ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
In order to combat the strap assembly taking on the wedge shape in the first place, epoxy was applied  beneath the clasps to prevent a downward translation of the assembly. While the original design called for  this feature, the location of those epoxy beads did not prevent the restraint band wedging near the clevis.  The epoxy bead was therefore moved to provide better support by interfacing directly with one side of the  clip while not interfering with the release pin. This epoxy bead also improved the tendency of the strap  assembly to slide as a removal force was applied to the release pin. Figure 14 also shows how a second  epoxy bead is placed 180 degrees from the first but on the top of the strap. This prevents the rear portion  of the strap from sliding upwards which also causes a wedging at the clasps.  Figure 14. Epoxy bead placement  These updates prevented the strap assembly from translating with respect to the hub which allowed for  consistent pull forces from the release pin.   Door bounce  The second significant find was the bouncing of the door during deployment. This was not noticed as an  issue until a cold thermal deployment test. The door opens quickly and in the original design, a Viton bumper  provided damping as the door hit a hard stop at ~90 degrees. At the cold temperature, the Viton did not  damp all return motion and during testing, re-contact with the hub assembly was identified. While this did  not prevent full deployment, it did have the potential to damage the sensitive RF element edges of the HSC  petals, especially since the re-contact location was adjacent to the extrusions of the ‘window’ feature on the  door. It was determined that a better bumper material was needed that would dissipate all energy and  prevent door bounce back. Figure 15 shows the contact zone and the motion of the components at the time  of contact.  Figure 15. Door bounce back illustration  The Viton bumpers were replaced with copper mesh to better absorb the energy generated during  deployment. These bumpers were a single use item that dissipated energy by crushing when contacted by  NASA/CP—2018-219887 347
the door. Careful setup was required to ensure that the bumpers would enable full motion of the door  eliminating the chance of the bumpers preventing the door from blocking / re-contacting the wrapped petal  assembly during deployment. This was done by installing the copper mesh in place and then opening the  door to pre-crush the bumpers to the minimum opening angle of the door. Figure 16 shows a pair of crushed  bumpers on the left and a pair of partially pre-crushed bumpers on the right. In the left image the layer of  copper sheet can be seen with copper mesh below it. The entire stack was encapsulated with Kapton tape.  Many tests were conducted to determine the ideal cross section area, height and buildup of these to  dissipate the proper amount of energy.  Figure 16. Copper mesh bumpers  Flight Assembly and Other Considerations  The two flight units were fabricated in March of 2017 and successfully passed qualification testing including  vibration and stow / deploy cycles at the extreme operational temperatures. The system was found to be  robust after multiple stow and deployment cycles however the high strain composite laminate did show  signs of wear-and-tear (abrasion marks, roughening of edges, etc.) after successive deployment cycles.  This was mainly due to the difficulty in handling / supporting the petals during the deployment testing and  the inter-petal rubbing during the stowage process. As such, the flight units were subjected to a limited  number of full unfurling tests as part of the acceptance testing checkout. It was also found that immediately  after deployment, creep effects from the high strain composite prevented the wings from deploying in fully  straight manner. As per the discussion defined in a separate publication [7], this deflection gradually  reduced in severity after deployment due to stress relaxation within the laminate. The Roccor team  performed a series of long-duration tests on the laminate architecture and determined that the petals would  deploy within the required precision parameters.  Figure 17. Stowed system during flight qualification vibration testing  NASA/CP—2018-219887 348
