Abstract: An electric propulsion module coupled to a spacecraft capable of providing thrust at a level required for multi-burn orbit transfer is disclosed herein. The electric propulsion system includes an electric propulsion thruster, a propellant tank and an energy storage device. In one form the energy storage device is a battery operable to provide sufficient power to maneuver the spacecraft quickly to avoid space debris and/or move to a different orbit through a multi-burn thrust procedure.

Abstract: Air-buoyant structures, and vehicles incorporating air-buoyant structures, are provided. Hollow, air-buoyant structures may include a shell of ultra-low density aerogel material, foam material, or vapor-expanded material that is strong and stiff enough to withstand atmospheric pressure and lightweight enough to achieve buoyancy in air under evacuation. The shell may be reinforced with a suitable reinforcing material, such as helical nanofibers. The air-buoyant structures may also include vacuum pumps and valves operably connected to or integrated with the hollow shell. The vacuum pumps and valves may be configured to pump air out of the hollow shell and allow air back into the hollow shell to control buoyancy.

Abstract: An unmanned aerial vehicle such as drone integrated with solar power unit, is disclosed. The solar power unit of the drone comprises a solar panel assembly positioned at a hull of the drone. The solar panel assembly is composed of a curtain-type shutter embedded with a plurality of photovoltaic cells. The shutter is formed of a plurality of slats having space between each of the slats, wherein the space facilitates for an effective dislocation of wind diffusion to maintain the stability of the hovering drone. The shutter further comprises a means for twisting and turning the shutter to provide balance for the hovering drone Further, the solar power unit comprises an energy storage unit operably coupled to the solar panel assembly for storage of the electrical energy from the solar panel. Further, the shutter is configured to spin cyclically diffusing the gust of wind harmlessly for stabilizing the hovering drone.

Abstract: A spacecraft includes a plurality of deployable module elements, at least one of the deployable module elements including a robotic manipulator, the spacecraft being reconfigurable from a launch configuration to an on-orbit configuration. In the launch configuration, the deployable module elements are disposed in a launch vehicle in a first arrangement. In the on-orbit configuration, the deployable module elements are disposed in a second configuration. The spacecraft is self-assembled by the robotic manipulator reconfiguring the spacecraft from the launch configuration, through a transition configuration, to the on-orbit configuration. The deployable module elements may be in a stacked arrangement in the launch configuration and may be in a side-by-side arrangement in the on-orbit configuration.

Abstract: In an on-orbit configuration, a spacecraft having a center of mass and a pitch axis passing through the center of mass includes a main body, a first solar array, and a first thruster is operable in a geostationary orbit with the first solar array deployed, proximate to a first north or south surface of the main body, such that a rotational axis of the deployed first solar array is substantially parallel to the pitch axis. The first thruster is disposed proximate to a second north or south surface of the main body, the first thruster having a thrust vector that is approximately coaligned with the pitch axis, the second surface being opposite to the first north or south surface. No solar array is proximate to the second surface.

Abstract: A method to control a sunlight acquisition phase of a spacecraft with a nonzero angular momentum of an axis DH. The spacecraft includes a solar generator configured to rotate about an axis Y. The spacecraft actuators are controlled to place the spacecraft in an intermediate orientation in which the axis Y is substantially orthogonal to the axis DH. The solar generator is controlled to orientate the solar generator towards the sun. The spacecraft actuators are controlled to reduce the angular momentum of the spacecraft. The actuators of the spacecraft engine are controlled to place the spacecraft in an acquisition orientation in which the axis Y is substantially orthogonal to the direction of the sun with respect to the spacecraft.

Abstract: A spacecraft operating in a low earth orbit having an altitude in the range of 160 to 800 km has a main body that includes heat dissipating electrical equipment and an earth-facing payload. Control surfaces on the spacecraft are articulated so as to: provide three-axis attitude control to the spacecraft main body using aerodynamic drag effects, such that the earth-facing payload is maintained in a selected orientation with respect to the earth; and control one or both of orbit altitude and period by articulating the control surfaces so as to regulate an amount of aerodynamic drag. The control surfaces include a first control surface disposed, in an on-orbit configuration, on a boom, the boom being mechanically coupled with the main body, and with one or both of a solar array electrically coupled with the electrical equipment and a thermal radiating array thermally coupled with the electrical equipment.

