TECHNICAL FIELD

[0001] The present subject matter relates, in general, to techniques for de-orbiting spacecraft.

BACKGROUND OF THE INVENTION

[0002] Advancement in space exploration and communication technologies has resulted in increase of launch of spacecraft, such as satellites. However, with proliferation in space exploration and communication technologies, launch of multiple space vehicles or spacecraft has resulted in increase in unwanted objects orbiting the Earth, commonly called as Space debris. Space debris is junk/waste fragments which orbit around the earth and has been accumulated due to human space activity. Space debris consists of fragments due to deterioration of spacecraft surfaces, used non- reusable stages of rockets, non-functional satellites, or the like. For example, spacecraft, such as satellites launched in an orbit around Earth to conducts dedicated operations, such as research, data collection, surveillance, or the like have a predetermined life for which the spacecraft remains in the orbit. Once the operation has concluded, the satellite remains in their orbit thereby acting as a trash in space. In other cases, satellite may suffer malfunction or failure rendering the satellite as trash in space. In yet another case, the satellite may break down into smaller fragments resulting in accumulation of debris in space.

BRIEF DESCRIPTION OF DRAWINGS

[0003] The detailed description is described with reference to the accompanying figures. In the figures, the left-most digit(s) of a reference number identifies the figure in which the reference number first appears. The same numbers are used throughout the drawings to reference like features and components.

[0004] Fig. 1 illustrates a drag enhancement apparatus for a spacecraft, in accordance with an implementation of the present subject matter.

[0005] Fig. 2a illustrates an undeployed drag enhancement apparatus, in accordance with an implementation of the present subject matter.

[0006] Fig. 2b illustrates a deployed drag enhancement apparatus, in accordance with an implementation of the present subject matter.

[0007] Fig. 3a illustrates a schematic of the drag enhancement apparatus, in accordance with an implementation of the present subject matter.

[0008] Fig. 3b illustrates a spool of the drag enhancement apparatus, in accordance with an implementation of the present subject matter.

[0009] Fig. 3c illustrates a tuff of wires of the drag enhancement apparatus, in accordance with an implementation of the present subject matter.

[00010] Fig. 3d illustrates a control of the drag enhancement apparatus, in accordance with an implementation of the present subject matter.

[00011] Fig. 3e illustrates an electrostatic charge generator of the drag enhancement apparatus, in accordance with an implementation of the present subject matter.

[00012] Fig. 4 illustrates another implement of drag enhancement apparatus, in accordance with an implementation of the present subject matter. [00013] Fig. 5(a)-(d) illustrates another design of the spool, in accordance with an implementation of the present subject matter.

[00014] Fig. 6(a)-(f) illustrates another implement of the drag enhancement apparatus, in accordance with an implementation of the present subject matter.

[00015] Fig. 7 illustrates yet another implement of the drag enhancement apparatus, in accordance with an implementation of the present subject matter.

[00016] Fig. 8(a)-(d) illustrates another type of drag enhancement apparatus installed within the spacecraft, in accordance with an implementation of the present subject matter.

[00017] Fig. 9(a)-(e) illustrates different designs of a drag device, in accordance with an implementation of the present subject matter.

[00018] Fig. 10 illustrates a method of deploying the drag enhancement apparatus, in accordance with an implementation of the present subject matter.

[00019] Fig. 1 1 illustrates a method of operating a lid 204 release mechanism of the drag enhancement apparatus, in accordance with an implementation of the present subject matter.

DETAILED DESCRIPTION

[00020] Proliferation in space launches has resulted in accumulation of trash in space, commonly called as space debris. Generally, most spacecrafts are launched in either low Earth orbits (LEOs) having an altitude ranging up to 2000 kilometres from Earth surface or in lower Medium Earth Orbits (MEOs) having an altitude ranging from 2000 kilometres and above from Earth's surface. As a result, much of the accumulation of space debris occurs in these regions. Space debris present in their orbit poses a threat to spacecrafts launched in space. For instance, space debris orbits at a speed in excess 10km/sec and thus possess high kinetic energy. As a result, when space debris collide with functional spacecraft the collision can result in either damage/failure of the spacecraft. In such cases, the collision also adds to the existing debris and increase the number of collisions.

[00021] Various techniques are devised to reduce accumulation of space debris. One such technique is generally known as de-orbiting of spacecraft. De-orbiting is a technique of bringing down the spacecraft from its orbit at a rate faster than a rate at which the spacecraft drops from the orbit naturally and causing the spacecraft to drop altitude. De-orbiting culminates with re-entry of the spacecraft in Earth's atmosphere causing the spacecraft to either disintegrate and burn mid-air due to atmospheric resistance or causing the spacecraft to crash into sea or ocean to prevent damages on Earth or both. One technique of de-orbiting is the use of small rockets called thrusters in the spacecraft that adjusts the trajectory of the spacecraft towards Earth. However, incorporation of thruster warrants addition of additional propellant which in turn results in increase in payload of the spacecraft. Further, increase in payload results in increase in costs associated with launching the spacecraft. Other techniques include use of a sail like structure to create aerodynamic drag. Aerodynamic drag is known as a resistance experienced by an object travelling in air. However, use of drag sails has many limitations. First, the drag sails are needed to be large enough to create sufficient drag to change the trajectory of spacecraft orbiting at high speed. Use of smaller drag sails does not provide adequate aerodynamic drag to de- orbit the spacecraft. As a result, use of drag sails increases the operational cost of the spacecraft. Moreover, drag sail are not effective in de-orbiting the spacecraft orbiting in MEO since air present at those altitudes is very thin to create drag needed to de-orbit the spacecraft. Moreover, deployment of drag sails requires complex mechanical mechanisms that are prone to failure.

[00022] Yet another technique is based on the use of electrodynamic tether device that includes a long current carrying conductor extending from the spacecraft. Further, the current carrying conductor extending from the spacecraft experience forces by Earth Magnetic field thereby causing the spacecraft to change trajectory. Use of electrodynamic tether requires high electric current in order to generate forces needed to change the trajectory. Moreover, use of electrodynamic tether does not work properly since Earth Magnetic field is not constant and vary from time to time and region to region. As a result, electrodynamic tether is less effective in de-orbiting the spacecraft.

[00023] To this end, a new concept of de-orbiting the spacecraft that alleviates the limitations of conventional de-orbiting methods for spacecraft is provided. The present subject matter relates to the concept of providing effective de-orbiting that provides effective deorbiting to a spacecraft in both MEO and LEO without causing substantial increase in payload of the spacecraft. The concept of present subject matter makes use of a combination of aerodynamic drag and Coulomb drag to de-orbit the spacecraft.

[00024] Various embodiments of the present subject matter disclose apparatus and method of enhancing drag for a spacecraft. In accordance with one embodiment, the apparatus for enhancing drag for the spacecraft includes a spool and tuft of wires wound around the spool and deployable to form a drag device of predefined drag to create aerodynamic drag for the spacecraft. Further, the tuft of ultra-thin wires offers more cross sectional area for experiencing drag than that of a drag sail of equivalent mass and size. As a result, the tuft of wires achieves better aerodynamic drag than the conventional drag sail and at the same time lighter thereby adding less to the payload.

