ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
433 Table 1. XSAS Mass Budget  Components  Mass (kg)   Array* 2.05   Release Mechanism * 0.13  Power Bus* 0.61   Ballast 0.71   Payload 1.50  TOTAL 5.00  * Includes 30% contingency 

It will be up to the continued design to modify XSAS fo r the new mass budget. Base d on the Table 1, this  could be achieved by incorporating the payload as  ballast rather than using empty weight.  

Design Requirements 

Top level requirements for XSAS are divided into three categories as de fined below. While most  requirements are vital to the success of this innovative design, some can be modified to satisfy mission specific constraints and needs. This ability to customize t he design makes XSAS a truly modular and  versatile power system, which is attractive to a va riety of researcher groups and industries. Specific  qualifications on design, performanc e, and operation are included within each major category to describe  the purpose of each requirement.  CubeSat Interface Constraints

The following list details the top level system requirem ents, followed by explanations of their application.  • XSAS shall have a stowed and deployed configuration  • Chassis shall include a 0.5U power bus  • Chassis shall include a 1U stowed solar array  • XSAS shall have a maximum mass of 5.0 kg   Due to their standardized geometry and deployment systems, CubeSats have strict requirements on their  stowed design. Maximum mass and volume constraint s include 5.0 kg and 3U geometry respectively.  Therefore, as listed, XSAS is required to be no more than 1. 5U when stowed. This size allows for a 0.5U  power bus for the power board, batteries, release mechanism, and payload interface, 1U for the stowed  16 panel array and structure, and the remaining 1.5U for payload and interfaces. A maximum mass of 3.5  kg is reserved for the XSAS system, leaving 1.5 kg for payload, reaso nable for many CubeSat  applications currently under consideration.  There are two flight configurations. The first is a stowed configuration where the panels are retracted like  an accordion and constrained in all degrees of freedom by a release mechanism. The second is a deployed position where the panels are extended and constrained by hinge mechanisms and a scissor structure.   XSAS Performance Requirements

Subsystems requirements shape the design and interact ion of the power system, structure, mechanisms,  attitude control system (ACS), and ground systems. The following list details the top level sub-system  requirements, followed by explanations of their application.  • XSAS shall provide at least 20 W beginni ng of life average continuous power  • Solar array shall have an active and redundant release mechanism  • Solar array shall have passiv e deployment and locking mechanisms  • Deployment and locking sequences shall apply zero shock to the power system and payload  • Once deployed, XSAS shall provide gravity gr adient stabilization for payload operation  • Ground system shall be able to ve rify fully deployed configuration  NASA/CP-2010-216272
436   ௦௣ൌቄ  ிೞ ௖ܣሻ , (2) 

where ௦௣ is the solar radiation pressure torque, ௦ is the solar constant, c is the speed of light, ௦ is the  surface area,ݍis the reflectance factor, ݅is the solar incidence angle, ௣௦ is the center of solar pressure,  and ݃ܿis the center of gravity; and 

௔ൌቄଵ ଶௗܣሻ , (3) 

where ௔ is the aerodynamic torque, ߩis the atmospheric density, ௗ is the drag coefficient, ܣis the  surface area normal to the direction of travel,ܸis velocity, and ௣௔ is the center of aerodynamic pressure.    From observation, the first term in Equations 2 and 3 is determined by the orientation and orbit of the spacecraft. Assuming these ar e constant, the second “sta tic margin” terms,  however, are controlled by the  mass distribution of XSAS. By controlling its deployed cg position with a ballast or secondary payload, these terms can be minimized to ensure the grav ity gradient stabilizing torque is dominant. Figure 2  displays the results from a simulation that estimate s the nadir pointing error for a 180° orbit between  poles at a range of LEO altitudes (500 to 2000 km). The simulation uses inputted mass and geometric values for an XSAS array, ballast, and payload. It assumes all restoring and disturbance torques are  balanced to determine the stable nadir orientation.  

Figure 2. Pointing Error from Balanced On-Orbit Torques:   (a) Current Mass Distribution; (b) Optimized Ma ss Distribution (Identical Axis Labels)   Figure 2a and 2b display the difference between an unbalanced and balanced deployed mass distribution for a XSAS 3U CubeSat and target 1. 5-kg payload. Condition 2a is dom inated by aerodynamic effects at  low altitudes, while 2b has both aerodynamic and sola r radiation components. In fact, 2a must fly above  900 km to have a 5 degree minimum pointing error. Unfo rtunately 2a refers to the current design, while 2b  represents an ideal, but over-budget, XSAS with an additi onal 0.7 kg of ballast. With this understanding,  the next XSAS design must implement a more effi cient distribution of ballast and payload mass. 