Abstract: A core element for use with an inflatable or expandable spacecraft is claimed. The core element has a plurality of panels connected by hinges so that the core element can be packed into a smaller volume for loading into a launch vehicle. Upon deployment in space, the core element can be unpacked and can enclose a larger volume than that in the packed state. Multiple core elements can be used to cover a spacecraft to a desired degree.

Abstract: A stackable satellite includes a satellite frame and at least one vertical pillar attached to the frame. The vertical pillar has an upper end and a lower end. The upper end is coupled to the lower end of the vertical pillar of the satellite above and the lower end is coupled to the upper end of the vertical pillar of the satellite below. The vertical pillar receives substantially all of the vertical load of the stackable satellite and any other satellites stacked above. Use of such vertical pillars removes the need for a dispenser or substantially lightens the dispenser mass to allow substantially more satellites to be carried in a payload.

Abstract: A method for localizing the source of air pollution that includes receiving pollution data from a network in communication with at least one air pollution sensor that is connected to at least one motor vehicle in a stationary position. Each stationary motor vehicle of the network is positioned at a different air pollution measurement location. The method may further include determining from the pollution data an origin of an air pollutant using an air pollution model provided by a pollutant source localization device including at least one hardware processor.

Abstract: Apparatus and methods for attitude control of a space vehicle. The space vehicle includes a reaction wheel subsystem and a propulsion subsystem for attitude control. A slew controller for the space vehicle detects a failure of the reaction wheel subsystem. To perform an attitude slew, the slew controller identifies a pair of thrusters in the propulsion subsystem that will rotate the space vehicle about a desired axis. The slew controller determines burn parameters for the pair of thrusters, and initiates the burns to perform the desired attitude slew.

Abstract: The present invention is realized by apparatus and methods for placing a large utility scale photovoltaic array in the low stratosphere of earth's atmosphere at an altitude of about 20 km, above clouds, moisture, dust, and wind. This is accomplished using a large light-weight, rigid, buoyant structure to support the large photovoltaic array. Long, strong and light tethers connect the buoyant structure to the ground and hold it in position against wind forces. The electricity output from the photovoltaic array is then coupled to high voltage transmission lines which connect from the platform to the earth's surface. The electricity is then transmitted through the high voltage transmission lines to the earth's surface where it is connected to the electrical supply grid and provides lower cost, more reliable electricity.

Abstract: A flexible space structure such as a solar array is composed of multiple solar cell modules (SCMs) each supporting an arrangement of solar cells on a frontside layer and incorporating a backside layer with a surface opposite from the frontside layer having a conductive coating. A selected portion of the SCMs have structural ground extension harnesses intermediate the frontside layer and backside layer. Conductive tapes secure vertically adjacent SCMs by attachment to the conductive coating and electrical jumpers interconnect the structural ground extension harnesses across gapped hinge lines of laterally adjacent SCMs.

Abstract: A directionally-controlled roll-out elastically deployable solar array structure is disclosed. The structure includes one or more longitudinal elastic roll out booms that may be closed section or open section to allow for efficient rolled packaging onto a lateral mandrel. A flexible photovoltaic blanket is attached to a tip structure and to a lateral base support structure, but remains uncoupled from the longitudinal booms. The solar array system may be stowed simultaneously into a rolled package comprised of the roll out booms and the flexible planar blanket together, or onto independent rolls. Alternatively, the system may be stowed by rolling the booms, and accordion Z-folding the hinged flexible photovoltaic blanket into a flat stack. Structural deployment is motivated by the elastic strain energy of the roll out booms, and several methods of deployment direction control are provided to ensure a known, controlled, and unidirectional deployment path of the elastically unrolling booms.

Abstract: A system for generating electrical power from solar radiation utilizing a thin film III-V compound multijunction semiconductor solar cell mounted on a support in a non-planar configuration is disclosed herein.

Abstract: Structures and methods are disclosed regarding deployable structures with expandable longerons adjustably coupled with supporting structures such that an angle between the supporting structures can be adjusted. Such structures can include and/or be used for solar arrays, bridges, support structures, and more. These structures can be easily transported to a new location and deployed from the stowed configuration into a larger functional structure. In some embodiments these structures can use one or more longerons that can have two resting states: deployed and rolled.

