ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
Figure 10. Solar Tracking Errors during Ground Testing  On-Orbit Solar Tracking Errors  As encouraging as the ground-testing results were, the true test of pointing performance was only possible  with the TPS operating on-orbit. During the early-orbit commissioning program, all of the TPS functionality  was exercised, which provided an opportune time to record high-rate 200-Hz measurements that are not  available during nominal operations and are able to reveal high-frequency pointing characteristics.  Figure 11 shows the azimuth and elevation solar tracking errors as measured by the HFSS at the native  control execution rate of 200 Hz. The measured pointing performance was slightly improved compared to  the ground-test measurements with 4 arcseconds 1σ in the azimuth axis and 2 arcseconds 1σ in the  elevation axis. Additional high-rate and low-rate pointing measurements from throughout the flight mission  are all consistent with these results. This performance easily satisfies the jitter requirement of  60 arcseconds 1σ and is a dramatic improvement compared to the stepper actuator approach from the  TSIS NPOESS-era in Figure 9. This significant improvement in pointing performance is a direct  consequence of using direct-drive BLDC actuators instead of stepper-based actuators that will allow the  TSIS scientific sensors to perform their critical measurements while pointed directly at the sun.  Angle (arcsec) NASA/CP—2018-219887 281
Figure 11. On-Orbit Solar Tracking Errors  Conclusion  Direct-drive BLDC actuators were selected and used for 2-axis solar pointing of the TSIS instrument  payload on the ISS and have provided improved performance compared to stepper-based actuators due  to a number of advantages. These advantages include continuous torque command authority, minimal selfinduced jitter, increased active control bandwidth, and structural separation, all of which produce a dramatic  increase in base disturbance attenuation and greatly reduced tracking errors.  As part of implementing the BLDC actuators for TSIS, many challenges were overcome and lessons were  learned regarding control electronics complexity, the importance of controller tuning, sensitivity to center of  gravity and bearing friction, how to account for unpowered free motion, the effect of structural coupling to  closed-loop control, and the importance of end-to-end testing.  Pointing performance was greatly improved compared to stepper-based values as evidenced by the  ground-test measurements of the disturbance transfer function that shows excellent broadband attenuation  and of the solar tracking errors that show significantly lower self-induced jitter. The culmination of this effort  are the on-orbit results that show solar tracking errors less than 4 arcseconds 1σ per axis, which provides  ample margin compared to the requirement of 60 arcseconds 1σ.  References  1. LASP Website for TSIS: http://lasp.colorado.edu/home/missions-projects/quick-facts-tsis/ 2. J. Budinoff, et al “Development of a Deployment and Latching Mechanism for a Pointing System on the Neutron Star Interior Composition Explorer (NICER) for Goddard Space Flight Center” Proceedings of the 43rd Aerospace Mechanisms Symposium , (May 2016), pp. 261-268 3. B. Oaida, D. Bayard, M. Abrahamson “On-orbit Measurement of ISS Vibrations during OPALS Extended Mission Operations” IEEE Aerospace Conference  (2017) Angle (arcsec) NASA/CP—2018-219887 282
Testing and Maturing a Mass Translating Mechanism for a Deep Space CubeSat  Alex Few*, Tiffany Lockett*, Richard Wilson**, David Boling** and Erik Loper*  Abstract  Near Earth Asteroid (NEA) Scout is a deep space satellite set to launch aboard NASA’s Exploration Mission  1. The spacecraft fits within a CubeSat standard 6U (about 300 x 200 x 100 mm) and is designed to travel 1 AU over a 2.5-year mission to observe NEA VG 1991. The spacecraft will use an 86-m2 solar sail to maneuver from lunar orbit to the NEA. One of the critical mechanisms aboard NEA Scout, the Active Mass Translator (AMT), has gone through rigorous design and test cycles since its conception in July of 2015. The AMT is a two-axis translation table required to balance the spacecraft’s center of mass and solar sail center of pressure while also trimming disturbance torque created by off-nominal sail conditions. The AMT has very limited mass and volume requirements but is still required to deliver a large translation range— about 160 x 68 mm—at sub-mm accuracy and precision. The system is constrained to operate in complete exposure to space with limited power and data budgets for mechanical and thermal needs. The NEA Scout team developed and carried out a rigorous test suite for the prototype and engineering development unit (EDU). These tests uncovered numerous design failures and led to many failure investigations and iteration cycles. A paper was previously presented at the 43rd Aerospace Mechanisms Symposia entitled, “Development of a High Performance, Low Profile Translation Table with Wire Feedthrough for a Deep Space CubeSat”. This paper will make note of specific lessons learned: manufacturing philosophy, testing ideologies for high-risk missions, thermal mitigation design for small, motor-driven mechanisms. Introduction  The Near Earth Asteroid (NEA) Scout flight mission discovered the need for an active mechanism for  controlling the location of center of pressure in relation to the center of mass during the early design phase  of the mission, post preliminary design review. The Active Mass Translator (AMT) was then conceived and  passed through Critical Design Review in July 2016 with minimum breadboard and bench top testing. The  engineering team planned and prepared test procedures, facilities and a rigorous test schedule to meet the  approaching flight hardware delivery date at the time. Due to resource constraints, the AMT design and  development was handled by the same team producing the NEA Scout solar sail subsystem (Figure 1).  *NASA Marshall Space Flight Center, Huntsville, AL ** Jacobs Technologies, Huntsville, AL Proceedings of the 44th Aerospace Mechanisms Symposium, NASA Glenn Research Center, May 16-18, 2018  NASA/CP—2018-219887 283
