ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
Because these tests would need sub-micrometer level precision over long durations of creep recovery,  position sensors with less than tens of nanometers scale errors were necessary. Temperature changes,  humidity changes and vibrations in the testing environment can significantly affect measurements at this  scale. Therefore, special attention was given to controlling and/or measuring the temperature change,  humidity change and vibration of the specimens during testing.   Test Environment  In order to minimize temperature and humidity variation during the creep recovery, the tests were performed  in a passively stabilized thermal chamber. With the doors sealed, the temperature variation within the  chamber over 24 hours was generally less than 0.1 degree Celsius. The chamber also provided passive  stabilization of humidity, but because the room housing the chamber was not humidity controlled the  baseline humidity level did change throughout the tests. Measurements show that over a 24-hour period  the humidity did not change by more than ±1 percent. The tests were performed on a 91.4 cm x 122 cm,  90-kg optics breadboard that was isolated with respect to the surrounding floor with four Barry Controls SLM-1A passive, inflatable, air-spring isolators. The isolators did not remove all disturbances from the surrounding room, but it did decrease the resulting response levels at the test apparatus. However, because the creep recovery after each deployment was at least 15 hours and data were sampled at 1 Hz, everything but extremely low-frequency vibration in the specimen could be averaged out of the data. The hinge deployment precision testing was performed on an Invar reference plate, as shown in Figure 4,  so that the sensors at the tip of the specimen would not move significantly with respect to the base mounting  point of the specimen during the tests. For the temperature variation seen in the sealed test chamber, the  thermal expansion of the reference plate over 24 hours should be less than 50 nm. The Invar reference  plate was mounted to the optics bench with one fixed and two flexured mounts. The flexured mounts  minimize the load that is applied to the reference plate due to the thermal expansion differential between  the Invar reference plate and the steel optics bench.  Figure 4. Hinge deployment precision test layout  The Invar reference plate also includes an adjustable stow restraint, shown in Figure 4. This provides a  fixed point to stow the specimen in a repeatable manner. In order to ensure stowed position repeatability,  the stow restraint contains one-half of a ThorLabs KB3X3 kinematic mounting base. At the mirror end of  NASA/CP—2018-219887 421
Figure 5. Test procedure flow chart  HSC Hinge Deployment Precision Results  An example of the post-processed results for all ten stow/deploy cycles for Specimen 4-6 is shown in  Figure 6. In this figure, the periods with no data are either stow periods or down time after a deploy collection  is complete and before the next stow. The blue points in the position plots are recovery periods, red points  are the final positions before the next stow. The results for each specimen is summarized by comparing the  absolute and consecutive change in the red final position data. For the tenth deployment, the final  60 seconds of recovery period data is used as the final recovered position for that cycle. A comparison of  the absolute and consecutive change for the results shown in Figure 6 is provided in Figure 7. The results  shown in Figure 7 highlight the fact that although the recovered position may drift over all ten stow and  deploy cycles, the test-to-test change remains relatively small. This is the real performance metric needed  for the planned application since the hinges will be folded up for launch then unfold a single time in orbit.  Two other characteristics were common in most of the specimen measurements. The first is the relatively  large amplitude of the noise in the pitch data when compared to the other axes, which is evident in  Figure 6. This is due to the small level of pitch error measured and the small baseline between the sensor  positions for this measurement, which amplifies the measurement noise in this result. The second is the  large position change that occurred after the first stow. While it is not as evident in the Specimen 4-6 data  shown above, it is evident in many of the other specimens. This may be indicative of a one-time setting of  the majority of the non-recoverable stress relaxation in the folded hinge. Therefore, in calculating the  deployment precision data for the specimens only the consecutive position error for tests 2-10 are used.   