ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
into a canister, but a CubeSat is a square. As such, there were four corners not being utilized. Four lead  screws could be located in the corners, driven by a motor (Figure 6). The main challenge: synchronization  of all four lead screws to ensure a steady deployment.   To synchronize, each lead screw was attached to a “planet” gear. Each “planet” gear interfaced with the  “sun” gear in the center of the antenna which kept all four lead screws in sync. While this is not a traditional  planetary gearbox, it is easiest to reference the design in these terms. Because of the waveguide in the  center of the antenna, the sun gear was mounted to a large diameter thin section bearing. A pinion attached  to the motor would drive one of the planet gears, in turn driving the sun gear, and then the other three planet  gears. While initially there was discussion of using two motors for redundancy in the system, if one of two  motors failed, extra torque would be required to back drive the non-functioning motor. Therefore, the  antenna was maintained with a single motor.  The lead screws attached to a brass threaded feature on the hub which would drive the antenna up and  down. A further advantage of the motorized system was realized when investigating adding launch locks to  the design. The lead screws could also be used to hold the hub down, in addition to deploying it. Therefore,  one system provided the launch lock and deployment capabilities.   Figure 6 . Motorized Deploymen t System Components  The prototype  antenna  was  retrofitted  with  the motorized  system. The  same mesh,  horn,  and secondary  reflector were used. This meant only the canister, base, and some components on the hub needed to be  replaced, along with the additional motorized drive system. This retrofitted designed was referred to as the  engineering model.   Engineering Model Testing  The antenna was then deployed with th e motorized deploymen t system , where all systems behaved  nominally. The antenna was taken to the RF range for testing after deploymen t to check the deployed  shape. It was found the motorized system could apply more preload to the system in the deployed state,  which resulted i n a better surface accuracy , and thus a gain of 42.7 dBi, onc e again exceeding  requirements.   © 2018 All Rights Reserved. Government sponsorship acknowledged.  NASA/CP—2018-219887 368
Figure 8. Flight KaPDA Installed with the Tyvak Spacecraft  Conclusions / Lessons Learned  KaPDA  has  provided  a number  of great  lessons  learned  with  regards  to mechanism  design,  given  it  combines motors, gear trains, lead screws, springs, and at one point pneumatics. There are a number of  separate actuating features, each of which have generated key lessons learned, detailed below.   Add as much Depth in a Deployable as Possible  It is common knowledge increasing the area moment of inertia improves cross section performance. While  this seems like a minor change, i t was a key instrumental factor in achieving the surface accuracy and  deployed stiffness. While i t is often challenging to add deep sections to a deployable because of stowed  size constraints, the KaPDA desig n achieved additional depth by placing i t where the bending momen t was  the highes t and reducing it where the bending moment was lower. The additional depth also made the  hinges less sensitive t o manufacturin g tolerances by allowin g hard stops to be placed far away from the  hinge pins.  Use Fixtures to Prevent Tolerance Stack up Issues  A key desig n decision i n the assembly process was to use fixtures to assemble components, and bond  them in place thereby achieving very accurate and consistent ribs . This effectively remove d the effects of  tolerance stack-up from assembly, suc h tha t the fixture was the key driver in achievin g the appropriate  geometry. This allowed both versions of th e antenna to be extremely precise and provide high RF  performance.   When to use Additive Manufacturing  Additive manufacturing was found to have advantages for buildin g small , complex components . This  allowed lower cost approaches than traditional machining for a par t that had better function . However , most  parts used in the antenna were still bes t implemented through traditional machining, and additive should  not be considered a replacemen t for traditional methods . Rather, it becomes a n alternate method in the  designer’s toolbox.   Use Deploymen t Methods Provide Control Authority  The original deployment method , using a gas powered pneumatic approach provided almos t no control  authority, other than deciding when t o start the deployment . Moving to a motorized system allowed specific  © 2018 All Rights Reserved. Government sponsorship acknowledged.  NASA/CP—2018-219887 371
