Abstract: Technology is disclosed herein for deploying stacked spacecraft. When the spacecraft are stacked corresponding z-axis magnetic torque rods of the spacecraft will align with each other along a z-axis. Thus, collectively the stack of spacecraft have one or more sets of magnetic torque rods aligned with the z-axis. Just prior to deploying the spacecraft the one or more sets of magnetic torque rods are operated to hold the stack of spacecraft together. For example, the north magnetic pole of the magnetic torque rod in one spacecraft may face the south magnetic pole of the magnetic torque rod in an adjacent spacecraft. To deploy the top spacecraft, the polarity of the z-axis magnetic torque rod(s) in the top spacecraft is/are reversed. After the spacecraft is clear of the stack the magnetic torque rod(s) in the deployed spacecraft may be de-activated. Then another spacecraft may be deployed in a similar manner.

Abstract: A satellite includes a plurality of thrusters disposed about the satellite, each of the plurality of thrusters having a minimum thruster firing time, and a control circuit connected to the plurality of thrusters. The control circuit is configured to identify violations of the minimum thruster firing time in a non-compliant thruster firing pattern selected to achieve a specified movement, generate a plurality of compliant thruster firing patterns by replacing each of the violations of the non-compliant thruster firing pattern by zero and a minimum time in different combinations, select a compliant thruster firing pattern from the plurality of compliant thruster firing patterns to produce a satellite movement that is within a predetermined range of the specified movement, and cause the plurality of thrusters to fire according to the compliant thruster firing pattern.

Abstract: A method for providing thermal balance of spacecraft electronics is provided. The spacecraft includes two or more electronic units wherein each electronic unit is capable of performing the same spacecraft operational task. The method for balancing the temperature of spacecraft electronics further includes providing each of the two or more electronic units with a temperature sensor for determining the temperature of that electronics unit. The electronic units and their respective temperature sensors are connected to a controller. In the event that the controller determines that the temperature of an activated first electronics unit has reached or exceeded a predetermined threshold, and the controller has determined that the temperature of a second deactivated electronics unit is below a predetermined threshold, the controller automatically deactivates the first electronics unit and activates the second electronics unit to perform the task previously being performed by the first electronics unit.

Abstract: A deployable clasping system is configured to be deployed from a component and securely clasp and release an object. The deployable clasping system includes a cable that is deployable from the transit vehicle. A clasp assembly is coupled to the cable. The clasp assembly is configured to securely clasp the object. A propulsion sub-system is coupled to one or both of the cable and the clasp assembly. The propulsion sub-system is configured to maneuver the clasp assembly to the object.

Abstract: In order to provide a temporary connection between two structure elements, such as two parts of a space launcher or of another type of spacecraft or aircraft, a connection device with controlled separation comprises a thermally breakable layer, two thermal containment layers, between which the thermally breakable layer is arranged, and an electrically conductive element arranged between the two thermal containment layers in such a way as to heat the thermally breakable layer using the Joule effect when an electric current passes through the electrically conductive element.

Abstract: A satellite that includes a housing, a circuit board containing circuitry and disposed in the housing, a battery disposed in the housing and electrically connected to the circuit board, one or more weights disposed in the housing, wherein the one or more weights are disposed away from a center of the housing, one or more solar panels on the housing and electrically connected to the circuit board, and an antenna electrically connected to the circuit board and including at least one segment that extends out of the housing. When multiple satellites are launched into orbit having different mass weights, they move away from each other at least partially because of the weight difference. In orbit, each satellite is subjected to aerodynamic drag torque and gravity gradient torque.

Abstract: Disclosed is a satellite including a cylindrical main body, the main body having an inner wall defining an inner space and an outer wall, and extending along a main axis between a lower end surface and an upper end surface, at least one of the lower end surface and the upper end surface including an interface mechanism intended for engaging with a complementary interface mechanism of another satellite or of a launcher, the satellite also including at least one external device attached to the outer wall of the main body, the outer device extending to project transversely from the outer wall relative to the main axis.

