ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
SCB Initiation Element Characteristics  Our SI design utilizes a Semiconductor Bridge (SCB) firing element manufactured by our SCB Technologies  division in Albuquerque, NM. SCB technology was originally developed by Sandia National Labs in the mid1980s as a low-energy replacement for HBW and Exploding Bridgewire devices. EBAD manufactures more  than 10,000 SCB based initiation devices each year used in aerospace and industrial initiation applications.  SCB firing elements offer a significant reduction in the all-fire energy over conventional HBW devices, while  meeting 1.0 A/1.0 W no-fire performance. The low-energy all-fire level of an SCB device is a direct result  of the much smaller mass of a semiconductor bridge as compared to a typical HBW initiator. The all-fire  energy for a typical SCB device is 200 microjoules, whereas the corresponding all-fire energy of an HBW  initiator is greater than 25 millijoules. While the all-fire energy is low, the application of a short-duration,  high-current pulse allows a SCB element to rapidly change from solid to plasma state in tens of  microseconds, providing a convective heat transfer into the surrounding pyrotechnic material.2 Figure 2. Semiconductor Bridge Construction  Smart Initiator Environmental Testing  Prior to application on the CST-100 program, EBAD successfully tested our SI to the following  environments. This provided confidence that our SI could hold up to the rigors demanded by a manned  spacecraft environmental qualification test program.  •Thermal Cycling: -54°C to 71°C •Random Vibration: 111 grms, 180 sec/axis •Pyrotechnic Shock: 14,000g peak •EMI: CE102, CS101, CS106, CS116, RE102, RS103, TT101 •Ionizing Radiation: To 8.00E+7 Protons/cm2-sec flux, 5.39 kRadSi total dose Prior Applications  Prior to CST-100, EBAD had significant test and fielding experience of our smart initiation system, having  produced nearly 6000 smart initiators. Most of the applications were for tactical systems, so our SI had a  Technology Readiness Level of 7 for manned space environments prior to our CST-100 qualification effort.  Prior papers described our technology being applied to derivative systems: an RF safe Mobility Unit3 and a  tactical missile divert and attitude control system4. Both applications demonstrated the technology could be  configured to the custom application needs of divergent systems, and could be readily scaled up to relatively  high production volumes. Metallized  Pads  Silicon Substrate  Doped PolysiliconActive  Bridge  Area NASA/CP—2018-219887 265
Application to CST-100 Starliner  EBAD’s Smart Initiation System was adopted for use on the CST-100 Starliner spacecraft for NASA’s  Commercial Crew Program. The benefit to the spacecraft was estimated savings of several hundred pounds  in cabling, firing electronics, and pyrotechnic batteries over a traditional initiation system using heritage  manned space HBW initiators. Our experience is that there are very few options on manned and exploration  spacecraft that can offer mass savings of this magnitude.    Radiation Environments  Prior to selection for the CST-100 platform, our Smart Initiator (SI) electronics were subjected to radiation  equivalent of 10X the estimated maximum mission life and worked flawlessly. Subsequently, EBAD has  implemented a Radiation Lot Acceptance Test (RLAT) screening process on EEE parts used in our design  to assure proper parts operation during radiation environments. Our design architecture has inherent  immunity to Single Event Upset since initiators remain in a powered-off configuration until immediately prior  to actuation when they enter a rapid Arm-Fire sequence. This sequence take place in less than several  seconds, so probability to upset is minimized.  Devices Initiated  Our Smart Initiators (SI) and Smart Detonators (SD) were used to initiate many unique devices used on the  CST-100 platform including separation nuts and bolts, cutters, frangible joints, and many other devices.  While the photographs shown in Figure 3 are not necessarily the exact devices, they offer a good  representation of the devices actuated on CST-100. EBAD provides the majority of the initiators, detonators,  and pyrotechnically actuated devices used on CST-100 and other current manned platforms in the U.S.   Figure 3. Types of Devices Initiated by Smart Initiators  Left to Right: Separation Nuts, Separation Bolts, Cutters, Frangible Joints  Environmental Validation and Qualification  EBAD successfully completed an extensive validation and qualification test program for the CST-100  platform. Testing was first done at the SI and SD level, and subsequently at the device level, joining the SI  and SDs to end effectors such as separation nuts. The total validation and qualification effort involved more  than 700 Smart Initiators and Detonators. The environments were roughly within prior test levels. Some  issues arose from self-induced shock, where there were high shock levels observed during actuation of the  various devices used on the spacecraft to redundant initiators. While there were a few bumps during the  test program, EBAD was able to fix these issues and successfully completed its qualification effort in 2017.  NASA/CP—2018-219887 266