Conclusion  In August of 2016, Roccor was approached to develop a CubeSat deployable antenna that would deploy  four large radial petals while housed in a small, 2U volume. It was found that High Strain Composites  coupled with a copper-beryllium alloy enabled a simplistic furled solution to stow the system with few  mechanical parts. A two-stage mechanized deployment scheme was selected to first expose the wrapped  petals to the outside of the spacecraft via a hinged joint and then separately trigger a dynamic unfurling of  the petals with a tether. This scheme was selected over other two-stage approaches due to the low risk of  binding, ease of securing the wrapped wings via external preload and overall mechanism simplicity. The  design was triggered by a single commercial GlenAir pin puller and utilized torsional springs to allow  rotation. The petals were unfurled via a restraint strap that was triggered and afterwards captured by a  series of three tethers. During the flight build, details such as the tether lengths, foam thickness and coaxial  wire routing were fine-tuned to ensure the system performed as expected. While there were numerous  lessons learned and small tweaks implemented during the flight build, qualification testing revealed two  significant issues that required action. The first was a modification to the restraint band clevis and addition  of guides to prevent binding during release. The second involved improving the energy absorption to  prevent bounce-back of the door and subsequent re-contact with the wrapped petal assembly of the door  during deployment. In May of 2017, Roccor delivered two, low-cost flight-qualified units to the customer  after a design, prototype, build and test campaign lasting a total of nine months. The hardware is slated to  fly in late 2018.   References  1. Doncaster, B., et. al., “SpaceWorks’ 2017 Nano/Microsatellite Market Forecast”, Proceedings of the AIAA/USU Conference on Small Satellites, Swiftly Session 2, SSC17, 2017 2. Adamowski, J., “SmallSat Market Forecast to Exceed $30 Billion in Coming Decade”, Space News , 9th, August 2017, Online 3. Murphey, T. et. al., “High Strain Composites,” Proceedings of the 2nd AIAA Spacecraft Structures Conference, 2015 AIAA SciTech Conference, Session on Composite Materials for Spacecraft Structures , AIAA 2015-0942, Kissimmee, FL. 4. Wikipedia page on “High Strain Composite Structure”, authored primarily by B. Davis 2014, https://en.wikipedia.org/wiki/High_strain_composite_structure 5. Storable Tubular Extensional Member Device, US Patent #3434674 6. Annon, “Tubular Spacecraft Booms (Extendible, Reel Stored),” tech. Prep. SP-8065, NASA, Feb 1971 7. Cox, K., et. al., “Flight Build of a Furled High Strain Composite Antenna for CubeSats,” Proceedings of the 5th AIAA Spacecraft Structures Conference, 2018 AIAA SciTech Conference, Session on Composite Materials for Spacecraft Structures , AIAA 2018-1678, Kissimmee, FL. NASA/CP—2018-219887 349
Figure 1. The NanoRacks CubeSat Deployers; Two ‘Quad Packs’ mounted to the MPEP (JAXA, 2013)  The NRCSD deployer has three major components: the baseplate assembly, the rails, and the door with  release mechanism. The baseplate assembly is located at the back of each deployer and consists of an  enclosing back plate, a spring that provides deployment velocity, and a pusher plate that interacts with the  satellite. The spring is a linear coil spring that is fabricated with a custom spring rate; the spring rate is  designed to deploy the satellites at 1.5 ± 0.5 meters per second in order to be compliant with the ISS jettison  policy. Additionally, the baseplate assembly has a ‘jack screw’, which is a screw that is installed through  the baseplate and against the pusher plate. It prevents the pusher plate from retracting as well as preloading  the system stack. The preload is required so that the satellite system is determinate and does not move  independently of the deployer during launch and ascent.  The rails are considered a mechanism and consist of features that are integrated into the NRCSD structure.  They are precision-machined flat features and serve as the interaction points to the enclosed satellites.  There are a total of eight rails that extend into the center volume of the deployer and serve to define the  payload volume. The satellites are in direct contact with the rails and slide along them as they are deployed.  The doors of the NRCSD are a ‘barn door’ design, consisting of two overlapping doors. This was done  instead of a single door so that the deployed configuration would remain inside the required volume of the  JEM Airlock upon return inside the ISS. Each door has a hinge on the exterior side, with the centers  overlapping. All doors and hinges, rails, and pusher plate mechanisms have remained the same in function  throughout iterations of the NRCSD and have been evaluated by NASA’s Mechanical Systems Working  Group. They have been determined to meet the requirements of MA2-00-057 and use the Design for  Minimum Risk designation as hazard control for the ISS program (Robertson, 2016). Note that this  evaluation was performed on later iterations of the NRCSD, and is now only applicable to the current design  in use.  NASA/CP—2018-219887 352