Abstract: A deployable structure that may include a slit-tube longeron and a flat panel coupled with the slit-tube longeron. The slit-tube longeron may include a tubular member having a slit that runs along the longitudinal length of the slit-tube longeron. The deployable structure may be configured to couple with a satellite. And the deployable structure may be configured to transform between a stowed state and a deployed state where the tubular member is substantially straight when the deployable structure is in the deployed state, and the tubular member is wrapped around the satellite when the deployable structure is in the stowed state.

Abstract: Collapsible solar array structures are disclosed that include a collapsible structure and detachable solar array. The solar array can be detached stowed separately from the collapsible structure. The collapsible structure can include a plurality of longerons and/or support structures. Longerons can have a slit along a longitudinal length of the longeron.

Abstract: A system and method for a self-charging battery cell are provided in which beta emissions from a Strontium-90 source are obtained by a sensor device and converted into electric energy. In embodiments, a scintillation device is used to intake emissions from a Strontium-90 source, and consequently emit a light or plurality of light flashes. A sensor device, e.g., a photodiode, is utilized to convert the light or plurality of light flashes into electric voltage, current and/or energy.

Abstract: An onboard solar cell array current and voltage characteristic determination method is preferably used on small spacecraft and determines the solar cell orientation relative to the sun by a comparison between prelaunch solar cell characteristics with on-orbit solar cell characteristics well suited for spin axis determinations and monitoring the degradation of on-orbit solar cells over the operational life of a picosatellite.

Abstract: Structures and methods are disclosed regarding deployable structures with expandable longerons adjustably coupled with supporting structures such that an angle between the supporting structures can be adjusted. Such structures can include and/or be used for solar arrays, bridges, support structures, and more. These structures can be easily transported to a new location and deployed from the stowed configuration into a larger functional structure. In some embodiments these structures can use one or more longerons that can have two resting states: deployed and rolled.

Abstract: A deployable structure is disclosed. The deployable structure may include one or more slit-tube longerons; and one or more flat sheets coupled with the one or more slit-tube longerons. The one or more slit-tube longerons and the one or more flat sheets may be stowed by rolling the one or more slit-tube longerons and the one or more flat sheets together into a roll. In one embodiment, at least a portion of the one or more slit-tube longerons may be exposed when stowed. In another embodiment, the one or more slit-tube longerons may be manufactured from a shape memory material. These slit-tube longerons unroll into to a straight configuration when exposed to heat.

Abstract: One embodiment of the invention includes a spacecraft. The spacecraft comprises at least one solar array panel comprising an array of solar cell sections. Each of the solar cell sections can be configured to generate electrical power from received solar radiation. The spacecraft also comprises a solar array selection controller configured to selectively deactivate a portion of the array of solar cell sections to generate a torque on the spacecraft based on a difference in at least one of solar and thermal radiation pressure between an activated portion of the array of solar cell sections and the deactivated portion of the array of solar cell sections.

Abstract: A wireless power transfer system includes: a plurality of power transmitters, each of which transmits a microwave; and a rectenna base station which receives the microwave to generate power. The rectenna base station includes: a rectenna; and control section which specifies an identification code for identifying each power transmitter and generates a command signal to change a phase of the power transmitter specified by identification code so as to increase a power value received at the rectenna. Each of the power transmitters comprises: a plurality of transmission antenna elements, each of which transmits the microwave to the rectenna base station; and a phase controller which makes phase change of the microwave based on the command signal from the phase monitor and control section of the rectenna base station if the identification code matches a stored identification code.

Abstract: A deployable structure is disclosed. The deployable structure may include one or more slit-tube longerons; and one or more flat sheets coupled with the one or more slit-tube longerons. The one or more slit-tube longerons and the one or more flat sheets may be stowed by rolling the one or more slit-tube longerons and the one or more flat sheets together into a roll. In one embodiment, at least a portion of the one or more slit-tube longerons may be exposed when stowed. In another embodiment, the one or more slit-tube longerons may be manufactured from a shape memory material. These slit-tube longerons unroll into to a straight configuration when exposed to heat.

Abstract: A flexible inflatable hinge includes curable resin for rigidly positioning panels of solar cells about the hinge in which wrap around contacts and flex circuits are disposed for routing power from the solar cells to the power bus further used for grounding the hinge. An indium tin oxide and magnesium fluoride coating is used to prevent static discharge while being transparent to ultraviolet light that cures the embedded resin after deployment for rigidizing the inflatable hinge.

