ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
447 Small Satellite Electronics & Considerations 

Very small cost-effective spacecraft such as CubeS ats are well known to take advantage and are almost  synonymous with employing the us e of commercial-off-the-shelf el ectronic components. The benefits  gained by designing from such volumes of commonly available and very inexpensive hardware, does not  come without some added risk. The demanding environment  of space in which all spacecraft are asked to  operate in, brings with it exposure to seve ral extreme factors, namely radiation.  

Figure 3: Details of drive train and power transfer sections. Stepper motor and multi-stage planetary transmission (A), idle r gear shaft (B), output shaft and gear  (C), and cable wrap (D).    Semiconductor based electronic components are inher ently susceptible to unw anted effects of space  radiation. These effects can be generally separated into two classes: cumulative and single event effects (SEE). The first class, cumulative effects are brought on gradually throughout the lifetime of exposure in a radiation rich environment. The upper limit on the cumulative amount of radiation a susceptible component can accept before failure is referred to as the total ionizing dose. Total ionizing dose is the measure of the cumulative dose of energy transferred into  the material by radiation in the form of ionizing  energy. Thorough testing and characterization of a com ponent can be used to confidently assess its total  ionizing dose. The second class, single event effects ar e quite different in that they can be attributed to  the energy transferred by a single particle interacting with an electronic device. A device susceptible to SEE can therefore exhibit failure at any moment.   There are different types of SEE. Single event upsets are non-damaging soft errors. Single event upsets  typically appear as transient signals, or as bit changes in memory stores. In contrast to single event upsets, which are only seen in software, there exists  a counterpart in hardware, single event latch-up.  Single event latch-ups can result in higher operat ing currents which can lead to component degradation  and possible failure.  

Radiation hardening, or “rad hard” for short, is a me thod of designing and testing electronic components  such that they are resistant to the unwanted effects associated with the environment of high altitude flights above Earth or in outer space. The typical co sts associated such “rad hard” devices are normally  D B  A C  NASA/CP-2010-216272
449   Figure 4: Block diagram of SADA drive and control electronics 

Figure 5: Inner workings of the prototype SADA shown on typical  CubeSat structure. The drive fits into the typically unused space at the end of a standard frame (e.g. Pumpkin v). 

Electrical Controls   This SADA ECU is designed to enable Sun tracking in two possible modes of operation. While the  spacecraft communication module was not included in the prototype, this first mode is intended to allow  the SADA ECU to receive commands over the spacecraft communication bus for purposes such as Sun tracking. Typical Sun tracking are:  “track”, “idle”, “stop” and “position.” If the ECU is not acting on  commands via a host, it will default to  an autonomous Sun tracking mode. In either mode of tracking, the  host may request data from the SADA at any time. For t he purposes of this first prototype, the device was  always run in autonomous Sun tracking mode.  The SADA ECU is capable of entering into a “stand-by” mode while the system is inactive. In stand-by  mode the system is still capable of  receiving and reacting to command s from the host. This powerful  NASA/CP-2010-216272
450 interrupt driven feature of the MCU unit chosen drastically reduces co mplexity of the software control  system and greatly reduces the power consumption of t he system. From a standpoi nt of consumption, the  system will operate on approximately 500 mW while actively tracking and less than 1 mW while in standby.   The solar array position is controlled in an open-loop fashion. The motor current is limited by the stepper  motor controller for thermal protection and power savi ngs reasons. The motor current is also monitored by  the MCU for increased perception from a controls standpoint. Motor steps are accumulated and limit switches (not shown in models) are used to reliably detect proximity of hard stops . The limit switches in  conjunction with hard-stops are approximately 360°  apart and allow for the angular position at the  respective extremes to be known by the MCU. One of the limit-switch/hard-stops will be designated as the  “home” position. This is the default starting point for rotation of the output shafts and attached solar  arrays. The shaft is free to rotate and track the S un (under command from either the host or via feedback  from the Sun sensor) while under control of the MCU fo r approximately one full rotation. At the end of this  rotation the next limit-switch/ hard- stop will be encountered. The system will recognize this limit and the  array will be rotated back the “home” position to pi ck-up where it left off and begin tracking again. The  process will be repeated on orbit, keeping the arrays poi nted at the Sun to maintain optimal power levels  on the solar arrays.   Sun Sensing