Figure 2. The Ball-Lock Mechanism Prior to Commanding (left) and After Commanding (right). A slight  retraction of the central plunger can be seen.  Concurrently an effort began to develop and fly a secondary latching system that was designed to prevent  inadvertent deployment in the case that commanding did not result in immediate satellite release. The  system was designed and handed over to NASA in 38 days and proved successful for its first mission.  Launched on the SpaceX-5 Crew Resupply Vehicle and installed on-orbit to the remaining NRCSD  deployers, the next Orb-2 NRCSD mission that occurred included the secondary latch system and operated  without incident. Later implementations of this mechanism by the manufacturer were not as successful,  leading to another, unrelated deployment anomaly. Current iterations of the NRCSD do not have a  secondary latch system.  Ground testing focused on the isolation of the variables that had been identified as possible failure causes.  During a series of deployment tests, the jack screw was torqued beyond specification and the signature  observed on-orbit was observed. The behavior proved to be repeatable, and the results were presented to  NASA and the Mechanical Systems Working Group. With this discovery, the analysis of the system began  and a mitigation plan was developed.  Assessment and Analysis  The baseplate assembly has threaded inserts that allow for screws to be inserted – these ‘jack screws’  pass through the external baseplate and touch the pusher plate. A total of four jack screws (Figure 3) were  used in the baseplate assemblies during the Orb-2 missions. The purpose is to apply a clamping force on  the system, namely the pusher plate-satellite-door system, and ensure that the stack-up is statically  determinate. The deployment spring in its compressed state exerts approximately 22.2 N (5 lbf) on the  pusher plate. Because of the satellite masses (typically 3-5 kg, 6.6-11 lb) and the launch vibration  environment, the jack screws were implemented as a design feature to prevent system gapping. Nominally  the jack screws are installed though the baseplate until they are touching the pusher plate, then torqued to  0.11 N-m (1 in-lb). A review of the Records of Assembly revealed that the installation torque had not been  recorded. Additionally, the manufacturer had not performed a risk assessment analysis identifying this as  a possible failure mode.  NASA/CP—2018-219887 354
Figure 3. NRCSD Baseplate Assembly (Orb-2 Mission)  An individual NRCSD deployer was analyzed using a simple Free Body Diagram. The load path was  identified from the baseplate assembly, through the satellites, into the door frame assembly, and eventually  into the ball-lock mechanism. The satellites were analyzed separately and all shown to be adequate for the  axial loads that they were possibly subject to. The analysis illustrated in Figure 4 identifies the system as  extremely sensitive to jack screw installation torque. One important thing to note is that the pin puller that  is used to retract the plunger is rated to 44.5 N (10 lbf) of retraction force. Additionally, all analysis cases  include the force of the linear spring in the compressed state, which exerts a force of 22.5 N (5 lbf) on the  system.  Figure 4. Load Path Direction of the NRCSD  The satellites are loaded into an individual NRCSD. Once loaded, with the doors closed and the baseplate  assembly installed, a preload is applied to the back of the system via the jack screws. As shown above, the  system transfers the jack screw loading through the satellites and into the doors. The satellites interact with  the door as identified in Figure 5 and transfer the loading through the steel ball bearings into the retractable  plunger. The free body diagram of the ball-lock system shows the force that is transferred to the plunger.  Note that it is assumed that the loading is symmetric – each door is loaded equally and thus the force on  each side of the plunger is equal.   NASA/CP—2018-219887 355