Abstract: A portable electrical motor-driven exoskeleton ankle joint robot with gear transmission and control system which is intended to provide walking assistance in different speed and walking conditions to persons with disability in walking or muscle weakness or joint problem.

Abstract: A method of guidance for placing a satellite on station comprises the following steps carried out during a predefined current cycle: A) determining on the ground a law of orientation of the thrust vector, and a history of state variables and of adjoint state variables of the satellite for the transfer from a starting orbit to a target orbit using optimal control theory, B) determining on the ground a law of evolution of the rotation of the satellite about the thrust vector, on the basis of the orientation law and of the history, C) representing according to a predetermined format the evolution of the state variables and adjoint state variables so as to obtain first parameters, D) representing according to a predetermined format a law of evolution of the rotation so as to obtain second parameters, E) concatenating the first and second parameters so as to obtain a guidance plan for the satellite, F) downloading onboard the guidance plan, G) periodically repeating according to a predefined period which is smalle

Abstract: Systems, methods, and apparatus for dual port ring compatible satellites are disclosed. In one or more embodiments, a method for removably attaching at least one spacecraft to a payload attach ring on a launch vehicle involves attaching a strongback of each of the spacecrafts to two respective payload ports of the payload attach ring on the launch vehicle. In one or more embodiments, the strongback of each of the spacecrafts is attached to each of the two respective payload ports via a respective beam mounted to a respective port adaptor plate. The method further involves separating the strongback of each of the spacecrafts from the two respective payload ports of the payload attach ring on the launch vehicle by using at least one mechanical actuator on each of the beams.

Abstract: A modular space vehicle chassis may facilitate convenient access to internal components of the space vehicle. Each module may be removable from the others such that each module may be worked on individually. Multiple panels of at least one of the modules may swing open or otherwise be removable, exposing large portions of the internal components of the space vehicle. Such chassis architectures may reduce the time required for and difficulty of performing maintenance or modifications, may allow multiple space vehicles to take advantage of a common chassis design, and may further allow for highly customizable space vehicles.

Abstract: The Dual Spacecraft Deployment System (DSDS) is a concept developed for deploying Dual Satellites into independent orbits from a single launch vehicle. The Deployment Concept, Deployment System and Spacecraft Design with Representative Instruments are presented herein. The design employs an innovative half-hex design for the spacecraft buses which allows the spacecraft to be accommodated in a parallel mounted configuration on the launch vehicle. The Spacecraft are arranged on the launch vehicle with a central deployment structure with a deployment mechanism which allows independent lateral deployment of two spacecraft. The Spacecraft Design can have a large volume, e.g., sized to accommodate Spacecraft Sub-systems and Instrument Systems. This concept allows the launch vehicle to accommodate two large Spacecraft deployed to different orbits with a single launch doubling the efficiency of the launch vehicle.

Abstract: A system and method for wirelessly communicating between a host bus of a spacecraft and a secondary payload. In one configuration, the host bus and secondary payload each include a wireless interface for establishing a radio or optical communications link thereby allowing for the elimination of a complex wiring harness connected between the host bus and the secondary payload.

Abstract: According to the invention, a link is established between the fixed part (4A) and the removable part (4B) of an enclosure (4) in which the second satellite is previously enclosed, said link being such that said removable part (4B) remains attached to said fixed part (4A) when said removable part (4B), initially in a first position for which said enclosure (4) is closed and encloses said second satellite, lies in a second position for which said enclosure (4) is open and able to release said satellite.

Abstract: Launch lock assemblies with reduced preload are provided. The launch lock assembly comprises first and second mount pieces, a releasable clamp device, and a pair of retracting assemblies. Each retracting assembly comprises a pair of toothed members having interacting toothed surfaces. The releasable clamp device normally maintains the first and second mount pieces in clamped engagement. When the releasable clamp device is actuated, the first and second mount pieces are released from clamped engagement and one toothed member of each retracting assembly moves in an opposite direction relative to the other one toothed member of the other retracting assembly to define an axial gap on each side of the first mount piece.

