ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
thickness set the groove depth. When testing began, it was found that a deeper groove was needed to fit  the tolerance ring. The surrogate ring was made from A513 mild steel with a height of 12.7 mm (0.500 in) and a wall thickness of 1.59 mm (0.0625 in). This material differs from the material of the flight rotor, Custom 455 stainless steel, but its low cost, accessibility, and ease of machining made it the best option. Table 4.  Rotor-mounting Options Press-Fit Epoxy Tolerance Ring Advantages• No additional parts  required • Simple to design• Flight heritage  (ACIS instrument, /g69/g82/g81/g71/g3/g75/g72/g79/g71/g3/g90/g76/g87/g75/g3/g507/g55/g3/g82/g73/g3 -145 °C)• Robust to thermal  expansion/contraction • Known torque and  radial load capacities Disadvantages• Loss of fit with thermal  variation • Requires additional  equipment for  installation• High CTE • Estimated elastic  modulus • Requires careful  application, curing• No prior experience • No (known) flight  heritage • Tests needed to get  groove fit right Testing Performed Thermal testing Torque testing Torque testing Figure 9. Dimensions of shaft/rotor test rig /g11/g19/g17/g22/g28/g19/g3/g76/g81/g3/g32/g3/g20/g19/g3/g80/g80/g12/g15/g3/g11/g19/g17/g19/g19/g24/g3/g76/g81/g3/g32/g3/g20/g3/g80/g80/g12/g15/g3/g11/g20/g17/g24/g23/g25/g3/g76/g81/g3/g32/g3/g22/g28/g17/g21/g3/g80/g80/g12/g15/g3/g11/g19/g17/g24/g24/g19/g3/g76/g81/g3/g32/g3/g20/g23/g3/g80/g80/g12/g3 During assembly of the test shafts, it quickly became apparent that the chosen groove depth was too  shallow to securely hold the tolerance ring. On the first assembly attempt, the tolerance ring was forced  out of the groove when even minimal force was applied to the surrogate rotor ring. Attempts to mate the  two pieces with excessive force only served to buckle the tolerance ring. A new surrogate shaft was machined for the tolerance ring unit, this time with a groove depth of 0.056 cm (0.022 in); this depth was selected to fully retain the 0.051-cm (0.020-in) base thickness of the tolerance ring along with 25% of the corrugated portion. The circumference of the ring was trimmed to maintain a gap of 0.15 cm (0.060 in) between edges, which fell within range of 0.10 to 0.20 cm (0.040 to 0.080 in) recommended by the  product engineers. A 1/2-13 bolt was secured through a bore through the center of the test shaft. Flat washers were placed  on either side of the bolt and the installation was torqued to 13.6 N-m (120 in-lb). The assembly was then  placed in a vise and the jaws were tightened around the ring. A torque wrench was then placed on the  bolt head and torque was applied to the assembly. The test procedure continued as follows: (1) Test both units to the peak torque provided by the motor during nominal operations (4 in-lb or 0.5 N-m); (2) If  successful, continue application of torque until failure occurs; (3) Note the torque at which the assembly failed. 531
/g57is a rotation adjustment factor, /g59is a radial load factor, /g60is an axial load factor, /g41/g85is the applied radial  load, and /g41/g68is the applied axial load. The factors /g57/g15/g3/g59 and /g60are tabulated for different values of the  quantity presented in Equation 5. /g3007/g3276 /g3027/g3005/g3118(5)  /g41/g68is again the axial load, /g61is the number of balls in the bearing, and /g39is the ball diameter. For the  bearing used in the scanner assembly, /g61= 42, /g39 = 1.59 mm (0.0626 in), and the dynamic load rating, /g38is  961 N (216 lbf).  The axial and radial loads were determined by assuming payload rotation in a 1- /g42environment, as the  force of gravity acting on the payload mass is higher than any loads that will be experienced on orbit.  