[00025] According to an aspect, the apparatus also includes a control module that deploys the tuft of wires to form the drag device. The control module, in one example, polarizes the tuft of wires and the spool causing a mutual repulsion thereby causing the tuft of wires to deploy away from the spool. In other words, the control module induces a charge of predetermined polarity to the spool 106 and the tuft of wires. Further, the control module allows the deployment of the tuft of wires resulting from the electrostatic repulsion to form the drag device that creates aerodynamic and Coulomb drag. Since, the deployment of the tuft of wires dues to electrostatic repulsion, a need of complex deployment systems is avoided. Moreover, polarized drag device interacts with the plasma/ ions present in high altitudes thereby creating additional drag, commonly called as Coulomb drag. Such drag is effective in MEO where the concentration of plasma/ions is comparatively more than in LEO. As a result, a synergistic effect of aerodynamic drag and Coulomb drag provides effective drag for the spacecraft.

[00026] These and other advantages of the present subject matter would be described in greater detail in conjunction with the following figures. While aspects of a drag enhancement apparatus can be implemented in any number of different configurations, the embodiments are described in the context of the following device(s) and method(s).

[00027] Fig. 1 illustrates an apparatus 100 for enhancing drag for a spacecraft 102, in accordance with one implementation of the present subject matter. The apparatus 100 includes a housing 104 attached to the host spacecraft 102. The apparatus 100 further includes a spool 106 and a tuft of wires 108 shown in Fig. 1 in deployed state. Further, the tuft of wires includes a proximal end 108-1 which is near to the spool and a distal end opposite to the proximal end 108-2. In one example, the proximal end 108-1 is connected to the spool by an anchor wire 1 10. The tuft of wires 108 once deployed forms a drag device 1 12 of a predetermined shape that creates a combination of aerodynamic drag and Coulomb drag for the spacecraft 102. The apparatus 100 can be employed for deorbiting spacecraft with their orbits entirely (both perigee & apogee) within 6000 km altitude range or those on elliptical/highly-eccentric orbits with their perigee lower than 6000 km altitude. During operation, the apparatus 100 is deployed when the spacecraft approaches the perigee. Further, deployment of the apparatus 100 of the spacecraft 102 to descend in successive revolutions around Earth resulting in orbital decay.

[00028] According to an aspect, at the distal end, a tail-like extension of wire can be present and also embedded within a simple flap/tag like structure as with the anchor wire 1 10. When in stowed state, the extension is held taut between the top flange of the spool 106 and curved plate 204-1 on the lid 204. This distal flap can serve 2 purposes - (1 ) avoid slippage of drag wire windings by holding them taut and securely in position to prevent the knotting or damage of tuft of wires 108 during launch. When the lid 204 opens, the flap gets released free and unwinding of wire tuft can happen; (2) Due to its large flat area, the tag like structure can experience better mutual repulsion force with the underlying windings around the spool barrel for a greater radially outward push. This facilitates easy initiation of unwinding and deployment of tuft of wires 108.

[00029] The apparatus 100 can be implemented in two ways. One of the way is a 'plug-and-play' deorbiting system, shown in Fig. 1 , that can be easily integrated to the spacecraft before launch, in particular for pico-, nano- and micro-satellites, for space debris mitigation to deorbit the spacecraft 102 either after their end of operational life (EOL) as a post-mission life disposal (PMLD) system or when the spacecraft becomes defunct due to on-board systems malfunction and is achieved by deploying the tuft of wires 108. The apparatus 100 module can be integrated with the spacecraft either externally or internally. Another implementation is by acting as a primary payload on- board a spacecraft that has micro-propulsion capabilities like electric- propulsion, etc. for removing space debris. In one example, the dedicated spacecraft can manoeuvre on orbit while in space and can reach hazardous space debris like spent-upper-stages of rockets, defunct satellites, etc., that is already present on orbit. Further, the dedicated spacecraft is provided with a grappling mechanism to grab on to these space debris objects using its arms/claws and then deploy the apparatus 100, thereby facilitating the rapid orbit-decay of the debris.

[00030] According to an embodiment, the apparatus 100 for enhancing drag provides Coulomb drag when the spacecraft 102 is orbiting in lower MEOs. As mentioned above, MEOs is a region that extends from an altitude of 2000 km and above. In such regions, density of air is very less and hence, aerodynamic drag is negligible. However, MEO is a region of plasma and highly charged particles called ions. Further, MEO also has a region called Inner Van Allen Radiation Belt (IVARB) that includes ions. In such orbits, the polarized tuft of wires 108, when deployed interact with the ions. In the illustrated embodiment, the tuft of wire is polarized with a charge same as the charge of ions present in MEO. As a result, when the tuft of wire is deployed from the orbiting spacecraft 102, the tuft of wires 108 repels with the like- charged particles of space plasma and ions in IVARB. Moreover, due to high speed of the tuft of the wires 108 travelling through the ions and the space plasma, there is a considerable amount of momentum loss of the tuft of wire thereby decelerating the spacecraft 102. As may be understood, greater density of space plasma and ion results in greater momentum loss of the tuft of wires, and in turn, of the spacecraft 102. Thus, the electrostatic resistance offered by the tuft of wire reduces the velocity of the spacecraft 102 that is travelling at orbital velocity. As the velocity drops, the spacecraft 102 starts descending from its orbit.

[00031] According to another aspect of the present subject matter, the apparatus 100 for enhancing drag provides Coulomb drag in addition to aerodynamic drag when the spacecraft 102 is orbiting in LEOs. As mentioned before, LEOs is a region that extends at an altitude of up to 2000 km from Earth surface. As the spacecraft 102 descends from MEOs to LEOs, the tuft of ultra-thin wires 108 experience drag from density of air gets thicker thereby providing aerodynamic drag to the tuft of wires 108. When the tuft of thin wires moves through the air, molecules in the air ram into the wires of the tuft thereby causing nanoscopic collision resulting in loss of momentum. The loss of momentum further exerts a tugging (pulling) force on the spacecraft in a direction opposite to motion of the spacecraft 102. Since the tuft of wires is anchored to the spacecraft 102, the tuft of wires forces the spacecraft to decelerate gradually resulting in deorbiting of the spacecraft 102. In addition to aerodynamic drag, the ions present in LEO also provide Coulomb drag further slowing the spacecraft 102.

[00032] It can be noted that aerodynamic drag effect will be dominant in the LEOs (up to 1000 km) while the Coulomb drag will be dominant in both LEOs and MEOs (from 1000km to 6000 km altitude). From above, it is understood, that the tuft of wires 108 makes use of both aerodynamic drag and Coulomb drag to de-orbit the spacecraft. Since the tuft of wires are electrostatically charged to a high potential, the tuft of wires 108 generates Coulomb drag force by ramming into ions present in space plasma in LEOs & MEOs, which accounts to plasma braking effect. This drag is incremented by addition of aerodynamic drag produced by their cross sectional / projected area perpendicular to the spacecraft direction of motion while in orbit. Both the aerodynamic drag & Coulomb drag act simultaneously as a 'hybrid drag' force to deorbit the spacecraft. Also, a magnitude of both the aerodynamic drag and Coulomb drag varies with changing altitude & inclination. Thus, aerodynamic drag & Coulomb drag can together cause greater orbit decay of the spacecraft in a given time than what each of them could accomplish individually & hence, results in the 'synergistic drag' effect. Thus, a combination of the aerodynamic drag and Coulomb drag provides a synergistic effect for rapid de-orbiting of the spacecraft 102 thus mitigating the issue of space debris.

