ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
256 ^ 

Figure 14 .  ISL APM  EM final Vibration Test Results  – Transfer Function   (analysis in blue vs. test in pink) 

Thermal Vacuum Test   The described non conformance for the vibration test delayed the Thermal Vacuum (TVAC) test . TVAC  test is about to be started, results will be part of the AMS presentation.  During the TVAC test , a nonoperational cycle will be performed, followed by a set of seven operational cycles.  During the seven  operational cycles the EM will be driven with a standard pointing profile extracted from the global pointing  budget. As the ISL will also be activated during transient phases in space,  this test is of major importance  to guarantee the functionality under ex treme thermal conditions. At dedicated inspection points the  starting current and the power consumption will be measured in hot and cold conditions. These  measurements will  also verify the bearing characterization breadboard tests.  

Temperature OutgassingTV-Chamber Evacuation>3h >1h TV-Chamber Repressurization+65°C +23°C -25°C Performance Test>1h >1h >1h >1h >1h Performance TestPerformance Test +75°C >1h >3h-45°C Performance TestPerformance TestPerformance Test >1h >1h >1h >1h >1hISL ASM HRM releasePerformance Test TV-Chamber Repressurization TV-Chamber EvacuationISL ASM HRM locking>1h >1h

Figure 15 .  ISL APM EM TVAC cycles   X-Axis Y-Axis  Z-Axis
257 Microvibration Test   The micro vibration test was successfully performed in the mechanisms group test facility . The test setup  was developed and continuously  improved during the last years  (principle shown in Figure 16). The micro   vibration test sensors  logs the forces acting over a time period. This data is subsequently analyzed using  a Fast Fourier Transformation algorithm to identify the frequencies and amplitudes of the emitted  vibrations. Details regarding the analysis method can be found in [ 2]. 

Figure 1 6.  ISL APM EM Microvibration Test Setup  

Both axes were tested separately  with different  velocities from 10 deg/sec to 90 deg/sec. The test result s  for the rotation of the a zimuth axis at 70 deg/sec can be seen in Figure 17 . A second test method was to  perform a motion profile with both axes operating in parallel . 

Test result s evaluati on showed that the required microvibration levels below 0.1 N and 0.1 N -m from   0-2000 Hz could not be met  with the current stepper motors and motion profiles.    Additionally , an impact from the COTS e lectronic s was seen. Tests comparing a COTS  electronics  and  the flight  electronics showed an improvement by a factor of 2 for the flight  electronics.  Dedicated micro  vibration reduction designs  were developed at Airbus DS [ 3] in the past . Those design improvements can  significantly reduce the microvibration. Further technical optimizations  to reduce microvibration at  spacecraft  level are ongoing at Airbus DS GmbH.  

Magnetic Moment Test   The Magnetic Moment Test is used to identif y the influence of the magnetic components installed within  the ISL APM EM on the satellite. R esults of the EM can also be used to optimize the position of the APMs  at spacecraft level . The two major sources of magnetic field emission are the stepper motors.    In order to identify the field emission the ISL APM EM was installe d in a Helmholtz Facility . Firstly , the  remanent (non- powered) magnetic field emission was measured (360 measurements around the ISL  APM EM). Secondly , the field was measured in powered condition (standard motion profile).  

Amplifier   EGSE    Control Laptop   APM   XPC Data  Acquisition   Kistler Table   Mechanical Isolation   Vibration Adapter
258   Figure 1 7.  ISL APM EM Microvibration Test Results for Azimuth Stage rotating at 70 deg/sec  