Assuming a payload mass of 1 kg (2.2 lbm), the axial load on the bearing is 10 N (2.2 lbf). After  performing the calculation for /g1832/g3028//g1852/g1830/g2870, it was found that /g57= 1.2, /g59= 0.0056, and /g60= 2.3. The equivalent  radial load was then calculated to be 23 N (5.17 lbf):  The default lifetime equation gives the expected lifetime that 90% of the bearings in a given application  will meet. To determine the expected lifetime that 99% of the bearings in a given application will meet, the  /g68/g20factor must be changed to /g68/g20= 0.21. This provides a conservative estimate of the expected bearing  lifetime. Finally, a factor of /g68/g21= 0.8 is used to account for the softer 440C stainless steel material that the  balls are made of. With all of the constants determined, the expected lifetime of the bearing could then be calculated, as  presented in Equation 6. /g2168=/g2183 /g2778/g2183/g2779/g4672/g2159 /g2172/g2200/g4673/g2780 =13068 [x 1 million] (6) The expected lifetime of the bearing under operation in a 1- /g42environment is 13 billion revolutions, which  is three orders of magnitude larger than the 12-month operational objective of 31.5 million revolutions. Thermal Validation Thermal-vacuum testing was also used to validate the hardware. A two-day test was conducted on the  flight-like Engineering Test Unit in order to characterize the performance of the full scanner assembly  over predicted operating temperature ranges, and also verify the workmanship of the individual parts  used in the assembly. Thermal modeling analyses, performed in Thermal Desktop, showed a predicted operating temperature  range of -5 to +25 °C for a circular 450-km orbit at 52° inclination. This conflicted with the 0 to +70 °C  operational temperature range of the MicroE 1500V encoder sensor as specified by the manufacturer. Therefore, a major goal of this thermal-vacuum test was to qualify the encoder for colder temperatures in order to gain margin from the manufacturer’s specifications. A procedure for the cold operating temperature qualification is shown in Figure 10.  533
Figure 10. Thermal-vacuum testing procedure The encoder performance was evaluated as a function of temperature by routing the encoder sensor’s  analog output to an oscilloscope outside of the thermal-vacuum chamber, and monitoring changes in signal strength (measured as voltage amplitude). As specified by the manufacturer, the signal strength is deemed nominal around 800 mV, and considered acceptable above 400 mV. During the test, the encoder was operational through a temperature range of -8.5 to +61.5 °C, adding  margin beyond the predicted operating temperature range. Though the encoder was subjected to temperatures colder than its specified operating range, the signal strength stayed within the acceptable range (shown in Figure 11). This provided the t eam with confidence in the enc oder’s performance at  colder-than-specified temperatures for short periods of time (< 3 hours, or 2 orbits). Figure 11. Encoder signal strength over expected operating temperature range 534
The overall scanner assembly performance was evaluated as a function of temperature by monitoring the  frequency of the encoder sensor’s analog output. The scanner assembly was set to spin at 60 rpm, or  7200 counts/sec (based on the 7200 gratings on the encoder disk). As set in our design requirements,  this frequency should be maintained within 5% of nominal, or 360 counts/sec at 60 rpm. If one of the parts were to fail, expand, contract, or degrade as a function of temperature, this could be seen in the frequency of the encoder signal. Over the tested temperature range of -8.5 to +61.5 °C, the encoder frequency stayed well within the 5% requirement even when tested beyond the predicted operating temperature range (shown in Figure 12),and did not show any trend as a function of temperature. This provided the t eam with confi dence in the  scanner assembly’s design and performance for operation in the predicted space environment.  Figure 12.  Scanner assembly performance over expected operating temperature range Workmanship, however, was deemed to be an issue after the epoxy bond between the glass encoder  disk and titanium mounting hub failed during bench testing several weeks after the thermal-vacuum test. Post-failure inspection determined that the epoxy did not successfully bond to the smooth glass surface of the encoder disk. As a result, greater care was taken in the surface preparation and epoxy application processes of mounting the encoder disk to the titanium hub, with the help of product engineers at MicroE Systems. The new unit was thermal-cycled and successfully tested for bond strength up to 111 N (25 lbf)prior to re-integration into the flight-model scanner assembly.  