Mechanisms   The overall layout of XSAS is defined by the af orementioned requirements and simulation results.  However, it is in the mechanical design and integrat ion of the mechanisms that  validate the reality of  XSAS beyond the simulation or theoreti cal concept. This however proved to be a non-trivial task with  contradicting tight volume and revolutionary powe r requirements. Careful design of the primary 050100150200 500100015002000-30-25-20-15-10-505 050100150200 500100015002000-2-1.5-1-0.500.51 Altitude (km) Orbit Position (deg) Pointing Error  (deg)  a. b. NASA/CP-2010-216272
437 mechanisms: the release and deployment  systems, was completed to ensure their successful integration.  Modeling using Solidworks computer aided design software proved to be a very useful integration tool as  it allows assemblies to move about realistic physical constraints. A three-pan el, fully constrained XSAS  CAD model was sufficient to capture all of the critical  mechanical details. Figure 3 is a detailed view of the  final April 2009 prototype CAD model. Each labeled component represents a sub-mechanism.   Table 3 lists the sub-mechanisms wi thin each primary mechanism. The XSAS chassis is also included in  the list because it interfaces directly with the depl oyment and release systems. The following subsections  will describe the design, application, and functionality of these features. 

Table 3. Sub-Mechanism Distribution 

Mechanism Sub-Mechanism   Chassis • Power Bus Structure • Ballast Plate    Release • Dyneema Cutter • Release Panels    Deployment • Hinge  • Array Panel • Scissor System • Rotary Damper 

Figure 3. CAD Model and Key Design Features     Ballast Plate  Release Panel  Scissor Structure  Array Panel  Rotary Damper  Power Bus Structure Primary Scissor Joint Hinge and Locking Latch  NASA/CP-2010-216272
438 Chassis   The XSAS chassis is composed of a power bus and ballast plate. The standardi zed geometry of COTS  chassis makes them unsuitable for use on these compon ents. As a result, the custom design of each  focuses on the versatility and machinability of the structure.   The power bus idealizes this criterion in the design of the structure and internal space. It houses the release mechanisms, batteries, and power board, and also serves as the attachment between the payload bus and the array system. With connections to a variety of components that are either unknown or constantly changing in design, the layout of t he bus will also need to chan ge. Therefore the structure  must be versatile and simple to modify. As a result, it is composed of removable panel walls fastened to aluminum posts. The individual panels are simple to machine, allowing single panels to be removed and modified without impact to the rest of the desi gn. Additionally, removable panels permit easy access  during assembly, integration, and testing. Finally, the solid aluminum posts allow for a continuous surface  for mounting internal components. This method is also  used for the fabrication of the ballast plate.   The ballast plate provides attachment volume for ballast mass and a standard interface to the P-POD deployment mechanism. The additional mass correlates the positions of the spacec raft’s center of mass  (c m) and center of drag (c d) to help control the passive ACS b enefits of the deploy ed structure. The  amount of mass is dependent on the geometry and mass distribution of the array and payload according to the gravity gradient stabilization theory describ ed in the Simulations section of this paper. 

Figure 4. Exploded View of the Power Bus Showing Removable Panels and Mounting Posts 

Release Mechanism

The release mechanism is composed of two hinged release panels and a cutter system. Motivations for  the design and integration of each feature focus on low cost, low profile, and high reliability.   The resulting cutter system successfully incorporates  each motivation. It is composed of 4 metal film  resistors wired in parallel and a single line of Dyneema wire. The Dyneema wire is used to tether the release panels to the power bus structure. To do th is, the wire is threaded between two holes on each  release panel creating parallel lengths of wire that span the bus' interior. Each length then rests against two resistors mounted to the top surface of the bus.  To deploy the release panels, an on-board 9-V source is  shorted across the resistor circuit. The resulting  heat generated by the resistors cuts the wire allowing the release p anels swing to open. Since one strand  of Dyneema wire is used, only one resistor within the quadruple redundant array is required for a  Removable Panels Ballast Mounting Posts  Power Bus Mounting Posts  NASA/CP-2010-216272
439 successful release. Additionally, with the placement of the panel attachments and positioning of the  resistors, the entire system requires a fraction of the available volume within the bus. Furthermore, the  use of redundant resistors and Dyneema was implem ented on the Delphi C3 nanosat mission [12] to  deploy its solar arrays, proving the re liability of the inexpensive system.  