© 2018 All Rights Reserved. Government sponsorship acknowledged. [3] J. F. Sauder, N. Chahat, R. Hodges, E. Peral, Y. Rahmat-Samii, and M. Thomson, “Designing, Building, and Testing a Mesh Ka-band Parabolic Deployable Antenna (KaPDA) for CubeSats,” in 54th AIAA Aerospace Sciences Meeting , American Institute of Aeronautics and Astronautics. [4] N. Chahat, R. E. Hodges, J. Sauder, M. Thomson, E. Peral, and Y. Rahmat-Samii, “CubeSat Deployable Ka-Band Mesh Reflector Antenna Development for Earth Science Missions,” IEEE Trans. Antennas Propag. , vol. 64, no. 6, pp. 2083–2093, Jun. 2016. [5] N. Chahat, J. Sauder, M. Thomson, R. Hodges, and Y. Rahmat-Samii, “CubeSat deployable Ka-band reflector antenna for Deep Space missions,” in 2015 IEEE International Symposium on Antennas and Propagation USNC/URSI National Radio Science Meeting , 2015, pp. 2185–2186. [6] M. Aherne, T. Barrett, L. Hoag, E. Teegarden, and R. Ramadas, “Aeneas -- Colony I Meets ThreeAxis Pointing,” AIAAUSU Conf. Small Satell. , Aug. 2011. [7] “UNFURLABLE SPACE REFLECTOR SOLUTIONS.” Harris Corporation, 15-Feb-2016. NASA/CP—2018-219887 373
selection. Table 1 provides a summary and more detail on some of the key requirements that drove and  constrained the mechanism’s design.  Table 1. Mechanisms Key Driving Requirements Summary  Category Requirement Text  CubeSat No space debris shall be created at any point in the mission.  CubeSat Pyrotechnics shall not be permitted.  CubeSat The 3U CubeSat shall be 100.0±0.1 mm wide. (X and Y dimensions) and be 340.5±0.3 mm tall. (Z dimensions)  CubeSat The only CubeSat structure that can contact the deployer are 8.5 mm wide rails and nothing can cross them.  CubeSat Components shall not exceed 10 mm normal to the surface of the 100 mm cube sides.  ALBus ALBus mechanisms are to be designed to increase reliability; e.g. releasing all four deployable SAs simultaneously.  ALBus The ALBus mechanisms are to be designed using SMAs.  ALBus The ALBus mechanisms shall be resettable so the flight hardware can be tested prior to flight and without  disassembly of the CubeSat.  ALBus The ALBus mechanisms shall be designed to structurally retain the deployable SAs during all mission phases prior  to being commanded to release.  R&R Upon command receipt, and only when desired, the R&R shall release the deployable SAs.  R&R The R&R shall be designed to accept and operate by electrical power provided by the ALBus electrical power  system which is 6 volt limited and 3.0 amps maximum.  Hinge Upon command and only when desired, the hinge shall deploy the deployable SAs to the deployed state and  prevent any detrimental damage to the deployable SAs, any other CubeSat structure, or to mission operations.  Hinge The hinge shall allow the deployable SAs to deploy to 90°-135° from the stowed configuration.  Hinge The hinge shall be designed to structurally support the deployable SAs during all mission phases.  Hinge The hinge shall allow two separate paths to conduct power from the deployable SAs to the spacecraft.  Mechanism The ALBus mechanisms are to use NASA-STD-5017 rev A as a guide to design the mechanisms.  Final Design  Figure 1 shows the overall architecture of the ALBus CubeSat which illustrates the mechanisms’ location.  Figure 1. ALBus CubeSat Architecture  Design Summary of the R&R  The R&R final design consists of a two-stage activated mechanism. The first stage is a pin-puller device  driven by an SMA linear actuator. The second stage is a hook and pin design that is released by a  compression spring loaded plate riding on plain bearings. The operation and design of the mechanism is  discussed in the following paragraphs. Figure 2 and Figure 3 illustrate details of the R&R component parts  NASA/CP—2018-219887 377