Conclusions  The MicroMAS scanner assembly is a novel new mechanism designed to provide satellites utilizing the  CubeSat form factor with dual-spin functionality. The utilization of multiple COTS components allowed for  successful development of a flight assembly that satisfies all of its subsystem requirements under a condensed timeline. The scanner assembly is capable of producing a continuous torque of 250 mN-mwhile providing sub-arcminute angular position knowledge, and allows for the transmission of 12 electrical  signals between the stationary and rotating hardware. Testing has shown that the scanner assembly can functionally operate in a vacuum environment and withstand temperatures in the range of -10 to +70 °C. The scanner assembly mechanism expands the possible science missions that can be accomplished with  the low-cost CubeSat platform. After successfully passing verification of the design and performance  requirements and undergoing final integration and testing within the full space vehicle, the scanner  535
NASA/CP--2000-209895 34thAerospace Mechanisms Symposium Compiled by:Edward A.Boesiger Proceedings ofasymposium heldat TheGreenbelt Marriott Hotel, Greenbelt, Maryland Hosted bytheGoddard SpaceFlightCenter andCo-sponsored bythe Aerospace Mechanisms Symposium Committee and Lockheed Martin SpaceSystems Company May10-12, 2000 May2000
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PREFACE TheAerospace Mechanisms Symposium (AMS)provides auniqueforumforthoseactive inthedesign, production anduseofaerospace mechanisms. Amajorfocusisthe reporting ofproblems andsolutions associated withthedevelopment andflightcertification ofnewmechanisms. TheNational Aeronautics andSpaceAdministration andLockheed MartinSpaceSystems Company (LMSSC) sharetheresponsibility fororganizing the AMS.Nowinits34thyear,theAMScontinues tobewellattended, attracting participants fromboththeU.S.andabroad. The34thAMS,hostedbytheGoddard SpaceFlightCenter(GSFC) inGreenbelt, Maryland, washeldMay10,11and12,2000.Duringthesethreedays,34paperswere presented. Topicsincluded deployment mechanisms, bearings, actuators, pointing and opticalmechanisms, SpaceStationmechanisms, release mechanisms, andtest equipment. Hardware displays duringthevendorfairgaveattendees anopportunity to meetwithdevelopers ofcurrentandfuturemechanism components. Thehighqualityofthissymposium isaresultoftheworkofmanypeople,andtheirefforts aregratefully acknowledged. Thisextends tothevoluntary members ofthesymposium organizing committee representing theeightNASAfieldcenters, LMSSC, andthe European SpaceAgency. Appreciation isalsoextended tothesession chairs,theauthors, andparticularly thepersonnel atGSFCresponsible forthesymposium arrangements and thepublication oftheseproceedings. Asincerethankyoualsogoestothesymposium executive committee atLMSSC whoisresponsible fortheyear-to-year management ofthe AMS,including paperprocessing andpreparation oftheprogram. Theuseoftradenamesofmanufacturers inthispublication doesnotconstitute anofficial endorsement ofsuchproducts ormanufacturers, eitherexpressed orimplied, bythe National Aeronautics andSpaceAdministration.
CONTENTS Symposium Schedule ..................................................................................................................... viii Symposium Organizing andAdvisory Committees ............................................................................. xii ALowPowerCryogenic Shutter Mechanism forUseinInfrared Imagers ................................................ 1 ScottSchwinger &ClaefHakun ACISDoorFailureInvestigation andMitigation Procedures ................................................................. 15 William Podgorski, NeilTice&PaulPlucinsky Designing Spacecraft Mechanisms forSand,Dust,WaterandFrost................................................... 29 KurtLankford Low-Cost Ejection Mechanism forProtective CoversonMeteosat Second Generation .......................... 39 JavierOrtega, Fernando Quintana &Gabriel Ybarra Misconceptions inMechanism Reliability .......................................................................................... 