ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
Titanium Bi-Pod  Flexure attaches  to spacer  Titanium Spacer SM Carrier Secondary  Mirror (SM) Beryllium Spacer bonded to SM  (The small post in middle of bond  area is not a load-carrying feature)  Figure 7. Secondary Mirror Mount  The beryllium spacers match the CTE of the mirror to minimize thermally induced strain in the bondline  and to maintain mirror figure. The bondline also pr ovides some isolation from  flexure fastener preloads,  which were light, to further minimize mirror distortion. Titanium spacers at the carrier interface were  machined in thickness and wedge to achieve ideal flexure alignment and for optimal mirror figure at 5K.  High load capacity was achieved with epoxy bonds to  beryllium (> three times better vs. titanium). Beyond  strength, a specific epoxy was also chosen for its desirable CTE and modulus of elasticity vs. other  adhesives, based on tests at room temperature to 5 K.   A challenge arose when finite element analysis (FEA) predicted high stress in the bond due to the epoxy  shrinkage when cycled to 5 K. The high stress occurred around the exposed periphery of the bond,  essentially independent of bond area. This raised concern micro-cracking could occur around the  periphery and cumulatively degrade bond strength over multiple 5 K cycles. However, we realized the analysis could be too pessimistic in predicting the adhesive stresses, due to assumed linearity and stress  infinity conditions at the bond edges often inherent in such FEA [1]. And BATC had successful use of  adhesives to 5 K, which further suggested the FEA did not provide for sufficient understanding of this issue. Testing was done to address this concern. Beryllium coupons were bonded and cycled ten times from  room temperature to 4 K. They were then shear and leverage strength tested at room temperature to  determine if any degradation resulted. There was essentially no difference in results between cycled and  un-cycled samples. In fact, all the cycled sample actually showed slightly higher strengths (a pleasant surprise).  Lesson learned #2 : Adhesive bonds at cryo-temperatures should be well understood and / or tested. For  the secondary mirror mou nt, FEA showed high stresses that suggested the bond joint strength could  cumulatively degrade by thermal cycling to 5 K. To better understand if this was an issue or not, tests  were done as described above. These tests showed no degradation in bond strength after thermal cycling  to 5 K and provided the understanding we needed to proceed.  408
Stepper Motor Component Testing  A stepper motor-gearhead assembly was tested early, as a component, because it was considered a  significant risk item to work at 5 K. We had one in house from a vendor that does a g ood job of matching  the CTE of materials in their units and that would have a good chance of working. The following briefly  summarizes the test of this motor.  Test Description Measure power-off backdriving torque and operating output  torques at room and cryogenic temperatures.  Figure 8.  Stepper MotorGearhead Motor Modifications The stepper motor-gearhead is shown in Figure 8. This unit was  disassembled and all its key bearing and gear surfaces were dry  lubricated with a BATC proprietary process. The unit also had non-metallic parts that were replaced with metallic ones, as  supplied by the vendor in another  version of this unit. It was  baked out to remove moisture and dry purged for testing.  Test Set-Up See Figure 10 on the next page for the set-up description (also  used for ball screw component testing).  Success Criteria /g120 Operate near 5 K.  /g120 Show no big increase in backdriving to rque (exhibit no to low binding).  /g120 Show starting (pull in) torque /g116 353 mN*m (50 oz*in).  Results The unit met all the success criteria, as shown by the results in Table 2. Stall (pull out) torque and winding  resistance were also measured for reference.  Table 2. Stepper Motor Component Test Results  Temperature of Unit Operated? Backdriving  Torque (mN*m) Starting (pull in)  Torque (mN*m) Stall (pull out)  Torque (mN*m)  Room (pre-test) Yes 212 - 233 466 - 480 508 - 530  7-14 K Yes 247 - 268 395 - 565 424 - 706  Room (post-test) Yes 212 - 247 438 - 480 480 - 530  (Tests were done at less than peak po wer, CW / CCW, and at multiple step rates)  An approximately 100:1 Relative Resistivity Ratio (RRR) was measured at 5 K, the RRR being the ratio of  winding resistance at room temperature vs. 5 K. This lower winding resistance at cryogenic temperatures was not an issue since the motor was current limited by its drive electronics.  The flight motor was similarly tested to higher acceptance torque values, since it is larger. Lesson learned #3(a) :The stepper motor-gearhead needed only simple modification to work near 5 K.  The modification was primarily dry lubric ation, using a BATC proprietary process.   409