Figure 1. AMT design  The late addition to the flight mission introduced a schedule, cost, and volume constraint that maintained  minimal margins for design development. The test approach for this high risk project was to perform as  needed development tests with prototype units and environmental tests on a high fidelity design unit. The  team utilized 3D printed ABS plastic prototypes, commercial parts and motors, and ambient environments  to perform first order functional tests.   For the prototype units, the project accepted the risk of an outside vendor using CAD models to produce  the mechanical parts. The cost and schedule approach was advantageous due to the low cost for each  piece part and the quick turnaround for product delivery. The out-of-house manufacturing shop imported  CAD models and produced hardware without needing a formal drawing. Navigating around the formal  drawing release process reduced schedule by 6 weeks.   To reduce risk for the flight unit, the team adhered to the Marshall Space Flight Center flight hardware  approved vendor list which did not include the manufacturing shop that produced the high-fidelity  engineering units. This decision cost the project time with reintroducing the hardware to a new  manufacturer. The flight hardware manufacturing company still operated with released drawings and  configuration control. The drawings up to that point in the project were only issued drawings, requiring little  oversight and signatures. The team had to build flight-quality release drawings and adhere to the quality  control process.   Class D high-risk missions have the ability to tailor center-level mandates for development flight hardware  build and manufacturing. The institutional understanding of that tailored approach was a technical and  programmatic challenge. A major lesson learned was to clearly communicate and repeat constantly the  project approach and risk posture for accepting lean manufacturing and acceptance processes. The  knowledge lost between the development hardware manufacturer and the flight hardware manufacturer  was not negligible. The team learned that using the same manufacturer for development and flight in the  future would greatly reduce the quality assurance concerns upon flight hardware acceptance.   NASA/CP—2018-219887 284
Figure 2. Size comparison of 6-mm stepper  Figure 3. 6-mm stepper motor temperature rise vs. time with varying current inputs to determine  sensitivity to conduction .  The motors did not have suitable mechanical joints to the structure to allow for effective thermal conduction.  Early in the design phase, the motor vendor provided mechanical fastening information. Their suggestion— to clamp the motors on the gearbox end near the output shaft—was typical for most applications, but we  soon learned that our vacuum environment required a very different design (Figure 4).   NASA/CP—2018-219887 286
Figure 4. Simple motor diagram showing where clamps should be applied (near left-most areas, away  from the motor). The motors had been clamped near the end of the gearbox, a location opposite of the motor coils  (Figure 5). Due to the construction of the motor, the gearbox and clamps were almost completely thermally  isolated from the heat source due to little contact areas between internal ball bearings, planetary gears,  and a magnetically powered rotor. Therefore, the design did not allow a conduction path between the coils  and the AMT structure.  Figure 5. Application of clamping suggestion, CAD on left, EDU on right.  The FDM0620 motors, not including the gearbox, are the same size as a mechanical pencil eraser and  weigh just 1 gram. With negligible surface area and thermal mass, the 0.5 W of power destroyed the motors  in seconds. FLIR images, as seen in Figure 3, show the case temperatures elevating just a few degrees  above ambient in atmosphere, even without any significant external conductive path. More accurate thermal  models estimated temperature deltas in the 200°C range for vacuum if the conductive path went  unchanged.   A new clamp design and thermal interface material was used to mitigate the issue. The new clamps covered  almost all of the motor’s cylindrical surface and a thin layer of conductive material added between the two  clamp halves and the motor and gearbox casings (Figure 6). Three interface materials such as conductive  greases, epoxies, and thermal springs were traded. Figure 7 shows the updated thermal analysis results  without thermal interface materials, then compares the three interface material options. The three bars for  each data set show the steady state coil temperature at 70°C, 20°C and -70°C.  NASA/CP—2018-219887 287