Figure 1 8.  ISL APM EM mounted in the Helmholtz facility  

In Figure 19 an exemplary AC field measurement of the powered status is shown. It can be seen that the  actual change of position of the motors towards each other has a more important effect on the AC field,  than the actual powered coils.  The motor current shows just a minor effect on the AC field (see e.g. ,  marked peak),  while changed motor orientations lead to significant  changes of the AC field of up to 280  nT.   10010210-5100Force in [N] Frequency in HzResidual Force Spectrum Fx 10010210-5100Force in [N] Frequency in HzResidual Force Spectrum Fy 10010210-5100Force in [N] Frequency in HzResidual Force Spectrum Fz10010210-610-410-2100Torque in [Nm] Frequency in HzResidual Torque Spectrum Mx 10010210-610-410-2100Torque in [Nm] Frequency in HzResidual Torque Spectrum My 10010210-610-410-2100Torque in [Nm] Frequency in HzResidual Torque Spectrum MzENERGY: RMS = 3.6708 N and STD = 2.4647 N and AMPLITUDE: Mean = 0.0051732 N and Max = 1.2224 N at 16.5998Hz ENERGY: RMS = 3.051 N and STD = 2.69 N and AMPLITUDE: Mean = 0.004955 N and Max = 1.0883 N at 60.7994Hz ENERGY: RMS = 45.8118 N and STD = 1.3815 N and AMPLITUDE: Mean = 0.0038868 N and Max = 1.8626 N at 276.6972HzENERGY: RMS = 3.4492 N and STD = 2.7444 N and AMPLITUDE: Mean = 0.0017662 N and Max = 0.59225 N at 16.5998Hz ENERGY: RMS = 4.7481 N and STD = 2.7929 N and AMPLITUDE: Mean = 0.0018038 N and Max = 0.76115 N at 16.5998Hz ENERGY: RMS = 0.73947 N and STD = 0.62162 N and AMPLITUDE: Mean = 0.0010917 N and Max = 0.85138 N at 276.6972Hz Helmholtz facility   APM on magnetic  moment test adapter
259  Field in nT   0     20     40     60     80   100   120  200 

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0 

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-400  Time in sec  To calculate the Magnetic Moment , the actual field measurements will be used to calculate ideal dipoles  at the position of the motors. Using this approximation enables the calculation of the Magnetic Moment of  these dipoles.          By Bz    

Figure 1 9.  ISL APM EM exemplary AC field measurement results  

Evaluating the results of the measurements, show that a non- powered Magnetic Momentum of 2.2 Am²  can be calculated. Variation of the magnetic moment is given over the motion of  the motors towards each  other. The maxi mum calculated Magnetic Moment is at 4.2 Am² and the lowest calculated Magnetic  Moment is at 2 Am².  

Measurement  and calculat ion does not meet the required values of 0.2 Am² . The high Magnetic Moment  is caused by  the motor size and the mass and force of the magnetic material used in the motor.  The  current situation could be improve by changing the motor size or materials as well as shielding.  But  further discussions with the contractor showed that the stated non-conformance report  can be accepted.  

Conclusions  

General   It can be stated that the chosen technical solution fulfills all major requirements. The required orbit lifetime  of 12 years could be proven by an accelerated life test. The pointing performance is well within the needed requirement an d the RF performance allows a successful communication. In order to fulfil the  torque budget a powerful stepper motor was chosen. This stepper motor leads to certain violations of requirements like the magnetic moment and the microvibration.    Conclusion af ter all tests is that the ISL APM EM is successfully qualified (status after vibration test) . All  non-conformance report s for the current Engineering Model phase have been closed successfully and  some further improvements for the APM have been identified to increase the performance and to gain some mass reductions.    Airbus DS Friedrichshafen delivers the ISL APM EM to Airbus DS Portsmouth after just 18 months  of  design, manufacturing and assembly/integration/test  activities including a dedicated life test . 

   Field in nT   0     20     40     60     80   100   120  50 

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261 Design Development of a Combined Deployment and Pointing System for the International Space  Station Neutron Star Interior Composition Explorer T elescope  

Jason Budinoff*, Keith Gendreau*, Zaven Arzoumanian*, Charles Baker*, Robert Berning**,   Todd Colangelo *, John Holzinger **, Jesse Lewis *, Alice Liu *, Alissa Mitchell *,   Chuck Monroe* , Richard Pugh**  and Roger Wil ley** 

Abstract    This paper describes the design of a unique suite of mechanisms that  make up the Deployment and  Pointing System (DAPS) for the Neutron Star Interior Composition Explorer ( NICER/SEXTANT )  instrument, an X -Ray telescope, which will be mounted on the Internati onal Space Station (ISS). The  DAPS system uses four  stepper motor actuators to deploy the telescope box, latch it in the deployed  position, and allow it to track sky targets . The DAPS gimbal architecture provides ful l-hemisphere  coverage, and is fully re- stowable. The compact design of the mechanism allowed the majority of total  instrument volume to be used for science. Override features allow DAPS to be stowed by ISS robotics.  

