ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
mirror. The polishing house is confident that these marks can be polished out of future mirrors.  Figure 11. Mounting Distortion Test, 0.68 N·m Torque Applied  The mirror was then cooled with liquid nitrogen. Testing the mirror over temperature showed that WFE is  dominated by the CTE mismatch of the VQ coating and the aluminum substrate producing bending in the  mirror at the low cryogenic temperatures. The aluminum shrinks more at cold temperatures due to its  larger CTE producing a slightly convex mirror. This bimetallic bending is much less than is typically seen with nickel-plated mirrors. Figure 12 shows how the mirror shape changed from concave at 292K to  convex at 186K. RMS WFE error measured at 292K and 186K were 0.039 and 0.106 waves HeNe  respectively.  Figure 12. Interferometric Measurements at 186K  Figure 13 shows the RMS, PV, and Power WFE from 292K to 141K. Using this data, the RMS WFE at the  required operating temperature of 210K is linearly interpolated to be nearly 0.083-wave HeNe.   The mirror retained a very good figure at 141K with a deformation shape nearly identical to the 186K  measurement and a 0.125-wave RMS WFE. At temperatures lo wer than 141K, the WFE did not increase  noticeably. Wave front measurements were measured as low as 92K with very similar performance as  seen at 141K. Transient thermal gradients affected the accuracy of this data. Quilting or rib print-through  was not detectable in the testing, even at temperatures as low as 92K. The mirror does however show substantial high order WFE.  394
The mirror surface figure exhibits  great stability under mounting and  cryogenic temperatures with some  expected and very small bimetallic bending of the optical surface. At the  210K operating temperature the RMS  WFE is near 0.083-wave HeNe.  Control System Performance Testing Considerable testing was performed  to quantify the control system performance. The real-time HIL tools  of the RPM process were used to test  and optimize the control laws for both  the elevation and azimuth control  loops. A graphical block diagram was used to monitor the performance and  change the control parameters “on the  fly” as the tests were running.  Tests performed included large and  small steps, and small-angle  frequency respon se. The frequency  response tests were performed using a control systems analyzer, which excited each axis with a sine  sweep of 100 arcsec amplitude and a frequency range from 1 Hz to 1000 Hz, and analyzed the resulting  waveforms for gain and phase shift. The analyzer was also used to test individual components and subsystems and thus obtain their open-loop transfer functions, which were then implemented in the  model.-0.6-0.5-0.4-0.3-0.2-0.100.10.20.30.40.50.6 0 50 100 150 200 250 300 350 Temperature (K)WFE Wa  HeNe (633nm) ves PV  RMS Power210 K Operating  Temperature Figure 13. Mirror WFE at Cryogenic Temperatures  As an example of model validation, Figure 14 shows the response of the azimuth axis to a 30-deg step  and compares it to the prediction from the simulation. The response is well behaved with little overshoot.  The average slew rate from 0 to 100% of the commanded angle is 164 deg/sec.  Performance Summary  Table 5 summarizes the performance  of the final prototype in key areas.   The mass total is only for the  mechanism and does not include the  electronics, since the electronics are  only breadboard at this stage of the  development.   The average power was determined  from the simulation by commanding  both axes to perform continuous  simultaneous slow scans, as would be  used for GMC. The peak power was  obtained by commanding both axes to  do a large-angle step simultaneously.   0   0.1 0.2 0.3 0.4 0.5 -505101520253035 Time (sec)muth Angle (deg)Azicommand data simulation Figure 14.  Azimuth Axis Large-Angle Step Response  395
-1200-1000-800-600-400-200020040060080010001200 11.4 11.5 11.6 11.7 11.8 11.9 12 12.1 12.2 12.3 Sensor Voltage (A+B, volts)Angle (arcsec )20.9 deg C 21.3 C 22.0 deg C 22.7 deg C 23.9 deg C 24.1 deg C 26.4 deg C 29.9 deg C 35.6 deg C 40.0 deg C Figure 16. Temperature Mapping  Launch Lock  Launch lock options included a permanent magnet, a  shape-memory alloy pin-puller, and a wax actuator. The  magnetic concept was chosen based on simplicity and a  6-year on-orbit heritage with another SDL scan mirror.  Figure 17. Magnetic Launch Lock The patent-pending launch lock is shown in Figure 17.  