ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
booms in assuming the final shape at the end of the deployment (Figure 11 C-E). The team noted that  increasing the boom length from 4 to 7.2 meters for the full-scale test uncovered potential issues with boom  buckling due to the shape dynamics during deployment.   The first area of focus while investigating boom buckling was the friction between the felt sliders and the  test surface. The full-scale tests were conducted at the flat floor facility. The felt sliders were placed along  the length of the boom at predetermined locations, shown in Figure 12. The sliders were assumed to put a  negligible force back into the system, which the booms could overcome. This was found to be untrue. Even  the small applied force of the sliders on the epoxy floor restricted the booms to follow their natural path out  of the deployer. The epoxy flat floor has a higher coefficient of friction than the waxed hardwood floor.  These forces had larger effects on the SSD and the boom root as the boom lengthened. The longer booms  had significantly more lateral loads to correct as the booms lengthened and the sail area increased (Figure  13).  Figure 12. Boom slider locations for ground testing.  Another contributor to the boom buckling was the location of the sail material at initial deployment. The 1G  environment pushes the sail material towards the floor, and as the sail deployed, friction and static electricity  pulled the boom tips towards the centers of the sail areas. If the material was on the wrong side of the  boom, the material would pull the boom in the wrong direction, thus starting a buckle like in Figure 13. In  the zero-G environment, this behavior would not be observed. To mitigate this during test, the team decided  that human intervention would be necessary if the sail material dragged over the booms and would be  placed in the proper quadrant before resuming testing.   Figure 13. Boom buckling in full-scale deployment test  The final contributor to boom buckling is the boom mass. On orbit, the booms would not need to compete  with gravity to support their own weight or deployment friction forces created by boom supports, sail friction,  or static electricity. However, during ground testing, friction and electro-static forces between the sail and  floor tended to cause boom buckling. As the sail deploys, the friction and static forces increase as the  NASA/CP—2018-219887 333
deployed sail area increases and the boom root’s restoring moment and stiffness lose any ability to correct.  Figure 14 shows that the friction forces furthest from the boom roots dominate boom deployment dynamics.  Figure 14. Deployed sail and deployment force diagram.  After observing this characteristic in many boom-only deployments, the team determined that any effort to  gravity offload the boom tips would remove 1G deployment characteristics and produce a more flight-like  test. Many offload options were traded, but tying balloons to the boom tips proved to be the simplest solution  (Figure 15). By removing some weight and friction forces from the regions furthest from the boom root, the  booms followed a more flight-like deployment path (similar to what was observed at the half-scale tests).  Furthermore, boom tips were more realistically affected by the sail edge tensioning created at the final  stages of deployment (Figure 11 D-E). The gravity offload and friction reduction methods gave the team far  more insight in the boom behavior on orbit.   During the ground tests, when buckling was initially observed, the team manually removed the buckling  from the system. As a risk mitigation activity, the team investigated whether or not the system could remove  a boom buckle autonomously. The new SSD motors could produce enough torque to retract the booms on  ground and on orbit. Knowing this, the team tested to see if strategic retraction operations during  deployment could remove buckles formed during off-nominal cases. The team learned that retraction steps  at 75% and 95% deployed would remove buckles while also reducing blooming in the boom spools. Though  this was a convenient operational change, the SSD has no method to determine the presence of a buckle,  nor that the buckles, if present, are removed.   NASA/CP—2018-219887 334
Figure 15. Boom tip gravity offloading.  Conclusion  Future solar sailing missions are already calling for sails 3 to15 times larger than the NEA Scout SSD. As  solar sails continue to increase in area, projects need to designate a considerable amount of effort into  developing a scaled test plan that can feasibly verify system-level requirements on the ground. The effects  of gravity and friction in the test set up can greatly influence the success of the deployment tests. Large,  thin film membrane structures are more susceptible to the effects of gravity and do not behave in a manner  that is easy to predict in flight.   Scaled testing in this project proved to be valuable and educational path for this system design. For future  cubesat missions with large deployable, thin-film structures, the team recommends allowing schedule and  resources for scaled testing. Facilities to test sails greater than 100 m2 will be difficult to find. Future projects  should reserve time to consider the efficacy of performing full-scale tests and determining the risk posture  of accepted scaled tests. Scaled testing with 3D prototype materials, motor sizing calculations, and sail  dynamics during testing were the main areas of lessons learned for the solar sail subsystem.   In summary, the mechanism encountered design challenges that stretched the team’s knowledge and  limited resources. The resiliency of the team and novel approaches kept the project aligned with the classD mission classification while keeping the overall success of the project in mind. The team hopes this paper  will provide valuable information for future solar sail mechanism designers.   References  1. McNutt, L.; Johnson, L.; Clardy, D.; Castillo-Rogez, J.; Frick, A.; and L. Jones. “Near-Earth Asteroid Scout.” AIAA Space 2014 Conference; 4-7 Aug. 2014; San Diego, CA; United States. 2. Sobey, A; Lockett, T. “Design and Development of NEA Scout Solar Sail Deployer Mechanism.” 43rd Aerospace Mechanisms Symposia; 4-6 May 2016; San Jose, CA; United States. 3. Dr. Fritz Faulhaber GMBH and Co. “Planetary Gearheads, Series 15/10”. www.micromo.com/media/pdfs/15-10_FMM.pdf NASA/CP—2018-219887 335