[00033] Fig 2 (a) and (b) illustrates the apparatus 100, in accordance with one implementation of the present subject matter. Fig. 2 (a) illustrates the apparatus 100 in undeployed state while Fig. 2(b) illustrates the apparatus 100 in a partial deployed state. With reference to Fig. (a) and (b), the apparatus includes a housing 104 that houses the spool 106 and the tuft of the wires 108. Further, the housing 104 may include a trunk 202 and a lid 204. In one example, the lid 204 can be hinged to the trunk 202 with a set of torsion spring to facilitate the opening of the lid 204. The apparatus also includes a lid 204 release mechanism (not shown) that allows the opening of the lid 204. The lid 204 release mechanism may include, but not limited to, heat destructible wires that holds the lid 204 closed until heated by a heating device of the lid 204 release mechanism.

[00034] According to illustrated embodiment, a base 206 of the trunk 202 can be attached to a chassis of the spacecraft 102 (not shown) by various means, such as, but not limited to bolts, rivets, welds, or the like. Although the housing 104 has a cuboidal structure, the structure of the housing can be of different geometric shapes, such as cubicle, trapezoid, cylindrical, hemispherical, polygonal and can be of different sizes according to a requirement of the spacecraft 102 on to which the apparatus 100 will be mounted. In case the housing is cylindrical, the housing can be attached by the base portion. Although not shown, the base portion 206 also allows for electronic interfaces, such as buses and cables between the spacecraft 102 and the apparatus 100.

[00035] As shown in Fig, the spool 106 is connected to the base of the housing 104 by one or more compressed springs 208-1 , 208-2, 208-3, commonly referred to as 208 hereinafter. In one example, the springs 208 can be helical spring. During the operation, the springs 208 pushes the spool 106 and the tuft of wires 108 out from the housing 104 as shown in Fig. 2(b) so that the housing 104 does not obstruct the unwinding of the tuft of wires 108. As shown in Fig. 2 (a), the tuft of wires 108 is wound around the spool in such a manner that the tuft of wires 108 can easily unwound from the spool 106 when deployed. Although the current implementation shows the use of springs 208, a spring loaded telescopic boom or a shape memory alloy based boom can also be to push the spool 106 and the tuft of wires 108 out from the housing 104.

[00036] The apparatus 100 also includes a control module 302 (not shown in Fig) that regulates all the operations of the apparatus 100. The apparatus 100 also includes an electrostatic charge generator (ECG) that, when activated, polarizes the spool 106 and the tuft of wires 108. In addition, the apparatus 100 includes a power system to power all the components of the apparatus. In one example, the power system and the ECG can be a part of the control module 302. The structural and operational details of the control module 302 and various other details will be explained in detail with respect to Fig. 3.

[00037] In another implementation, the apparatus 100 may not include the control module 302 for polarizing the tuft of wires 108. Such system is called passive charging. In such variant, there is no dedicated ECG for electrostatic charging and the deployment happens entirely on a passive basis on the interaction of the apparatus 100 with space plasma in LEOs & MEOs, in particular, the inner Van-Allen radiation belt (IVARB). Here, the initial activation of UWDES is based on the lid-release command from either the spacecraft or the RTC on-board. Placement of solar panels in this passive UWDES variant is optional. When the activation of ECG for charging the tuft of wires + spool is not possible due to any on-board failure in an active- charged variant of UWDES, then if the lid 204 opens based on a fail-safe mechanism, the charging & deployment of tuft of wires + spool can still happen in this passive charging mode. Considering a polar orbit for a spacecraft, on each pass through the polar regions it interacts with the inner Van-Allen radiation belt (IVARB) that has high concentration of positive charges and the tuft of wires most likely will acquire the positive charges that causes the mutual-repulsion of wires for their deployment as well generate Coulomb drag. The charging of the wires here takes place passively and does not rely on the spacecraft's power or ECG to get polarized.

[00038] Fig. 3 a-e illustrates various components of the apparatus 100, in accordance with one implementation of the present subject matter. Fig. 3(a) illustrates a schematic of the apparatus 100, in accordance with one implementation of the present subject matter. As mentioned before, the apparatus 100 broadly includes a control module 302 and the tuft of wires 108. The control module 302 may also include a power system 304 that further includes power sources, such as batteries 306 and solar panels 308. In one example, the solar panels 308 may be mounted on lid 204 and housing or on the spacecraft 102 or both. Moreover, the apparatus 100 preferably coupled to one face of the spacecraft 102 towards one corner of it to facilitate greater exposure to the Sun and to avoid any obstruction to lid 204 opening. The power system 304 may also include a power regulator and distributer 310 that regulates the power from the power sources to the control module 302 and the spacecraft 102. In one example, the apparatus 100 includes a timer 312 coupled to the power sources that triggers the activation of a lid 204 release mechanism 314. The timer 312 is a module that operates independent from the controller 326 to facilitate the deployment of the tuft of wires 108, in case the controller 326 fails to deploy the tuft of wires 108. The apparatus 100 may also include component, such as angle sensor/ gyro sensor 316 that measure an orientation of the lid 204 to monitor the opening of the lid 204.

[00039] In one example, the control module 302 may be hermetically sealed. All the subsystems of the control module 302 can be co-located on the same board or the ECG 318 may be on a separate board to protect the other two components from charges leaked out from the ECG 318. The power system 304 draws power from the solar panels 308 and supplies it to the host satellite during its mission and to the apparatus 100 after the end of life of the host satellite. In the apparatus 100, it powers the sensors and the control module 302. The controller 326 receives data from all the sensors onboard the payload module and processes it and uses it for housekeeping and health monitoring purposes. Main function of controller 326 is to activate the ECG 318 based on the data from the sensors.

[00040] The apparatus 100 can act autonomously i.e. having self- sufficiency and can sustain by itself and operate independent of the spacecraft. Additional to the ECG 318 module, it houses a dedicated OBC board and APS (auxiliary power subsystem) card. It acts like a stand-alone plug-and-play module and does not depend on the spacecraft for power or operating commands. Further, the controller 326 along with the timer 312, can be used to activate & deploy tuft of wires 108 when there is a failure of the controller 326 of apparatus 100. The controller 326 of control module 302 can also be used to process the data received from various sensors present in the apparatus 100. The advantage of autonomous ability is that even if the spacecraft suffers with multiple subsystem failure, the apparatus 100 can still operate & perform its duty of reliably deorbiting the spacecraft 102.