The final deployment precision data for all test specimens is shown in Table 4. Results from Specimen 4-2  are not included because it failed due to creep rupture during the third stow period. Results from Specimen  6-3 are not included because it was not tested based on the poor results for this material demonstrated in Specimen 4-3. The hinge testing showed that two of the six composite layups demonstrated piston precision within 0.53 µm, pitch and roll precision within 5.9 µrad and axial precision within 0.36 µm. This level of piston and axial deployment precision with a passive hinge is similar to the best measurements presented by Domber et al after thermal bias was removed. Initial analysis of the current data to remove thermal bias show marginal improvement in deployment precision. These results will be included in future publications. NASA/CP—2018-219887 424
Table 4. Hinge deployment precision results  Measurement RSS  Error 4-1 4-3 4-4 4-5 4-6 6-1 6-2 6-4 6-5 6-6 Piston (µm) 0.04 1.26 19.91 1.32 0.53 0.27 1.38 0.80 1.58 0.38 0.39  Roll (µrad) 1.01 6.48 50.39 3.47 3.05 5.83 7.04 2.05 5.86 5.16 3.76  Pitch (µrad) 1.67 5.94 18.69 5.3 4.34 1.49 5.73 2.25 12.89 4.43 2.36  Axial (µm) 0.02 0.17 0.64 0.36 0.26 0.27 0.31 0.14 0.51 0.09 0.25  HSC Frame  After sub-micron deployment precision of HSC hinges had been demonstrated, the question remained as  to whether sub-micron deployment precision could be achieved in a complex structure assembled from  multiple HSC hinges. Although frames using HSC tapes have been developed before13,19, HSC hinge based  frames with sub-micron level deployment precision have not been demonstrated in the literature. To  investigate the potential precision of a deployable frame, a deployable frame test article was designed that  uses HSC hinges at each hinge point. A diagram of the frame and test apparatus is shown in Figure 8. The  deployable frame is made up of low Coefficient of Thermal Expansion (CTE) carbon fiber tubes connected  to HSC hinges. The hinges used in the frame are smaller than those presented previously in this paper.  The hinges have a 12.7-cm free-edge length and are made from tapes with 5.04-cm-diameter curvature  and 2.54-cm flat width. The material used in these hinges are made with the -6 layup defined in Table 2.  The hinges are made with aluminum end fittings and are bonded to the low CTE tubes with Hysol EA 9394  epoxy. At these bond joints, the composite tapes and the low CTE tubes overlap by approximately 12 mm,  so there is no portion of the frame expanse that does not contain composite materials. This is important  because if a portion of the frame expanse was aluminum it would have much higher thermal expansion.  The frame is bonded to the metrology frame through Invar joints. The mounting frame is fixed to the optical  bench in the test chamber using thermal expansion flexures so that shape change in the optical bench is  not transferred to the metrology frame. The final frame test article and metrology frame are shown in the  stowed and deployed configurations in Figure 9.  Figure 8. HSC frame deployment precision test setup  NASA/CP—2018-219887 426
Figure 9. HSC frame in test chamber deployed (left) and stowed (right)  The frame deployment precision testing used the same sensors, test chamber and test sequence as  described previously for the hinge. The initial round of deployment precision testing was performed in the  orientation shown in Figure 9, with gravity pulling the tip of the frame down in the positive piston direction.  The frame was held in a fixture when stowed, also shown in Figure 9, with the goal of having a repeatable  stowed position for each test.   HSC Frame Deployment Precision Results  The deployment precision data obtained for the frame is shown in Figure 10. Based on these results the  piston deployment precision is 2.3 µm, the pitch and roll precision is less than 6.3 µrad and axial precision  is 0.24 µm. Two potential sources of unintended experimental error can be seen in the data from   Figure 10. First, the piston creep-recovery data showed continuous drift over time that can most likely be  attributed to gravity. Because the test-to-test deployment precision was determined from the final position  before the next stow in the test sequence and the duration of recovery was varied from 15 hours to 180  hours, gravity shift over time potentially increased the deployment precision error during these tests.  Second, a significant bias in piston position can be seen after stow 7, which is the creep recovery period  between approximately 240 hours and 340 hours in Figure 10. A relatively large bias in piston position like  this was not present in any of the hinge data. One potential source of the bias may be a variation in the  stowed shape in the eighth stow. When the frame is folded, it is fairly compliant and has many possible  stable configurations. It could be that the stowed frame was in a slightly different stowed configuration  during this stow periods which led to a different stress-relaxation state than other stows. If this test is  removed from the data, the deployment precision error decreases from 2.3 µm to 1.7 µm. In order to  address these potential unintended error sources, a new round of testing is being performed with the frame  oriented vertically, as shown in Figure 9. The results from these tests will be in a future publication.  NASA/CP—2018-219887 427