This rigid structure had a diameter of 2.6 in (6.6 cm) and contained finely tuned RF elements and electronic  components, requiring delicate care. This geometry is shown as the light green components in Figure 2.  The RF petals extended beyond the interlocking plates and were geometrically required to sustain a  ~0.25-in (6.4-mm) bend radius as they transitioned from the central hub structure to the larger, wrapped  geometry required of the 2U CubeSat form factor. This immediate region, defined as the transition area,  contained the highest strain within the system design. Finally, the petal length and spacecraft envelope of  ~3.1-in (7.9-cm) square required each petal to be wrapped just over two full rotations around the central  supporting structure with a maximum petal thickness of 0.020 in (0.5 mm). Further considerations were  imparted to maximize volume such as a non-circular wrapping geometry and the clocking of the central hub  with respect to the space envelope to maximize the transition area bend diameter as seen in Figure 2.  Figure 2. Basic geometry of deployed CubeSat antenna  The initial petal laminate design consisted of a homogenous approach where the central RF conducting  element was sandwiched with reinforcing lamina across the entire petal length. Here the RF element was  baselined to be Copper Beryllium (CuBe) due to its excellent electrical conductive performance and ability  to withstand high strains. Several lamina materials were considered, however, E-glass was set as the  baseline due to the material’s high strain capability, availability and low cost. Upon early modeling and  coupon fabrication, it was quickly determined that a variation of laminate designs across the length of the  petal was required to optimize performance. For example, it was found that reinforcing the outer 1/3rd of the  petal geometry did not increase deployed stiffness as the added mass of the lamina at the distal end  overpowered the benefits of the improved localized stiffness resulting in an overall lower frequency. It was  also found that this added distal mass reduced the 1g buckling resistance, a property that made ground  testing more difficult. The finalized petal architecture of the petals is shown in Figure 3 with the R0-R3  regions representing different levels of high strain composite reinforcement. The R3 region is exposed  CuBe while the R2 and R1 regions have one and two layers of glass reinforcement respectively. The R0  region is in the transition area, per Figure 2, experiences the highest strain during packaging and hence  maintains only a single layer of reinforcement. This reduces the overall deployment energy and effects of  creep. The length of each region, L A, LB and L C were optimized to increase the petal’s deployed frequency  and 1g buckling resistance. The images on the bottom of Figure 3 show an early prototype of the petal  design in the deployed and furled states.   NASA/CP—2018-219887 339
Figure 3. Schematic of petal geometry (top), photo of structural   prototype deployed (bottom left) and stowed on the rigid hub (bottom right)  Mechanical Exposure and Deployment of Wrapped Wings  The initial mechanical deployment concepts consisted of a single-phase process whereas the unfurling and  translation out of the enclosed envelope happened simultaneously. This approach however was deemed  high risk due to uncertainties of blooming / binding of the petals during initial release. In addition, risk  reduction testing of this highly coupled and dynamic deployment would require detailed analysis and a  complex offloading setup. As such, a two-phase deployment was baselined where the wrapped assembly  would initially move out of the stowed volume as a rigid body and then as a separate function, unfurl the  wings. It was determined that the best way to restrain the wrapped system was with a single circumferential  band applied at the mid-plane of the system. Upon exposure of the wrapped petal assembly outside of the  spacecraft, the band could be released with a pin allowing for a dynamic unfurling of the four wings.  Alternative approaches such as utilizing radial pins or a breakable wire that would penetrate through the  wings, to provide continuous or sequential wing deployments, were deemed as technically feasible however  not within the schedule or budget of this program.  To expose the wrapped petal assembly, several architectures were traded with the two strongest  candidates focusing on 1) a linear sliding deployment and 2) a pivoting architecture, shown in Figure 4 &  Figure 5 respectively. The linear translation approach utilized a sliding mechanism to shoot the wrapped  petal assembly out in an axial fashion. This system was restrained with two release doors and triggered by  a single hold down release mechanism. Once the petals were cleared from the cavity, the restraint band  would be released and captured by a single tether. This architecture had several advantages such as the  co-alignment with the spacecraft envelope, ease of the wiring harness management and similar design to  mechanisms associated with p-pod deployment canisters. The challenges however were identified such as  the complexity of multiple sliding interfaces, limited space at the region near the door where release device  was housed and the opportunity of binding along the winged surfaces with particular concern focused on  the restraint strap. The pivoting architecture consisted of a single door that preloaded the wrapped petal  NASA/CP—2018-219887 340