51 WalterHolemans &DonaldGibbons Lessons Learned fromtheManufacturing andTestoftheInternational SpaceStation DriveLockAssembly ...................................................................................................................... 61 CurtisAIImon Development ofanAutomated BoltDriverfortheSpaceStationTrussAssembly .................................. 75 Richard Meinhold &KenSeidner DesignofaPower-Assisted Spacesuit GloveActuator ....................................................................... 89 Russell Howard EnergyAbsorber fortheInternational SpaceStation........................................................................ 103 BobWilkes&LoraLaurence Friction DriveCharacteristics fortheSpaceStationMobileTransporter ............................................. 117 Richard Hughes, DanielHoyt&JuanCarreras Development ofaCryogenic Nanometer-Class Repeatability LinearActuator ...................................... 131 RubenNalbandian &AlsonHatheway Application ofMagnetic SmartMaterials toAerospace MotionControl............................................... 137 ChadJoshi&BruceBent ASeparation System Solution fortheInterimControl Module ........................................................... 143 BrianWhalen TheDevelopment ofanUltraLowShockPyrotechnically Actuated Separation Nut............................. 151 FredSilverman &JerryO'Quinn V
Piezoelectric PumpDevelopment ................................................................................................... 157 FranklinSager MiniatureLinearActuator .............................................................................................................. 163 Arrnond Asadurian TheSandmeier FieldGoniometer: AMeasurement ToolforBi-Directional Reflectance ........................ 167 MarkTurner&JeffBrown MotorRedundancy Concept: Efficient orReliable? .......................................................................... 175 EricFavre Development ofaMiniature Two-Axis, Triple-Helmholtz-Ddven Gimbal............................................... 189 BozSharif,EdJoscelyn, BrianWilcox&Michael R.Johnson CoverDriveandLockRingMechanisms forGenesis ........................................................................ 199 LouiseJandura Landsat 7SolarArrayTesting Experiences ..................................................................................... 213 DanielHelfrich Deployment Dynamics ofSolarArrayswithaLink-in-Slot Hinge/Latch Mechanism ............................. 225 WalterDaniel&GoMichael Bidinger Inflatable BoomControlled Deployment Mechanism for theInflatable Sunshield InSpace(ISIS)FlightExperiment ................................................................ 231 George SapnaI!1,JohnFolke,Charles Sandy&DaveCardogan ANovelMechanism usingShapeMemory AlloytoDriveSolarFlapsoftheINSAT-2E Satellite ............ 239 N.Viswanatha &T.Murali Evolution ofanActuator Family...................................................................................................... 251 PaulStigell,PekkaKyrenius, TomiYlikorpi &KaiViherkanto Coaxial CableFailureinaSpacecraft Mechanism ........................................................................... 265 Michael Chiu TheSeaWinds Scatterometer Antenna Subsystem ontheQuikScat Spacecraft ................................ 275 BrianWorkman &EricSchwartzbaum Antenna Pointing Mechanism forSteerable SpotBeamAntennas ..................................................... 287 Manfred Schmid EarthScanner Bearing Accelerated LifeTest.................................................................................. 303 BrianDietz,StevenVanDyk &Roamer Predmore Measurements ofFriction Coefficients between Oxidized and Contaminated Surfaces inVacuum andinAir.................................................................................. 317 PeterFrantz&Stephen Didziulis Marshall SpaceFlightCenterHighSpeedTurbopump Bearing TestRig............................................. 331 Howard Gibson, ChipMoore&RobertThorn vi