Abstract: A device is disclosed for sequencing the deployment of a deployable structure comprising a first body, at least one second body mobile relative to the first body, and at least one third body mobile relative to either the first body or the second body and disposed, in an at least partially folded position of the structure, against or between the first and second bodies. The device comprises a first member fixedly mounted on the third body and preventing its deployment when it is immobilized and a second member mounted on the second body and immobilizing the first member until the second body has effected a selected portion of a kinematic that drives it from an initial position to a final position.

Abstract: A space-based power system. The system maintains proper positioning and alignment of system components without using connecting structures. Power system elements are launched into orbit, and the free-floating power system elements are maintained in proper relative alignment, e.g., position, orientation, and shape, using a control system.

Abstract: A solar cell for a solar generator panel (10, 10?). According to the invention, the cell is coupled to a reflector (70, 71, 700) in such a manner that together with the reflector the cell forms an individual component (20, 21, 20?), while the other end (E2, E2?) of the reflector remains free, the mechanical flexibility properties of the reflector being determined in such a manner as to enable it, in the absence of any vertical pressure, to stay upright in a first position in which its free end points towards outer space, thereby defining a “upper” first face (701, 711) of the reflector facing outer space, while the “lower” opposite face (702, 712) faces the panel, and in such a manner that in a second position, in response to the application of vertical pressure, it is capable of presenting its upper face facing towards the plane of the panel. The invention is particularly applicable to space vehicles having local-concentration solar panels.

Abstract: A device is disclosed for reducing the risk of primary and secondary electrostatic discharges occurring in particular in the solar generators of spacecraft. A solar energy concentrator device for spacecraft includes a reflector for reflecting solar radiation onto a photovoltaic cell for converting solar energy into electrical energy and a heat transfer arrangement for transporting to a cold area heat energy stored by the cell following reception of solar radiation. One particular application of the device is to solar panels for satellites.

Abstract: A solar generator panel includes at least one reflector for reflecting solar radiation onto at least one photovoltaic cell. The reflector takes the form of a flexible material blade, a first end of which is fixed to the panel and a second end of which remains free. The material has mechanical properties such that, in a first position, in the absence of vertical pressure, it maintains itself erect so that its free end points to outer space, thereby defining an upper face that faces outer space and an opposite lower face that faces the panel, and, in a second position, in response to vertical pressure, it orients its upper face toward the plane of the panel. One particular application is to high-power solar panels for satellites.

Abstract: A space solar concentrator based on deployable reflectors attached to a panel base is described. Each of the reflectors includes sides coated with reflective coating on surfaces facing away from the base. In one embodiment, the sides of the reflector have thicker portion to provide stiffness for deployment absent compressive force, and thinner portion to reduce weight and storage energy. In another embodiment, both reflective sides are unfolded when the reflectors are in stowed position. During deployment, a stop side will limit the deployment of both reflective sides and ensure the proper reflecting position. In addition, these embodiments reduce the storage energy.

Abstract: A solar panel for a spacecraft has a base with a face, and at least one row of solar cells and at least one elongated reflector are mounted on the face of the base. The reflectors and the rows are mounted generally parallel to each other in an alternating fashion. The reflector has first and second reflecting sides when the reflector is in a deployed position. The reflector is mounted so that the first side of the reflector is adjacent to a row of solar cells and reflects radiation incident on the first side onto the adjacent row of solar cells when the reflector is in a deployed position. Preferably, a plurality of rows and a plurality of reflectors are mounted on the face of the base, with at least one of the reflectors being disposed between two adjacent rows of solar cells.

Abstract: A lightweight structure for space application is deployable from a compact bundle of interconnected struts into a plurality of tiled substantially rectangular bays. Each bay preferably comprises four or six hinged strut members having substantially rectangular cross sections, such that a solar or other suitable blanket may be compactly nested within the bundled struts. In one aspect of the invention, the blanket is attached to two opposed strut members of a six-member strut bundle, such that the blanket is unfolded during deployment of the strut bundle into a substantially rectangular bay. In another aspect of the invention, the blanket is attached to, and held against, a single strut in a strut bundle. After the strut bundle is deployed into a substantially rectangular bay, the blanket is spread across the bay via a cable mechanism or other deployment mechanism.