that go along with the description. Table 2 summarizes the functional performance capability specifications  and requirements. These specifications are driven to have positive margins per design guidance from  NASA-STD-5017 [8]. Figure 2. R&R Component Parts and Design  Figure 3. R&R Mass Estimate and SMA Linear Actuator Details  ALBus’s initial temperature requirements for a safe SA deployment were set to be >100°C, which exceeded  any commercial alloy capability. Therefore, the linear actuator consists of an alloy with an atomic  composition of Ni 19.5Ti50.5Pd25Pt5 resulting in high transition temperatures above 100°C, work output  exceeding 15 J/cm3 [9], and capability to process into small diameter wire. Thus, rods were drawn into a  0.508 mm diameter wire which was trained, cut into segments, and then installed on a custom linear  actuator produced by MIGA Motor Company [10]. A total of five SMA wires are connected to guide rails.  Once heated past the transition temperature using direct current (joule heating), each SMA wire contracts  to pull its associated guide rail. The summation of the five SMA wires yield a cumulative displacement of  7.1 mm travel to pull the pin and release the second stage.  The pin-puller consists of a 17-4 PH H900 stainless steel pin put in a double-shear configuration between  three bushings. The bushings are made of a polyimide plastic that has impregnated graphite.  Once the pin-puller releases the release plate, four compression springs move the plate, unlatching all four  deployable SAs. The compression springs were installed concentrically over linear guide bushings. These  bushings are made of the same polyimide plastic as the pin puller bushings. The guide bushings ride on  guide rods that are made of 17-4 PH H900 stainless steel. The release plate unlatches the arrays using a  hook and pin latch. There are two latches per SA panel for a total of eight hooks. Both the hooks and pins  are made of 17-4 PH H900 stainless steel. A thin film of a MoS 2 lubricant is applied on the hooks and pin  to mitigate any binding concerns from friction, although the joint is lightly loaded and lubrication is probably  not needed.  NASA/CP—2018-219887 378
Design Summary of the Hinge  The final design of the hinge consists of two aluminum hinge knuckles that pivot over a hinge pin, two  superelastic SMAs, and a latch to keep the SA in the deployed state. The operation and design of the  mechanism is discussed in this section. Figure 4 and Figure 5 illustrate details of the hinge component  parts that accompany the description. Table 2 summarizes the functional performance capability  specifications and requirements. These specifications are driven to have positive margins per design  guidance from NASA-STD-5017 [8]. Figure 4. Hinge Component Parts and Design  Figure 5. Hinge Component Electrical Interfaces and Design Details  After the R&R releases the SAs, they are free to rotate and each array is driven open by two preloaded  superelastic SMAs per array. In this design, a Ni-rich Ni 50.7Ti49.3 (atomic %) superelastic alloy was selected  to serve a dual purpose of (i) a spring load to open the arrays and ( ii) a current carrying conductor to transmit  power from the SAs. The superelastic material was rolled into a 0.2-mm-thick sheet with a transition  temperature (i.e., martensite start temperature) below 0°C. At room temperature, the sheets exhibited a  superelastic plateau between 200 and 300 MPa, depending on the heat treatment used. This superelastic  plateau denotes the effective start of the materials’ stress-induced transformation from the stiffer phase  known as austenite to the more compliant phase known as martensite. The superelastic sheets were  machined into a flat-shape profile and then shape set to a specific U shape with a custom jig. After several  NASA/CP—2018-219887 379
configuration as seen in Figure 6, which should be the lowest stress state. This was verified by attempting  to perform 15 wear-in cycles on the flight design, which resulted in more SMA spring failures in early cycles.  The SMA hinge springs needed to be disassembled, redesigned, and replaced. The corrective action  consisted of moving the solder attachment to cover the bottom half of the SMA hinge spring only and away  from the rivet. The SMA was redesigned to remove a narrowed-down region that was no longer needed  since the solder is only being applied at the very bottom edge. This increase in area was estimated to  reduce the stress on the part by more than 20%. Finally, the washer under the rivet, which was still needed  to prevent the rivet from potentially cracking the SMA, was bent upward to allow the SMA to have a  smoother curve onto the PCB. These changes can be seen in Figure 7. New SMAs were then fabricated  with the same processes that were conducted on the original batch such as torque testing each new spring.  The new design was then installed onto the EDU and underwent a 4-times cycle life test (130 cycles) to  gain confidence that the changes would correct the issue, which it passed. The redesign was then  reinstalled on the flight hardware and underwent a 15 cycle wear-in test to ensure good workmanship and  no other issues. Figure 7 shows the new SMA spring design and attachment method.  