Ball Screw Component Testing  Ball screw testing was performed early because there was concern its running friction torque might  change appreciably when rotating at 5 K. The following briefly summarizes the testing of the ball screw.  Test Description Measure starting and running torque at room temperature and  near 5 K. Two ball screws were tested, a smaller, preloaded,  carbon steel unit, shown in Figure 9, and a larger un-preloaded  440C version.  Ball Screw Modifications The ball screws were disassembled and dry lubricated with a  BATC proprietary process. Some internal plastic parts were  changed to metal parts, for flight but not for this testing. The units were baked out to remove moisture and dry purged for  testing. Figure 9.  Ball Screw  Success Criteria Max torque of 35.3 mN*m (5 oz-in), but a goal of 17.6 mN*m  (2.5 oz-in) is preferred.  Results The test was a success, as shown in Table 3. The larger, un-preloaded scre w met the torque limit and  was near to the preferred goal. The preloaded units torque was shown to be undesirable. So, a smaller  un-preloaded screw was chosen for the FM-EM, which ultimately met the preferred goal.   Table 3. Ball Screw Component Test Results  Temperature at  test Smaller, preloaded screw  max torque* (mN*m) Larger, un-preloaded screw  max torque (mN*m)  Room (pre-test) 21.2 - 35.3 8.8 At 4.8 – 5.2 K 35.3 - 67.1 22.0 Room (post-test) 21.2 - 28.2 11.3        * 6 balls missing (2 per track), lost in lubrication process  Dynamometer (shown)  or torque watch  LHe LevelItemunder test Dewar Tubes Test Set-Up (used for ball screw and motor) The items were tested in a dewar with liquid  helium, per Figure 10. They were held with a  low thermal conductance outer tube and  turned with or by another tube. A torque watch was used for ball screw testing, while  the dynamometer shown was used for the  motor testing.   Lesson learned #3(b) : The ball screw  needed only minor modification to work near 5 K. The modification was primarily dry  lubrication, with a BATC proprietary process.   (Dewar not shown)  Figure 10.  Component Test Set-Up  410
Focus Mechanism Assembly Testing  The FM-EM and FLT-FM assemblies were tested as described in this section, to evaluate how well the  focus mechanism met its requirements over ambient and cryogenic temperatures,  before and after  vibration testing, and after life testing of the FM-EM.   Challenges overcame included precisely measuring the “mirror” (simulator) motion in a cryogenic  environment. Also, the flight mechanism survived a severe over-test condition,  as described  herein. And  relevant items that contributed to successful testing are also noted.  Test Description All the key requirements per Table 1 were evaluated, and are summarized as follows:  /g120 Focus range, step size, and repeatability  /g120 De-center and tilt over the focus range  /g120 De-center, tilt, and focus shift after launch vibe  /g120 Vibration testing and determination of mode frequencies  /g120 Life testing to 1X life and for up to 4X life  FM-EM performance testing was don e in vacuum at cryogenic temperatures and in ambient conditions,  both before and after vibration testing, and after life testing.   FLT-FM performance testing was done at ambient conditions, but it was later cryo-tested at the telescope  assembly level. The FM-EM and FLT-FM are similar, except for the notable differences shown in Table 4.  Table 4.  Differences between the FM-EM and FLT-FM  FM-EM FLT-FM SM mass simulator Flight SM and mount  Smaller motor and single wound Larger motor and dual wound  Flexures sized for 70 G in all axes Flexures sized to 125 G in focus axis  Carbon steel ball screw 440C induction hardened ball screw  No limit sensors Limit sensors  Test Set-Up Description The FM-EM was tested in the dewar shown in Figure 11. The test measured mirror motion relative to the  mechanism interface, using a mirror simulator. A brief description of the test set-up is as follows:  /g120 The mechanism was mounted on a stable reference plate that was used for cryogenic or ambient  testing and for vibration testing. It had position sensors on it that monitored the mirror simulator  motion relative to the plate, hence, the mechanism interface.  /g120 Legs (not shown) mounted the reference plate to the dewar and prov ided a conductive path to the  dewar cold plate. The legs were removed so the plate could be used in vibration testing too.  /g120 VITs (Variable Impedance Transducers) monitored the mirror simulator motion along the focus  axis, in two axes of tilt, and in two lateral axes for decenter. The VITs were initially calibrated at room temperature and to 80 K, but not to lower temperatures.  /g120 Anticipating the VIT calibration could change below 80 K, a secondary focus measurement  technique was provided by using two laser interferometers on a stable base. One measured  mirror simulator motion and the other measured the reference plate motion. The difference between these measurements represented the desired focus motion, relative to the plate.  411