For purposes of the prototype, a very simple Sun sensor was created by using two inexpensive COTS  photovoltaic (PV) cells. A “fence” or intra-bank barrier was positioned between the PV cells. In Figure 6,  the action of the intra-bank barrier in determining Sun po sition is illustrated. In th is diagram, it can be  seen that as the angle of the light source (or Sun in this case) changes relative to the array it casts a  shadow onto the cell bank furthest from it, causing a reduced output from t hat bank (or string). A  photograph of the prototype sensor can be seen in Figur e 7. A simple algorithm was written to interpret  the overall effect of the system with an incident light source. 

Figure 6: Illustration of Sun angle to obscured portion of cell banks. Note that of the total available  area (represented by ‘A’) the obscured area (‘B’ and ‘C’) decreases as the light source becomes normal to the panel surface. By comparing the output of the two banks shown, we are able to determine gross array position relative to the Sun. 

NASA/CP-2010-216272
451 The algorithm was written and executed on the MSP430 MCU to accomplish a simple but effective Sun  tracking mode.  The 12-bit ADC on-chip peripheral pr ovided by the MCU was used to sample and convert  the loaded PV cell bank voltages. The ADC was called upon to sample both PV cell bank banks at  approximately 1 kHz; this data product was then used by the algorithm to determine position. A “dead  band” variable was defined and used to account for acceptable difference between the PV cell bank voltages. When the relative voltage between the two PV cell banks deviated by a value greater than the  dead-band, the light-source’s angle of incidence is infe rred and the direction in which the PV array is to be  rotated is established. A sample of  this algorithm is shown here: 

GetDIRECTION( ) {  IF (SUNSENSE1 > (SUNSENSE2 + dead_band))  return CLOCKWISE;  IF (SUNSENSE1 < (SUNSENSE2 - dead_band))  return COUNTER_CLOCKWISE;  ELSE  return NO_MOTION;  } 

Figure 7: Prototype Intra-bank Barrier Test   Because the drive itself is so thin (6.5-mm thick) and so much is packe d in the available space, there is  little room for power transfer. Typically, devices used for transfer of power generated in the solar arrays  into the body of a spacecraft have been either slip rings or twist capsules. Slip rings are devices that use sliding contacts on rotating rings to allow continuous rotation of the ring (typically it is this part that is connected to the arrays themselves). In this drive, the option exists to use either a slip ring or twistdevice. Both are limited however, as the number of sl ip ring circuits and their current-carrying capacities  are limited by the space available. Because of this, the method described above was developed to  perform course sun tracking without additional sensors on the wing.  

   NASA/CP-2010-216272
452  

Figure 8: Characterization of breadboard solar array configuration with raised, intra-bank  barrier. The top graph shows the output of the individual strings as a function of Sun (gimbal) position; the bottom graph shows the normalized voltage ratio . 

  0.80.850.90.9511.05 - 9 0 - 7 5 - 6 0 - 4 5 - 3 0 - 1 5 0 1 53 04 56 07 59 0Output Voltage, Single String, V/V o L R 0.850.900.951.001.051.101.15 -90 -75 -60 -45 -30 -15 0 15 30 45 60 75 90 Angle of Incidence, degNoralized Voltage, V L/VR NASA/CP-2010-216272
453 Conclusions 