[00041] Fig. 3(b) illustrates the spool 106 of the apparatus 100, in accordance with one implementation of the present subject matter. The spool 106 is the central body of the apparatus and performs two functions. First, the spool 106 acts as a support structure 320 and hold the tuft of wires 108 securely in a wound position till the tuft of wires 108 is deployed. Second, the spool 106 facilitates the deployment of the tuft of wires 108 by acting as a gaussian surface to accumulate static charge when the spool 106 is polarized. Once polarized, the spool 106 creates a strong electric field causing repulsion to the like-charged tuft of wires 108 thereby unwinding the tuft of wires 108. According to the illustrated aspect, the spool 106 includes a support structure 320 having an external surface -1 to support the tuft of wires 108. The support structure 320 can be a hollow or solid structure. The spool 106 also includes a top plate 322 mounted on top of the support structure 320 to prevent unwanted unwinding of the tuft of wires 108. In addition, the top plate 322 also accumulates static charge with the support structure 320. In one example, the top plate 322 can have a bi-convex shape. The spool 106 also includes a base plate 324 that mounts the support structure 320 to the housing 104 (not shown in Fig). In one example, the base plate 324 can be convex-o-concave surface such that a concave surface of the base plate 324 attaches to the housing 104 and the convex part mounts the support structure 320 thereon. The convex portion also acts as the gaussian surface for accumulating the charges. However, the concave portion remains uncharged. The base plate 324 (flange) can be a bi-convex hollow disc structure with a large central hole on a bottom face of the base plate 324 for the springs to pass through the halo & connect to the concave interior of the base plate 324. This reduces the charges at the outer rim and also due the mutual repulsion force prevents the tuft of wires 108 coming closer to the springs or the concave face of bottom flange. In one example, for greater the charge accumulation & resultant stronger electric field, in this spool design the bottom flange can be larger than the top flange and also it is curved enough such that the potential build up will be greater.

[00042] According to the illustrated aspect, the support structure 320 can have different shapes, such as spindle shape, hourglass shape, round plane cylinder or a cylinder made of multiple rings placed one above the other. In one example, the support structure 320 can have a convex contour to provide maximum surface for accumulation of charge. In addition, the support structure 320 should be large enough to reduce/ if not prevent, shape memory acquired by the tuft of wire wound around the spool 106.

[00043] According to another embodiment, the spool 106 may just include the support structure 320 and the bottom plate without the presence of the top plate 322. In this case, when the lid 204 of apparatus 100 is closed, either the lid 204 or the curved plate 204-1 attached to the lid 204 may act as a temporary removable top cover, that securely holds the tuft wires in position until the lid 204 opens. In yet another embodiment, the convex portion of the base plate 324 and the top plate 322 acts as stops for securing the tuft of wires 108 thereby preventing a need of the support structure 320. In yet another embodiment, the spool 106 may just include the base plate 324 such that the tuft of wires 108 may be arranged in a donut shape and the curved plate 204-1 on the lid 204 secures the tuft of wire inside the housing 104.

[00044] According to an aspect, the spool 106 can be made from a variety of material, such as, but not limited to, electrically conductive metal/alloy/composite or can be made of a polymer/composite/ shape- retaining material (metal, etc.,) with a thin outer layer/coat of an electrically conductive material (metal, etc.,) or with an electroplated coating on outside which acts as a hollow metallic structure and builds electrostatic potential on its planar/convex exterior surface when polarized.

[00045] As shown in Fig. 3 (c), the tuft of wires 108 are made of multiple strands of ultra-thin wires attached together to form a web. In one example, the tuft of wires includes multiple strands of thin wires such that one end of each wire is tied together to form the proximal end 108-1 (shown in Fig. 1 ) and other end of each wire is tied to form the distal end. In another example, the tuft of wire includes a single long strand of wire folded/pleated to form the web structure by clamping ends alternate ends together, thereby creating two ends with each end node comprising a set of alternate ends tied/fused together. The portion of the tuft of wires running (stretching) between the two nodes is an inter-nodal region.

[00046] Further, a length of the tuft of wire can range from about 0.5 meters to about 5000 meters depending upon an amount of aerodynamic drag to be generated. The tuft of wires can be made from a variety of material, such as, but not limited to Carbon fibre (T300, T800H, T800S, T1000G, Toray's MJ and M series), unsized carbon fibre, precursor PAN fibre, partially-carbonized carbon fibre, carbonized carbon fibre, partially graphitized carbon fibre, Pitch type carbon fibre, Poly-acrylo nitrile (PAN) fibre, Acrylic fibre, Glass fibre (S-glass, R-glass, D-glass, E-CR-glass, A- glass, C-glass, T-glass), fibreglass, Aramid fibre, Para-aramid fibre ( Kevlar fibre, Technora, Twaron, Heracron), Meta-aramid fibre (Nomex, Teijinconex), Innegra S, Vectran Precision wires/fibres of metal/metal alloys (copper, Be- Cu alloy, gold, Tungsten, etc.,) [00047] Apart from the aforementioned materials, the wires can be made from by nanowires or nanotube. Nanowires and nanotubes (carbon nanotubes and others), given their extremely low thickness and moderate density, when carried on-board even in minute quantities can provide enormous drag area on deployment contributed by their maximum cumulative cross-sectional area. This allows more drag area to be generated by a given mass of payload material to be carried for, thus allowing the spacecraft 102 to experience huge drag force and hence very rapid deorbiting of spacecraft in comparison to other drag-enhancing devices of similar mass and size (form- factor) characteristics. This also facilitates smaller spacecraft (pico, nano and micro-satellites) to carry on-board drag-enhancing payloads with sizeable drag area for effective deorbiting capability.

[00048] According to an aspect, the tuft of wires provides better aerodynamic drag than the conventional drag sail as the tuft of wires provides more cross-sectional area than the conventional drag sail of equivalent mass. Following example illustrates the advantage of tuft of wires over the conventional drag sail made from a given material:

[00049] Considering a given amount of material having a mass 'm' and a volume V, the material is formed into a square shaped drag sail of dimensions 'x' and a minimal thickness y. For the drag sail:

volume of the drag sail (v) = x ² y (1 ) maximum possible cross-sectional area (MPCA) a _max = x ² minimum possible cross-sectional area a _min = xy

[00050] Now, taking same volume of material is used to form 'n' number of strands of wire each having a length Ί _W ire' and minimal thickness y. Further, the strands of wires are used to form the tuft of wires. For the tuft of wires: volume of the drag wires (v) = ^n7iy ^ ^wire (2) from the equation (2): l _wire = (3) [00051] Considering a longitudinal axis of the tuft of wires is oriented perpendicular to the direction of motion of the spacecraft 102, at any instant of time the effective area experiencing drag (EAED) is the cumulative MPCA of the tuft of wires 108. Hence, the total drag-area (a_drag) from all drawn wires is given by.

(4) equating the value from l _wire from equation (3)

^drag ~ ⁼ 1·2732 ^χ (5)

^drag 1·2732α _7ηαχ

[00052] From the above it can be concluded that the cumulative MPCA exposed (available) to produce drag by the wires is 1.2732 times (i.e., 27.32% more) more than the MPCA exposed to produce drag by conventional drag sail of same thickness & drawn from the same amount of material.