Conclusion  The new focus mechanism was completed and shown to provide repeatable focus positioning to 5 K. Its  design and test challenges were successfully overcome, on schedule. The mechanisms concept and implementation have been sh own to be simple, effective, and very robust. It is currently in service in the  Spitzer Space Telescope, one of NASA’s “Great Observatories”.  After launch on August 25 th, 2003, the new focus mechanism has performed as planned in flight. The  observatory has surpassed its expected 2.5-year life and is now approaching a 5-year life goal. (The  mechanism only needed operation twice during this time; once to verify its successful operation, and then to achieve the final focus position that is currently in use.)  Testing of the stepper motor and ball screw components significantly mitigated risk early in the program.  The build and test of the mechanism engineering model further mitigated risk and allowed for verification  of performance over and above the required life.  Lessons learned, resulting from this successful mechanism development, are summarized below.  Lessons Learned Summary 1. Titanium 6Al-4V (ELI) is useful as a flexure material to liquid helium temperature s since it has  sufficient plane strain fracture toughness (K IC) to < 5 K.  2. Adhesive bonds at cryo-temperatures should be well understood and / or tested. For the mirror  mount, FEA did not provide sufficient understanding of the bond. Testing with coupons was  required and showed the bond strength was acceptable after multiple cycles to 5 K.  3. The geared-stepper motor and ball screw components needed only slight modification to work at  5 K, which was primarily dry lubrication.  Subsequent Developments  This mechanism has provided heritage for othe r programs at Ball Aerospace, as follows:  /g120 The James Webb Space Telescope (JWST):  Similar geared-motors are planned for use in the  cryogenic nano-actuators used to position its primary mirror segments.   /g120 Kepler (another space-borne telescope):  A derivative of this mechanism is planned to move the  primary mirror for focus adjustment.  Acknowledgements  This device was developed with funding and oversight from JPL under contract 960669.  Thanks to Robert M. Warden (BATC) for developing the mechanism conce pt and to the BATC team and  suppliers that supported this effort. And in memory of Mike Rice, an outstanding technician on the team.  References  1. MIL-HDBK-17-3F, section 6.2.3.6.  2. T. D. Scharton. Force Limited Vibration Testing Monograph. NASA Reference Publication RP-1403  (May 1997).  414
Development of a Spacecraft Antenna Pointing Gimbal Charles Monroe* and Peter Rossoni*  Abstract  The development of the pointing gimbal in the high-gain antenna system (HGAS) of the Solar Dynamics  Observatory spacecraft is described. The gimbal was designed for 5 years of service in Geo-Synchronous  orbit. The hardware incorp orates multiple levels of redundancy, allows harnessing and waveguide along  its full length across its two axes of rotation and points with an accuracy of better than 0.065°. Significant  issues with actuator alignment, Electrical Contact Ring noise, pointing budget, and waveguide failures are  described, along with their respective resolutions.    Introduction  This paper outlines requirements, design and development activities of the SDO gimbal. Several  hardware anomalies and their resolution are described. The critical reliability level was a driver for most of  the issues uncovered during the gimbal development.   Significant design areas include the actuator and contact-ring mechanisms and waveguide. Unique  events and lessons-learned include the encoder alignment to the actuators, noise during component-level testing, replacing flex waveguide and accommodating the harness.  Background  The Solar Dynamics Observatory (SDO), shown in Figure 1, is a NASA spacecraft that will collect data  from the Sun during its 5-year life. The spacecraft was designed by and is being integrated at NASA Goddard Space Flight Center in Greenbelt, MD. Universities and industry provide its science instruments.   