assembly within the canister. Upon release of the door via pin puller, the petal assembly would pivot out,  via torsional springs, a full 180 degrees in the opposite direction with respect to the door and hard stop  outside of the spacecraft. Half-way through this rotation, the restraint band would be triggered and captured  via tethers. Unique to this architecture is the load path and intentional contact of the wrapped petals and  the spacecraft envelope. This allowed the petals to be further compressed and preloaded when stowed,  forming a more direct and benign load path to the structure. This architecture was ultimately selected due  to the simplistic nature of the rotational mechanism, the clean preload between the furled wings / stored  envelope and lower concerns of petal binding.  Figure 4. Translational architecture showing: 1- the stowed system, 2- post translation   of the wrapped petal assembly outside of the storage volume, and 3- post petal unfurling  Figure 5. Pivot architecture showing: 1- the stowed system, 2- the door fully opened and rotation   of the wrapped petal assembly, part way through the 180-degree motion and 3- post petal unfurling  HDRM, Electrical and RF Considerations  The system deployment was initiated by a 067-011 Pin Puller provided by GlenAir of Glendale, California.  This hold down release mechanism (HDRM) restrained the door during launch and upon command,  enabled free motion of the door and subsequent wing assembly. The door and wing were preloaded with  torsional springs sized to enable swift motion to ensure unobstructed rotation of the wing assembly. While  the release of the wing assembly preload provides a strong kick-off, the motion of these two components  NASA/CP—2018-219887 341
are dominated by springs along the hinge axes. Hard stops on both the lid and wrapped wing assembly  were placed to prevent bounce-back and potential re-contact during deployment.  The GlenAir Pin Puller operates via a fuse-wire system whereas an electrical current causes a preloaded  wire to fuse, enabling a pre-loaded mechanical bolt to translate under considerable force. For this effort,  the GlenAir team fabricated and qualified a custom size, delivering a series of flight mechanisms within  three months of the request shown in Figure 6.  Figure 6. GlenAir 067-011 pin puller hold down release mechanism  The wing assembly required a rigid coaxial cable that would need to dynamically deploy with the system.  To resolve this, the cable was laid along the base of the canister with a series of bends incorporated. During  the full rotation of the wing assembly, the bends would allow for the cable to flex and twist with modest  parasitic force. The clean nature of the canister walls eliminated potential snag hazards of the cable.  Detailed Flight Mechanism Design  An early prototype was fabricated in November of 2017 for design feedback and risk mitigation. The focus  of this effort centered on verifying the stowed petal assembly shape and ensuring clearance during rotation.  It was found that the wrapped, square-like shape shown in Figure 2 was hard to achieve and instead the  petals, upon an application of preload from the system, formed an oblong shape with a defined and  consistent elastic response. This result was built into the design of the flight system. Other details pertaining  to the strap, tethers, incorporation of foam and wire harnessing are described in further detail below.   Strap and Tether Design  The stowed petals are held into a cylindrical shape with a circumferential restraint strap, this is shown in  Figure 7 and as the pink component in Figure 8. The strap has clasp features bonded on either end which  is temporarily held together with a removable release pin. This pin is connected to a tether that removes  the pin from the clasp as the system deploys allowing the antenna elements to unfurl.  NASA/CP—2018-219887 342
Figure 7. Strap assembly shown in the stowed configuration  The restraint pin pull tether works in tandem with two other tethers designed to restrict the location of the  deployed strap assembly and to capture the hardware post deployment to prevent interference with the  hardware and orbital debris. The sequence of system deployment can be seen in Figure 8. Phase 1 shows  the system in the stowed configuration with everything constrained and preloaded by the door (green).  Once the release is initiated, the door begins to open which allows the stowed hub assembly to begin exiting  the stowed volume. Phase 2 shows how the door opens faster than the hub assembly to prevent contact.  Both items are driven by torsion springs which are sized to provide adequate torque and sequence the  deployment. The detail view of phase 2 shows the release tether (left side of image in turquoise) attached  to the release pin. During this phase the tether is still loose. This view also shows the band tether (right side  of image in turquoise) in a slack state.   Phase 3 shows the release tether becoming tight which is accomplished by the cam shaped feature near  the root pivot of the hub assembly (bright yellow). As the hub assembly deploys the distance between the  release tether and base increase, thus providing a pulling motion on the release pin. Once the hub assembly  rotates a few degrees beyond the position shown in Figure 3 the release pin will be removed. It is important  to note that the system was carefully sequenced such that the petals on the stowed hub would not begin to  deploy until there was no chance of contact with the spacecraft. The springs are also sized so that the pin  will be removed in the absence of kinetic energy.  Phase 4 shows the restraint band assembly completely removed from the stowed hub assembly which will  allow the petals to begin unfurling. This image also shows how the band tether will restrict the location of  the deployed strap assembly. This is critical for the system so that there is no interference in the deployed  state. Also notice the release pin is still attached to the strap assembly with a third tether. The length of the  three tethers is critical for consistent performance and to prevent tangling. Tether length was set at final  assembly by hand deploying the hub assembly in order to guarantee successful performance.   NASA/CP—2018-219887 343