The 8 columns are divided into 4 pairs. Each pair describes the maximum temperature for each clamp  interface material option. The first column is for when both AMT motors are powered, producing 427 mW  of residual heat, and the second column describes one motor under power, producing 213 mW of residual  heat.  Requirements to avoid offgassing, have a very high thermal conductivity, and simple disassembly directed  the team towards the indium thermal spring. This material has high thermal conductive properties, is  compressible (allowing for machining tolerances and clamping applications), and is common in space flight  electronics. After taking a few weeks to redesign the AMT chassis to include the larger clamps and indium  interface material, the team then set to properly understand the motor’s internal conductive path (Figure 8).  Figure 8. New clamp design showing compressed indium. Notice the textures. This perforated surface is  characteristic of the uncompressed material. The indium sheet will flatten when compressed.  Misunderstood Motor Internal Design  The third issue uncovered during the failure investigation was the team’s lack of knowledge of the motors  internal design. Though it is common for vendors not to present all design data for a customer, we asked,  and asked often. The vendor was responsive, uncommonly understanding and provided all data that was  available.   Figure 9. X-ray of failed motor (left) and new motor (right). Note the irregularities in the failed motor most  likely created from temperature spikes.  In addition to consulting the vendor for internal design information, the team conducted non-destructive  evaluation using X-ray (Figure 9). The team took X-rays of 4 failed vacuum test motors and 10 untested  motors. The team compared the internal images to datasheets to determine that even our understanding  of the heat path even after weeks of collaboration with the vendor was still inaccurate. It became clear that  NASA/CP—2018-219887 289
the motors were not as readily suitable for vacuum as expected. The conductive paths were miniscule,  even compared to the motors small stature. Heat was conducting through multiple small junctions before  ever reaching the motor casing. This was proof that the coils were reaching incredible temperatures before  the motor surface measured any change.   Thermal analyses were updated with the new conductive path information and produced results more  consistent with the data from the first test. The analysis referenced in Figure 7 showed that coil  temperatures rise about 100°C above their maximum survival temperature to nearly 230°C. During that  time, the gearbox was seeing minimal thermal changes. The TVAC test thermocouple data confirmed that  this updated analysis was more accurate. Figure 10 compares the thermal paths required under the old  and new clamp designs.  Figure 10. Thermal Path Explained: old clamp design vs. new clamp design.  Further complicating the issue, we had been relying on thermocouple data coming from the motor case and  clamp to determine motor health. The thermal disconnect between the coils, case, and clamp voided our  previous data and left us unsure how we could properly monitor motor health during a TVAC test.  Thermocouple data from the motor surfaces had limited value due to poor internal conduction and the  motors size did not allow internal monitoring with any test equipment. The team had no way to determine  the motor’s health with any available test equipment and any data received was of little value to determining  the motor’s coil status.  Method to Determine Motor Health  Eventually, the team realized that the motor coils were essentially a pair of resistance temperature detectors  (RTDs). Any temperature change in the copper coils would correlate to a measurable impedance change  and we could use this principal to determine the motor’s internal temperature. The principal equation is  shown in Eq. 1.  NASA/CP—2018-219887 290
test the motors in the correct thermal and vacuum environment are difficult to find on center without a cost  and schedule burden.   Figure 11. AMT motor coil temperature vs. vacuum chamber ambient temperature. Data collected using  coil impedance measurements and temperatures calculated using Equation 1. Development of the AMT flight unit was an excellent opportunity for the young NEA Scout team to learn  about flight hardware testing. The team learned of the false sense of accomplishment bench top testing  can provide. Blind dependence on vendor information proved to be a cost and schedule hit that fortunately  the project was able to afford at the time. Do not take any analytical data from the family-oriented  documentation provided by a vendor without application confirmation or performing inspections.   In summary, the project accepted the risk of utilizing a lien approach to hardware development to help save  cost and schedule. The assumptions made early in the project were with the best information gathered at  the time. Without the anomalies experienced during thermal vacuum testing  The project made the right decision with reserving a large portion of the project schedule to design and  development. Leveraging rapid prototyping, out of house manufacturing shops, and communicative  vendors allowed the project to be successful and absorb all of the lessons learned while still delivering flight  hardware. Use of the lien development process allowed to team to get familiar with the hardware and  confident in the design. This gave the project confidence in accepting the flight hardware mechanism  design.   References  1. McNutt, L.; Johnson, L.; Clardy, D.; Castillo-Rogez, J.; Frick, A.; and L. Jones. “Near-Earth Asteroid Scout.” AIAA Space 2014 Conference; 4-7 Aug. 2014; San Diego, CA; United States. 2. Few, A.; “Development of a High Performance, Low Profile Translation Table with Wire Feedthrough for a Deep Space CubeSat.” 43rd Aerospace Mechanisms Symposia; 4-6 May 2016; Santa Clara, CA.; United States. NASA/CP—2018-219887 292