Application to Future Platforms  With the successful qualification for the CST-100 platform, this technology is being adopted on other  platforms within the space community. One of the applications is coupling this technology with a nonexplosive Split Spool Release Device (SSRD) offered by our NEA Electronics subsidiary. NEA SSRDs are  widely used in the space industry for solar array, antenna reflector, payload release, and spacecraft  separation and are well known for high preload capability while offering low shock separation. The weight  and mass benefits can be similar for systems that use various one-shot technologies, although they were  very pronounced for the CST-100 application due to the large number of pyrotechnic devices employed  during the launch and subsequent re-entry events.  Figure 4. Example of an NEA SSRD Device  References  1. Boucher, C., and Novotney, D. B., “Performance Evaluation of an Addressable Integrated Ordnance System,” AIAA 2001-3636, 37thAIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit , Salt Lake City, UT, July 8-11, 2001.  2. Boucher, C., and Novotney, D. B., “Next Generation Semiconductor Bridge Initiators,” AIAA 2000-3731, 36th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit , Huntsville, AL, July 16-19, 2000.  3. Barglowski, M, and Boucher, C., “Advanced Ordnance Initiation System Development for a Unique Set of Requirements,” AIAA 2002-4333, 38thAIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit , Indianapolis, IN, July 7-10, 2002.  4. Novotney, D. B., and Kochanek, A., “Intelligent Initiation Systems for Divert and Attitude Control Applications,” AIAA 2005-4502, 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit , Tucson, AZ, July 10-13, 2005. NASA/CP—2018-219887 267
Use and Advantages of Direct-Drive Brushless DC Actuators for Precision  Instrument Pointing of the Total and Spectral Solar Irradiance Sensor  Patrick Brown*, Andrew Engelmann * and Ryan Lewis* Abstract  This paper describes the use of direct-drive brushless direct current (BLDC) actuators for precision solar  pointing of the Total and Spectral Solar Irradiance Sensor (TSIS) instrument that is currently flying on the  International Space Station (ISS). Advantages of direct-drive BLDC actuators compared to stepper-based  actuators are described with an emphasis on the ability to provide excellent disturbance attenuation to base  motion jitter. Lessons learned from incorporating direct-drive BLDC actuators into the TSIS system are  presented to provide a context of the challenges that are associated with this technology. Finally, measured  results from ground testing and on-orbit solar tracking are presented that demonstrate superior pointing  performance compared to stepper-based actuators with solar tracking errors of less than 4 arcseconds 1σ  per axis.  TSIS Instrument  TSIS is a NASA instrument designed, built, and operated by the Laboratory for Atmospheric and Space  Physics (LASP) at the University of Colorado that continues the 39-year-long record of total solar irradiance  (TSI) and 14-year-long record of spectral solar irradiance (SSI) [1]. Both TSI and SSI are critical to longterm climate monitoring. TSI is measured by the Total Irradiance Monitor (TIM), and SSI is measured by  the Spectral Irradiance Monitor (SIM). These two sensors comprise the TSIS instrument along with the  pointing system that will be discussed in this paper.  TSIS was launched to the International Space Station (ISS) on December 15, 2017 aboard a SpaceX  Falcon 9 rocket in the Dragon trunk. It was then installed on the zenith side of the ISS at Site 5 of Express  Logistics Carrier (ELC) 3 as shown in Figure 1. After robotic arm installation, the TSIS deployment system  rotated the gimbaled instrument suite into its operational position above the ELC. TSIS’s operational  position above the ELC allows for a hemispherical view of the sky and solar viewing every orbit during its  5-year mission. In order to perform solar measurements, TSIS must be pointed directly at the sun with less than 0.017 deg  (60 arcseconds) 1σ precision while subject to movement by the nadir-pointed ISS. To decouple the TSIS  instruments from ISS motion, the Thermal Pointing System (TPS) was developed. The TPS provides both  thermal and pointing control for the TSIS scientific sensors; however, this paper is only concerned with the  pointing aspects of the TPS.  *Laboratory for Atmospheric and Space Physics, University of Colorado, Boulder, CO Proceedings of the 44th Aerospace Mechanisms Symposium, NASA Glenn Research Center, May 16-18, 2018 NASA/CP—2018-219887 269