[00053] Taking another example to better illustrate the above calculations, if a drag sail of 1 m side and 1 pm thickness is used, a total surface area (TSA) and MPCA of the drag sail will be 2 m ² & 1 m ² respectively. Now, if the same square sheet is drawn into a round wire of same thickness, and keeping the volume constant at 1 cm ³ then, a total surface area (TSA) and MPCA both rise up to 4.23 m ² and 1.414 m ² respectively for the wire from the initial 2 m ² TSA & 1 m ² MPCA for the drag sail. From above calculations, it is understood, the long ultra-thin wire has greater TSA (1 1 1.5% rise) and MPCA (27.32% rise) than the drag sail of same thickness and similar volume. Ultimately the ultra-thin round-wire has a significant 27.32% increment in MPCA when drawn from a square sheet of same thickness and mass. Thus, thin wires when employed as tuft of wires for drag enhancement of a spacecraft and oriented with their length normal to direction of motion of the spacecraft contribute to greater drag area (EAED) and the resultant greater aerodynamic drag effect in comparison to drag sails/gossamers having membranes of same thickness & same mass as tuft of wires and are also oriented with their plane normal to the spacecraft's direction of motion (maximum drag-generating orientation).

[00054] Also, from the above calculations it is evident that there's one order of magnitude rise in drag area (MPCA) provided by tuft of wires for every order of magnitude decrease in their thickness, keeping the total volume of the wire(s) constant. Hence, for a given amount drag-wire payload material, the thinner the wires are, the greater is the drag area (MPCA) presented by the tuft of wires 108.

[00055] Fig. 3(d) illustrates the control module 302, in accordance with one implementation of the present subject matter. The control module 302 controls all the operation of the apparatus 100. In one example, the control module 302 may include a controller 326 that triggers the deployment of the tuft of the wires 108. In one example, the control module 302 includes the electrostatic charge generator (ECG 318) that provides electrostatic charge to polarize the spool 106 and the tuft of wires 108. In addition, the control module 302 may include the power system 304 that provides power supply to all components of the apparatus including the controller 326 and the ECG 318. In one example, the power system 304 may receive power from the solar panels 308 and redirects some of the power to the spacecraft 102 when the apparatus 100 is not deployed and the power system 304 can redirect all the power to the controller 326 and the ECG 318 to polarize the tuft of wires 108 and the spool 106. The operational details of the control module 302 will be explained in subsequent embodiment.

[00056] Fig. 3(e) illustrates a schematic of the ECG 318, in accordance with one implementation of the present subject matter. The ECG 318 includes a voltage source 328 that derives power from the power system 304 to the ECG 318. The ECG 318 also include a high voltage generator 330 that is coupled to an emitter 332. The high voltage generator 330 receives the voltage from the voltage source 328 and steps the voltage up to sufficient high voltage for the emitter 332. The emitter 332 provides charges to a collector 334 by drawing charges from a reservoir, that can be, in one example, the spool 106 and the tuft of wires 108. The ECG 318 along with its supporting electronics & circuitry can be housed within a hermetically-sealed electrically-insulated case designated as ECG 318 module. In active charging module system, the tuft of wires 108 are preferably charged positive for better performance of drag device 1 12 in LEOs & MEOs that have a dominant number of positive-charged particles/ions but, can also be charged negatively, if required. An electron emission device (electron gun) can be mounted externally (body-mounted) on the payload module in case of the active charging module and operated to shoot off back into space any excess charges acquired from there by the charged tuft of wires and thus, maintain the spacecraft system neutral.

[00057] Although the Fig. 2 and 3 illustrates the spool 106 being deployable from the housing 104, another implementation of an apparatus 400 has the spool fixedly attached to the chassis of the spacecraft 102 as shown in Fig. 4. Further, a lid 204 402 may be hinged to the spacecraft to cover the apparatus 400 a housing may not be needed. In such case, the lid 204 402 may include side walls 404 that provide all round cover the apparatus 400. During the deployment, the lid 204 release mechanism 314 releases the lid 204 402 such that the lid 204 402 does not provide any obstruction in the deployment of tuft of wires 108.

[00058] Fig. 5 (a)-(d) illustrates an implementation of another implementation of fixed spool 500, in accordance of the present subject matter. The spool 500 may include a container 502 and a support structure 504. In one example, the container 502 can have funnel like structure having multiple facets 506 forming an internal surface 508 of the container 502. Further, each facet 506 has a curved profile and merges with adjacent facet to the form the internal surface 508. In one example, the facets 506 acts as the gaussian surface for accumulating the charges. According to the illustrated aspect, the support structure 504 has a frustum shape having a tapered surface from bottom of the support structure to the top of the support structure. In one example, the support structure 504 can be placed at a centre of the container as shown in Fig. 5 (d). Further, the support structure 504 can be solid or hollow structure an external surface 510 of the support structure can also act as the gaussian surface for accumulating the charges. In one example, the support structure 504 may be installed inside the container 502, such that a space 512 is formed between an external surface of the support structure and the internal surface of the container 502. This space 512 is used to stow the tuft of wires 108. In addition, the combination of the external surface and the internal surface of the support structure 504 and the container 502 offers more gaussian surface than the gaussian surface of the spool 106 (shown in Fig. 2) thereby providing better deployment of the tuft of wires 108. Further, the container 502 and the support structure 504 are designed such that the support structure 504 can be detached and removed out of the container 502 for winding the tuft of wires 108 around the support structure 504 and after winding the tuft of wires, the support structure 504 along with the wound tuft of wires 108 can be placed back into the container 502 and fixed (fastened) to the centre of the container 502. Further, to prevent the slippage of tuft of wires 108, the support structure 504 has a flange towards the widest end of the tapered structure and fixed permanently thereto. On the opposite side, i.e. towards the narrowest end of the tapered support structure 504, the lid 204 of payload module is initially attached to the core during winding of the tuft of wires 108 thus, acting as a flange temporarily. Once the support structure 504 with its wound tuft of wires 108 together with the permanent flange and the lid 204 are placed inside the payload container and fixed firmly to its floor, the lid 204 can now be hinged (coupled) to the payload module and the link between it and the support structure 504 can be removed. Thus, the lid 204 stays in place holding the wound tuft of wires 108 securely in position. When the module's lid 204 opens during deployment, the lid 204 is free to move away from the support structure 504 facilitating free movement of the tuft of wires 108 out of the container 502.

[00059] During operation, when the lid 204 is released, the controller 326 activates the ECG 318 to provide electrostatic charge to the tuft of wires and the spool 500 thereby deploying the tuft of wires 108. As both the curved multi-faceted walls of the container 502 as well as the tapered support structure 504 act as gaussian surfaces with their convexity facing towards the stowage space of the container 502, with the electrostatic charges accumulated on the surfaces they generate an electric field directed towards this residual/stowage space, thus charging as well as mutually repelling the tuft of wires 108 stowed therein. This facilitates the deployment of tuft of wires 108 by forcing them gently out of the container 502, due to mutual-repulsion. On getting charged electrostatically, the top most coils/windings of tuft of wires 108 repel mutually and move out of the container pulling along with the subsequent coils/windings, thus gradually unwinding, unfurling and deploying the whole length of tuft of wires 108 out of the container 502 and so completing the deployment phase.