The combination of a magnet with a cone-shaped  receptacle provides restraint in all directions. The restraint force is adjustable by changing the air gap,  which is done by screwing the cup holding the magnet in  or out and locking it with the set screw. The holding force is high enough to restrain motion during external loads,  but is low enough that the actuator can overcome the  holding force to release the lock. For this particular  prototype, the latch force can be about 2.2 N maximum,  since it is limited by the maximum current (~1 amp) and torque of the voice coil actuator. Analysis was performed  to validate the design by applying the launch vibration  spectrum to the dynamic model of the FSM shown in Figure 9. Vibration Testing  A vibration test was  conducted to verify the  performance of the launch lock and the survivability of  the FSM. The 3-minute test  was based on proto-qualification levels for the  planned STP-SIV spacecraft, Table 6. STP-SIV Spacecraft Vibration Spectrum  Frequency (Hz) PSD – Acceptance (g2/Hz) PSD – Protoqualification  (g2/Hz)  20 0.005 0.01 100 0.012 0.024 800 0.012 0.024 2000 0.005 0.01 Overall (grms) 4.27 6.04 397
shown in Table 6, with the first minute at reduced power and two minutes at full power. The test consisted  of the following duty cycle:  /g120 -12 db for 15 sec  /g120 -9 db for 15 sec  /g120 -6 db for 15 sec  /g120 -3 db for 15 sec  /g120 0 db for 2 min.  The launch lock magnetic force was set to about  1.5 N for the first vibe test. The launch lock  released after 105 se c., but no violent mirror  motions were observed. Measurement of the latch force after the test showed it to be about 1.4 N, so  the set screw may have been loose. The latch  force was increased to about 2.2 N for the second test. This test was successful. Before and after  each test, the FSM was powered up and exercised  with its nominal duty cycle of a Lissajous figure, an “S” figure, and slow scans. The performance  appeared nominal before and after each test. The  FSM was also repeatedly commanded to latch and  unlatch positions, verifying that the voice coil could  overcome the latch force.   Figure 18.  FSM Mounted in Thermal Vacuum  Chamber  Thermal-Vacuum Testing  Although the FSM prototype was not intended for harsh te sting, it was mounted in one of SDL’s thermalvacuum chambers (Figure 18) and subjected to 55 duty cycles with pressures as low as 1e-7 torr and  temperatures to 164 K (the lower limit of the chamber’s capability). The azimuth axis worked for 26 runs as the temperature dropped, until the thermocouple on the az imuth motor housing reached 227 K. This  axis then ceased working as expected due to the bearing lube freezing up, since its published range is  233 K to 423 K. The elevation axis operated for all runs, down to 164 K. Figure 19 shows  how the elevation axis “rings” at cold  temperatures due to a change in the plant resulting in improper gains.  The feedthrough  cables add considerable noise, as shown for  small angles in Figure 20. The azimuth noise was noticeably worse than the  previous year, possibly due to bearing  abuse during the vibe test, or loss of lubricant during vacuum testing. The  encoder optics could also be impaired due  to outgassing on the lens. Figure 21 shows  how the “before” and “after” performance of  the elevation axis at ambient for large angles is nearly identical. Some outgassing  was evident from the residue collected in the  chamber’s cold trap, although it was not analyzed after testing.   9 9.2 9.4 9.6 9.8 10 10.2 10.4 10.6-6-4-20246 Time (sec)Elevation Angle (deg) Command Case 45, 1.2e-7 torr, 254 K Case 43, 1.1e-7 torr, 165 K Figure 19.  Effect of Temperature on Elevation Step  Response  398
Lessons Learned and Future Work  The FSM prototype has been shown  through testing to have the potential for  robust performance in the harsh environment of space. The elevation axis  operates at hard vacuum and cold  temperatures down to 164 K. The “before” and “after” ambient performance of the  elevation axis is virtually the same. The  azimuth axis works down to 233 K until the current bearing lube freezes up. The launch  lock successfully handles a typical launch  environment.   Obviously, some materials changes would  need to be made for space and/or cryogenic applications. The azimuth lubricant would  need to be changed. Also, the gains for both  axes should be scheduled based on  temperature. The cause and corrective action for the cable noise in both axes would need to be identified. 1.95 2 2.05 2.1 2.15 2.2 2.25-100-50050100150200250300350400 Time (sec)Elevation Angle (arcsec)Command Original 2006 data, ambient Case 52, ambient, original cables Case 51, ambient, feedthrough cables Case 35, 1.e-6 torr, 198 K, pump off Figure 20.  Effect of Cable Noise on Elevation SmallAngle Response  The elevation axis circuit design could be modified to allow a slightly higher current limit of about 2 amps  to provide additional latch force and stall torque margins.  