Figure 5. Free Body Diagram of the Ball-Lock Mechanism  The hinge screws, contact points, and doors were evaluated for structural adequacy, and all components  had appropriate margin, even for extreme loading cases. Contributions to the system by the torque spring  on the front door hinges were not considered because of the very low spring constant (0.027 N-m/rad) and  corresponding low force (<<4 N). Additionally, the spherical contact surfaces were evaluated for Hertzian  stress and were determined to not experience deformation under load; this was later confirmed by postflight metrology. A coefficient of static friction was applied based on the material properties of the plunger  and steel ball fishines, and no deformation was considered.  Figure 6. Cross-Section of the NRCSD Ball-Lock Mechanism  At approximately 44.5 N (10 lbf) of static friction, the pin puller does not have enough force to retract the  ball-lock plunger. As can be seen in Table 1, this occurs at around 0.45 N-m (4 in-lb) of installation torque.  Using linear interpolation, this value occurs at 0.42 N-m (3.69 in-lb).   NASA/CP—2018-219887 356
Table 1. Installation Torques vs. Static Friction Force on the Ball-Lock Plunger  Conclusion  Due to jack screw torque variability, it is reasonable that Orb-2 NRCSDs that experienced anomalies were  installed with a jack screw installation torques greater than or equal to 0.42 N-m (3.69 in-lb). This would  lead to ball-lock mechanism binding. The installation procedure, which did not record the installation torque,  which was uncontrolled in some cases, further indicates this as a cause.  The anomaly signature was repeatable in ground tests. If installation torques were higher than 0.68 N-m  (6 in-lb), the entire system would seize and no deployment would occur. If the installation torque was  between 0.45 and 0.68 N-m (4-6 in-lb), the signature observed on-orbit was recreated. In these cases,  delayed deployment was observed in approximately 50% of cases, which is also consistent with the  behaviors observed on-orbit.  Implementation of New Processes  A number of new processes were implemented based on the conclusions of the anomaly investigation. A  notable early solution to the safety risk of delayed deployment, the secondary latch, actually led to additional  anomalies. The primary solution was the removal of the jack screws on-orbit. Additional NRCSDs were  onboard the ISS during the time of the anomalies and investigation; the jack screws were removed by  astronauts onboard, and the mission was continued. All deployments were successful, and this operational  procedure was implemented for all subsequent NRCSD missions.  Ground procedures were also modified to include the recording of the jack screw installation torque. The  design of the baseplate was modified to include only a single screw, which provided a much more  predictable control of installation torque. Tests were performed to measure the effect of installation torque  variation, which validated the previous analytical models.  The anomaly, the secondary latch issues, and other factors led to the decision to perform a complete  redesign of the NRCSD door and release mechanism. The new design was performed by NanoRacks  engineering (previous design had been done by the manufacturer). The new design included the removal  of the ball-lock mechanism and an overall reduction in part count. A tolerance assessment was performed  as part of the design process, the results of which were presented to the Mechanical Systems Working  NASA/CP—2018-219887 357
Group. Additionally, the Aerospace Corporation was hired as a third-party reviewer to ensure design  functionality. The new door was first implemented within a year of the on-orbit anomalies and has had a  100% success rate.  References  1. JAXA. (2013). JEMRMS Multi-Purpose Experiment Platform/Payload Standard Interface Control Document.  Japan Aerospace Exploration Agency. 2. Robertson, B. (2016). Design for Minimum Risk (DFMR) of the NanoRacks CubeSat Deployer - Internal (NRCSD-I).  Houston, TX: NASA. Figure 7. The NRCSD on the MPEP, grappled by the JEM RMS  NASA/CP—2018-219887 358
The antenna was chosen to be optimized for Ka-band, specifically the frequency range from 32 to 35 GHz,  as this is a frequency used by the Deep Space Network for communications and is also a frequency for  precipitations radars. To ensure high efficiency at this frequency, the antenna had to deploy to a surface  accuracy of 0.4-mm RMS or greater. A trade study on antenna diameters revealed a 0.5-meter antenna  would be large enough to offer a major advancement in capabilities. Combining these three aspects into  one antenna would multiply data rates by 100 times what the APDA antenna would have achieved.  