Figure 1.  NICER/SEXTANT  in the stowed and deployed configurations      NICER/SEXTANT  Instrument   The Neutron Star  Interior Composition Explorer (NICER) / Station Explorer for X -Ray Timing and  Navigation (SEXTANT) is a deployable, clustered array of 56 soft X -ray telescopes that will be installed  on the I nternational Space Station . The telescope cluster will track astronomical targets with a 3σ error of  <0.6 milliradian  (2 arcminute) of pointing error . NICER/SEXTANT  will also demonstrate pulsar -based  celestial navigation, and will be used as a receiver for experimental X -ray communications.                                                               * NASA Goddard Space Flight Center, Greenbelt, MD   ** Moog, Inc., Chatsworth, CA  

Proceedings of the 43rd Aerospace Mechanisms Symposium, NASA Ames Research Center, May  4-6, 2016
262 ISS Mission Implemen tation  

NICER/SEXTANT  will be delivered to ISS via the SpaceX Dragon vehicle in 2016, and has a mission  lifetime of at least 2 years. During that time, the instrument will require 2- axis, full hemisphere pointing at  >95% duty cycle, and many stow/ deploy cycles . The instrument utilizes the Active Flight Releasable  Attachment Mechanism (AFRAM) interface and will b e robotically installed at Express Logi stics Carrier #2  (ELC2), Site 7, shown in Figure 2. 

Figure 2. NICER/SEXTANT  will be mounted to ELC2, approximately where the EVA   crew member is shown.  

Deployment and Pointing System (DAPS) Mechanism   As an AFRAM -based payload, the NICER/SEXTANT  instrument was subject to a limited launch envelope.  It was important for the mechanism architecture to be compact, which allowed the majority of volume to be dedicated to science. An optimal architecture utiliz ing four actuators was devised that combined the  deployment and pointing functions; this suite of mechanisms is called the Deployment and Pointing System (DAPS). It is similar to the architecture used to deploy and point the communications antenna on the J apanese Experiment Module [1].  

Architecture   The DAPS system is comprised of a deployment actuator, a deployment latch actuator, with an azimuth actuator and elevation actuator arranged in a biaxial gimbal. The gimbal actuators are oriented in an elevation over azimuth configuration. The gimbal  actuators and deployment actuators are separated by a  275-mm boom. The deployment actuators, at the base of the boom, are mounted to the AFRAM plate.  The gimbal  actuators are connected to the large X -Ray Timing Instrument (XTI) Box, which houses the X - Ray teles cope array. The XTI has a mass of 172 kg, and DAPS must point it at rates of approximately   1 degree/sec during slew and less than 0.4 deg/sec during science target tracking. Four frangible bolt  launch l ocks constrain the XTI box directly to the AFRAM/Inst rument Plate, which greatly reduces the  launch loads that the DAPS will be subjected to.  

Deployment   After the NICER/SEXTANT  instrument is installed on the ELC, the launch locks are released and the  DAPS deploys NICER/SEXTANT  from the stowed configuration into the operational configuration. The  deploy actuator rotates the DAPS boom, pointing actuators, and XTI 84 degrees, clearing the launch lock towers and placing the XTI box above the ELC2 structure. The deploy latch actuator then inserts a pawl
263 into the rotor of the deploy actuator, mechanically latching it in place. The azimuth actuator then slightly  rotates the XTI box to clear a hardstop feature, called the elevation hook. The hook is a simple linear bar  connected to the boom. When the XTI is stowed, the hook is within a small ring on the XTI. This mechanically constrains the elevation from inadvertently shifting while the deploy actua tor rotates  due to  the large center  of gravity  offset of the XTI relative to the elevation axis. While the elevation actuator  detent has positive margin to resist this inertial torque, the hook was implemented to provide positive mechanical constraint resulting in an additional layer of safety. It ens ures that the XTI box cannot swing  out during deployment and potentially cont act the ISS solar arrays, less than 0.5 m away.    With the hardstop feature cleared, the elevation actuator rotates the XTI box up in elevation, resulting in NICER/SEXTANT  fully de ployed and ready to begin science operations . The azimuth and elevation  actuators in the pointing gimbal then slew to and track various astronomical targets, typically observing 2 to 3 different targets per orbit. Continuous tracking of targets through zenith is interrupted through the   10-degree keyhole; relatively rapid azimuth slewing enables these target s to be immediately  re-acquired  as they cross out of the keyhole.  The volume swept out by all degrees of freedom is shown in Fig.  3.  

Figure 3.  The operational swept volume NICER/SEXTANT  is shown. The clearance between the  swept volume  the starboard solar array is also illustrated.  