Abstract: A spacecraft may include a module structure having a plurality of module sides. The spacecraft may include a central cylinder extending through a center of the spacecraft. The central cylinder may be the only closed cross-section extending along a longitudinal axis of the spacecraft.

Abstract: The autonomous space vehicle video monitoring system is an electrically and mechanically powered camera system for video and audio monitoring of space vehicles. The system allows for autonomous monitoring and implementing corrective actions, providing telemetry and communications to keep the health and wellbeing of a space vehicle. It utilizes a Modular Mounting formed to exact dimensions on any type space vehicle. It has an arm like device to enable complete viewing of the entire 360 degree surface of a space vehicle. The autonomous space vehicle video monitoring system also utilizes rotor joints allowing maximum angle flexibility for viewing the space vehicle. The system has electrically conductive tracks and wheels that have a camera mounted on it with an arm like device in turn enabling 360 degree viewing of the space vehicle surfaces. Delivered in pre-designed sections then fit to the specific structure and assembled on the space vehicle.

Abstract: An electrical connector assembly for suppressing a back electromotive force (“EMF”) spike originating from an electromagnetic device is provided. The electrical connector assembly includes a pin-side electrical connector coupled to the electromagnetic device. In one example, the pin-side electrical connector has a diode opening surrounded by an insulating material. The electrical connector assembly further includes a socket-side electrical connector for mechanically mating with and electrically coupling to the pin-side electrical connector. The socket-side electrical connector includes a suppression diode for suppressing the back EMF spike from the electromagnetic device. The suppression diode is positioned within the diode opening when the socket-side electrical connector is mechanically mated with and electrically coupled to the pin-side electrical connector.

Abstract: A method for simulating the movement behavior of a fluid in a closed moving container is provided. The simulation is based on the determination of the potential movement path of the center of gravity of the volume of the fluid as an elliptical trajectory situated in a disturbance plane having certain semi-axes.

Abstract: The invention relates to a load bearing interface ring (1) for a space craft. The load bearing interface ring (1) comprises a first part (3) and a second part (4). The first part (3) is arranged to carry compression loads and is mounted on an inside surface (11) of the second part (4). The second part (4) is arranged to carry shear and global bending loads. The second part (4) comprises at least two laminate plies (8), where the laminate plies (8) are oriented in a normal plane (12) of a jacket surface (5) of the load bearing interface ring (1).

Abstract: A spacecraft may include an upper core structure or a lower core structure. The upper core structure may include an upper cylinder for supporting an upper spacecraft of a dual-manifest launch configuration. The lower core structure may include a lower cylinder for supporting a lower spacecraft with the upper cylinder mounted on top of the lower cylinder. The upper cylinder may have an upper cylinder inner diameter that may be substantially similar to the lower cylinder inner diameter.

Abstract: A rocket is provided and includes booster stages at a rear of the nose cone, the booster stages being configured for propelling the nose cone in a propulsion direction and a divert control system housed entirely in the nose cone for controlling an orientation of the propulsion direction.

Abstract: A launch vehicle payload that includes at least two spacecraft is disclosed. The launch vehicle includes a single payload adapter. Each spacecraft has a launch vehicle adapter structure providing a respectively coplanar structural interface directly with the single launch vehicle payload adapter. The spacecraft share a launch vehicle payload fairing volume substantially side-by-side and are detachably coupled together such that a positive clearance is provided between adjacent, non-abutting structural body surfaces of each spacecraft.