Figure 7. SMA Hinge Spring Flight Design Change, Attachment Corrective Action, and Repair  Once assembled, similar to the EDU testing, the flight subcomponents underwent functional and  performance testing to ensure they were functioning correctly. Then, the units were integrated into the flight  CubeSat assembly and underwent environmental testing. This included a random vibration test to 10 gRMS  for 1 minute in three axes and a thermal/vacuum bake out. The CubeSat was functional before and after  the vibration test and after the thermal/vacuum test. It was then stored for shipment to the launch service  provider with a target launch date of April 2018. Images of the flight hardware are provided in Figure 8.  Figure 8. ALBus CubeSat R&R and Hinge Mechanisms Flight Hardware  NASA/CP—2018-219887 387
testing. Because of the extreme difficulty of implementing full solar array tests in compliance with the Gold  Rules, most large solar array testing has to be scaled back via a waiver—true of these two arrays as well.  S-NPP Solar Array Development Figure 1 depicts the S-NPP and JPSS-1 satellites with their highly similar solar arrays deployed, showing  their zenith-facing and Earth-facing sides, respectively. Three solar panels are tied together by two interpanel hinge (IPH) lines consisting of a passive and powered hinge and a main deployment hinge (MDH) at  the interface between the inner panel and the rotational drive assembly (RDA), which rotates the entire  array 360 degrees, to track the sun throughout each orbit. Each of the powered hinge lines are driven open  one way by a constant-torque spring and rate controlled with eddy current dampers.  Figure 1. S-NPP and JPSS-1 Satellites (fully deployed)  Hinge Torque Margin Troubles  The solar array for S-NPP was designed to be compliant with the MMA requirements in the beginning of  the program showing large torque ratios, but as the program reached the testing phase and the use of the  GEVS torque margin analysis was followed, torque margins dropped. There were multiple issues that  eventually drove torque margins negative which are explained below. Early in the design phase, the torque  margins were required to be greater than zero using phased factors of safety applied to the sum of known  torques (ΣΤ K) and unknown torques (ΣΤ U) according to the following formula:  	 = 	 (∙Σ+∙Σ)−1	must	be > 0.0 Table 1. GEVS Factors of Safety through program life  Typical known resistive torque are things like inertia, or magnetic coulomb friction and unknown resistive  torques are things like friction within the mechanism or the resistive torque of a harness.  Shown in column 2 of Table 2, the initial revision of the torque margin analysis, in 2003, was determined  using analytical values for resistive torques and used the Critical Design Review (CDR) factors of safety  from Table 1. (Take note that IPH1 is the hinge line closest to the spacecraft body.) Revision A of the  analysis was an update based on initial testing of the hinges and the supporting harness as well as  NASA/CP—2018-219887 390
measured values for damper resistive torques. The harness resistive torque changed from an anticipated  value of 0.06 N-m (0.5 in-lb) to a measured value of 0.36 N-m (3.2 in-lb) for the IPH, and the MDH harness  resistive torque changed from an anticipated value of 0.23 N-m (2 in-lb) to a measured value of 0.17 N-m  (1.5 in-lb). Revision B resulted from a NASA project request to analyze the margin using a mix of factors of  safety. Revision C of the analysis addressed a reduction in the available torque of the system (which was  reduced) and another (and final) change in the factors of safety: the IPH and the MDH margins were both  determined using Acceptance/Qualification values from Table 3. Available torque was measured in testing  in 2004 and 2005 with a torque watch which did not provide an accurate torque versus angle measurement,  and minimum and maximum available torques were not available either.  