[00060] Fig. 6 a-f illustrates another configuration of the apparatus, in accordance with one implementation of the present subject matter. Fig. 6(a) illustrates the apparatus 600 in undeployed states whereas Fig. 6(b) illustrates the apparatus 600 in a semi-deployed state. Now, the spool 602 of the apparatus 600 is similar to the spool 106 shown in Fig. 2 having a support structure, a top plate and a base plate. However, in the illustrated embodiment, the apparatus 600 includes a tape spring 604 wound around the flanges of the spool 602. In one example, one end 604-1 of the tape spring 604 is attached to the base of the housing 606 while the other end 604-2 is attached to the top of the top plate by, but not limited to, fasteners. The tape spring 604 can be a double/multi-element boom having a shape memory. The whole assembly of spool 602 with both the tuft of wires 108 and the tape spring 604 wound about it is placed under tension inside the module, using a compressed ejection helical-spring similar to the springs 208 positioned at the floor of the housing and the lid 204 is closed to hold the spool and its contents securely in position.

[00061] When the lid 204 606 is commanded to open, the whole spool 602 along with the tape spring 604 wound about the spool 602 is to be pushed out of the housing 608 by ejection spring. Later the tape spring 602 will unwind (deploy) and straighten exposing the spool 602 with the tuft of wires 108 windings. Finally, the ECG 318 is powered on, which transfers the charges to the spool 602 and the tuft of wires 108 via the tape spring 604, facilitating the charging and unwinding of the tuft of wires 108. [00062] Fig 7 illustrates another implementation of an apparatus 700, in accordance with one implementation of the present subject matter. The apparatus 700 is mostly similar to the apparatus 600 shown in Fig. 6 (a)-(f), but with a different design of the spool 702. The spool 702 is an elongated structure (dumbbell shaped) and is positioned horizontally in a housing 704. In addition, the flanges 706 of the spool 702 are metallic/metal coated and are spherical & equal sized. In the illustrated example, the spool 702 is stowed under tension by having ejection helical springs compressed against the base plate & held in position with the lid 204 closed. This helps in unobstructed deployment/ejection of the spool on the opening of lid. The metallic core/barrel is spindle/cylinder shaped about which the tuft of wires are wound. Further, the tape spring 708 is wound about the spool 702 over the flanges perpendicular to the tuft of wires 108 windings. In the illustrated embodiment, the horizontal positioning of the spool 702, its elongated structure and the broader mouth of the housing 704 together makes the deployment easier.

[00063] Fig. 8 a-d illustrates an apparatus 800, in accordance of an implementation of the present subject matter. As shown in Fig. 8, the apparatus 800 is formed inside the spacecraft 102. In addition, a lid 204 802 is also built into the spacecraft 102 and allows release of the tuft of wires 108. As shown in Fig. Fig. 9(a), the apparatus 800 is in undeployed state and Fig. 8(b) illustrates a state of the apparatus 800 just at the time of deployment. Further, Fig. 8(c) illustrates the tuft of wires 108 deployed from the spacecraft 102 and Fig. 5(d) illustrates the formation of the drag device 1 12. The operation of the different types of apparatuses will be explained with respect to Fig. 10.

[00064] Fig. 9 a-e illustrates different forms of drag devices 1 12, in accordance with one implementation of the present subject matter. The drag device 1 12 can attain different shapes (configurations) which is determined based on an amount of mutual-repulsion force (due to electrostatic charging) acting on the drag device 1 12 against counter-balancing forces acting on the wires i.e. drag effect or reaction of aerodynamic & Coulomb drag & gravity gradient effect. On deployment of the apparatus 100, the tuft of wires into 3D web, as one end (proximal node) of the drag device 1 12 is anchored to the spacecraft through the anchor wire, the rest of its structure tends gradually towards Earth due to its microgravity acting upon it (3D drag device 1 12) in the form of gravity gradient effect, if the drag force acting upon it is comparatively lower. Drag device 1 12 can form into the following shapes or configurations: 1. boat, 2. spindle or pear, 3. globe/orb, 4. flower/ biconvex disc/ hub-of-spokes, 5. bunch-of-balloons.

[00065] Fig. 9 (a) illustrates the boat configuration of the drag device 1 12. This will be acquired when the electrostatic mutual-repulsion force between the tuft of wires is just enough to prevent their contact, but allows them to freely move around, thus behaving as a flexible structure [Fig.6(A)]. With the drag force acting gently upon it, the drag device 1 12 structure settles down into an elongated-boat shape due to the bellowing-effect caused by the drag force.

[00066] In the boat configuration, if it's (boat's) long-axis is nearly perpendicular to the velocity-vector of spacecraft & the tuft of wires are mostly straight, the drag-area (cumulative cross-sectional area of ultra-thin tuft of wires) contributed by drag device 1 12 is roughly above 85% of their MPCA. When the drag force acting on it rises, the boat-shaped drag device 1 12 tilts/inclines backwards away from the plane normal to velocity vector of spacecraft causing a decline in drag-area presented by the boat-shaped web. This stabilizes the boat structure at a certain angle to the local vertical on orbit with the drag force balancing against the decline in drag area that in turn reduces the drag on it.

[00067] Fig 9 (b) illustrates the spindle configuration of the drag device 1 12. If the tuft of wires is charged further than the boat configuration, they move further apart from each other while arranging into a spindle or pear configuration that behaves like a semi-rigid entity. Most of the wires in this arrangement try to settle down in the outermost region of the spindle or pear, making it essentially into a hollow-spindle or pear-shaped drag device. In the pear configuration, if it's (spindle or pear) long-axis is nearly perpendicular to the velocity-vector of spacecraft, the drag-area contributed by drag device 1 12 is roughly 80-95% of the cumulative MPCA of tuft of wires.

[00068] Fig 9 (d) illustrates the globe configuration of the drag device 1 12. On charging the drag device 1 12 beyond that of spindle configuration, it will attain a near-globular shape as the wires further repel mutually and move apart from each other radially outwards while counter-balancing against the gravity gradient & drag effects. In this globe/orb-configuration, the drag-area contributed by drag device 1 12 is around 70-85% of the cumulative cross- sectional area of tuft of wires available. In this globe/ orb-configuration, irrespective of its orientation with respect to the velocity-vector of spacecraft, the drag-area projected (presented) by it will be nearly constant for a given drag device.

[00069] Fig 9 (e) illustrates the flower configuration of the drag device 1 12. On charging the wire-web to a high potential, the wires finally settle down into this flower- configuration pushing the inter-nodal regions of the tuft of wires radially outwards as far apart as possible while pulling the two nodes of the wire-web nearer. Here, the mutual-repulsion between all the tuft of wires with their adjacent ones is balanced by the mutual-repulsion between proximal and distal halves of the drag device 1 12. Therefore, the higher the numbers of tuft of wires present in the wire-web and the higher they are charged electrostatically, the more radially outward they spread (Flower- configuration) and the closely pulled will be the two nodes of wire-web (i.e., the proximal and distal halves of the drag device 1 12 will come closer to each other) making it into an even flatter biconvex structure. Here, anchor-wire acts as 'stalk' of the flower-configuration through which it is towed by the spacecraft.

[00070] This is the most stable configuration of the drag device 1 12 possible. When the wire-web attains this configuration, and is being towed by spacecraft in lower altitudes where the density of space atmosphere is considerably higher, it undergoes a change in its shape to a conical structure and this effect is known as 'coning'. The angle of the cone depends on the drag force acting upon it and the electrostatic mutual repulsion between the tuft of wires. The more acute (angle of the cone) the flower-cone shaped drag device 1 12 becomes, the less will be the drag-area presented by it. Thus, greater electrostatic-charging of tuft of wires is necessary to maintain a flat/plane shaped flower-configuration of drag device 1 12 in lower LEOs, where the drag force is relatively high.