Solar Array Drive Assembly Qualification Lessons Learned  Jonathan Wood* and Timothy Pargett** Abstract  Being handed the role of responsible engineer for qualification of an already completed design provides a  unique challenge. Budget and schedule will not permit a complete review of every design detail or reperforming every analysis, so it is up to the knowledge and experience of the responsible engineer to  anticipate which environments or design features pose the highest risk to a successful qualification. This  paper presents two test failures that occurred during the Common Solar Array Drive Assembly (CSADA)  qualification. Several potential issues were successfully averted, but this paper is focused on the design of  the CSADA shock test apparatus and loss of step stability during hot thermal vacuum (TVAC) testing as  both of these issues provide useful lessons learned or reinforce good design practice. Both issues were  navigated successfully, resulting in a fully qualified design with several flight units successfully built and  tested to date.  Introduction  Recently, Lockheed Martin developed a modernized version of the A2100 Bus with increased payload  capacity, requiring higher power output from the solar arrays. This required a higher capacity Solar Array  Drive Assembly. The new SADA, called the “Common” SADA or CSADA, had undergone a CDR in 2010  and was deemed ready for qualification testing.   Figure 1. CSADA Assembly *Lockheed Martin Space, Sunnyvale, CA ** Moog Space and Defense Group, Mountain View, CA This document consists of general capabilities information that is not defined as controlled technical data  under ITAR Part 120.10 or EAR Part 772.  Proceedings of the 44th Aerospace Mechanisms Symposium, NASA Glenn Research Center, May 16-18, 2018 NASA/CP—2018-219887 293
Figure 2. Jumped Slip Ring Brush Based on the high-speed video results, the shock pulse duration was measured via the accelerometer time  history data. The measured shock pulse duration was 15 ms, which is 3X longer than the worst-case onorbit pulse duration of 5 ms. This is a limitation of MIPS table shock testing – typically, to meet the high  frequency shock level, excessive energy has to be applied in the low frequencies, resulting in over-test in  the lower frequency range, and a longer than desired shock pulse duration with larger displacements.  Based on these observations, a more flight-like shock test setup was needed.  Revised Shock Test Setup  After the shock failure investigation demonstrated the sensitivity to over-testing the unit in the low frequency  range, especially near 200 Hz, Moog was contacted to perform a mechanical impact shock test tailored to  mitigate the low frequency over-test. Typically, shock level overages at lower frequencies are not a concern,  but this was a critical parameter for the CSADA. Moog had developed shock test fixtures and techniques  to provide this kind of pulse shaping capability previously and were able to adjust and tune an existing  shock fixture to provide an acceptable shock dose to the test article using the Lockheed Martin-provided  vibration test fixture.  The shock test method still relied on a mechanical impact to create the shock event, but several techniques  were used to shape the frequency content of the shock environment to minimize the low-frequency energy  while still achieving the required minimum levels across the full spectrum. The shock is created using a  pneumatic gun and a shock test fixture. The pneumatic gun uses compressed air to propel a metal projectile  that impacts the shock test fixture. The projectile and shock test fixture are specially designed and tuned to  recreate a specific Shock Response Spectrum (SRS) upon impact. Once tuning of the fixture is complete,  shock levels are controlled by the pressure in the air tank when the gun is fired; this results in highly  repeatable shock levels from one shock event to the next. A simplified schematic representation of such a  shock test is shown in Figure 3 and an example of a shock test setup with actual hardware in the Moog lab  is shown in Figure 4.  NASA/CP—2018-219887 295
Figure 3. Schematic Representation of Shock Test  Moog used a metal-to-metal impact for imparting energy into the shock fixture. The key characteristics of  the impact event are controlled by a combination of several variables. The size and shape of the metal  bullet (See Figure 5), the design of the steel strike plate on the shock fixture, and the sacrificial target  material placed between them during the impact are all used to shape the duration and frequency content  of the impact. Moog has developed a set of bullets based around the four-inch (10-cm) diameter bore of  their custom pneumatic gun that spans the range of input excitation typically required for pyroshock  simulation. Similarly, a set of steel strike plates have been developed to be used in conjunction with these  bullets. The combination of bullet and strike plate selection are the primary means used to adjust the  excitation pulse width imparted to the shock fixture. This initial pulse then propagates through the shock  fixture, exciting modes of that structure that further shape the resulting shock environment. Figure 4. Example Shock Test Setup Shock Gun SlugAccelerometers UUT DAQShock  Test  FixtureStrike Plane NASA/CP—2018-219887 296