Figure 8. Deployment sequence with tether details  Strap Presence in the Door Window  A key requirement of the design involved verifying that each payload stow process was consistent and  would remain in the proper configuration throughout testing and launch. A major concern was the placement  of the restraint strap with respect to the hub assembly and placement in the stow volume. As described  above if the restraint strap assembly were to shift it could impact the timing of the tethers, increasing risk  of binding and/or premature petal unfurling. This risk was mitigated by designing a window into the door  which trapped the clasp and release portion of the strap assembly. This entrapment prevents the release  pin from shifting and being unintentionally removed from the clasps via direct interference with the door as  shown in Figure 9. Once the deployment sequence is initiated, the door swings open first allowing the rest  of the system to deploy unimpeded.   Figure 9. Release pin shown in door window NASA/CP—2018-219887 344
Incorporation of Foam  Polyimide foam was placed on the bottom surface of the chassis and inside the door to control the amount  of deflection on the stowed antenna, thus generating preload to stabilize the assembly for launch. This style  foam was chosen for its low outgassing and damping properties. The foam is also available in various  thicknesses which allowed the fine tuning of preload on the antenna by simply scaling the thickness.  Preload was measured and fine-tuned in the testing phase of the program.   Figure 10. Encapsulated polyimide foam on the system door  The foam must be preconditioned in order to function in a reliable manner. To do this Roccor essentially  used a large rolling pin to flatten the foam into a consistent thickness. During this process it was noticed  that the foam lost roughly 50% of its original height. Once in this state the amount of preload generated  remained consistent throughout testing.   The foam was positioned in the deployer and held in place using pressure sensitive adhesive. Due to the  likelihood for the foam to generate particulate, it was covered with a polyimide film and seamed around the  perimeter, this essentially encapsulated the foam. One small area of the seam was left unsealed to allow  venting of the cavity. This layer of foam and polyimide also had the secondary benefit of providing abrasion  protection on the antenna elements by preventing direct contact with the aluminum surfaces inside of the  deployer.   Wiring Harness Guidance  In order to electrically connect the deployed hub to the spacecraft, careful electrical harness routing was  required. Due to the 180-degree rotation of the hub and baseplate, significant slack was allowed in the  harness, and epoxy staking locations were limited. The only option for staking was a single location near  the top of the antenna. The smooth walls of interior volume prevented in any snagging that would prevent  deployment. Careful routing also aided in providing deployment force as opposed to developing a parasitic  torque.  Flight Development and Lessons Learned  In Q2 of 2017, two systems were fabricated and flight qualified. Due to the fast nature of this 9-month  program, many of the design features discussed above were tailored during flight build to ensure the system  functioned as expected. This included the exact sizing of tether lengths, implemented foam thickness to  control stowed preload, and the final geometric coax cable routing scheme. While numerous minor  unexpected issues presented themselves during the testing phase of the program, two were considered  significant with the possibility of preventing or degrading successful deployment on orbit. Both of these were  addressed with minor modifications to the design and are presented below.  NASA/CP—2018-219887 345
Strap Performance  The first significant find during testing was the release pin jamming in the clevises and not being extracted  during deployment. This prevented the restraint band from releasing causing a failed deployment. It was  found that this issue was caused by a wedging effect created by the tether pull force on the strap. Figure  11 shows the ideal location of the strap represented by the green line and the as tested location in red.   Figure 11. Wedging in the Release Strap Assembly  Upon further design review it was found that the geometry of the machined clasps with the tang offset from  center caused wedging when the assembly was under tension. Future designs will incorporate a centered  tang which will balance the tensile forces. Figure 12 shows the component names and geometry of the  strap assembly to better explain the situation.   Figure 12. Release strap assembly details  Redesigning the parts was not an option due to schedule, so other modifications were implemented. First  the sharp edges of the tang were chamfered to prevent them from digging into the pin. This decreased the  amount of pull force needed to extract the pin. Figure 13 shows the chamfer added to the upper surface of  the hole, this was also completed on the lower surface.  Figure 13. Tang chamfer  NASA/CP—2018-219887 346