ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
Figure 10 Harness Testing Apparatus and Results  A typical data plot on the right side of Figure 10 shows the reduced data, accounting for the parasitic effects  of the apparatus and the resulting resistive torque of the harness over the range of motion. These results  allowed engineering to determine the number springs that were required to meet the torque margin  requirements. Once the springs were installed onto the mechanism, it was again attached to the apparatus  used to characterize the drag torque of the hinge to measure the motion of the hinge and to directly measure  the output torque of the hinge and validate the torque margin analysis.  Figure 11 shows a typical data plot and the slope of the data shows how the constant-torque spring has  some variation over the range of motion that the hinge rotates through. The friction in the system due to the  applied load of the springs is the difference between the curves divided by two, and the output torque of  the spring is the average value of the curves. The minimum output torque is considered the available torque  to the system for determining the torque margin and the minimum value was found to always be at the end  of travel.  Figure 11. Typical Torque Data  A similar approach was used to characterize the inter-panel hinges and to determine the number of springs  necessary to meet torque margin requirements. Running the hinges through this level of characterization  was a significant improvement over the approach taken during the S-NPP effort and provided a high level  of confidence in how well the hinges would perform at the solar array subsystem level.  Solar Array Subassembly Testing  To be acceptable for delivery to observatory integration, the solar array subsystem had to pass a suite of  tests in the sequence shown in Figure 12 (which includes procedure numbers for reference only).  NASA/CP—2018-219887 397
Figure 12. JPSS-1 Test Flow  Manual deployments were agreed upon at this level of testing since powered deployments were performed  at the observatory level pre/post vibration/acoustics testing. Those powered deployments demonstrated  the end-to-end capability of the system to power the release mechanism (TiNi actuator) and break the  retention bolt.  Deployment testing was a challenge on S-NPP and formally recognized as a risk during the JPSS-1  program. There were no plans to change the approach in deployments from how S-NPP had done them; it  was agreed that there were practical limitations with the air bearing gravity-negation system.  Due to test fixture limitations, during the S-NPP subsystem testing, it was not possible to reach the correct  test levels during vibration testing; this testing had to be deferred to the observatory level. For JPSS-1, a  thorough review of the MGSE led to the decision to create a handling fixture that the array could stay  integrated to for deployment, vibration, and thermal testing. This new fixture allowed for easier handling  and transportation of the entire array and is shown in Figure 13.  Figure 13. Solar Array Assembly Manual Deployment Test – Stowed Position  The setup for the alignment of the hinge axis is shown in Figure 14. The deployment floor was aligned to  gravity and it was possible to align the main deployment hinge axis to the deployment floor. This was done  by placing two laser tracker nests onto the panel and rotating it 90 degrees. The resulting arc measurements  allowed the hinge axis orientation to be determined. It was not possible to account for the orientation of the  inter-panel hinges to the deployment floor.  NASA/CP—2018-219887 398
Figure 14. Setup for Alignment of MDH Hinge Axis  During preparation for JPSS-1 deployments, a problem that arose was that the S-NPP polyolefin membrane  was wavy and would not lay flat. This was not acceptable and the material was discarded in favor of an  approach that NASA had developed successfully during the GPM program. That approach consisted of two  layers, a layer of Poron ™ foam against the deployment floor and a layer of Melinix on top of that. Ball  worked with NASA to understand the implementation of this system and it was successfully installed onto  the deployment floor. Prior to the first manual deployment run-for-the-record some engineering tests were  performed to gain confidence that the deployment would be successful. This was needed as the air bearing  offload system was not instrumented to show the amount of load that each air bearing was carrying. The  engineering deployments allowed evaluation of the air bearings to determine if they were adjusted correctly  and operated over the full range of motion. One thing that was observed was that if an air bearing was  overloaded it would “plow” through the Poron ™ support and it was easy to note which bearings were not  adjusted correctly.  The air bearing support was supplied with air through a single hose to each panel and can be seen in  Figure 15. It became apparent that the manual handling of these hoses could potentially impact how the  array deployed and care had to be taken in how the hoses were managed. If they were twisted the wrong  way, then a retarding torque could be applied at the hinge line and impact deployment time or the sequence  in which the panels locked out. The first manual deployment took 4 minutes and 30 seconds to complete  without requiring intervention.  NASA/CP—2018-219887 399