Abstract: A spacecraft coupling system includes an integral component that can be deformed and placed into a stable state that locks one component into a mating component, and can easily be released from the deformed state, decoupling the two components. The integral component may include a central ring, a plurality of leaf elements arranged at the perimeter of the component, and a plurality of fins that extend outward from the ring to the plurality of leaf elements. A rotation of the ring element while the component is held stationary causes the fins to urge the leaf elements toward the receiving surface areas and to subsequently tension the leaf elements against surfaces on the mating component. To reduce cost and complexity, the integral component is a metal, such as titanium, that can be formed using an additive manufacturing process, commonly termed a 3-D printing process.

Abstract: A space probe including a descent module with a mobile exploration vehicle, wherein the descent module is a landing module inside which the mobile exploration vehicle is fastened, the landing module being provided with landing legs which can be deployed under the lower level of the mobile exploration vehicle.

Abstract: A system and corresponding method for a launch vehicle or spacecraft for releasably connecting a first body (100) with a first interface (101) and a second body (200) with a second matching interface (201), said system comprising: a belt (300) for releasably securing said first body (100) to said second body (200); a circumferential trench (203) at said second interface (201) for receiving said belt (300) in order to provide for a fixation of said belt (300) with relation to said second body (200); mechanical fixation means (301) attaching said belt (300) to said first body (100), said mechanical fixation means (301) being provided to enable a circumferential shifting movement of said belt (300) with respect to said circumference of said first body (100) in order to initiate a separation of the first body (100) from the second body (200), wherein said circumferential shifting movement induces a relative movement of said belt (300) with respect to said circumference of said first body (100) in order to remove

Abstract: A side-by-side dual-launch spacecraft arrangement is provided. The arrangement may include a dual-launch adaptor, a first spacecraft, and a second spacecraft. The first spacecraft and the second spacecraft may be mounted on the dual-launch adaptor and may be arranged side by side on the dual-launch adaptor. An aspect ratio of each of the first and second spacecraft may be within a range of 0.55 and 0.8.

Abstract: An adaptor system includes a space vehicle separation plate operably coupled to a space vehicle, a launch vehicle adaptor plate operably coupled to a launch vehicle capable of carrying the space vehicle into space for release of the space vehicle from the launch vehicle, an actuator release mechanism assembly, a bend/shear restrain assembly that is non-coaxial with the actuator-release connector, and a biasing element. The actuator-release mechanism assembly may be configured to separably couple the space vehicle separation plate to the launch vehicle adaptor plate. The actuator-release mechanism assembly may pass through the launch vehicle adaptor plate to engage the space vehicle separation plate to hold the space vehicle separation plate substantially parallel to the launch vehicle adaptor plate prior to release of the space vehicle separation plate.

Abstract: A gravity acceleration station for producing gravity acceleration and creating conditions for living under a permanent effect of gravity acceleration more than 1 g for prolonged periods of time. The station comprises a base and a hollow torus, rotating around a central vertical axis. A support of the station and motors for rotation of the station are located peripherally, along with the perimeter of the torus. That feature allows variable size of the station with diameter more than 100 meters, larger area for location of objects, and gradual increase of gravity acceleration from the center of the station along the radius. Due to a mechanism for altering the angle of deviation of the premises of the station, the value of the net acceleration can be changed according to the needs while keeping direction perpendicular to the floor of the premises. The station can be located on the ground or underground.

Abstract: A multiple space vehicle launch system that may be adapted to be disposed within a payload region of a launch vehicle fairing. The launch system may include a first space vehicle, a second space vehicle releasably attached to the first space vehicle and oriented relative to the first space vehicle such that, when placed within the fairing, a launch load of the first space vehicle is transmitted to and borne by the second space vehicle. In certain embodiments, the first and second space vehicles each may include one of an electrical propulsion unit and a hybrid chemical and electrical propulsion unit. Use of electrical or hybrid chemical and electrical propulsion units enables the second space vehicle to bear all or a significant portion of the launch load of the first space vehicle, thereby eliminating the need for additional support structure.

