Technology is disclosed herein for dispensing stacked spacecraft. A stack of spacecraft (102-1,...,102-9) may be joined together in an accordion configuration by disengageable links (240). A disengageable link (240) may join two adjacent spacecraft at one edge of the spacecraft. Some of the links are on one side of the stack with other links on an opposite side of the stack to provide the accordion configuration. After the tie-down mechanism (210. figure 6a) releases the stack of spacecraft from a launch adaptor (220) the stack unfolds. Initially, the disengageable links continue to hold the spacecraft together in the accordion configuration with the angle between each pair of adjacent spacecraft increasing. After a pair of adjacent spacecraft have unfolded a sufficient amount to prevent collision, the disengageable links release one of the spacecraft to thereby dispense the spacecraft.
An example of an apparatus includes a spacecraft body (1402) and a solar array (1404) that is attached to the spacecraft body (1402). In addition, a solar array spring element (1410,1414) is configured to provide a force between the solar array (1404) and one or more components of a neighboring spacecraft (1418) in a stack of spacecraft.
An example of an apparatus includes an inlet to connect to a propellant source (712) a pressure regulator (RG2) connected to the inlet to reduce propellant pressure from a first pressure at the inlet to a second pressure. The apparatus includes a manifold (730) connected to the pressure regulator (RG2) to receive propellant from the pressure regulator (RG2) at the second pressure and includes a plurality of manifold outlets. The apparatus further includes a plurality of gas lines (GLl-n), each gas line extending from a corresponding manifold outlet for connection to a corresponding satellite propellant tank (PTl-n).
A Radio Frequency Identification (“RFID”) sensor harvests power from a RFID interrogation signal received from a RFID reader and harvests power from surrounding RF energy which is at a different frequency than the RF interrogation signal. The harvested power is used to operate the sensor and transmit data sensed by the sensor.
H01Q 1/22 - SupportsMounting means by structural association with other equipment or articles
G06K 19/07 - Record carriers with conductive marks, printed circuits or semiconductor circuit elements, e.g. credit or identity cards with integrated circuit chips
5.
MICROCONTROLLER BASED SOLAR ARRAY ENERGY TRANSFER BATTERY CHARGE CONTROL
Technology is disclosed herein for a power control and distribution unit, PCDU, (206) of a spacecraft that has a microcontroller (208) to control battery charging from solar arrays (104). Using a microcontroller (208) within the PCDU (206) reduces the complexity of the PCDU (206). The microcontroller (208) may be programmable and reprogrammable, which allows the charging of the battery (216) to be adapted to various conditions. For example, the microcontroller (208) can be programmed in accordance with the mission to optimize battery charging for that mission.
Described herein is a power processing unit (PPU) for use with a Hall Effect Thruster (HET) and a Propellant Management Assembly (PMA) of a spacecraft. The PPU comprises an anode and ignitor supply subsystem that provides anode and ignitor signals to an anode and an ignitor circuit of the HET. The PPU also comprises a valve control subsystem that provides valve control signal(s) to valve(s) of the PMA. The anode and ignitor supply subsystem and the valve control subsystem are each coupled to a low voltage (LV) bus of an electrical power subsystem of the spacecraft. The anode and ignitor supply subsystem includes a step-up DC-DC converter having a transformer that steps-up a voltage of the LV bus to a higher voltage used to produce the anode and ignitor signals. The valve control subsystem is devoid of a transformer. An Electric Propulsion System (EPS) includes the PPU, HET and PMA.
An apparatus includes an integral, additively manufactured, actuation device having a rigid portion comprising a shaped structural member and a flexible portion comprising a helical torsion spring. In a spacecraft application, a spacecraft appendage may be coupled with a deployment mechanism, the deployment mechanism including at least one integral, additively manufactured, actuation device having a rigid portion comprising a shaped structural member and a flexible portion comprising a helical torsion spring.
Described herein are apparatuses and methods for use therewith that can be used to remove dust and other types of particulates from a solar array of a spacecraft, a lander, a rover, or the like. Such an apparatus can include a main body and a solar array extending from the main body. One or more piezoelectric devices is/are attached to the solar array. The piezoelectric device(s), when activated, is/are configured to vibrate at least a portion of the solar array to thereby loosen particulates adhered thereto. The apparatus also includes one or more linear actuators that when actuated is/are configured to at least one of bump against, push on, or pull on at least a portion of the solar array to thereby jettison from the solar array at least some of the particulates that were loosened by the one or more piezoelectric devices.
B08B 7/02 - Cleaning by methods not provided for in a single other subclass or a single group in this subclass by distortion, beating, or vibration of the surface to be cleaned
B08B 7/04 - Cleaning by methods not provided for in a single other subclass or a single group in this subclass by a combination of operations
B08B 13/00 - Accessories or details of general applicability for machines or apparatus for cleaning
B06B 1/06 - Processes or apparatus for generating mechanical vibrations of infrasonic, sonic or ultrasonic frequency making use of electrical energy operating with piezoelectric effect or with electrostriction
Described herein are apparatuses and methods for use therewith that can be used to remove dust and other types of particulates from a solar array of a spacecraft, a lander, a rover, or the like. Such an apparatus can include a main body and a solar array extending from the main body. One or more piezoelectric devices is/are attached to the solar array. The piezoelectric device(s), when activated, is/are configured to vibrate at least a portion of the solar array to thereby loosen particulates adhered thereto. The apparatus also includes one or more linear actuators that when actuated is/are configured to at least one of bump against, push on, or pull on at least a portion of the solar array to thereby jettison from the solar array at least some of the particulates that were loosened by the one or more piezoelectric devices.