Abstract: Solar array assemblies and systems and methods for deploying solar cell arrays from a spacecraft. A solar cell panel assembly comprises a first flexible solar panel and a rotational member. A first extension assembly is disposed proximate to a first end of the rotational member and a second extension assembly is disposed proximate to a second end of the rotational member. The solar cell panel assembly further comprises a first support member and a second support member. The first end of the first support member is coupled to the first end of the rotational member and the first end of the second support member is coupled to the second end of the rotational member. The second ends of the support members are coupled to the first flexible solar panel. A tether assembly couples the extension assemblies to the support members.

Abstract: A space-based power system. The system maintains proper positioning and alignment of system components without using connecting structures. Power system elements are launched into orbit, and the free-floating power system elements are maintained in proper relative alignment, e.g., position, orientation, and shape, using a control system.

Abstract: The satellite solar generator structure comprises a plurality of solar panels (PR, PS) each having a plane surface, the panels being disposed in superposition in said structure (ST), the structure including at least one rod (TI) holding said panels (PR, PS) pressed against one another, two successive panels being in contact via bearing points (AL) disposed at the peripheries of the panels. In this structure, spacers (CA) are distributed over the surfaces of the panels (PR, PS) and are disposed between pairs of successive panels in order to improve contact between two successive panels so that the structure presents better mechanical strength during the satellite launch stage. With this arrangement, a fraction of the forces is taken up by the spacers instead of being taken up by the panels, thus making it possible to optimize the solar generator, for example by reducing the weight of the panels constituting it.

Abstract: Compounds are expressed by general formula of AxBC2−y where 0≦x≦2 and 0≦y<1, and have CdI2 analogous layer structures; A-site is occupied by at least one element selected from the group consisting of Li, Na, K, Rb, Cs, Mg, Ca, Sr, Ba, Ti, V, Cr, Mn, Fe, Co, Ni, Cu, Zn, Zr, Nb, Mo, Ru, Rh, Pd, Ag, Cd, Hf, Ta, W, Re, Ir, Pt, Au, Sc, rare earth elements containing Y, B, Al, Ga, In, Tl, Sn, Pb and Bi; B-site is occupied by at least one element selected from the group consisting of Ti, V, Cr, Zr, Nb, Mo, Hf, Ta, W, Ir, and Sn; C-site is occupied by at least one element selected from the group consisting of S, Se and Te; the compounds exhibit large figure of merit so as to be preferable for thermoelectric generator/refrigerator.

Abstract: The object of the invention is to provide a solar panel and a method for manufacturing the same, wherein the manufacture of the solar panel for use in space and repair work thereon can be performed easily and in little time. An entirety of the solar panel is configured by detachably linking a plurality of unit solar cell modules, which include a plurality of solar cells and connection wires for connecting the solar cells, and electrically connecting the unit solar cell modules to one another. Thus, a solar panel can be easily fabricated by combining standardized unit solar cell modules. Moreover, by simply mechanically connecting the unit solar cell modules, they can be electrically connected, and thus the time necessary for manufacture or repair of solar panels can be shortened.

Abstract: A solar generator panel includes at least one reflector for reflecting solar radiation onto at least one photovoltaic cell. The reflector takes the form of a flexible material blade, a first end of which is fixed to the panel and a second end of which remains free. The material has mechanical properties such that, in a first position, in the absence of vertical pressure, it maintains itself erect so that its free end points to outer space, thereby defining an upper face that faces outer space and an opposite lower face that faces the panel, and, in a second position, in response to vertical pressure, it orients its upper face toward the plane of the panel. One particular application is to high-power solar panels for satellites.

Abstract: The present invention relates to an apparatus and method for the design and manufacture of foldable integrated stiffeners. These stiffeners may be used in, for example, thin-film arrays of electrochemical devices.

Abstract: The solar panel for space satellites has a plane surface comprising a plurality of photovoltaic cells (CE) disposed side-by-side and substantially forming a grid, said cells (CE) being electrically connected into at least one serpentine (SE1), each serpentine comprising cells (CE) connected in series two-by-two forming parallel cell segments (SG), two successive segments (SG) of a serpentine being electrically connected at their ends, and each serpentine comprising positive terminals (+) and negative terminals (−), said positive (+) and negative (−) terminals being distributed along each serpentine (SE1) and over the surface of the panel so that the positive terminals are far away from the negative terminals, to reduce the risk of electrical arcing between two terminals. With this arrangement, the risk of electrical arcing is reduced without increasing the fabrication cost of the panel.