Because this system was targeted for a CubeSat, it had several key dimensional constraints. CubeSats are  modular satellites, built on around a 1-unit (U) system. One “U” is 10 cm by 10 cm by 10 cm. CubeSats  have been launched in sizes of 1U, 1.5U, 2U, 3U, 6U, and 12U. (6U systems are approximately 10 cm by  20 cm by 30 cm, whereas 12U systems are 20 cm by 20 cm by 30 cm). To accommodate the CubeSat unit  system, the antenna had to stow in a 10 cm by 10 cm cross-sectional square, with a goal of keeping the  height as short as possible, at approximately 15 cm. This would allow the antenna to consume only half of  a 3U spacecraft or a quarter of a 6U spacecraft.   Key Subsystems and Components  An overview of the key subsystems and components is beneficial before discussing the development of the  KaPDA antenna. The subsystems are illustrated in Figure 1. The canister and hub make up the primary  deployment actuation  sub-system. The canister encircles the antenna when stowed. Near the bottom of the  canister is the hub, to which the ribs and the horn mount. The ribs are divided into two parts, the root ribs ,  which attach to the hub and are the closest to the center of the antenna, and the tip ribs , which are the  outermost ribs when the antenna is deployed. The tip ribs are attached to the root ribs via the mid-rib hinge .  The horn is primarily an RF component, but the exterior walls serve to guide and position the sub-reflector and position the ribs when stowed.   Figure 1. Key KaPDA Components  Overview of the KaPDA Development Sequence  To deploy, the hub is first driven upwards. (Figure 2, A-B). As the hub nears the top of the canister, the root  ribs begin to bloom, opening (B-C). When the tip ribs reach the point where they become free of the horn  interference, they are free to actuate at the mid-rib hinge (Image C). The hub continues to travel upwards  until the root ribs fully deploy (image D). After the ribs are mostly deployed, the sub-reflector is allowed to  telescope along the horn, and reach its final fully deployed location (C to D).  © 2018 All Rights Reserved. Government sponsorship acknowledged.  NASA/CP—2018-219887 362
Figure 2 . KaPDA Deployment Sequence Early Development of KaPDA  Characterization of the Aneas Antenna  The development of KaPDA began by researching the Aneas antenna. The team firs t met with the Aneas  team at USC/ISI to capture lessons learned, and was able to borrow the antenna from USC/ISI for  metrology. A theodolite measured th e accuracy of the deployed shap e of APDA and foun d tha t the surface  error was an average of 2.4-mm RMS . While this is perfectly adequate for an S-band antenna , requirements  for a Ka-band antenna are tighter, a t 0.4-mm RMS maximum error . APDA was designed with thin ribs,  which helps to reduce storage space , but also impacted the surface accuracy possible . Torsion springs  actuated the hinges and setscrews adjusted the position . These were key elements requiring redesign to  improve accuracy.   A second issue when going from an S-band design to a Ka-band design is the mesh. At S-band, 10 opening  per inch (OPI) mesh is adequate, which also requires a low amount of tension to achieve its shape. The 40  OPI Ka-band mesh requires a much greater tension of 17.5 N/m. This  means the antenna must be designed  to achieve  greater  values  of preload  upon  deployment;  approximately  250  N. The  APDA  deployment  architecture  only  achieved  a fraction  of this,  and  therefore  the deployment  approach  needed  to be  completely redesigned.   The third  and  most  significant  issue  is the RF design.  The  Aeneas  antenna  used  a splash  plate  feed  connected to a co-ax cable. At Ka-band this would create far too much loss, removing any gains achieved  through surface accuracy. As a result, an entirely new RF design was required for the antenna to operate  at Ka-band with minimal loss.   RF Design Effort  The first approach to create a system tha t would operate a t Ka-band was to develop th e RF design . While  the idea of usin g a parabolic dish t o reflec t RF energy remains the same as APDA, the res t of the system  had to be completely redesigned, t o the poin t which th e KaPDA RF design and the Aeneas RF desig n share  no heritage. In order to achieve high frequency communications with low loss, th e RF energy mus t be kept  in the electro-magnetic wav e form all the way through exiting the antenna . Three subsystems were used to  achieve this : the secondary reflector, horn and waveguide . The secondary reflector collects the RF energy  from the parabolic dish , and reflects it into the horn. It also has a critical feature where it corrects for the  geometric errors i n the mesh which occur because of the finite number of ribs . The horn concentrates RF  © 2018 All Rights Reserved. Government sponsorship acknowledged.  NASA/CP—2018-219887 363