Design,Development andTestingofUmbilical SystemMechanisms forthe X-33Advanced Technology Demonstrator ....................................................................................... 343 AlanLittlefield &Gregory Melton Development ofaFlywheel forEnergyStorageandAttitudeControl ofanExoatmospheric Interceptor ................................................................................................... 359 DougHavenhill, FrankNimblett, SteveNichols&JohnYoung v_
SYMPOSIUM SCHEDULE TUESDAY, 9MAY2000 6:30-8:30 CHECK-IN Greenbelt Marriott Annapolis Room WEDNESDAY, 10MAY2000 8:00 Wednesday Presenters' Breakfast 8:00 CHECK-IN ANDREFRESHMENTS 9:OO INTRODUCTORY REMARKS Stewart Meyers, HostChairman NASAGoddard SpaceFlightCenter, Greenbelt, MD StuartH.Loewenthal, General Chairman Lockheed MartinSpaceSystems Company, Sunnyvale, CA CENTER WELCOME AIDiaz,CenterDirector NASAGoddard SpaceFlightCenter, Greenbelt, MD 9:30 SESSION I-INSTRUMENTS/RELIABILITY DickCasper, Session Chair Aeroflex Laboratories, Farmingdale, NY •ALowPowerCryogenic Shutter Mechanism forUseinInfrared Imagers ScottSchwinger &ClaefHakun, NASAGoddard SpaceFlightCenter, Greenbelt, MD •ACISDoorFailureInvestigation andMitigation Procedures William Podgorski &PaulPlucinsky, Smithsonian Astophysical Observatory, Cambridge, MA; andNeiiTice,Lockheed MartinSpaceSystems Company, Denver, CO •Designing Spacecraft Mechanisms forSand,Dust,WaterandFrost KurtLankford, Starsys Research Corp.,Boulder, CO •Low-Cost Ejection Mechanism forProtective CoversonMeteosat Second Generation JavierOrtega, Femando Quintana &GabrielYbarra, SENER Ingenieria ySistemas, S.A.,Spain •Misconceptions inMechanism Reliability WalterHolemans, Planetary Systems Corporation, Washington, D.C.&Donald Gibbons, Lockheed MartinSpaceSystems Company, Sunnyvale, CA I2:00 LUNCH BREAK 1:15 SESSION II-SPACE STATION Manfred Schmid, Session Chair DaimlerChrysler, Friedrichshafen, Germany •Lessons Learned fromtheManufacturing andTestoftheIntemational SpaceStationDrive LockAssembly CurtisAIImon, Lockheed MartinSpaceSystems Company, Sunnyvale, CA •Development ofanAutomated BoltDriverfortheSpaceStationTrussAssembly Richard Meinhotd, Moog,Inc.,EastAurora, NY;&KenSeidner, TheBoeingCompany, Huntington Beach,CA •DesignofaPower-Assisted Spacesuit GloveActuator Russell Howard, University ofMaryland, College Park,MD
•EnergyAbsorber fortheInternational SpaceStation BobWilkes,Lockheed MartinSpaceOperations, Houston, TX,&LoraLaurence, NASA Johnson SpaceCenter,Houston, TX •FrictionDriveCharacteristics fortheSpaceStationMobileTransporter RichardHughes&DanielHoyt,TRWAstroAerospace, Carpinteria, CA;andJuanCarreras, The BoeingCompany, Huntington Beach,CA 3:45 BREAK 4:00-5:00 SESSION III-POSTER PREVIEWS Christopher Hansen, SessionChair NASAJohnson SpaceCenter •Development ofaCryogenic Nanometer-Class Repeatability LinearActuator RubenNalbandian, Moog,Inc.,Chatsworth, CA,&AlsonHatheway, AlsonE.Hatheway, Inc., Pasadena, CA •Application ofMagnetic SmartMaterials toAerospace MotionControl ChadJoshi&BruceBent,Energen, Inc.,Billerica, MA •ASeparation SystemSolution fortheInterimControlModule BrianWhalen, U.S.NavalResearch Laboratory, Washington, D.C. •TheDevelopment ofanUltraLowShockPyrotechnically Actuated Separation Nut FredSilverman &JerryO'Quinn, Hi-Shear Technology Corporation, Torrance, CA •Piezoelectric PumpDevelopment FranklinSager,Oceaneering SpaceSystems, Houston, TX •Miniature LinearActuator ArmondAsadurian, Moog,Inc.,Chatsworth, CA 6:00-10:00 RECEPTION &DISPLAYS ATTHEGREENBELT MARRIOTT PosterPapersandinvitedcomponent andsoftware vendorsdisplaycurrentproducts and providetutorials. THURSDAY, 11MAY2000 7:00 Thursday Presenters' Breakfast 8:00 SESSION IV-INSTRUMENTS &ACTUATORS WillieBlanco,SessionChair NASAGoddard SpaceFlightCenter,Greenbelt, MD •TheSandmeier FieldGoniometer: AMeasurement ToolforBi-Directional Reflectance MarkTurner&JeffBrown,NASAAmesResearch Center,Mountain View,CA •MotorRedundancy Concept: EfficientorReliable? EricFavre,ETELAerospace Inc.,Boulder, CO •Development ofaMiniature Two-Axis, Triple-Helmholtz-Driven Gimbal BozSharif&EdJoscelyn, Aeroflex Laboratories, Farmingdale, NY;andBrianWilcox&Michael R.Johnson, JetPropulsion Laboratory, Pasadena, CA •CoverDriveandLockRingMechanisms forGenesis LouiseJandura, JetPropulsion Laboratory, Pasadena, CA 10:00 BREAK ix