Table 2. Changes in Margins during S-NPP  Hinge Type Rev - Rev A Rev B Rev C Pre  Spacecraft  Integration Pre-Ship  Torque Margin Pre-Ship  Torque Ratio  IPH1 0.695 0.791 0.361 0.279 0.067 -0.016 1.58  IPH2 0.695 0.791 0.361 0.279 0.067 0.06 1.94  MDH 0.306 0.780 0.439 0.780 0.09 -0.22 1.48  IPH Torque – known  N-m (in-lb)0.73  (6.50) 0.65  (5.76) 0.65  (5.76) 0.65  (5.76) 0.73  (6.42) 4.98  (IPH 1)  5.55 (IPH 2)  IPH Torque –  variable  N-m (in-lb)0.61  (5.41) 1.28  (11.32) 1.28  (11.32) 1.28  (11.32) 1.20  (10.65) 15.25  (IPH 1)  10.50  (IPH 2)  IPH available torque  N-m (in-lb)4.97  (44) 6.33  (56) 6.33  (56) 4.52  (40) 3.73  (33) 31.9  (IPH 1)  31.1 (IPH 2)  MDH Torque –  known N-m (in-lb) 1.39  (12.33) 1.28  (11.37) 1.28  (11.37) 1.28  (11.37) 0.89  (7.87) 7.85  MDH Torque –  variable N-m (in-lb) 0.86  (7.62) 1.48  (13.1) 1.48  (13.1) 1.48  (13.1) 1.88  (16.6) 33.1  MDH Available  torque (in-lb) 6.10  (54) 8.70  (77) 8.70  (77) 8.70  (77) 5.54  (49) 60.5  FS known 1.5 1.5 1.25 1.5 1.5 1.5  FS unknown 3.0 3.0/2.0 3.0 2.0 2.0 2.0  Table 3. Updated Factor of Safety Values  In 2009, after hardware storage, concerns were raised about the variability of output torques in the hinges  and it was determined that better characterization of the hinge lines was required. IPH testing was carried  out to measure the output torque across the full range of motion in both directions and the available torque  dropped from previously measured values thanks to better measurement technique. It also found that the  resistive torque from the pawl was variable due to the cam shape of the hinge. The MDH was also recharacterized and again the minimum torque measured during the range of motion resulted in a significant  drop in the torque margin at spacecraft pre-integration.  NASA/CP—2018-219887 391
GSE consisted of a number of thick honeycomb panels that were attached to each other to provide a surface  that the solar array could be deployed on through the use of air bearings. This configuration presented a  number of challenges as at each panel joint there often arose a discontinuity that resulted in the air bearing  “dropping out” so that the solar array deployment would stop. A number of fixes were attempted to close  out this gap and prevent the array from stopping during deployment. The final solution used a continuous,  smooth, and flexible roofing material that was laid out over the top of the deployment fixture. This roofing  material was a 1.52-mm (0.060-in) thick polyolefin membrane and the resulting smooth table surface is  shown in Figure 4.  Figure 3. First Motion “Pop and Catch” Testing (Overview and GSE Closeup)  Figure 4. Subsystem Deployment Testing with Polyolefin Membrane Topping the Raised   Deployment Floor  NASA/CP—2018-219887 393
Figure 6. MDH in Deployed Orientation  This IPH hinge design shown in Figure 7 also went through similar evolutions as the MDH. The output drum  was increased in size to accommodate the larger constant-torque spring. As with the MDH the take up roller  was increased in diameter to minimize the deflection and accommodate the stronger spring sized for JPSS1. The biggest change for the IPH design was the change in the design of the cam profile. The right side of Figure 7 shows the change from S-NPP, in blue, to JPSS-1 shown in yellow. Figure 7. Powered IPH Shown in Stowed Position and Cam Profile Change (JPSS-1 changes in gold;  dimensions in inches)  Figure 8 shows how the S-NPP cam profile produced a varying drag due to changes in the compression of  the pawl spring across the range of motion. The modification to the cam profile for JPSS-1 achieved a  constant resistive torque value, making it consistent with the non-varying profile for the MDH.  NASA/CP—2018-219887 395
Figure 8. S-NPP Variation in Pawl Drag Torque Over Range of Motion  Hinge Testing  The JPSS-1 MDH and IPH hinge designs were both characterized for sources of drag torque in a much  more consistent and repeatable method than S-NPP. Figure 9 shows the test set up used to characterize  the MDH hinge. This apparatus consists of a drive motor in-line with a torque sensor and the unit under  test. This test drove the hinge in both directions and had the pawls riding on the surface in both directions,  but the apparatus did not go through the full range of motion to prevent the pawls from locking out.  Figure 9 also shows an example data plot for the MDH hinge resistive torque test. In comparison, S-NPP  characterized drag torque manually, with a torque watch which was not as repeatable nor as controllable  over the range of travel.  Figure 9. MDH Drag Torque Testing Apparatus and Results  In parallel, the harness resistive torque was characterized using a similar apparatus, shown in Figure 10.  This apparatus allowed the harness to be articulated over the full 90 degrees in both the deploying and  stowing direction to allow the peak resistive torque to be measured. Measurements are made at ambient,  hot, and cold extremes. First the apparatus itself was cycled over the thermal environment and driven  across the range of motion to characterize the effects of the apparatus and then testing was repeated after  the harness was installed.  NASA/CP—2018-219887 396