Re-Stowage  The NICER/SEXTANT instrument must occasionally return to the stowed configuration for ISS maneuvers, crew  Extra -Vehicular Activity (EVA), local Extra- Vehicular Robotic (EVR) operations, and  visiting vehicle proximity operations. The instrument also must be stowed prior to robotic removal and disposal. The DAPS stow process is the exact reverse of the deployment process. To increase the stowed stiffness after the initial (and permanent) release of the XTI launch locks, shear -bearing indexing  conical surfaces on the AFRAM plate must engage their corresponding cups on the XTI. The alignment of these mating surfac es is critical to NICER/SEXTANT meeting its  stowed stiffness requirement. The DAPS  ensures that these surfaces align correctly prior to mating by using a precision guide channel which forces the relatively coarse positioning ability of the actuators into a repeatable precision trajectory which  guarantees proper alignment upon contact. A guide pin on the XTI engages a precision slot on the DAPS  bracket near the deploy actuator as the DAPS deploy actuator rotates through TBD degrees, nearing the  final stowed hardstop. Azimuth position errors are corrected by the stow alignment slot. The torque resulting from engagement of the guide slot will backdrive the azimuth actuator into the correct orientation for engagement. The elevation axis is limited by hardstops w ithin the actuator at 0 (stowed, end of travel)  and 172 (maximum elevation, end of travel) degrees.
265 Detent Torque   The DAPS vendor recommended a reduction in the detent torque in order to further reduce the pointing  errors . The detent torque was reduced by increasing the air gap between  the rotor and stator . This also  reduced the motor constant, and more power would be required to maintain torque output, but this was acceptable. This was tested on a development motor, which was used to determine the air gap size which balanced detent tor que and motor constant . The same air gap modification procedures demonstrated on  the test motor were used to manufacture the NICER/SEXTANT engineering test unit and flight actuators .  The NICER/SEXTANT pointing actuator detent torque levels were lowered to less than 0.03 N -m at the  output . The detent torque reduction had reduced the overall stepper -induced pointing error . Unfortunately,  when the stepper actuator rates crossed the low frequency structural modes, the pointing error level still exceeded the requirement .   

Tuned Mass Dampers   One of the techniques for reducing vibratory motion is the tuned mass damper (TMD) . The TMD is a  resonant device designed to dissipate vibrational energy from the structure. They are passive mechanical  devices that do not re quire power or electronics . The TMDs are particularly effective for  NICER/SEXTANT, as they suppress low -frequency vibration excited by both stepper and ISS -induced  disturbances . The NICER/SEXTANT TMDs were designed to provide 1- 3% damping ratio for the fir st  three flexible- body modes . They are oriented to reduce vibrations along the axes where the vibrational  displacement occurs. The mounting locations of TMDs are shown in Figure 4. A sample of stepper - induced jitter results for the performance with and w ithout TMDs are shown in Fig.  5. 

Figure 4.  Tuned Mass Dampers, shown in green , were added to the corners of the XTI box  

Figure 5.  The elevation actuator exceeds the 66 arcsecond requirement  when step rates   are near 10 Hz  

0 2 4 6 8 10 1210-1100101102103 Step rate [Hz]Detector LOS [RSS(x,y), asec 0-P]CDR Runs    El w/TMD Radial  43.84 asec 0-P El no TMD Radial 201.63 asec 0-P 66.0 asec 3-sigma
266 As shown in Figure 5 , The TMDs were critical in meeting the pointing performance requirement . They  greatly reduced the structural vibrations when the step rates pass through the structural mode  frequencies . Similar results are shown in Figure . 6 for the ISS -induced jitter, where th e TMDs not only  reduced the peak jitter, but also stabilize the pointing performance due to ISS disturbance frequency variations.  

Figure 6.  Addition of the TMDs reduced the overall predicted pointing  error (Radial LOS in the  graph) by a factor of 2.7  

A combination of relatively low -cost techniques that included microstepping, low -detent torque, and TMDs  were used to improve the NICER/SEXTANT pointing performance and meet the NICER/SEXTANT pointing requirement with adequate margin. Several tests (see Fi g. 10) are currently underway to validate  the pointing performance on the ground prior to launch.  

ISS Requirements Effects    The DAPS had to meet additional requirements from the space station program. C ontingency  considerations required that the NICER/SEXTANT  instrument have interfaces to the ISS Special Purpose  Dexterous Manipulator (SPDM) teleoperated robot. These interfaces would allow the SPDM to return the instrument to the stowed configuration if DAPS failed.  This is necessary as the payload must be in the  stowed configuration to fit within the disposal vehicle.  