The focus of future efforts are being centered more on the development of space-type electronics, as  described in the following paragraphs.  The control algorithms and filters need to be further refined and optimized to reduce the noise, especially  in the elevation axis. This may be done by moving the derivative term to the feedback loop, using optimal  state-space methods with observers, etc.  The algorithms need to be implemented in low-cost space-qualified processing hardware, such as an  FPGA, DSP, or microprocessor. This may necessitate conversion to fixed-point math, as was done for the  radiation-hardened fixed-point microprocessor used to control a FSM in another SDL space payload (Ref.  3). A protocol or set of operational codes needs  to be defined to communicate digitally with  the FSM, such as commanding it to move to  a position, querying it for the current  position, etc.  The D/As, A/Ds, wedge sensor electronics,  and quadrature encoder functionality need  to be implemented using space-rated components.  1.95 2 2.05 2.1 2.15 2.2 2.2501234567891011 Time (sec)Elevation Angle (deg)Command Original 2006 data, ambient Case 33, 1.9e-7 torr, 198 K Case 53, ambient, original cables Case 50, ambient, feedthrough cables Figure 21.  Ambient Data for Elevation Angle LargeStep Response The autocoding feature of the RPM process  needs to be evaluated to determine how much of the code automatically generated  for the control algorithms can be used in the  final product.  399
The input power interface needs to be defined to either use the spacecraft power to drive the motors or  else have a customer supply the motor power.   Off-the-shelf space-rated three-p hase motor amplifiers need to be evaluated.  It would also be desirable to develop an automated procedure for mapping the wedge sensor output  against temperature and angle using optical techniques and precise temperature control.  Summary  A low-cost two-axis FSM has been successfully developed and demonstrated for air- and space-based  sensors. The 75-mm aperture mirror is lightweight, isolated by flexures, and maintains a surface figure of  less than 0.1-wave RMS down to a temperature of 210 K. The drive system uses an innovative  combination of off-the-shelf components to achieve large angles, high slew rates, high bandwidth, and relative position error less than 1 arcsec. A rapid prototyping methodology has been used to develop the  control laws. Temperature mapping has been employed to improve the absolute pointing accuracy. A  passive patent-pending launch lock has been developed and demonstrated. Vibration testing has been  performed to show the survivability of the FSM and the successful operation of the launch lock. The  mechanism has been subjected to many duty cycles in a hard vacuum and at low temperatures. Areas for improvement include a lower-temperature azimuth lubricant, gains scheduled with temperature, noise  reduction in the cables, a higher elevation axis current limit, and development of space-worthy  electronics.  Acknowledgments  The authors express their sincere appreciation to the following individuals for their significant  contributions: Richard Sanders and Melissa Draper, mechanical design; Trent Newswander, mirror design, analysis, and testing; Brent Jensen, structural analysis; Brian Thompson, trade studies and  conceptual design; Steve Dans ie, Andrew Little, and Scott Schicker, me chanical technologists; Duane  Miles, optical technologist; Jeff Blakeley, analog and power electronics; James Cook and Zach Casper, electronics and software; Adam Shelley and Quinn Young, thermal design; Aaron Gilchrist, temperature  mapping; and Jim Herrick, vacuum testing. The authors also appreciate the funding and support from the  Research Division of SDL, directed by Dr. J. Steven Hansen and Dr. Scott Jensen.  References  1. Steven R. Wassom, Morgan Davidson, Trent Newswander, James Cook, Zach Casper, Adam  Shelley, “Fine Steering Mirror for Smallsat Pointing and Stabilization,” 20th Annual AIAA/USU Conference on Small Satellites, Paper #SSC06-VIII-7, 17 August 2006.  2. Ahmad, Anees, Optomechanical Engineering Handbook , CRC Press, 1999.  3. Steven R. Wassom, Chad Fish, Mitch Whiteley, Dave Russak, Joel Nelsen, Brian Thompson, Glen  Hansen, Jason Wooden, Larry Gordley, John Burton, Mark Hervig, Paul Cucchiaro, Dan Hammerle,  “SOFIE Pointing Control System,” SPIE Proceedings Vol. 6297, Infrared Spaceborne Remote  Sensing XIV , 7 Sep. 2006.  400