The development of an application specific low-profile solar array drive assembly for CubeSats was  successfully developed and demonstrated at SmallSat 2009. The problem of addressing the demand for increasingly more power on a CubeSat bus can be rea lized through implementation of this type of SADA.    The prototype ECU developed for the SADA described he rein possessed only basic functionality and was  done so intentionally for simplicity and demonstration-friendly purposes. It will be critical that future iterations of the ECU incorporate key sub-system s such as a communication bus, external memory  stores, etc. Even further, these future ECU's shoul d also include features also described herein;  specifically for radiation hardening.  While such as pects were decidedly foregone during the development  of this first prototype, it is imperat ive that they be included in order for the electrical system to be tried as a  viable flight unit.  A simplistic and cost-effective method for determining the angle of incidence of a light source (such as the Sun) was developed herein. Autonomous Sun tracking was realized by using PV cells (a common component to any satellite) and simple addition of a shadow-casting intra-bank barrier. While the method  proved to be useful, it is acknowledged that a relati vely large amount of error in achieving normal is  unavoidable with such an effort. Higher accuracy COTS Sun sensors are available for the same purpose, with of course added but likely acceptable costs. Fu ture work should include and effort to down-select  and implement a COTS Sun sensor, while further work using the cost-effective 

References                                                    i  Toorian, A. et al. “The CubeSat Approach to  Space Access” Aerospac e Conference, 2008 IEEE,  March 2008.  ii  Blocker, A., et.al., ”TINYSCOPE – The Feasib ility of a 3-Axis Stabiliz ed Earth Imaging CubeSat  from LEO”, 22nd Annual AIAA/USU Conference on Small Satellites, 2008.  iii  Lee, E., et.al., “The *.Sat CubeSat bus: When Three Cubes Meet”, 19th Annual AIAA/USU  Conference on Small Satellites, 2005.  iv  Dowd, M. “How Rad Hard Do You Need? T he Changing Approach to Space Parts Selection”,  Maxwell Technologies Microelectronics, 2002.  v  http://www.cubesatkit.com/  