Figure 1. TSIS Instrument Shown at ELC-3 Mounting Location on ISS  TPS Hardware Configuration  The heart of the TPS is a two-axis, elevation over azimuth gimbal that points the instrument optical bench  (including the SIM and TIM sensors) relative to the TSIS base plate as shown in Figure 2. The optical bench  dimensions are 1.0-m wide by 0.9-m deep by 0.38-m tall with a moving mass of 83 kg and a rotational  inertia of 10.2 kg •m2 about the azimuth axis and 3.8 kg •m2 about the elevation axis. BLDC actuators are  used for both rotational axes and consist of the following key components: BLDC direct-drive motors, 2speed (1x,16x) resolvers, brakes, twist capsules for signal/power feed-thru, and output duplex pair  bearings. The TPS has a dedicated control electronics box that includes the drive electronics, a flight  computer, power handling, and interfaces to the various TPS actuators and sensors. For closed-loop  pointing control with respect to the sun, the TPS has redundant High-rate Fine Sun Sensors (HFSS) that  are mounted on the optical bench near the TIM and SIM sensors. Finally, as already mentioned, the TPS  includes a deployment system that rotates the optical bench to its operational position and holds it rigidly  in place. Incidentally, the primary deployment actuator was derived directly from the NICER DAPS actuator  that is also flying on the ISS [2].   NASA/CP—2018-219887 270
Figure 2. TSIS Instrument Components  Gimbal Motion Profile  TSIS follows approximately the same motion profile every orbit as shown by the example flight data in  Figure 3. The azimuth axis provides the primary motion of the gimbal over the course of the orbit as it tracks  the sun overhead from sunrise to sunset from +110 to -110 deg at a rate of approximately 4 deg/min. The  elevation axis accounts primarily for variations in the solar beta angle that varies periodically every 60 days  with annual maxima/minima of ±75 deg. Additionally, both axes account for motion due to static and  dynamic offsets of the ISS flight attitude relative to the Local Vertical-Local Horizontal reference frame.  Also, the TPS points to zenith during every eclipse to support dark space measurements.   Figure 3. TPS Gimbal Motion Profile  ISS Jitter Environment  When TSIS was directed in 2014 to fly on the ISS, the ISS disturbance environment was poorly understood  from a pointing perspective, which led to significant uncertainty in the effect that base motion-induced jitter  NASA/CP—2018-219887 271
motion as possible (i.e., ISS jitter). These simulations were performed by exciting the base of a 1 degree  of freedom actuator with a range of sinusoidal frequencies at fixed amplitude and comparing the output  motion of the load when operating in both open-loop and closed-loop cases.  The stepper-based actuator results are shown in Figure 4 for open- and closed-loop cases. The y-axis  shows the transmissibility, which is defined as the absolute value of the output magnitude to the input  magnitude expressed in decibels. The x-axis shows the base excitation frequency in hertz on a log scale.  Examining the open-loop response, three regimes are present in different frequency ranges that are  characteristic of a spring-mass-damper system subjected to base excitation. For low frequencies (< 5 Hz),  the output motion tracks the input motion identically (0 dB) as a rigid body, which implies that any base  motions are replicated exactly in the output. For mid-frequencies near the natural frequency (~11 Hz) of the  load/actuator system, the output motion is amplified compared to the input motion by up to 14 dB (5x  amplification). And finally, for higher frequencies above the natural frequency, the output motion is less than  the input motion, which represents attenuation that is also referred to as seismic isolation. It is worth  mentioning that for this characteristic, the larger the output inertia, the lower the natural frequency and the  larger the frequency range of passive seismic (inertial) attenuation.  