Abstract: Latching or locking deployment hinges are provided that include a shape-memory alloy (SMA) pre-load device configured to induce a pre-load across a contact interface of the latching hinge in a direction substantially perpendicular to the hinge pivot axis when the latching hinge is in a closed condition. The hinges may be sprung or actuated using a powered drive mechanism. Some implementations may feature a sprung latch/hook that is tensioned using the SMA pre-load device. Some other implementations may feature a multi-bar linkage that is tensioned using the SMA pre-load device.

Abstract: Systems and methods for interconnecting dual manifested spacecraft for launch by a launch vehicle are disclosed. Different motive forces are utilized to couple a first spacecraft to a second spacecraft and to restrict demating of the second spacecraft from the first spacecraft. Some systems and methods utilize pneumatic pressure to permit mating of a first spacecraft to a second spacecraft and utilize spring force to restrict demating of the second spacecraft from the first spacecraft.

Abstract: A spacecraft may include an upper core structure or a lower core structure. The upper core structure may include an upper cylinder for supporting an upper spacecraft of a dual-manifest launch configuration. The lower core structure may include a lower cylinder for supporting a lower spacecraft with the upper cylinder mounted on top of the lower cylinder. The upper cylinder may have an upper cylinder inner diameter that may be substantially similar to the lower cylinder inner diameter.

Abstract: A high capacity broadband service is provided from an Earth orbiting satellite having a payload that includes multiple large antennas, the satellite having an inverted configuration when deployed on orbit. The satellite has an aft surface proximate to a launch vehicle structural interface, a forward surface opposite to said aft surface, and a structure disposed therebetween. In a launch configuration, two or more deployable antennas are disposed, undeployed, on or proximate to and forward of the forward surface; on-orbit, the payload is operable when the satellite is disposed with the aft surface substantially Earth facing while the forward surface is substantially facing in an anti-Earth direction. Each deployable antenna is disposed, so as to be Earth facing, when deployed, in a position substantially forward and outboard of the forward surface.

Abstract: Payload adapters for launch vehicles include a ring structure having an opening formed therein and at least a portion of an antenna assembly coupled to a circular sidewall of the ring structure at least partially within the opening of the ring structure. Satellite assemblies may include a payload adapter comprising a ring structure having a circular opening formed therein and an antenna assembly including a parabolic reflective dish formed within the circular opening. Launch stack systems may include a primary payload and a secondary payload adapter including an antenna dish integrally formed with a ring structure of the secondary payload adapter. Methods of forming a satellite assembly include positioning at least a portion of the antenna assembly within an opening formed in a ring-shaped payload adapter for a launch vehicle and coupling the at least a portion of the antenna assembly to the ring-shaped payload adapter.

Abstract: A platform and launch initiation control system for secondary spacecraft for a launch vehicle includes an aluminum honeycomb core sandwiched between top and bottom structural aluminum skins and strengthened by a spider-like stiffener. Spool inserts for engaging the secondary spacecraft are arranged on the platform. A computer processor based auxiliary payload support unit (APSU) is provided on the platform and receives power and enable signals through a single cable from the LV to permit the APSU to control the release of the secondary spacecraft transparently to the LV without impacting LV operations.

Abstract: An adapter assembly for interconnection with a launch vehicle is provided. Generally, the adapter assembly is positioned between two separable components of a launch vehicle along a longitudinal axis of the launch vehicle to structurally interconnect the components. In an embodiment, a solar panel having a stowed state and a deployed state is interconnected to the adapter assembly. In a stowed state, the solar panel can be positioned within the interior space of the adapter assembly to utilize space within the launch vehicle that would otherwise not be fully utilized.