NASA/CP-2010-216272
NASA/CP–2014-217519 MAY 2014Edward A. Boesiger and Claef Hakun, Editors   National Aeronautics and   Space Administration  Goddard Space Flight Center   Greenbelt, Maryland 2077142nd Aerospace Mechanism Symposium  Proceedings of a symposium held atHyatt Regency, Baltimore, MDHosted by NASA’s Goddard Space Flight Center andLockheed Martin Space Systems Company Organized by the Mechanisms Education Association May 14–16, 2014
National Aeronautics and   Space Administration  Goddard Space Flight Center   Greenbelt, Maryland 20771NASA/CP–2014-217519 May 2014Edward A. Boesiger  Mechanisms Education Association and  Lockheed Martin Space Systems Company, Sunnyvale, CA Claef Hakun  Goddard Space Flight Center, Greenbelt, MD42nd Aerospace Mechanism Symposium  Proceedings of a symposium held at Hyatt Regency, Baltimore, MDHosted by NASA’s Goddard Space Flight Center andLockheed Martin Space Systems Company Organized by the Mechanisms Education Association May 14–16, 2014
/g36/g89/g68/g76/g79/g68/g69/g79/g72/g3/g73/g85/g82/g80/g29/g3/g3  /g49/g36/g54/g36/g3/g38/g72/g81/g87/g72/g85/g3/g73/g82/g85/g3/g36/g72/g85/g82/g54/g83/g68/g70/g72/g3/g44/g81/g73/g82/g85/g80/g68/g87/g76/g82/g81/g37115 Standard Drive /g43/g68/g81/g82/g89/g72/g85/g15/g3/g48/g39/g3/g21/g20/g19/g26/g25/g16/g20/g22/g21/g19/g49/g68/g87/g76/g82/g81/g68/g79/g3/g55/g72/g70/g75/g81/g76/g70/g68/g79/g3/g44/g81/g73/g82/g85/g80/g68/g87/g76/g82/g81/g3/g54/g72/g85/g89/g76/g70/g72/g3 /g24/g21/g27/g24/g3/g51/g82/g85/g87/g3/g53/g82/g92/g68/g79/g3/g53/g82/g68/g71/g3/g54/g83/g85/g76/g81/g74/g191/g72/g79/g71/g15/g3/g57/g36/g3/g21/g21/g20/g25/g20/g3/g51/g85/g76/g70/g72/g3/g38/g82/g71/g72/g29/g3/g36/g20/g26Level of Review: This material has been technically reviewed by technical management and  professionally reviewed by NASA’s Goddard Space Flight Center./g55/g85/g68/g71/g72/g3/g81/g68/g80/g72/g86/g3/g68/g81/g71/g3/g87/g85/g68/g71/g72/g80/g68/g85/g78/g86/g3/g68/g85/g72/g3/g88/g86/g72/g71/g3/g76/g81/g3/g87/g75/g76/g86/g3/g85/g72/g83/g82/g85/g87/g3/g73/g82/g85/g3/g76/g71/g72/g81/g87/g76/g73/g76/g70/g68/g87/g76/g82/g81/g3/g82/g81/g79/g92/g17/g3/g55/g75/g72/g76/g85/g3/g88/g86/g68/g74/g72/g3/g71/g82/g72/g86/g3/g81/g82/g87/g3/g70/g82/g81/g86/g87/g76/g87/g88/g87/g72/g3/g68/g81/g3/g82/g73/g73/g76/g70/g76/g68/g79/g3/g72/g81/g71/g82/g85/g86/g72/g80/g72/g81/g87/g15/g3/g72/g76/g87/g75/g72/g85/g3/g72/g91/g83/g85/g72/g86/g86/g72/g71/g3/g82/g85/g3/g76/g80/g83/g79/g76/g72/g71/g15/g3/g69/g92/g3/g87/g75/g72/g3/g49/g68/g87/g76/g82/g81/g68/g79/g3/g36/g72/g85/g82/g81/g68/g88/g87/g76/g70/g86/g3/g68/g81/g71/g3Space Administration./g55/g75/g76/g86/g3/g49/g36/g54/g36/g3/g70/g82/g81/g73/g72/g85/g72/g81/g70/g72/g3/g83/g88/g69/g79/g76/g70/g68/g87/g76/g82/g81/g3/g70/g82/g81/g87/g68/g76/g81/g86/g3/g69/g82/g87/g75/g3/g56/g17/g54/g17/g3/g42/g82/g89/g72/g85/g81/g80/g72/g81/g87/g3/g90/g82/g85/g78/g86/g3/g68/g81/g71/g3/g81/g82/g81/g16/g74/g82/g89/g72/g85/g81/g80/g72/g81/g87/g3 /g90/g82/g85/g78/g86/g17/g3/g49/g36/g54/g36/g3/g75/g68/g86/g3/g69/g72/g72/g81/g3/g74/g85/g68/g81/g87/g72/g71/g3/g83/g72/g85/g80/g76/g86/g86/g76/g82/g81/g3/g87/g82/g3/g83/g88/g69/g79/g76/g86/g75/g3/g68/g81/g71/g3/g71/g76/g86/g86/g72/g80/g76/g81/g68/g87/g72/g3/g81/g82/g81/g16/g74/g82/g89/g72/g85/g81/g80/g72/g81/g87/g3/g90/g82/g85/g78/g86/g3 /g70/g82/g81/g87/g68/g76/g81/g72/g71/g3/g75/g72/g85/g72/g76/g81/g3/g90/g76/g87/g75/g3/g68/g79/g79/g3/g82/g87/g75/g72/g85/g3/g85/g76/g74/g75/g87/g86/g3/g83/g85/g72/g86/g88/g80/g68/g69/g79/g92/g3/g85/g72/g87/g68/g76/g81/g72/g71/g3/g69/g92/g3/g87/g75/g72/g3/g68/g88/g87/g75/g82/g85/g86/g3/g68/g81/g71/g18/g82/g85/g3/g70/g82/g83/g92/g85/g76/g74/g75/g87/g3/g82/g90/g81/g72/g85/g86/g17/g3/g3/g3 /g41/g82/g85/g3/g56/g17/g54/g17/g3/g42/g82/g89/g72/g85/g81/g80/g72/g81/g87/g3/g90/g82/g85/g78/g86/g15/g3/g81/g82/g3/g70/g82/g83/g92/g85/g76/g74/g75/g87/g3/g76/g86/g3/g70/g79/g68/g76/g80/g72/g71/g3/g76/g81/g3/g87/g75/g72/g3/g56/g81/g76/g87/g72/g71/g3/g54/g87/g68/g87/g72/g86/g3/g88/g81/g71/g72/g85/g3/g55/g76/g87/g79/g72/g3/g20/g26/g15/g3/g56/g17/g54/g17/g3/g38/g82/g71/g72/g17/g3  /g36/g79/g79/g3/g50/g87/g75/g72/g85/g3/g53/g76/g74/g75/g87/g86/g3/g53/g72/g86/g72/g85/g89/g72/g71/g17
iii PREFACE  