The release panel design mimics the motivations used on the cutter system. Each panel is composed of a custom formed aluminum sheet and a single spring hi nge. The geometry of the panels fully constrains the  stowed array with minimal impact on the array panel dimensions. To accomplish this, each release panel has three folded edges called ribs. The side ribs will be referred to as ribs 1 and 2, and the bottom rib will be rib 3. This labeling is displayed in Figure 5. Ribs  1 and 2, and the lower portion of each panel fit within  recessed areas on the power bus. This constrains the lateral translational degrees of freedom (DOF) of  the array and rotational DOF about the vertical axis , z. Rib 3 is held within a slot in the power bus  constraining the vertical translational DOF and rotational DOF about the x axis. The final rotational DOF about the y axis is constrained by the tension of t he cutter system's Dyneema wire and the resulting  interference of the panel and power bus. Once the system is released the spring hinge quickly rotates the  release panels to the ballast plate. This prevents any interference with the deployment sequence of the  array and also adds additional mass the ballast plate.  The panels are also designed to protect the solar array mechanisms and structures while stowed. The ribs create a constant surface along the edges of the stowed array to act as  rails within the P-POD  deployment mechanism. According to standard CubeSat requirements, the rails must cover at least 75% of the CubeSat length [5]. The release panels provide a nearly complete surface to satisfy this requirement.     

Figure 5. Release Panel Geometry and Constraint Method   Deployment Mechanism

The deployment mechanism is made up of the hinge system, array panel structure, scissor system, and rotary damper. Like previous mechanisms, the design of the deployment mechanism was non-trivial, needing numerous iterations to best satisfy the XSAS requirements. Each sub-mechanism is inspired  from simple known concepts that have been creativity applied to the CubeSat geometry.   The hinge system is a compound mechanism that passiv ely controls, deploys, and locks the solar array at  each panel joint. The challenge was minimizing the ve rtical height of each compound hinge within the  stowed 10cm length, as the height det ermined the number of solar panels in the array. The original design  concept (Figure 6a) employed separate spring, damping, and locking hinges. This design did not meet COTS and size requirements. As a result, this arrangement would need custom designed components,  which proved prohibitively expensive. The next it eration (Figure 6b) removed the damping hinge and  z  y x 1  2 3 Power Bus  Recessed Areas  NASA/CP-2010-216272
440 replaced the locking hinge with a plunger bolt. Again, failed size requirements and tolerance issues forced  another redesign. The final two designs (Figure 6c,d) employ a modified COTS spring hinge mated with a  custom locking latch. The end of each latch arm is  bent at a 90 degree angle creating teeth. While  stowed, these teeth rest on the outside of the hinge barrel, which pretensions the latch arms. During  deployment, the teeth ride along the out side of the rotating barrel until they fall into slots machined into  each hinge. This not only locks the hinge, but orients t he open array at a predefined angle. In the case of  XSAS, it is 180 degrees. Manufacturing constraints in fluenced the final redesign displayed in Figure 6d. 

Figure 6. Hinge Design Evolution   The scissor system manages the orientation of the a rray during the deployment sequence. This concept  incorporates inspirations from a co mmon scissor lift. It is composed of two parts, the scissor structure and  primary scissor joint. The structure coordinates a consistent deployment of each array panel and is fastened to only one side of the array, minimizing losses to the array panel area. The joint restricts the movements of the structure to a single vertical DOF.  To control this vertical motion, the final joint  connection slides within a channel cut into the bus panel wall. Using simulations from the CAD model, the joint arms are placed as close to the bus wall as poss ible as shown in Figure 7a. As a result, the joint  does not impact the geometry of the array and has minimal impact on the volume within the power bus.     Additionally, the scissor joint provides a single point at which array can be controlled. Since the damping  component of the hinge system was re moved from the final design, a rota ry damper is placed at the base  of the array to control the deployment velocity. The da mper uses viscous effects to counteract the velocity  of the array as the damping constant increase with an increased rotational velocity.         

Figure 7. Scissor Joint Composition: a. Front Internal View; b. Side View   Successful assembly of each compone nt is made possible by the geometry  of the panel structure. Proper  cutouts and hard point connections within each p anel manage all connection an d clearance requirements.  This was designed confirmed through CAD simulations.  Solar cells assembled on printed circuit board  (PCB) can then be plated to the primary panel structure. 

   a. b. c. d. Scissor Joint   Battery   Power  Board   Control Channel  Pivot Axis  Latch  Hinge  NASA/CP-2010-216272
441 Hardware and Prototype 

The primary goal of this first phase of the XSAS de sign was to create a protot ype to demonstrate and  study the manufacturing methods, in tegration, and functionality of t he high risk mechanisms. Therefore  the prototype serves primarily as a technology demonstrator rath er than flight hardware. It was  successfully tested using the af orementioned release and deploym ent system. The prototype and  deployment sequence is displayed in Figure 9. This section describes the deta ils of the hardware used. 

 a.  b.   c.  d. 

