公司 一种被动离轨立方体卫星
CN110304274A (zh) * 2019-06-14 2019-10-08 苏州展驭长空空间技术有限公司 一种用于被动离轨的充气增阻球
EP3770352A1 (fr) 2019-07-24 2021-01-27 Instytut Podstawowych Problemów Techniki Polskiej Akademii Nauk Concept de la structure sdt (tensegrite auto deployable) pour le levage rapide et precis des aerostats a helium, en particulier dans la stratosphere
CN111619828A (zh) * 2020-04-20 2020-09-04 中国卫通集团股份有限公司 一种同步轨道卫星离轨的方法及装置
CN111619828B (zh) * 2020-04-20 2021-12-07 中国卫通集团股份有限公司 一种同步轨道卫星离轨的方法及装置
CN117022685A (zh) * 2023-08-04 2023-11-10 南京理工大学 一种空间碎片无源驱动增阻离轨装置

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