The Aerospace Mechanisms Symposium (AMS) provides a unique forum for those  active in the design, production and use of aerospace mechanisms. A major focus is the  reporting of problems and solutions associated with the development and flight  certification of new mechanisms. Organized by the Mechanisms Education Association,  responsibility for hosting the AMS is shared by the National Aeronautics and Space  Administration and Lockheed Martin Space Systems Company (LMSSC). Now in its  42nd symposium, the AMS continues to be well attended, attracting participants from  both the U.S. and abroad.  

The 42nd AMS, hosted by the NASA Goddard Space Flight Center (GSFC) in Baltimore,  Maryland, was held May 14, 15 and 16, 2014. During these three days, 43 papers were  presented. Topics included payload and positioning mechanisms, components such as  hinges and motors, CubeSats, tribology, and mechanism testing. Hardware displays  during the supplier exhibit gave attendees an opportunity to meet with developers of  current and future mechanism components.  

The high quality of this symposium is a result of the work of many people, and their  efforts are gratefully acknowledged. This extends to the voluntary members of the  symposium organizing committee representing the eight NASA field centers, LMSSC,  and the European Space Agency. Appreciation is also extended to the session chairs,  the authors, and particularly the personnel at GSFC responsible for the symposium  arrangements and the publication of these proceedings. A sincere thank you also goes  to the symposium executive committee who is responsible for the year-to-year  management of the AMS, including paper processing and preparation of the program.   

The use of trade names of manufacturers in this publication does not constitute an  official endorsement of such products or manufacturers, either expressed or implied, by  the National Aeronautics and Space Administration.
vii Magnetic Gearboxes for Aeros pace Applica tions ................................................................................. .... 365  Jose Luis Perez-Diaz, Efren Diez-Jimenez , Marco Antonio Alvarez-Valenzuela,   Juan Sanchez-García-Casarrubios, Christi an Cristache & Ignacio Valiente-Blanco   Primer – Stepper Motor Nomenclature, Definiti on, Performance and Re commended Test Methods ...... 375  Scott Starin & Cutter Shea 

Development Testing and Subsequent Failure Investigation of a Spring Strut Me chanism ..................... 391  Jared Dervan, Brandan Robertson,  Lucas Staab, Michael Cul berson & Joseph Pellicciotti 

Development of a Tendon-Act uated Lightweight In-Space MA Nipulator (TAL ISMAN) ............................ 405  William Doggett, John Dorsey, Thomas Jones & Bruce King 

The Use, Evolution and Lessons  Learnt of Deployable Static Solar Array Me chanisms .......................... 421  Mark Ferris & Andrew Haslehurst 

Deployment Mechanism for Ther mal Pointing System ............................................................................. 435  Kraig Koski 

ISS Ammonia Pump Failure, Recovery, and Lesson Learned- A Hydrodynamic   Bearing Perspective ........................................................................................................... ....................... 451  Robert Bruckner & Richard Manco II 