The closed-loop response assumes a control bandwidth that is much lower than the fundamental frequency,  which is a reasonable assumption given the need to avoid active excitation of the fundamental mode. The  response is nearly identical to the open-loop response, except that the closed-loop control is able to actively  attenuate low-frequency base motion. This is an intuitive result given the successful implementation of solar  arrays steppers to track the sun in the presence of slowly varying spacecraft base motion.  The key take-away of this plot is that a range of frequencies exist that can pose a risk to precise pointing  where base motion jitter is either passed to the output in a rigid body manner or, even worse, amplified to  magnitudes larger than the input frequency. If the base motion jitter is well understood and the predicted  output is within allowable levels, this uncompensated jitter regime might be acceptable. But in the case of  the poorly understood ISS jitter environment and with little margin in the pointing budget, this stepper motor  characteristic posed a significant risk to the TSIS TPS jitter performance.   Figure 4. Stepper Frequency Regimes in Response to Base Excitation  NASA/CP—2018-219887 274
Next, open-loop responses are compared between the stepper-based actuators that were just discussed  and direct-drive BLDC actuators as shown in Figure 5. Because of the direct-drive aspect of the BLDC  actuators that were described earlier as structural separation, the BLDC response has no significant  resonance and the rigid body coupling is only present at low frequencies (< 1 Hz). In the same manner as  the stepper-based actuators, seismic (inertial) isolation/attenuation is present but at a much lower frequency  of ~1 Hz compared to ~15 Hz. This is a significant difference between the two actuator types that is capable  of providing significant additional passive attenuation to base motion that increases with increasing load  inertia.  Figure 5. Comparison of BLDC to Stepper Open-Loop Responses  Finally, closed-loop responses are compared between the stepper-based actuators and direct-drive BLDC  actuators as shown in Figure 6. The increased active control bandwidth of the BLDC actuator is able to  complement the passive attenuation to produce the robust, broadband jitter attenuation that is shown in the  plot. The BLDC actuator clearly provides increased attenuation across all frequencies compared to the  stepper-based actuator with none of the significant rigid body and resonant susceptibility. For this reason,  BLDC actuators were selected to provide a robust pointing solution for TSIS in the presence of the uncertain  ISS jitter environment.  NASA/CP—2018-219887 275
Figure 6. Comparison of BLDC to Stepper Closed-Loop Transmissibility Responses  Lessons Learned  Although direct-drive BLDC actuators can provide superior pointing performance compared to stepperbased actuators for precision pointing applications, there are associated challenges that were learned firsthand during TSIS development and will be shared here.  Control Electronics Complexity  Stepper control electronics are well understood and relatively easily matched to actuators for most  applications. A prime reason is that the electronics operate open-loop with respect to the actuator motion.  On the other hand, BLDC actuators require closed-loop control in order to function properly for pointing  applications. Properly developing the closed-loop system that consists of the actuator, the feedback  sensors, the drive electronics, and the control law is much more complicated than open-loop stepper  electronics.  In particular, for TSIS, the electronics and motors were developed concurrently which often resulted in a  case of the “chicken or the egg” for development decisions. In order to minimize this effect, prototype  versions of the actuators and control electronics were developed and tested together to ensure mutual  compatibility. This prototype approach is highly recommended.   Additionally, BLDC pointing performance is directly correlated to the feedback sensor performance, so the  feedback sensor electronics are much more critical than they are for stepper-based actuators. Finally,  having a configurable control law in software is much more complex and performance critical than most  stepper commanding approaches. For all of these reasons, it is critical that sufficient resources are  allocated for development of control/feedback sensor electronics and associated software if BLDC  actuators are used.  Performance Dependent on Tuning  As originally mentioned, steppers use a simple set-it-and-forget-it control approach that does not require  any tuning after they are built. BLDC actuators, on the other hand, require precise and frequent tuning in  order to operate properly and meet requirements.  NASA/CP—2018-219887 276