Figure 9. XSAS prototype deployment sequence: a. Stowed configuration;   b. Metal film resistors positioned within power bus burn Dyneema wire restrains;   c. Scissor structure evenly deploys panels; d. The completely deployed XSAS structure   All materials used on the XSAS protot ype were purchased from stock supplies. Table 4 lists the materials  used for each primary mechanism. Aluminum 7075 and aluminum 6061 are used for all chassis panels  and posts respectively. The most obvious reason for this selection is aluminum’s low density, corrosion resistance, and relatively high st rength, which makes it widely used on spacecraft chassis. Additionally,  each type of aluminum is approved fo r CubeSat structures by the CubeSat Requirements document [5].  Aluminum 5052 is used for the rele ase panels because of its good form ability properties. It is used by  COTS CubeSat Kit structures, which are bent to optim ize the stiffness and mass of the structure. Quasiisometric 2-mm carbon fiber plate, manufactured by Dragonplate, is us ed for the solar panel structure to  optimize the stiffness of the deployed panels. The prof ile of each material is design used CAD software  and cut using a CNC waterjet cutter. Flat-head, counte rsinked machine screws are also used to join all  exterior structural panels. This keeps a smooth exterior surface, prev enting any interference with the PPOD deployment mechanism.      NASA/CP-2010-216272
443 hardware is much more complicated. Stainless steel sleeves with a #2 inner diameter made it possible to  stagger the locations of each arm. Again 2-56 hardware and brass washers are used.  Rotary dampers produced by ITW De lpro for the automotiv e industry where donat ed to the XSAS project.  Unfortunately it was discovered through testing that these rotary dampers acted as a friction joint rather than a viscous damper. Although no COTS rotary dampe r was ever found for this application, it was  included in the design to demonstrate an integration method for future development   Testing   Testing was limited to the observed bench-top performance and condition of each mechanism. To accomplish this, low-friction test carts, made of acrylic panels and captured ba ll bearing castors, were  fabricated to provide the best ben ch-top method to observe  the behavior of the system. XSAS is mounted  on the carts with the hinge axis perpendicular to the table surface. Figure 11 displays the assembled test  cart.  

Figure 11. Assembled Test Cart (Bottom View) 

This setup allows for minimally constrained 2DOF motion along a table surface isolating the controllability of the deployment system. To furthe r validate the test conditions, the motion in and out of the test plane  should be negligible even in orbit due to the inhere nt bending stiffness of the st ructure against the hinge  axis. Additionally, to test the release mechanism th e cutter resistors were wired to a switch and a 9-V  battery on the bottom of the test cart. The switch and del ay in the cutter allows the tester to activate the  circuit and step back from the system, prevent ing them from impacting its performance.    The release system performed flawlessly during all at tempts and the scissor system worked as planned to  provide a consistent expansion of all array panels. Improvements focused on the redesigned of the hinge  and locking system. Fatigue from multiple test trials permanently deformed the barrel hinges and locking prongs negating the required tolerances to ensure a locked hinge. Additionally, due to the strain and  unidirectional layup of the scissor joint members, one of  the arms split in half along its length. This failure  points towards the use of metallic components for the joint structure since an improved layup could not be  found in COTS components. 

Conclusions and Future Work 

While the concept of a deployable solar array such as XSAS is simple, its implementation is nontrivial,  especially under the constraints of the CubeSat form. The analysis and simulations presented in this paper yielded a functional prot otype of the XSAS mechanism. Addi tionally, analytical trades have  characterized its power and ACS capabilities. Results proved extrem ely successful for a concept of this  infancy. Successful, integration, deployment, and control of the array were achieved, proving the  hardware could function within CubeSat volume and mass constrains. Moreover, estimated average and  maximum power capabilities at 23 and 32 W, respectively, exceed initial design requirements.   XSAS  Ball Casters Caster Panel Resistor Battery  Resistor Switch  NASA/CP-2010-216272
446 

Figure 1: 6.5mm “slice” of CubeSat bus where SADA will integrate on the end of typical  CubeSat structure. 

Mechanical Description  

The solar array drive assembly performs key system functions, rotating the solar arrays to keep them  optimally oriented with respect to the Sun and providing a path for power transfer from the arrays to the  CubeSat bus. The prototype system is shown in Figur e 2. This prototype was specifically developed to  make use of off-the-shelf technology to minimize the cost and complexity of the design.   Solar arrays are attached to either end of the SADA onto output shafts protruding from the unit. Cable-twists are used to carry power from the arrays and each wing has a dedicated cable-twist assembly. Cable-twists offer several advantages, but from an o perational standpoint they need to be protected from  over winding. To prevent over-winding, mechanical hard stops are implemented to limit the rotational travel of the output shafts. The SADA is designed to accommodate slip rings for continuous rotation if  required, but the total number of circuits available for both power transfer options is limited due to the small height of the device. This became one of t he major challenges in the electrical design. 

Figure 2: Prototype unit shown with cover on (left) and off (right) on top of Honeybee Robotics headquarters in Manhattan. 

   NASA/CP-2010-216272