Wide Range Vacuum Pumps for the SAM Instru ment on the MSL Curi osity Rover /elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis /elipsis463  Paul Sorensen, Robert Klin e-Schoder & Rodger Farley 

SLM Produced Hermetically S ealed Isolation Valve ................................................ /elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis463  James Richard 

DRAGON - 8U Nanosatellite Orbital Deployer /elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis../elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis /elipsis/elipsis487  Marcin Dobrowolski, Jerzy Grygorczuk, Barto sz Kędziora, Marta Tokarz & Maciej Borys 

Mechanism Design and Testing of a Self-Depl oying Structure Using Flexible Composite   Tape Spri ngs/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis /elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis ..... /elipsis..497  Joseph Footdale & Thomas W. Murphey 

Innovative Escapement-Based Mechanism for Micro-Antenna Boom D eployment ./elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis /elipsis/elipsis511  Marta Tokarz, Jerzy Grygorczuk, St anisław Jarzynka & Henryk Gut 

Design and Functional Validation of a Mechani sm for Dual-Spinning CubeSats /elipsis /elipsis/elipsis/elipsis/elipsis/elipsis/elipsis/elipsis /elipsis/elipsis523  Eric Peters, et al
viii SYMPOSIUM SCHEDULE 

WEDNESDAY, 14 MAY 2014   7:00 Wednesday Presenters' Breakfast - Pi sces Room, 15th fl oor, Hyatt Regency   8:00  CHECK-IN AND REFRESHMENTS  - Ballroom Foyer, Hyatt Regency 

8:15  INTRODUCTORY REMARKS  - Constellation Ballr oom, Hyatt Regency  

Joseph P. Schepis, Co-Host Chairman, NASA Goddard Space Flight Center, Greenbelt, MD  Stuart Loewenthal, Deputy Chai rman, Lockheed Martin Spac e Systems, Sunnyvale, CA  Christopher Scolese, Director, NASA Goddar d Space Flight Center, Greenbelt, MD   8:45  SESSION I – PAYLOAD & DEPLOYMENT MECHANISMS  James Wells, Session Chair  NASA Langley Research Center, Langley, VA 

• An Innovative Aperture Cover Mechanism used on SDO/EVE and MMS/SDP  Stephen Steg, William Vermeer, Scott Tucker & H eather Passe, Laboratory for Atmospheric and  Space Physics, University of Colorado, Boulder, CO 

• Metering Wheel-Wire Track Wire  Boom Deployment Mechanism  Mark Granoff, University of New Hampshire, Durham, NH 

• Development of Cryogenic Filt erwheels for the HERSCHEL P hotodetector Array Camera &  Spectrometer (PACS)  Christian Körner & Dirk Kamp f, Kayser-Threde GmbH, Munich, Germany; Albrecht Poglitsch &  Josef Schubert, Max Plank Institut für extraterrestrische Physik, Garching, Germany; U. Ruppert  & M. Schoele, Freie Universitat Berlin , Berlin, Germany 

• Development and Preliminary Testing of a High Precision Long Stroke Slit Change Mechanism  for the SPICE Instrument Gabriel Paciotti, Fabrice Rottmeier & Luc Bl echa, Almatech, Lausanne, Switzerland; Martin  Humphries, SpaceMech Ltd ., Bristol, United Kingdom 

• High Gain Antenna System Deployment Mechanism  Integration, Characterization, and Lessons  Learned Fil Parlong, Vantage Systems Inc., Lanham, MD; Blair Russell, Walter Garcen, Chris Johnson &  Craig Huber, Stinger Ghaffarian Technologies Inc., Greenbelt, MD; Chris Rose, NASA Goddard  Space Flight Center, Greenbelt, MD 

• In-Orbit Performance of the MWRI Scanning Mechanisms  Manfred Schmid, Airbus Defenc e & Space, Friedrichshafen, Germany; Miao Jun & Yu Shuang,  Aerospace System Engineering, Shanghai, China   11:45  LUNCH   Lunch for AMS Attendees in the Atrium and Harborview, Hyatt Regency