Figure 10. HSC frame creep recovery data versus calendar time (blue lines are recovery, red circles are  final position before next stow)  Conclusions  This work has demonstrated the level of deployment precision achievable for several HSC hinges and a  frame built from multiple HSC hinges. The deployment precision for the best performing hinges, the frame  and several other precision deployable structures from the literature are presented in Table 5. As can be  seen from this table, the current HSC hinge and frame measurements show a significant improvement over  the JWST DOTA and other non-HSC deployable structures and are the first to include both roll and pitch  precision. The level of precision demonstrated by the structures in this paper are still not sufficient for final  alignment for visible-wavelength optical systems. Nonetheless, metrology and position actuation to correct  for sub-micrometer position error can be much simpler that what is needed for less precise systems, such  as JWST. To support the development of future HSC precision deployable telescopes, MITLL is currently  developing actuation and metrology methods that take advantage of the precision enabled by HSC hinges.  Table 5. Previous optical deployable structure deployment precision measurements  Test Article Piston (µm) Axial (µm) Roll (µrad) Pitch (µrad)  JWST DOTA1020 30 - 44 IFH Specimen 2120.85 0.31 - - EDSM112.97 2.97 - - Lidar Prototype118.59 0.81 - - MITLL HSC Hinge 6-6 0.39 0.25 3.76 2.36  MITLL HSC Frame 2.27 0.24 4.01 6.28  NASA/CP—2018-219887 428
The Synchronization Mechanism for Solar Array with a Three-Stage Deployment  Zehong Yan*, Hongjun Shang*†, Guowei Zhao**†, Qiaolong Yang*, Jingya Ma*and Jianjun Wang* Abstract  Cable-and-pulley system (also called Closed Cable Loop, CCL) is a common type of synchronization  mechanism to coordinate the motion of a multiple panel solar array. Typically, in applications the  transmission ratio of a cable-and-pulley system is constant during the deployment. For other applications,  the synchronization characteristic can be changed by varying the configuration of the pulley. The solar  array of GF-3, a satellite operating on dawn-dust orbit, required a particular kind of two-stage deployment:  during the first stage the four-panel stack rotates 90° around the root hinge; during the second half the  root hinge rotates from 90° to 180° and each panel hinge rotates from 0° to 180° synchronously. The  unique feature of the GF-3 solar array synchronization mechanism is that the root pulley consists of two  half pulleys: a fixed pulley and a swing pulley. The design of such a root pulley changes the ratio between  the root hinge and panel hinge. Although the deployment of the solar array contains three stages actually,  the trajectory approximately meets the requirement. The detailed working principles and considerations of  the synchronization mechanism are provided. The synchronization mechanism has been proven both in  ground test and flight.  Introduction  GF-3 is a satellite operating on dawn-dust orbit, equipped with two deployable solar array wings[1]. There  is a particular layout of the solar array, characterized by the axis of rotation of the solar array drive  mechanism parallel (rather than the more common normal) to the mounting surface of the solar array on  the satellite. The root hinge must deploy 180 degrees, but more importantly, the trajectories of all panels  have to be confined to a limited area that requires a two-stage deployment movement of the panels as  shown in Figure 1.  Figure 1. The two-stage deployment  There are several ways to achieve the trajectory, such as adding a special latch at 90°stroke of the root  hinge. However, varying the synchronization characteristic is a better choice for this application due to the  key advantages, such as simplicity, reliability, light weight, and more importantly, continuous controllable  deployment and ability to reverse at any position. On the basis of the traditional cable-pulley system, the  trajectory can be changed by varying the roll-on and roll-off rules of the root cables, and the key is how to  *Beijing Institute of Spacecraft System Engineering, Beijing, China **Beihang University, Beijing, China  † denotes corresponding authors  Proceedings of the 44th Aerospace Mechanisms Symposium, NASA Glenn Research Center, May 16-18, 2018 NASA/CP—2018-219887 431