Figure 15. Deployed Array on Air Bearing Support Equipment  After the deployment was completed, each hinge line was inspect to verify that the pawls were engaging  with the cams correctly. It can be seen in the left-side ovals in Figure 16 that the lower pawl of the MDH  engages into the cam further than the upper pawl. This is a result of minor variations in the deployment  floor height and the fact that the air bearings cannot be controlled to equally off-load the panel.  Figure 16. MDH Pawl Engagement Issue  Once the array was in the deployed position it was possible to perform electrical and Large Area Pulsed  Solar Simulator (LAPSS) testing. Once the panels were secured in a flight stowed configuration to the  MGSE the entire barrel assembly that the array was attached to was transported to perform environmental  NASA/CP—2018-219887 400
testing. Thermal vacuum testing was performed on the entire array for four cycles, to make up for the four  cycles missed at the observatory level when the wing is removed from the spacecraft for thermal balance.  This test is limited in the ranges to which the array subsystem can be exercised. It was not possible to cycle  to the hot operational temperature of the array as it would exceed the transition temperature of the shape  memory alloy in the TiNi actuators and the limits of the wet lubrication in the eddy current dampers.  The solar array test set up was vibration tested with a solar array mass simulator prior to integration of the  flight solar array to the MGSE to validate the modes of the MGSE prior to flight vibration testing. The array  was exposed to sine sweep and sine burst to qualify all of the hardware for quasi-static loading and the  low-frequency sine transient or sustained sine environments that were anticipated during launch. The test  setup is shown in Figure 17. Random vibe was also performed as risk reduction; the array was to be  qualified at observatory acoustic testing.  After environmental testing the array was transported back to the deployment floor and configured for  manual deployment. With minimal adjustment done to the array and no verification of hinge line orientation  the array was deployed and took 9 minutes and 30 seconds to complete. Each panel successfully locked  out and completed its range of motion but this significant difference in time could have been an indication  of damage to the hardware due to environmental testing. After adjusting the MGSE to orient the handling  fixture in a similar manner as the pre-environmental tests, it deployed in 5 minutes. This was much more in  line with the previous test and, after completing electrical and LAPSS testing, the array was delivered for  observatory integration. Once mounted to the satellite bus, the array passed all of its electrical tests.  Figure 17. Solar Array Subsystem Vibration Testing  Satellite-Level Testing  At the Observatory level, the array was to see more vibration testing, after which first motion testing was  required to demonstrate it was going to operate properly after launch. Then, while JPSS-1 was in TVAC  testing, the solar array was removed and mounted back to the MGSE and a final full deployment test was  completed. This testing was complicated by the fact that dampers were removed after the array was in the  partially deployed condition, after a first motion release, also known as a “pop and catch,” so that the array  could be safely stowed to the bus for removal. It was decided to start the deployment test where the pop  NASA/CP—2018-219887 401
and catch test had ended to test the dampers thru their full range of motion without adjusting where on the  gear train it had stopped. After the array had been attached to the air bearing hardware, the array was  deployed. This deployment took longer than anticipated, 10 minutes and 27 seconds, without manual  intervention. Without adjusting the test set up, the array was stowed and allowed to deploy from its stowed  position again, and that took 8 minutes and 10 seconds, without manual intervention. Since this was still  out of family with the subsystem results, an investigation was started to understand the difference. It was  determined that since the air bearings were used and adjusted for the observatory-level test, and then  utilized untouched for the subsystem-level test, they were now improperly adjusted for the subsystem  testing. After adjusting the air bearing set up, by manually walking out the array with eddy current dampers  removed, the deployment was repeated and completed in 4 minutes and 16 seconds, which matched the  results from subsystem-level testing. With this success, all that remained before shipment and launch was  to remount the array and complete a final pop and catch.  Summary and Lessons Learned  The design and development of the S-NPP array resulted in many lessons learned during the first  implementation of this style of array by Ball Aerospace, in the following areas:  •Number of leaf springs used within each drive stack to limit inter-spring friction. •Inspection of parts and verification of part processing. •Cam profile of hinge surfaces to provide constant resistive torque. •Inadequate resistive torque characterization. •Inadequate output torque characterization. •Agreed upon requirements for resistive torque characterization. •Design of vibe fixtures to achieve required test spectrum. •Deployment fixturing design and verification. The JPSS-1 solar array design updates—changing the profile of IPH cam, updates to the constant-torque  spring and the spring mounting, and removing potential harness hang-ups—and the follow-on tests were  very successful, but some issues along the way made it necessary to repeat tests. Among these were:  •Deployment floor material changes from polyolefin membrane to Poron™ and Melinix. •Re peatability of MGSE alignment. •Maintaining configurations of air bearing MGSE between different test set up usage. •Difficulty repeating air bearing alignments and determining proper offloading. The issues associated with the air bearings and determining whether the offloading is being accomplished  correctly has driven the undertaking of a new research and development effort to address these troubles  found during S-NPP and JPSS-1.  Acknowledgements  The authors wish to acknowledge Mr. Joseph Pellicciotti for his crucial support with the S-NPP solar array  flight qualification, and for the portions of his presentations which made it into this paper. The authors  especially want to thank the engineers, technicians and support personnel of Ball Aerospace who turned  the designs into hardware that made these two missions possible.  NASA/CP—2018-219887 402