Robotic Interfaces   Each of the four actuators has a standard 7/16- inch (1.1-cm) EVR torque bolt coupled to the motor shaft.  These bolts serve as an interface through which torque can be applied to drive  the motor shaft, which is  amplified by the harmonic drive gearing and rotates the actuator output. The torque bo lts reside within  Micro -Conical Fixture (MCF) Standard D exterous Grapple Fixtures  with their associated Modified  Truncated Cone (MTC) visual alignment targets. These  EVR interfaces  are shown on the engineering test  unit (ETU) actuator in Figure 7.   These bolts and EVR interfaces are on the motor faces of all DAPS actuators except the Azimuth pointing  actuator. This actuator has the motor face buried within the boom and required a 90 degree gear box  extension to allow the EVR bolt, MCF, and MCT  to be accessible by SPDM. These interfaces are sho wn  in Figure 8.
267  

Figure 7.  Robotic interfaces are shown on the DAPS ETU actuator.  The black MTC target sits  above the MC F grapple fixture on the actuator face.  The EVR bolt is buried within the MCF.  

Figure 8. DAPS ISS Robotic Interfaces. Each actuator has a MCT  grapple fixture, a MTC visual  target, and external labels.    To accommodate an accidental  runaway impact of the SPDM robot, DAPS had to demonstrate fracture  margin against an impact of 556 N ( 125 lbf) , and an EVA kick load of 556 N ( 125 lbf).  

Torque Limiting   The minimum input from the SPDM torque driver is 6.7 N -m (5 ft-lbf). Should this amount of torque drive  a  DAPS actuator into a hardstop, internal damage may occur . A torque- limiter w as built into the EVR  interface to prevent accidental damage. T orque levels of 0.73 N -m (~6.5 in- lbf) are transmitted through  the tor que limiter.  

External Markings   To facilitate the contingency robotic restow operation and meet Design for Minimum Risk  requirements,  external status markings and indicators  (arrows on moving tabs) were implemented which indicated the  displacement position of each axis. Each actuator had gradations labelled  with return to stow position
268 EVR turn counts, and arrows indicated the direction to rotate the EVR bolt to return each axis to the  stowed position. The location of these gradations and markings were unique for each axis. To meet EVA legibil ity requirements, the text was ~6.7 mm (18 pt. ) or larger.  

Verification    The DAPS is being developed as a protoflight unit, with a single ETU actuator to validate low detent  modifications  and EVR interfaces. The ETU actuator is shown in Fig.  7. 

Figure 9 .  DAPS flight unit as delivered to the NICER/SEXTANT  instrument.    As of this writing the DAPS has been delivered to GSFC and is  undergoing performance testing as shown  in Fig. 10. This is to be followed by integration into the NICER/SEXTANT  instrument  and environmental  testing at the instrument level.  

Figure 10.  DAPS flight unit during performance testing with an XTI inertia simulator floating  on air   bearings.  

References  

1. Stevenson, J., Katsuyama, Y., Fukatsu, A., Kurihara, K., Saito, T., Zoren, M., Mallonee, S., (1999).  Development of the JEM ICS Antenna Pointing System In Space Mechanisms and Tribology,  Proceedings of the 8th European Symposium (Ed. D. Danesy), ESA SP- 438, European Space  Agency, Noordwijk, The Netherlands
270        

‘DARPA Phoenix  Camera Boom ’ ‘NASA Cube Sat Magnetometer  Boom ’  • 2.5” (6.35 -cm) diameter, circular cross  section   • 60” (1.5-m) deployed length  • Embedded wires allowing power and  data transmissions for distal camera  • External motor  • 0.75” (1.9-cm) diameter, circular cross  section   • 50 inch (1.3-m) deployed length   • Near -zero CTE design, high precision  distal end knowledge   • Cube Sat form factor   • Magnetically clean  

‘DARPA Phoenix HIMast  Boom ’ ‘NASA A steroid Redirect Mission’ (ARM)   • 1” (2.5-cm) diameter with extended cross  section   • 180” (4.6-m) deployed length   • Low CTE design, high precision distal  location   • Embedded high current power supply  wires   • Fits the NovaWurks  ‘HISat ’ envelope  with an  external drive  motor  • 1.4” (3.6-cm) diameter, circular cross  section with overlap   • 80 inch (2 -m) deployed length   • Low CTE design   • Internal motor  

Figure 1.  The Roccor High Strain Composite Slit- Tube Deployer Family.