derive the profile of the root pulley. Yannick Baudasse presented a new synchronization concept of solar  array deployment used in Cosmo SkyMed [2]. The root pulley is changed into two parts: an unwinder arm  and a winding with special contour. At the first 90° of deployment, root cables neither roll on nor roll off  and just overpass the axis of rotation. After 90°, the arm begins to swing and the cable which it is  attached to rolls off. Then, another root cable begins to wind on the winding, which means roll on. So,  there is a nonlinear synchronization ratio between root hinge and panel hinge. Frans Doejaaren  presented another solution used in Sentinel 1 [3]. The root pulley consists of two coupled pulleys that are  offset when installed and use different radii. This adapted design changes the roll on and roll off  characteristics of the cables such that the wing deploys approximately as required. The above two  methods are simple enough, but resulted in the actual trajectory with more than acceptable deviation from  the ideal one.  In the development of the GF-3 solar array, we found a synchronization method that could reduce the  deviation remarkably. The method is an improvement on Yannick`s design. The improved deployment of  the solar array contains three stages. During the first stage, the four-panel stack rotates about 70°driven  by the root hinge. During the second stage, the root hinge rotates from 70° to 110°, and each panel hinge  rotates from 0° to 40° following a changing ratio (from null to 2:1) with respect to the root hinge. During  the third stage, the root hinge rotates from 110° to 180°, and each panel hinge rotates from 40° to 180°  following a 2:1 ratio. Although the synchronization ratio during the second stage is nonlinear, the process  is transient. Therefore, the actual trajectory is much closer to the ideal trajectory.  The unique feature of the GF-3 solar array synchronization mechanism is that the root pulley consists of  two half pulleys with same radii: a fixed pulley and a swing pulley, which is shown in Figure 2. In the  folded state, the center of the swinging pulley is away from the center of the fixed pulley. During the first  deployment stage, the root cables show the same performance as Yannick`s. During the second stage,  the swinging pulley gradually approaches the fixed pulley. The cable attached to swinging pulley rolls off.  At the same time, the other root cable rolls on the fixed pulley. At the end of the second stage, the center  of the swinging pulley coincides completely with the center of the fixed pulley to form a complete root  pulley. During the third stage, the synchronization characteristic is identical to the traditional cable-pulley  system.  Figure 2. The root pulley of GF-3 solar array  NASA/CP—2018-219887 432
Description of Synchronization Method  Overview  According to the requirement for solar array deployment, the angular velocity ratio between the panel  hinge and root hinge is 0:1 during the first stage, and 2:1 during the second stage. In a cable-and-pulley  system, only with zero radii of root pulley could the ratio be 0:1, which means that the cables should be  placed across the axis of the root hinge. Nevertheless, due to the ratio 2:1, the radii of root pulley should  be twice than the panel pulley in phase 2. That is to say, there is a “snap action” of the root pulley  between phase 1 and phase 2, which is the main design challenge of synchronization mechanism.  It is hard to achieve that the radii of the root pulley rises suddenly from 0 to twice of the panel pulley.  Meanwhile, the angular velocity of the panel hinge will increase immediately from 0 to twice of the root  hinge even if the radii of root pulley can be changed successfully. This would cause a shock that is  detrimental to deployment. The more important aspect of the requirement is that the trajectories of the  mass centers of all panels have to be confined to a limited area. Therefore, it is a significant design  challenge to add a transition between phase 1 and phase 2. The ratio needs to change smoothly from 0:1  to 2:1 with no jerk, and the trajectories must meet the requirement.  In a cable-and-pulley system, the rotation of the hinges is coordinated by the cables rolling on and rolling  off the pulleys. If the cable rolls on the pulley during rotation, it must roll off on the other side due to the  constant length of cable. If there is a reference point located on an intermediate position of the cable, the  distance between the point and pulley decreases when rolling on, and increases when rolling off.  Therefore, the mechanism that can alter the distance will perform the same function as the pulley. In  GF-3, the root pulley is substituted by two half pulleys: a fixed pulley and a swing pulley, with the result  that the ratio can change smoothly from 0:1 to 2:1.  