Spacecraft Common Deployable Boom Hinge Deploy and Latching Mechanisms  Paul Lytal* and Marcel Renson**  Abstract  JPL is developing deployable radio frequency (RF) reflector booms for spaceflight usage on two different  Earth orbiter projects scheduled to launch in the early 2020’s. The static alignment and thermal stability  requirements for these RF reflector booms are challenging. A common mechanism design has been  developed for the deployment and latching of boom hinges that avoids significant parasitic loads on the  alignment critical structures after completion of mechanism function. The mechanism design includes a  high-aspect-ratio torsion spring with viscous damper for hinge closure as well as an actuator-driven flexured  hook and roller latch for joint preloading. This paper describes noteworthy mechanism design details, test  results, challenges, and lessons learned during the development effort.  Introduction  The RF reflector boom deployment sequence is shown in Figure 1 for each project. On a given boom, only  one hinge is deployed and then latched at a time. After completion of hinge closure and latching for all  boom hinges, the passive RF reflector deploys.   Figure 1. RF Reflector Boom Deployment Sequences  The boom hinge deploy and latching mechanism design includes a spring/damper mechanism for hinge  closure and an actuator-driven latching mechanism for hinge joint preloading (Figure 2). The spring/damper  mechanism automatically proceeds with hinge closure immediately following mast launch restraint  separation. Hinge latching does not commence until successful confirmation of hinge closure to within the  latching capture range from redundant hinge closure sensors.  *Jet Propulsion Laboratory, California Institute of Technology, Pasadena, CA ** D.E.B. Manufacturing Inc., Lakewood, NJ Proceedings of the 44th Aerospace Mechanisms Symposium, NASA Glenn Research Center, May 16-18, 2018 NASA/CP—2018-219887 403
Figure 2. Hinge Deploy & Latching Mechanisms (NISAR Hinge Shown)  Project Background  This common boom hinge deploy and latching mechanism design is being developed for two different Earthorbiter projects: the Surface Water Ocean Topography (SWOT) Project and the NASA-ISRO Synthetic  Aperture Radar (NISAR) Project. The SWOT Project is an international collaborative effort between NASA  and the French Government Space Agency. The earth science objectives of the SWOT Project include  high-definition temporal and spatial mapping of all fresh and salt water bodies around the globe for a  minimum 42-month time period. NISAR is an international collaborative effort between NASA and the India  Space Research Organization (ISRO). The earth science objectives of the NISAR Project include land and  ice mass observation to improve natural disaster prediction, deforestation modeling, and polar ice cap  reduction modeling, amongst other objectives. Precision alignment and stability of the RF reflectors to the  RF feeds is crucial to mission success of each project. Figure 3. NISAR (Left) and SWOT (Right) Earth Observation Spacecraft  Common Boom Deploy & Latching Mechanism Development  The SWOT and NISAR Projects elected to pursue a common deployable boom design to the maximum  extent practical to reduce overall development cost and schedule. While detailed mast geometry  necessarily differs due to mission architecture and configuration differences, the basic design and  construction of the booms is common between the two projects.   NASA/CP—2018-219887 404
The deploy and latching mechanisms are highly similar between the two projects, only differing as required  by the structure that the mechanisms mount to and differences in flight environments. Noteworthy  differences between the two projects are highlighted in the sections that follow as relevant to the  mechanisms described in this document.  Driving Design Considerations  In addition to typical spaceflight mechanism design constraints (limited mass and volume, launch  environments, on-orbit environments, ground testing considerations), the avoidance of significant and  variable loading from mechanisms hardware onto the alignment critical structure was a primary design  driver. The approach taken was to mechanically disconnect all possible mechanisms hardware from the  mast after completion of mechanism function as show in Figure 4. In addition, the mechanical attachment  of all mechanisms hardware to the mast uses flexures or non-preloaded pin/slot interfaces. Specific  implementation of the mechanism mechanical decoupling is shown in Figure 4.   Figure 4. Mechanism Mechanical Disconnection from Alignment Critical Mast Structure  Another critical design driver for deploy and latching mechanisms hardware was conductive thermal  isolation of mechanisms hardware from the mast structure. This enables the control of mechanism  hardware to tighter temperature ranges than the parent structure as necessary. Isolation was accomplished  using small contact areas and low thermal conductivity Ultem 2300 isolators. An example of this  methodology is shown in Figure 5.   NASA/CP—2018-219887 405