Abstract: The satellite solar generator structure comprises a plurality of solar panels (PR, PS) each having a plane surface, said panels being disposed in superposition in said structure (ST), said structure including at least one rod (TI) holding said panels (PR, PS) pressed against one another, two successive panels being in contact via bearing points (AL) disposed at the peripheries of said panels. In this structure, spacers (CA) are distributed over the surfaces of the panels (PR, PS) and are disposed between pairs of successive panels in order to improve contact between two successive panels so that the structure presents better mechanical strength during the satellite launch stage. With this arrangement, a fraction of the forces is taken up by the spacers instead of being taken up by the panels, thus making it possible to optimize the solar generator, for example by reducing the weight of the panels constituting it.

Abstract: A backup power supply apparatus. The apparatus includes a light absorbing device which consists of at least one light absorbing element. The light absorbing elements are pivotally engaged with each other and are foldable for being housed inside a casing for portability. The light absorbing elements may be unfolded to absorb light from a light source to generate electric power output required by a load.

Abstract: High-output transmitters are applied to a power generation satellite and a transmission antenna apparatus and a transmission antenna having an array antenna with a small number of element antennas and a reflecting mirror antenna is applied, thereby providing a power generation satellite and a transmission antenna apparatus in a space photovoltaic generation system comprising transmission antennas each having performance equivalent to that of a phased array antenna with a smaller number of transmitters and a smaller number of element antennas than those of a transmission antenna implemented as a phased array antenna.

Abstract: A power generation satellite has a photoelectric conversion unit for converting sunlight into electric energy, a transmission frequency conversion unit for performing frequency conversion of the electric energy to a microwave, a microwave control unit for controlling the amplitude, the phase, or the amplitude and the phase of the microwave, and a transmitting antenna for radiating the microwave. A plurality of the power generation satellites are placed in space to form a power generation satellite group and an array antenna having the transmitting antennas of the power generation satellites in the power generation satellite group as element antennas is formed.

Abstract: In a solar array paddle having panels which are connected to a main body of a satellite and rotatably connected to one another, elastic insertion members for prevention of vibration are disposed in a clearance between each of the panels and between the main body and the panel adjacent to the main body when the panels are housed, folded, and layered. Each of the insertion members for prevention of vibration is inserted into the clearances and between the main body of the satellite and the panels adjacent to the main body such that one surface contacts a first panel and the other surface is adhered to a second panel.

Abstract: Provided are methods of manufacturing an electrostatically clean solar array panel and the products resulting from the practice of these methods. The preferred method uses an array of solar cells, each with a coverglass where the method includes machining apertures into a flat, electrically conductive sheet so that each aperture is aligned with and undersized with respect to its matched coverglass sheet and thereby fashion a front side shield with apertures (FSA). The undersized portion about each aperture of the bottom side of the FSA shield is bonded to the topside portions nearest the edges of each aperture's matched coverglass. Edge clips are attached to the front side aperture shield edges with the edge clips electrically and mechanically connecting the tops of the coverglasses to the solar panel substrate. The FSA shield, edge clips and substrate edges are bonded so as to produce a conductively grounded electrostatically clean solar array panel.

Abstract: The present invention provides a solar cell comprising a substrate, a first buffer layer disposed above the base layer, a second buffer layer disposed above the first buffer layer, a first boron compound layer disposed above the second buffer layer, a second boron compound layer disposed above the first compound layer, and a window layer disposed above the second compound layer, wherein the first compound layer comprises a first type of doping, wherein the second compound layer comprises a second type of doping, wherein the second buffer layer comprises a higher energy bandgap than the first compound layer, and wherein the first buffer layer and the second buffer layer permit a boron content in the first compound layer and the second compound layer to be greater than 3%.

Abstract: This invention relates generally to method and apparatus for solar-boost assist and solar-sail assist by a mini-optics light concentrator system utilizing a dynamic ensemble of mini-mirrors. In the boost phase of a rocket ship launch, the system focusses solar energy into the rocket chamber to raise the temperature of the propellant and increase the impulse. In the mid-course phase, the system focusses solar energy onto the rocket's solar sail to increase the thrust. In both cases, the system reduces the weight of the rocket by providing a weightless source of thrust.