Abstract: One embodiment of the invention includes a spacecraft system. The system includes a spacecraft payload system coupled to a spacecraft frame. The system also includes a plurality of spacecraft panels disposed about the spacecraft frame. Each of the plurality of spacecraft panels can be communicatively coupled together via a network and configured substantially identically with respect to each other, and can include a processor and associated spacecraft control components. The processors of each of the spacecraft panels controlling the respective spacecraft control components independently to cooperatively and autonomously implement spacecraft control functions to implement a common mission objective.

Abstract: Systems, apparatuses, and methods for removal of orbital debris are provided. In one embodiment, an apparatus includes a spacecraft control unit configured to guide and navigate the apparatus to a target. The apparatus also includes a dynamic object characterization unit configured to characterize movement, and a capture feature, of the target. The apparatus further includes a capture and release unit configured to capture a target and deorbit or release the target. The collection of these apparatuses is then employed as multiple, independent and individually operated vehicles launched from a single launch vehicle for the purpose of disposing of multiple debris objects.

Abstract: Embodiments of a launch lock assembly are provided, as are embodiments of a spacecraft isolation system including one or more launch lock assemblies. In one embodiment, the launch lock assembly includes first and second mount pieces, a releasable clamp device, and an axial gap amplification device. The releasable clamp device normally maintains the first and second mount pieces in clamped engagement; and, when actuated, releases the first and second mount pieces from clamped engagement to allow relative axial motion therebetween. The axial gap amplification device normally residing in a blocking position wherein the gap amplification device obstructs relative axial motion between the first and second mount pieces. The axial gap amplification device moves into a non-blocking position when the first and second mount pieces are released from clamped engagement to increase the range of axial motion between the first and second mount pieces.

Abstract: A system, method, and apparatus are disclosed for agile dedicated spacecrafts for spinning microwave imagers and sounders. In one or more embodiments, the system, method, and apparatus involve an agile, zero net-momentum, spinning space vehicle, which includes a body and an instrument package. In one or more embodiments, the instrument package is mounted directly onto the body such that the space vehicle can point the instrument package not only along the space vehicle's orbital velocity vector, but in any direction within the space vehicle's field of regard. The space vehicle's spin axis is aligned with the instrument package's scan axis. The space vehicle experiences zero net-momentum on orbit by including a counter-rotating momentum storage device. In one or more embodiments, the instrument package is a scanning microwave imaging/sounding instrument, which is utilized as an Earth climate and weather sensor.

Abstract: A platform and launch initiation control system for secondary spacecraft for a launch vehicle includes an aluminum honeycomb core sandwiched between top and bottom structural aluminum skins and strengthened by a spider-like stiffener. Spool inserts for engaging the secondary spacecraft are arranged on the platform. A computer processor based auxiliary payload support unit (APSU) is provided on the platform and receives power and enable signals through a single cable from the LV to permit the APSU to control the release of the secondary spacecraft transparently to the LV without impacting LV operations.

Abstract: A first part, a second part, and a seal ring are fastened together to form a pressure structure configured for a spacecraft structure of a spacecraft, wherein the first part and the second part are each formed from a single workpiece.

Abstract: An apparatus and method are disclosed for dual Evolved Expendable Launch Vehicle (EELV) Secondary Payload Adapter (ESPA) port small satellite designs. The apparatus and method provide payload volume for larger satellites. In one or more embodiments, the apparatus and method include a plurality of small satellite components, a payload adaptor ring, and at least one pivoting hinge system. In at least one embodiment, the plurality of small satellite components includes at least one payload, one bus, and/or one solar panel. The small satellite components are mounted on the payload adaptor ring. At least one pivoting hinge system connects together at least two of the small satellite components. Upon deployment of the small satellite components from the payload adaptor ring, at least one pivoting hinge system combines together at least two small satellite components, thereby creating at least one single larger satellite.