Figure 5. Example of SWOT and NISAR Mechanism Thermal Isolation Design  Hinge Deployment Mechanism  The hinge deployment mechanism is designed to close boom hinges from the launch position through the  deployed/closed hinge configuration. A separate analogous hinge deployment mechanism is used at each  hinge. The hinge deployment mechanism was designed to overcome worst-case torque opposing  deployment associated with hinge-crossing wire harness, coulomb friction sources, and ball bearing rolling  friction losses over the full range of thermal environments that could take place in flight. Challenges  encountered during prototype hinge testing (Figure 6) in worst case thermal environments are described in  the sections that follow.  Figure 6. NISAR Prototype Thermal Test Hinge  A spring/damper mechanism was developed to perform hinge closure. Several different options were  considered for the boom hinge closure mechanism. A serpentine wire rope spooler mechanism (as used  on the SMAP Project RF boom) was considered but rejected primarily due to the inability to test  mechanisms in a flight-like manner prior to full mast assembly. Various actuator-driven linkages were  considered and discarded due to higher estimated mass than the spring/damper mechanism option. Direct  actuator driven hinge closure was considered as well, but was also deemed more massive and a higher  risk development than the spring/damper option.  The following sections provide information regarding unique design and testing challenges with the  spring/damper mechanism as well as the deployable hinge crossing harness.  NASA/CP—2018-219887 406
Spring Design Overview  A pair of torsion springs is used at each hinge, each capable of independently transmitting torque to deploy  the hinge for robustness. Each spring is supported by an aluminum outer mandrel which in turn rotates on  glass-filled Teflon bushings on an inner mandrel (Figure 7).  Figure 7. Spring Location in Hinge (Left) and Independent Spring Mandrels (Right)  A custom 17-7 CH900 torsion spring solution was deemed necessary given the limited volume and high  output torque needed for the Hinge Deployment mechanism, particularly for the smaller volume available  in the NISAR hinge configuration (116-mm [4.567-in] length and 43-mm [1.693-in] diameter for spring  assembly). Elgiloy was briefly considered and dismissed due to long lead time as well as spring vendor  cautionary guidance that Elgiloy spring development programs frequently encounter challenges, often with  little to no improvement in measured strength performance over 17-7 CH900. The projects elected to stick  with the 17-7 CH900 material which the spring vendor had the most experience working with.   A high-aspect-ratio rectangular cross section (3.8 to 1) is being used for the spring to increase the spring  wire moment of inertia per unit spring body length. This resulted in lower bending stress and higher spring  constant than a round wire spring with the same wire width and number of spring turns.   There are six types of hinges for the SWOT and NISAR booms with different stow angles. It is worth noting  that the relaxed spring arm angle of the springs for each hinge were selected to ensure maximum torque  across the range of motion of the hinge without violating NASA-STD-5017A advisory that Mission Critical  Springs maintain positive margins above a 1.5 factor of safety to yield at the maximum deflection (fully  stowed) configuration [1].  Spring Arm Twist  In early testing with spring arms contacting round pins with no other support, undesirable arm longitudinal  twisting was observed at spring windup angles greater than 180 degrees, resulting in significant reduction  in spring stiffness and higher-than-intended stress in the spring arms. Destructive characterization testing  at room temperature demonstrated the onset of yielding at approximately 445 degrees as opposed to  approximately 730 degrees predicted by analysis with no spring arm twist. Guide vanes were added to the  spring mandrel design to minimize this arm twist phenomenon. Follow-on destructive spring testing with the  guide vanes demonstrated onset of yielding at approximately 740 degrees, which more closely matches  analytical predictions (Figure 8).  NASA/CP—2018-219887 407
Figure 8. Destructive Spring Characterization Test Results, Spring Arms Supported  Spring Assembly Thermal Testing  Developmental testing was conducted to evaluate the impact of spring wet lubrication on spring  performance across temperature. Torsion springs were grease plated with a thin film of Braycote 601EF  and tested in the same configuration as identical clean, unlubricated springs. Test results indicated that  lubricated spring performance was nearly identical to unlubricated spring performance at and above room  temperature. Unlubricated spring torque performance was significantly better than lubricated spring  performance in the worst case cold environment (-100 °C) as shown in Figure 9. The average improvement  in spring output torque was up to 1.7 N•m [15 in•lbf] at maximum spring deflection in usage. This is expected  to be the result of wet lubricant thickening at cold temperature. The flight torsion springs were unlubricated  for both projects.  Figure 9. Prototype Spring Assembly Thermal Test, Lubricated versus Not Lubricated  Prototype spring assembly testing (consisting of a spring, mandrel, hinge pin, and bushings) was also  conducted at the hot and cold extremes of the expected flight operational temperature range. Test results  01020304050607080 0 50 100 150 200 250 300 350 400Torque (in•lbf) Spring Windup Angle (deg)Spring/Hinge Pin Assembly Output Torque vs Spring  Windup Angle (With/Without Braycote 601EF) -100degC, Lubricated -100degC, Not Lubricated NASA/CP—2018-219887 408