B64G 1/44 - Arrangements or adaptations of power supply systems using radiation, e.g. deployable solar arrays
B64G 1/52 - Protection, safety or emergency devicesSurvival aids
B08B 7/02 - Cleaning by methods not provided for in a single other subclass or a single group in this subclass by distortion, beating, or vibration of the surface to be cleaned
An orbital satellite has a pair of multi-axis booms including both thrusters for course/attitude adjustment and an end effector for grappling payloads and manipulating other tools and objects. The satellite may launch with a primary payload affixed to a bus and one or more secondary payloads affixed to an ESPA ring. Once in orbit, the end effector may be used to grapple the primary and/or secondary payloads and rearrange them on the bus. In further aspects, the end effector may be used to make bus repairs or take measurements, or hold tools that are used to make bus repairs or take measurements.
A distributed computer system for a spacecraft is disclosed. The system has multiple computer nodes, each controlling a different aspect of a mission of the spacecraft. Each node includes a control circuit(s) that controls a set of components, a router processor, and a programmable processor. The programmable processor of each respective computer node issue commands to the control circuit(s) of the respective computer node to carry out an aspect of the mission associated with the respective computer node. Upon failure of the programmable processor in a particular computer node, a healthy programmable processor sends commands to the router processor in the particular computer node. The router processor of the particular computer node routes the commands received from the remote programmable processor to the control circuit(s) in the particular computer node to control the set of components to carry out the aspect of the mission associated with particular computer node.
G06F 11/20 - Error detection or correction of the data by redundancy in hardware using active fault-masking, e.g. by switching out faulty elements or by switching in spare elements
Technology is disclosed herein for a spacecraft launch restraint and dispensing structure (100). The dispensing structure (100) has a number of trusses (110) and a central structure (104). When the trusses (110) are in a support position, each spacecraft may be supported at one point by the central structure (104) and at two points by one or more of the trusses. Therefore, each spacecraft may be supported at three points, thereby providing a stable support for each spacecraft. The spacecrafts do not touch each other and do not bear the weight of other spacecrafts. In a deployment position, the trusses (110) extend away from the satellites and do not support the satellites; however, the satellites initially remain connected to the central structure (104). In the deployment position, the trusses (110) are out of an ejection path such that the satellites can be ejected in a desired sequence from the central structure (104).
Technology is disclosed for a spacecraft launch restraint and dispensing structure. Stacks of spacecrafts (102) may be arranged around a central post (104). The dispensing structure has primary tie-down mechanisms (120) that axially clamp the stacks of spacecrafts (102) when in a stowed position. Each primary tie-down mechanism (120) may have a rod (110a, 110b) located between two adjacent stacks, such that the rod (110a, 110b) tensions two stacks. In a deployment position, the primary tie-down rods extend away from the stack such that an ejection path is cleared. The dispensing structure also includes secondary tie-down mechanisms (122) that radially connect the spacecrafts to the central post (104). After the primary tie-down rods are moved to the deployment position, the secondary tie-down mechanisms still hold the spacecrafts (102). The spacecrafts may be deployed by issuing control signals to the secondary tie-down mechanisms when the primary tie-down rods are in the deployment position.
A passively damped mechanical system is disclosed, for example for use in aerospace applications where vibration can adversely affect navigational and operational instruments. In one example, the passively damped mechanical system includes an end fitting (110) of a strut (106) used to connect a structural element (102) to a payload (104). The end fitting (110) may include outer and inner cylindrical hubs (120, 122), with a space between the outer and inner cylindrical hub (120, 122) at least partially filled with a viscoelastic material (136). In a further example, the passively damped mechanical system includes legs (204) used to connect a structural element (102) to a bracket (200) configured to support a payload. Each leg (204) may include a hollow interior having a lattice (208) structure to add strength and a viscoelastic material (210) to provide passive damping.
A method of roll steering of a spacecraft to align an aspect of the spacecraft, such as the surface of solar arrays carried by the spacecraft, to the sun, is described. The roll steering occurs only when the sun is at an angle relative to the orbital plane of the spacecraft and when the spacecraft is not eclipsed by a body it is orbiting. A spacecraft may include a controller which causes an attitude control subsystem to steer the spacecraft about a roll axis to position the surface of the solar array such that an axis normal to the surface of the solar array is aligned with the direction to a sun when the sun is visible to the spacecraft, and maintain a fixed orientation of the spacecraft about the roll axis when the sun is not visible to the spacecraft.
Techniques for testing a component for compatibility with a depressurization profile associated with a launch vehicle payload fairing during ascent are disclosed. A pressure of air or other gas within a pressure vessel containing the component is raised to a first value substantially higher than one atmosphere absolute pressure. The air/gas pressure within the pressure vessel is lowered, by venting into the ambient atmosphere, at a rate simulating or demonstrating margin with respect to the launch vehicle payload fairing depressurization profile. The component is inspected for damage. The component may be a panel including a honeycomb core sandwiched between two faceskins, the panel having a planar area in excess of twenty five square feet.
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