and that there is no substitute for direct measurements because they may differ dramatically compared to  predictions.   End-to-End Testing  Appropriate end-to-end testing is invaluable for any space-borne system, and it definitely holds true for  BLDC-based systems because of all the lessons that were previously described. In addition to the prototype  testing, frequency response measurements, and modal testing, TSIS performed a rigorous end-to-end test  program that exercised every pointing mode in a flight-like manner that even included physical simulations  of the rotating ISS with an inertially fixed sun as shown in Figure 7. As the following section will attest, all  of this testing contributed positively to the successful operation and performance of the TSIS TPS when it  reached orbit.  Figure 7. TSIS TPS Test Configuration During Subsystem Characterization Results  In spite of the challenges that were associated with using direct-drive BLDC actuators, excellent results  were achieved during the test program and during early on-orbit operations that confirmed the many  benefits of BLDC actuators for precision pointing. A small sampling of results is shown in this section.  Stepper Performance Baseline  For comparison, correlated simulation results from an earlier, NPOESS-era of the TSIS TPS design are  presented in Figure 8. These results show the residual azimuth axis error of 33 arcseconds 1σ when solar  tracking at ~3.6 deg/min without any additional structural error (10 arcseconds 1σ) or spacecraft base  motion (41 arcseconds 1σ). This residual motion profile was a consequence of commanding steps at ~8  NASA/CP—2018-219887 278
steps/s which repeatedly excited the first mode (~11 Hz). When considering all of these error sources, the  60 arcseconds 1σ requirement was only marginally satisfied.   Figure 8. TPS Residual Jitter Motion for NPOESS-era Design  Disturbance Transfer Function Measurements  As previously described, frequency response measurements of the BLDC actuators were performed  frequently throughout the test program, and these measurements included different transfer functions  including the transfer function shown in Figure 9. This plot shows the modeled and measured “input”  disturbance transfer function that is an indication of how well the closed-loop BLDC actuator is at  attenuating torque disturbances. This disturbance transfer function was selected because it can be readily  measured in the lab, whereas the previously discussed BLDC closed-loop transmissibility transfer function  in Figure 6 would require a shaker-table approach for direct measurement that was beyond the scope of  the TSIS program. The as-built system agrees extremely well with the modeled dynamics and shows robust,  broadband disturbance attenuation that far exceeds a stepper/harmonic drive-based actuator and serves  as the best measurement of robust disturbance attenuation for this system.  NASA/CP—2018-219887 279
Figure 9. Measured Input Disturbance Transfer Function Ground-Testing Solar Tracking Errors  After the BLDC control loops were properly tuned using the frequency response measurement approach  previously discussed, steady-state solar tracking tests were performed using the test apparatus shown in  Figure 7 that included ISS base motion simulated by a rotational stage and a fixed solar simulator.   Figure 10 shows the steady-state solar tracking errors (self-induced jitter) for the azimuth axis during ground  testing as measured by the High-rate Fine Sun Sensor (HFSS). The HFSS errors are shown in  magenta/red, and an independent angular measurement was provided by a ring laser gyro (RLG) shown  in blue that tracks the HFSS errors throughout the dataset. The measured HFSS error is 6 arcseconds 1σ  over this dataset, and the RLG error is slightly less at 4 arcseconds 1σ because it is only measuring the  mechanical motion and not the optical noise from the solar stimulus. This dataset was of high importance  during the test program because it demonstrated that the BLDC actuators were capable of providing very  low self-induced jitter compared to the 60 arcseconds 1σ requirement and compared to the stepper motor  results from Figure 8.  Magnitude (dB) NASA/CP—2018-219887 280