Abstract: Payload adapters for launch vehicles include a ring structure having an opening formed therein and at least a portion of an antenna assembly coupled to a circular sidewall of the ring structure at least partially within the opening of the ring structure. Satellite assemblies may include a payload adapter comprising a ring structure having a circular opening formed therein and an antenna assembly including a parabolic reflective dish formed within the circular opening. Launch stack systems may include a primary payload and a secondary payload adapter including an antenna dish integrally formed with a ring structure of the secondary payload adapter. Methods of forming a satellite assembly include positioning at least a portion of the antenna assembly within an opening formed in a ring-shaped payload adapter for a launch vehicle and coupling the at least a portion of the antenna assembly to the ring-shaped payload adapter.

Abstract: Embodiments of a launch lock assembly are provided, as are embodiments of a spacecraft isolation system including one or more launch lock assemblies. In one embodiment, the launch lock assembly includes first and second mount pieces, a releasable clamp device, and an axial gap amplification device. The releasable clamp device normally maintains the first and second mount pieces in clamped engagement; and, when actuated, releases the first and second mount pieces from clamped engagement to allow relative axial motion therebetween. The axial gap amplification device normally residing in a blocking position wherein the gap amplification device obstructs relative axial motion between the first and second mount pieces. The axial gap amplification device moves into a non-blocking position when the first and second mount pieces are released from clamped engagement to increase the range of axial motion between the first and second mount pieces.

Abstract: A missile includes a separable shroud that covers a nose portion of the fuselage of the missile. The shroud covers and protects a seeker window and a seeker at the nose of the missile. The shroud is configured to remain coupled to the missile during and immediately after launch of the missile, and to separate during flight under the action of aerodynamic forces. Toward that end parts of the shroud are initially coupled together by a retainer, which allows the parts to separate during flight. The retainer may include one or more tension bands that break at a certain tension, and/or one or more weakened parts of the shroud, which break during flight. Parts of the shroud may include inward protrusions that make contact with an ogive portion of the nose of the fuselage.

Abstract: A method for implementing a satellite fleet includes launching a group of satellites within a launch vehicle. In an embodiment, the satellites are structurally connected together through satellite outer load paths. After separation from the launch vehicle, nodal separation between the satellites is established by allowing one or more of the satellites to drift at one or more orbits having apogee altitudes below an operational orbit apogee altitude. A satellite is maintained in an ecliptic normal attitude during its operational life, in an embodiment. The satellite's orbit is efficiently maintained by a combination of axial, radial, and canted thrusters, in an embodiment. Satellite embodiments include a payload subsystem, a bus subsystem, an outer load path support structure, antenna assembly orientation mechanisms, an attitude control subsystem adapted to maintain the satellite in the ecliptic normal attitude, and an orbit maintenance/propulsion subsystem adapted to maintain the satellite's orbit.

Abstract: A payload launch lock mechanism includes a base, a preload clamp, a fastener, and a shape memory alloy (SMA) actuator. The preload clamp is configured to releasibly restrain a payload. The fastener extends, along an axis, through the preload clamp and into the base, and supplies a force to the preload clamp sufficient to restrain the payload. The SMA actuator is disposed between the base and the clamp. The SMA actuator is adapted to receive electrical current and is configured, upon receipt of the electrical current, to supply a force that causes the fastener to elongate without fracturing. The preload clamp, in response to the fastener elongation, either rotates or pivots to thereby release the payload.

Abstract: An apparatus and method are disclosed for dual Evolved Expendable Launch Vehicle (EELV) Secondary Payload Adapter (ESPA) port small satellite designs. The apparatus and method provide payload volume for larger satellites. In one or more embodiments, the apparatus and method include a plurality of small satellite components, a payload adaptor ring, and at least one pivoting hinge system. In at least one embodiment, the plurality of small satellite components includes at least one payload, one bus, and/or one solar panel. The small satellite components are mounted on the payload adaptor ring. At least one pivoting hinge system connects together at least two of the small satellite components. Upon deployment of the small satellite components from the payload adaptor ring, at least one pivoting hinge system combines together at least two small satellite components, thereby creating at least one single larger satellite.