Fixed Pulley and Cable  The schematic of the fixed pulley and cable during the deployment is shown in Figure 3. The fixed  position of the cable passes through the spin axis of the root hinge. In phase 1, the cable rotates around  the spin axis, but the distance between the reference point and fixed pulley is constant during the  deployment. Therefore, the cable need not roll off the panel pulley on the other side. In phase 2, the cable  rolls on the fixed pulley gradually during the rotation, and the distance between the reference point and  pulley decreases at the same time. At the end of phase 2, the cable makes tangential contact with the  circular contour of fixed pulley. The ratio between panel hinge and root hinge becomes 2:1 for the radii of  the fixed pulley is twice of the panel pulley. In phase 3, the cable rolls on the fixed pulley continually, and  the process is consistent with a traditional cable-pulley system.  Figure 3. The schematic of fixed pulley and cable  NASA/CP—2018-219887 433
Swing Pulley and Cable  The schematic of the swing pulley and cable during the deployment is shown in Figure 4. The cable winds  on the swing pulley at the stowed configuration. In phase 1, the swing pulley tends to rest on the right  stop by the tension of cable. The cable rotates around the axis of root hinge, where it contacts with the  swing pulley. The distance between the reference point and swing pulley is constant during the  deployment. At the end of phase 1, the extended line of cable between guide pulley and swing pulley just  crosses the axis of the swing pulley. In phase 2, the swing pulley rotates counterclockwise and is apart  from the right stop by the tension of cable. The distance between the reference point and swing pulley  increases at the same time. At the end of phase 2, the swing pulley rests on the left stop, and the cable  just makes tangential contact with the circular contour of swing pulley. The center of swing pulley’s  circular contour overlaps the axis of the root hinge at the same time. In phase 3, the cable rolls off the  swing pulley continually, and the process is in accordance with a traditional cable-pulley system.  Figure 4. The schematic of swing pulley and cable  Deployment of Solar Array  Both the cable-fixed-pulley system and cable-swing-pulley system have a three-stage synchro process,  which is the underlying premise for the deployment function. Furthermore, although they are installed  separately, they must be kept consistent during deployment. The deployment of the solar array with two  panels coordinated by such a cable-pulley system is shown in Figure 5. It illustrates an approximate  synchronization principle. In phase 1, the outer panel keeps stacked with the inner panel because both  cable 1 and cable 2 cross the axis of root hinge during deployment. In phase 2, the reference point 1  moves away from the panel pulley along with the cable 1 rolling on the fixed pulley gradually. Meanwhile,  the reference point 2 approaches the panel pulley due to the swing pulley rotating off center of the axis of  root hinge. The outer panel deploys against the inner panel. At the end of phase 2, the center of the  swinging pulley coincides completely to the center of the fixed pulley to form a complete root pulley, and  both of the cables make tangential contact with the root pulley. In phase 3, all panels deployed  synchronically.   Phase 2 is the key process of the deployment. There are several problems that must be solved including:   When should it start and finish?  In such a cable-pulley system, the length of the cables rolling on the fixed pulley and rolling off the swing pulley must match with each other well. However, there is an obvious structural difference between the fixed pulley and swing pulley. How can they adapt to each other? NASA/CP—2018-219887 434
Figure 5. Deployment of solar array  Design and Analysis  Mathematical Model  Although the synchronization mechanism is not significantly bulkier than the standard one, the motion is  much more complex. The baseline concept must be optimized by the mathematical model. The length of  the cables is the only connection between the fixed pulley and swing pulley. Therefore, the mathematical  models of the fixed pulley and swing pulley should be established respectively.  The mathematical model of the fixed pulley in phase 2 is shown in Figure 6. The displacement of the  reference point on the cable is equal to the length variation of the cable between the guide pulley and the  center of fixed pulley where the cable is fixed. The length variation is an important parameter that is  derived from the following set of equations.  ∆= ++−+ (1)  NASA/CP—2018-219887 435