[00071] The drag device 1 12 in flower-configuration, if it is mostly flat with minimal coning and if the plane/disc of the flower-shaped wire-web is normal to the velocity vector of spacecraft, presents a drag-area that is about 80-95% of the cumulative MPCA of tuft of wires. The drag area (EAED) in this case also depends on how radially outward the tuft of wires spread/bend because the more they spread radially outwards pulling the two nodes of wire-web closer, the greater will be the projected length of (both halves of) each individual drag-wire strand normal to the velocity vector of spacecraft and hence, the higher cumulative EAED of the 3D flower-shaped wire-web. [00072] Multiple tuft of wires can be deployed using either a single deployable tape spring or a single long anchor wire as shown in Fig. 8(d). All these drag device 1 12s are held together like a bunch-of-balloons by anchoring (fastening) all their individual anchor-wires to the spool or to a central anchor wire. Main advantage of this configuration is that, if many ultra- thin drag wire strands have to be accommodated into the drag device 1 12, as they all can't be crowded into a single drag device 1 12 that may render them ineffective, they can be distributed and placed into multiple individual drag device 1 12s that spread around evenly and thereby increasing the effectiveness of the apparatus 100.

[00073] Once the drag device 1 12 is deployed, two factors that counterbalance each other affect its orientation in space with respect to the spacecraft's velocity vector as follows:

Drag force acting on the drag device; and

Gravitational pull of Earth

[00074] The microgravity exerted by earth on the drag device 1 12 causes it to tend towards earth similar to the working of a gravity-boom used by various spacecraft. When the torque due to drag force is minimal and the pull due to earth's microgravity is significant enough, the drag device 1 12 orients itself such that the anchor wire's long axis aligns mostly with the local vertical (zenith- nadir axis) in orbit. If the resultant torque due to drag force acting upon it rises, the drag device 1 12 tilts (inclines) backwards (with respect to velocity vector) away from its local vertical. If this torque due to drag force is much higher in comparison to the downward (towards nadir) pull due to gravity gradient effect, then the whole drag device 1 12 will be oriented such that it is towed at the back of the spacecraft with the anchor wire nearly aligning with the spacecraft's velocity vector. [00075] In the boat, spindle/pear-shaped configurations of 3D drag-drag device 1 12, when the drag force acting on it rises, it tilts/inclines backwards away from the plane normal to velocity vector of spacecraft causing a decline in drag-area (EAED) presented by it. This stabilizes the long axis of the 3D drag device 1 12 at a certain angle of inclination to the local vertical with the drag force balancing against the decline in drag area that in turn reduces the drag on it.

[00076] Irrespective of the orientation of entire drag device 1 12, the amount of drag force acting upon each individual drag wire strand depends solely on the orientation of its long axis with respect to the velocity vector of the spacecraft. Thus, for each drag wire strand the EAED is directly proportional to its projected length normal to the velocity vector. Thus, the drag force acting on the entire drag device 112 depends on the number of tuft of wires and their projected length in the plane normal to the velocity vector of spacecraft.

[00077] Given the shape of boat and elongated spindle or pear configurations, when they align (orient) themselves to the local vertical (vertically towards earth) they present maximum length of the wires normal to the velocity vector and thus, experience the near maximum drag force for their MPCA. Considering the sphere/globular configuration, irrespective of its orientation with respect to the velocity vector as well as local-vertical, the drag-drag device 1 12 always produces almost constant drag force as the cumulative projected length of all tuft of wires in the wire-web remains nearly constant in any plane passing through its centre.

[00078] In flower, when the drag force and resultant torque acting upon it are high enough, the drag device orients such that the disc/ plane in which tuft of wires are present aligns with the plane perpendicular to the direction of motion of the spacecraft and the anchor wire aligns opposite to the velocity vector as in passive aero stabilization. With this orientation, depending upon their radially outward spread caused by the mutual-repulsion, the tuft of wires have their EAED close to their MPCA and experience the corresponding drag.

[00079] Preferably the tuft of wires is positively charged electrostatically and so, when moving through space plasma while on orbit, these charged tuft of wires interact with the ionospheric plasma. When the tuft of wires is moving at orbital velocity in LEOs, due to their high relative velocities with respect to the particle/molecules/ions present there, on collision (interaction) with these particles there will be momentum transfer between both. If charged particles like ions and electrons collide with the charged tuft of wires, these polarized tuft of wires may either gain or lose charges based on the polarity (positive or negative) of charged particles, their energies (potential) and their masses.

[00080] Depending upon the loss or gain of charges by the charged ultra-thin tuft of wires on collision with particles (charged/ neutral) in orbit, the potential of the charged tuft of wires may need to be regulated by powering on the ECG 318.

[00081] Irrespective of their orientation, when the drag-drag device 1 12 is charged adequately to acquire the globe/orb-configuration, it presents nearly a constant amount of EAED, which is around 80% of the MPCA for a given set of ultra-thin drag-drag device 1 12. Considering a drag-sail of equal MPCA is deployed in upper LEOs where passive aero-stabilization is not much effective, its attitude might be such that the plane of the sail is not normal the spacecraft's velocity. This is because the aero drag essential for passive aero stabilization is lacking at those altitudes & hence the EAED for drag sail might not be corresponding to its MPCA in comparison to the drag- wires drag device 1 12. [00082] Whereas the drag enhancing device 1 12 deployed in upper LEOs has relatively higher EAED which is almost equal to its MPCA. Therefore, the drag enhancing device 1 12 device (apparatus 100) can be used for deorbiting spacecraft in upper LEOs with appreciable performance unlike drag-sails that require active attitude-control of spacecraft in those orbits and so are not very effective. Apparatus 100 doesn't require passive aero stabilization for optimum drag generation but, can still make use of it. The time taken for a spacecraft (with or without drag-enhancing devices) is longer to deorbit, using drag force, from 1200 km to 1000 km than from 1000 km to km. This implies a drag-enhancing device like the tuft of wires web apparatus 100 can perform reasonably well in upper low-Earth orbits is extremely beneficial to bring down the total-time taken to deorbit the spacecraft orbiting at higher altitudes.

[00083] This, when coupled with the Coulomb drag, allows the apparatus 100 to deorbit spacecraft from much higher altitudes of about 1000 km - 6000 km where conventional aero drag-enhancing devices are ineffective. This synergistic effect helps in bringing down the total/cumulative time-period of deorbiting, while necessitating much less mass of drag- enhancing payload material to be carried on-board. Thus, apparatus 100 acts as a hybrid deorbiting system with higher efficiency than conventional deorbiting that employ only a single kind of drag force.