Cryogenic Focus Mechanism for the Spitzer Space Telescope  William C. Schade* Abstract  A new focus mechanism was developed, tested, and flown for the Spitzer Space Telescope (“Spitzer”),  one of NASA’s “Great Observatories”. Figure 1 shows the Flight Focus Mechanism (FLT-FM), now in  Spitzer. The mechanism uniquely provides robust support and precise focus adjustment for the Spitzer secondary mirror, from 300 K to a 5 K cryogenic environment.  This paper summaries the requirements, performance, description, and testing of the focus mechanism,  including key component level tests of a geared-stepper motor and ball screw. Also, a secondary mirror  mount is described that minimizes mirror distortion and supports high loads. Several design and test  challenges were overcome and lessons l earned from this successful development include:  /g120 Titanium is useful as a flexure material to liquid helium temperatures.  /g120 Adhesive bonds at cryo-temperatures should be well understood and / or tested.  /g120 Geared-stepper motor and ball screw components were simply modified to work to < 5 K.  Figure 1. Flight Focus Mechanism (FLT-FM)                                                             * Ball Aerospace & Technologies Corp. (BATC), Boulder, CO  Proceedings of the 39th Aerospace Mechanisms Symposium, NASA Marshall Space Flight Center, May 7-9, 2008  401
Introduction  The Spitzer observatory shown in Figure 2 includes the Cryogenic Telescope Assembly (CTA) that directs  infrared signals to various instruments by means of a beryllium primary and secondary mirror. Ball Aerospace & Technologies Corporation (BATC) supplied the CTA as shown in Figure 3, with funding and  oversight provided by NASA-JPL. Early in the program it was decided a focusing capability was desired,  providing the cost and complexity to achieve it was reaso nable and providing it could be developed on  time. Fortunately these objectives were met, so the focus mechanism is included in the CTA as shown,  mounted on a beryllium metering structure.  Spitzer was launched on August 25 th, 2003 with the focus mechanism and other key telescope  components at ambient temperature. The telescope components were later cooled in space to less than  5.5 K. This space-assisted cooling approach was beneficial in helping to preserve Spitzer’s cryogen.  Consequently, the observatory has surpassed its expected 2.5-year life and is approaching a 5 year life  goal. Additional Spitzer facts are given on the next page, for reference.  0.85 m dia  PrimaryMirrorFocus Mechanism  (& Secondary Mirror)  Instrument  Chamber  Cryostat  (Liquid  Helium) Metering Structure Dust Cover  (ejected after deployment)  Solar Panel  and Sun  ShieldCryogenic Telescope  Assembly  (CTA)  Spacecraft  Bus  Figure 2.  The Spitzer Observatory                                                                                  Figure 3. Cryogenic Telescope Assembly (CTA)  402
Spitzer “Fast Facts” For reference, this page includes Spitzer “Fast Facts” from the California Institute of Technology web site  http://www.spitzer.caltech.edu/about/fastfacts.shtml . The Spitzer Space Telescope is a space-bor ne, cryogenically-cooled infrare d observatory capable of  studying objects ranging from our Solar System to the distant reaches of the Universe. Spitzer is the final element in NASA's, Great Observatories Program and an important sc ientific and technical cornerstone of  the Astronomical Search for Origins Program.  Launch Date: 25 August 2003  Launch Vehicle/Site: Delta 7920H ELV / Cape Canaveral, Florida  Estimated Lifetime: 2.5 years (minimum); 5+ years (goal)  Orbit: Earth-trailing, Heliocentric  Wavelength  Coverage:3 - 180 microns  Telescope: 85-cm diameter (33.5 inches), f/12 lightweight Beryllium, cooled to less 5.5 K  Diffraction Limit: 6.5 microns  Science  Capabilities:Imaging / Photometry, 3-180 microns  Spectroscopy, 5-40 microns  Spectrophotometry, 50-100 microns  Planetary Tracking: 1 arcsec / sec  Cryogen / Volume: Liquid Helium / 360 liters (95 Gallons)  Launch Mass: 950 kg (2094 lb) [Observatory: 851.5 kg, Cover: 6.0 kg, Helium: 50.4 kg, Nitrogen  Propellant: 15.6 kg]  Major Innovations  /g120 Choice of Orbit   /g120 Warm-Launch Architecture   /g120 New Generation of Large-Format Detector Arrays   /g120 Lightweight, cryogenic optics   The Spitzer Team  /g120 Jet Propulsion Laboratory   /g120 Spitzer Science Center, California Institute of Technology   /g120 Ball Aerospace and Technologies Corporation   /g120 Lockheed Martin Space System Company   /g120 Smithsonian Astrophysical Observatory   /g120 NASA-Goddard Space Flight Center   /g120 Cornell University   /g120 University of Arizona  403