This observatory transfers 150Mbps (millions of bits per second) of solar imagery (with overhead) per day  from its 28.5° inclination, geosynchronous orbit at 36,000 km (22,400 mile) altitude, to the ground station  in White Sands, New Mexico. The gimbal geometry that is most conducive to this end is a two-axis azimuth/elevation configuration. The azimuth axis will rotate once per orbit (once per day), and the  elevation axis will rotate up to ±65 degrees to allow the antenna to point to the desired Earth coordinates  at the SDO Ground Station. To avoid exce ssive spacecraft roll maneuver s, a dual HGAS approach was  taken, with antenna systems on opposite sides of the spacecraft, allowing selection of the optimum  gimbal for downlink via scheduled hand-offs.  * NASA Goddard Space Flight Center, Greenbelt, MD  Proceedings of the 39th Aerospace Mechanisms Symposium, NASA Marshall Space Flight  Center, May 7-9, 2008 415
Figure 1. Solar Dynamics Observatory with One of Two High-Gain Antennas and Gimbals Circled  Driving Requirements  An important driver for the gimbal system is the downlink requirement. A rate of 150 MB/second is  needed with 99.99 percent reliability over a 99 percent duty cycle.   The 99.99 percent reliability requirement reflects the transmission error rate. To achieve this percentage  during periods of transmission, the azimuth axis must be able to rotate continuously without downtime for  “rewinding” of the harness about the axis of rotation. This drives the need for an electrical contact ring  assembly (ECRA), a slip ring or roll ring, to pass power and signal through the axis.   The 99 percent duty cycle addresses periods of fog and rain at the ground station. The spacecraft itself  will occasionally occlude the view from a single antenna. A continuous downlink capability dictates two antennas—one on either side of the spacecraft. During portions of the year, a daily hand-off between  antennas will be required.     These requirements lead to a highly reliable, 100% duty-cycle design, with no planned datalink  interruption. The characteristics of the antennas and the power av ailable for transmission drive the  need to keep RF  throughput loss low—the gimbal itself was allocated a loss of less than 1.45 dB. To meet this  requirement, an all-waveguide RF system was selected as opposed to the simpler coaxial cable approach. This necessitated waveguide rotary joints at each axis of rotation.    Also, minimizing loss drives the need for having a pointing capability of + 0.30° to the ground station for all  error sources, including spacecraft position and orientation. Of this amount, there is 0.14° budgeted for  random and calibration errors of the gimbal.  416
Gimbal Design  The overall configuration is shown in Figure 2. The azimuth axis has unlimited rotation and the elevation  axis has a + 69° range. The continuous azimuth rotation is made possible by having the power and signal  transferred through the ECRA. For the elevation rotation, a rotary cable wrap wherein the cable is  carefully spiraled through the center of the elevation actuator manages the harness.    There are two rotating sections of waveguide on this two-axis gimbal. An azimuth section rotates with the  azimuth axis and extends from the azimuth actuator up to the elevation axis. An elevation section rotates  with the elevation axis and extends from the elevation actuator up to the antenna.  Elevation Axis High Gain Antenna  Azimuth RF  Slip Joint Elevation RF Rotary Joint  Electrical Contact Ring Assembly (ECRA)  Azimuth Axis Azimuth Waveguide  Azimuth Rotary Actuator Azimuth RF Rotary Joint  Figure 2. Gimbal Cross-section  417
Rotary Actuators  The rotary actuators for azimuth and elevation are identical except for hard stops on the elevation  actuators limiting travel to + 69 degrees. Each commercial actuator consists of a stepper motor, hybrid  optical encoder, and harmonic drive gear reducer in a titanium housing. Once delivered, the units were  tested for compatibility with the GSFC-designed control ele ctronics, and characterized for settling time,  torque margin, encoder output and alignment. Then thermal hardware was applied, as shown in Figure 3, and the harnessing prepared for integra tion to flight hardware.     