[00084] Fig. 10 illustrates a method 100 for providing navigation, in accordance with one implementation of the present subject matter. The method 400 can be implemented by the apparatuses described above. The exemplary method may be described in the general context of computer executable instructions embodied on a computer-readable medium. Generally, computer executable instructions can include routines, programs, objects, components, data structures, procedures, modules, functions, etc., which perform particular functions or implement particular abstract data types. The method may also be practiced in a distributed computing environment where functions are performed by remote processing devices that are linked through a communications network. In a distributed computing environment, computer executable instructions may be located in both local and remote computer storage media, including memory storage devices.

[00085] The order in which the method is described is not intended to be construed as a limitation, and any number of the described method blocks can be combined in any order to implement the method, or an alternative method. Additionally, individual blocks may be deleted from the method without departing from the spirit and scope of the methods, systems and devices described herein. Furthermore, the method can be implemented in any suitable hardware, software, firmware, or combination thereof.

[00086] The process begins at block 1010 by receiving a signal from the control module 302 for deploying the tuft of wires 108. In one example, the signal can be a signal for the lid 204 release mechanism 314 to release the lid 204 allowing the deployment of the tuft of wires 108. Further, the signal for deployment can sent either from the spacecraft 102, or a ground based controller 326, or an autonomous function that checks if the spacecraft 102 is operational, or from the timer 312. In either condition, the control module 302 activates the lid 204 release mechanism 314 and the ECG 318 to trigger the deployment of the tuft of wires 108. Once the actuation of the ECG 318 occurs, the operation moves to the next block.

[00087] At block 1020, the tuft of wires 108 is deployed. The tuft of wires 108 is deployed when the lid 204 release mechanism 314 releases the lid 204 for easy and unobstructed deployment of the tuft of wires 108. In case of the apparatus with deployable spool, once the lid 204 is released, the spool 106 is ejected out of the module by ejection springs and is held firmly by them at a specified distance from the spacecraft 102. On activation of apparatus 100, the lid 204 opens exposing the spool entirely following which the ECG 318 is activated. On lid 204 opening, which is assisted partly by the election spring present between the curved plate 204-1 & the lid, the curved plate 204-1 settles at the brim of the lid 204 with help of spring attached thereto. As the curved plate 204-1 is also charged by ECG 318, the curved plate 204-1 repels mutually the like-charged tuft of wires away from the lid, thus, preventing them from coming closer to the lid. The unwinding & deployment of tuft of wires takes place on being charged by ECG 318. Once the ECG 318 is powered on, the charges generated by the ECG 318 are transferred to the spool 16 and the tuft of wires 108 via the springs and initiate deployment of tuft of wires through the unwinding based on their mutual repulsion with the spool & the windings beneath. Alternatively, in case of fixed spool, the tuft of wires 108 and the spool 106 is exposed to space. Now, once the lid 204 is released, the controller 326 actuates the power system to actuate the ECG 318. On ECG 318 activation, the electrostatic charges are transferred to the spool 106 via the springs. During operation, when charging of the spool 106 and tuft of wires 108 begins, the charges starts accumulating both on the spool and the tuft of wires 108, the charges start settling on both the spool's exterior and the tuft of wires 108. Here, the outermost winding/coil of the wires directly repels with the layer of windings immediately beneath it. The outermost layers of windings together act (temporarily) as a cylindrical core/barrel of the spool and are responsible for generating the electric field necessary for the mutual repulsion based unwinding of wires instead of the support structure present beneath the innermost layer of wire windings. Therefore, the support structure may not hold much significance with respect to unwinding of wire tuft till the innermost layers of tuft of wires 108 start undergoing the process of unwinding & deployment. [00088] In case of passive charging, once the lid 204 opens, an electrical contact gets established between the outer of the housing 104 (that acts as the ground of module and also as a plasma collector 334 on interaction with space plasma while in orbit) and an interior of the housing. With this, there will be transfer of excess charges from the module's frame/chassis to the interior of payload container, thus electrostatically charging the tuft of wires stowed inside the container causing their deployment.

[00089] For effective deployment of the tuft of wires 108, activation/ release of the lid 204 is important. Even with the failure of various subsystems on-board the spacecraft 102 or the controller 326, if the lid 204 releases and opening take place, the spool if present can get ejected making way for the passive-charging based deployment of tuft of wires. Fig. 1 1 illustrates a flowchart 1 100 representing following modes of operation of releasing the lid 204 mechanism and the activating the ECG 318.

[00090] Feedback based activation: In this mode, after launch into space, the spacecraft 102 sends a signal to the controller 326 of the apparatus 100 on a regular-basis at pre-defined time intervals. That time interval can be daily or weekly or monthly indicating an operational condition of the spacecraft. In one example, the operational condition can be the spacecraft's health and performance & to let the controller 326 know that the spacecraft 102 is still in control. Further, the controller 326 also verifies the receipt of the feedback signal. Once the spacecraft 102 shutdowns after its functional life time or EOL- End of Life or the spacecraft 102 fails unexpectedly before EOL, the feedback signal from the spacecraft 102 stops coming indicating a failure on-board the host spacecraft. As soon as the controller 326 detects a failure, it records the event as no-feedback period and register the same. In one example, the controller 326 is programmed to starts the countdown 'n' corresponding to the number of no-feedback. In one example, the controller 326 waits for 3 to 5 no-feedback periods (No feedback count) to determine the either the EOL or failure of the spacecraft 102. This count provides sufficient time for the debugging in case the failure can be fixed to prevent accidental loss/de-orbiting of spacecraft. Once the no feedback count reduces to zero, the controller 326 actuates the power system to activate/trigger the lid-release mechanism & to execute follow-on deployment steps.

[00091] Ground command based activation: Another mode of activation is by sending ground signals to the spacecraft 102 indicates the deployment of the tuft of wires 108. Despite the proper functioning of spacecraft 102 in orbit in the scenarios like collision avoidance, potential system failure etc., a command can be sent from the ground control station to the spacecraft 102 which in turn instructs the controller 326 to activate/trigger the lid 204 release mechanism 314 and deploy the tuft of wires 108 for immediate deorbiting.

[00092] Timer 312 based activation: The above mentioned two methods are either dependent on feedback or ground based command which are prone to fail due to various reasons, such as failure in feedback system or failure in telecommunication system. Therefore, a third method maybe employed that acts as a failsafe system, in case either of the above to method fails. The third method employs the timer 312 positioned near the lid- release device. Prior to the launch, the timer 312 is programmed to activate the lid 204 release mechanism 314 at a predetermined time (e.g. 5yrs exactly from the moment of launch of Satellite) after the launch of the spacecraft 102. Further, from the moment of launch into space, the timer 312 keeps running clocking until the predetermined time as arrived and once the predetermined time comes, the timer 312 activates or gives command to the lid-release mechanism 314. As mentioned before, the timer facilitates the power supply from the solar panel directly to the lid 204 release mechanism 314 bypassing all other systems like controller 326 or the power system. As may be understood, the timer 312 is the primary fail-safe mechanism for apparatus 100 deployment based on lid-release followed by tuft of wires 108 deployment through passive charging.

[00093] Although the subject matter has been described with reference to specific embodiments, this description is not meant to be construed in a limiting sense. Various modifications of the disclosed embodiments, as well as alternate embodiments of the subject matter, will become apparent to persons skilled in the art upon reference to the description of the subject matter. It is therefore contemplated that such modifications can be made without departing from the spirit or scope of the present subject matter as defined.