Focus Mechanism Key Requirements and Performance  The purpose of the focus mechanism is to maintain the Spitzer’s secondary mirror position and, if desired,  move it along the CTA optical axis for focus adju stment. As a single axis device, it must maintain mirror  alignment with minimal de-center or tilt motion of the mirror due to launch, cool-down, or focus operation.  This basic functionality is reflected in the key requirements and tested performance of the mechanism, as  summarized in Table 1. And, since it must be controlled reliably and remotely, redundancy is required.  The performance results are for ambient and cryogenic temperatures, before and after exposure to  launch vibration levels, and after life testing of a Focu s Engineering Model (FM-EM) to 1X life. The FMEM was built and tested to reduce risk and for life testing up to 4X life, prior to making the Flight Focus  Mechanism (FLT-FM). The tests for the FLT-FM were more abbreviated, but performan ce was similar.   Table 1. Focus Mechanism Key Requirements and Performance   Key Requirements Tested Performance (Meets all)  Range /g116/g114 0.25 mm   /g100/g114 0.50 mm /g116/g114 0.25 mm (soft limits)  /g100/g114 0.33 mm (hard stops)  Step size /g100 2.5 /g80m /g100 1.3 /g80ma Repeatability/g100/g114 1.25 /g80m (unidirectional) /g100/g114 0.81 /g80ma (bi-directional) b De-center over range c/g100/g114 5.0 /g80m /g100/g114 3.61 /g80m Tilt over range c/g100/g114 58 /g80rad /g100/g114 41 /g80rad Shift after launch d /g100/g114 12.5 /g80m de-center  /g100/g114 116 /g80rad tilt /g100/g114 9.3 /g80m de-center  /g100/g114 110 /g80rad tilt  Operating temperature 300 to 2.5 K Meets Operating pressure Ambient to 10-6 Torr Meets Launch acceleration  70 G lateral  125 G axial Meets First mode frequency > 150 Hz 330 Hz  Clear aperture of SM /g135 120 mm min /g135 123.8 mm min  Mass (with mirror) /g100 4 kg Meets Notes pertaining to Table 1 : a) Mean and standard deviation values for these are given in the assembly test section.  b) Repeatability was met bi-directionally, while only unidirectional was required.    c) De-center and tilt over range refer to mirror motion al ong and rotation about any axis normal  to the optical axis, respectively, over the mechanism’s full range.  d)Shift after launch refers to mirror motion due to temperature repeatability and launch  vibration combined. The former is any difference in mirror position from ground alignment at 5 K and after cooling again on orbit. The latter is any permanent shift due to launch. Aside from the requirements, a goal was set to limit the focus shift of the mirror after launch to /g100 2.5 /g80m. This was a goal only since the mechanism can correct for focus after launch and it was anticipated  performance might slightly exceed this. Performance for this was measured at /g100 3.2 /g80m, which was just  over the goal, as was expected might happen.  The FM-EM also demonstrated a viable flight mechanism could be made on time. The FM-EM design was  started in November 1997, it was built, and cryo-testing was started by June 1998, as required. This  success paved the way for the build of the FLT-FM.    404
SM and carrier  mass simulator  NutFlexure Hard Stop  (1 of 2) LeverArm SM Cruciform  Flexure Drive  Flexure PivotFlexure LeverArmFM-EM Hardware  (with cover)  Drive  Housing  SM Flexure (not flight  design)  Stepper MotorGearhead Ball Screw  and Nut Main Housing  Figure 5. Focus Mechanism Engineering Model (FM-EM)  406
Flight Focus Mechanism (FLT-FM) Figure 6 shows the Flight Focus Mechanism, which is  much like the FM-EM except for some notable  developments. Most notable is the addition of the flight secondary mirror and its mount, shown in the next  section. Also, the motor is dual wound for redundancy.  The drive and pivot flexures were also significantly strengthened because the axial (focus) load  requirement was changed from 70 G to 125 G and a “low risk fracture part” methodology was voluntarily  imposed to enhance reliability. This methodology is defined in a JSC memo (June 1992), TA-92-013 and  included designing the flexures to /g116 10X their required fatigue life (including all vibe cycles).  Other developments included: The addition of the mirror mask and stray light baffle. A larger motor was  provided for more torque margin. And Variable Impedance Transducer (VIT) sensors were added for  “soft” limits.  VIT (1 of 2) SM Mask / Stray  Light Baffle  SM Figure 6. Flight Focus Mechanism (FLT-FM)  Secondary Mirror Mount The secondary mirror mount approach is shown in Figure 7. The mirror is a convex hyperbolic front  surface on a beryllium substrate. Beryllium spacers are bonded to the substrate and titanium bi-pod  flexures are then conventionally attached. This approac h helps to minimize mirror distortion and provides  high load capacity. It also presented challenges due the bonds being used at cryogenic temperatures.  407