Figure 3. Rotary Actuator Shown During Thermal Hardware Application  Actuator Description The three-phase stepper motors are redundantly wound. There are two redundant encoders:  a coarse  encoder on the output that determines hemisphere and home, and a fine encoder on the input that counts  each step taken. The actuator details are summarized below:  Actuator Parameter Value Output step 0.0075 degree  Harmonic drive gear ratio 200:1 Motor step 1.5 degrees  Unpowered detent torque 34 N-m (300 inch-pounds)  Max required slew speed (under the following  conditions at qualification temperatures) 30 degrees/min (66.7 pulses/sec)       Driven inertia 2 kg-m2      Driven offset load 28 N-m (250 inch-pound)       Driven friction load 2.5 N-m (22 inch-pound)  Step Settling of Bearings On orbit, a motor step will be taken roughly every two seconds. For the actuator life test, time constraints  drive the need for more frequent steps, but the period between steps should be no less than the time  required for bearing balls to settle. It was decided that after the ball motion  decreased to  a point where  the magnitude of the oscillations is less than the width of the Hertzian contact patch, the bearings would  be considered essentially settled. We believe this settling criterion to be consistent with ball-pass analysis  for lubricant tribo-degradation. The time  required to reach this point is 35 msec, so the actuator life test  could be run at ~28 pulses per sec, which is an acceleration factor of ~57.    Actuator Encoder Alignment Actuator position is determined by internal optical encoders. Alignment of the encoder is inferred from its  output. Even though settling time and torques were within requirements, so me actuators had marginal  418
alignment of their encoder discs to the step detents. The easiest way to test for alignment is by the quality  of the encoder output during settling in its detent. A well-aligned encoder disc would have its LightEmitting Diode (LED) centered in the disc aperture. Even at the beginning of settle, when the oscillations  are highest, little or no light would be occulted by either edge, as shown in Figure 4a (each window is two  steps wide). Encoders that are not optimally aligned would shadow a portion of the light while settling in a  step detent. Of 10 potential flight actuators, 6 were aligned with less than optimal performance, as shown  in Figure 4b by the encoder light output “hash” during characterization testing. The actuator specification called out static alignment only; as a result, all 10 flight actuators satisfied the specification requirements.     Just prior to delivery to the spacecraft, two of the flight actuators were damaged beyond repair by  excessive heat in a Goddard thermal vacuum chamber. Two of the less-than-optimal actuators were  brought to flight status. These passed the static alignment specification but during a high-rate slew  operation could incur positioning errors of one step. This error is reset in the control electronics when the  actuator passes through the “home” position, so the condition is tolerable during slews. Under normal  tracking, there is adequate time for the encoder output to settle and the output to be verified. The only remaining issue is diode output over the mission life. As the diode response decays due to radiation  effects and the normal degradation due to operation, the partial occulting could reduce margin on the  encoder output.    Figure 4a. Good Encoder Alignment – Two-Step Window  Figure 4b. Less than Optimal Encoder Alignment  419
Electrical Contact Ring Assembly (ECRA)  The elevation axis harness, which consists of 14 power circuits and 26 signal circuits, is routed through  the ECRA. The outer structure of the ECRA, shown in Figure 5, is stationary and is hard-mounted to the same bracketry as the azimuth actuator stator. The inner structure of the ECRA is fully supported in the  stationary section by a duplex set of angular contact bearings as well as a trailer bearing. A tooling ball  mounted in the base of the rotating portion of the ECRA mates with a slot in the azimuth output shaft allowing the gimbal to transmit the rotary motion while permitting slight angular misalignment between the  ECRA and the output shaft.   Drive  Tooling  Ball  Figure 5. ECRA with Thermal Hardware Attached (prior to lead tape over wrap)  The structure consists primarily of aluminum and titanium components. The ECRA uses gold/silver/nickel  alloy mono-filament brushes in gold-plated brass grooves. Two outer brush blocks that are part of the  stator support brushes that span the gap between the stationary and rotary portions of the assembly. Each groove accommodates two brushes, one leading and the other trailing. The power circuits utilize a  three-groove design while the lower-powered signal brushes utilize a two-groove design. The final result  is a current-carrying margin in the power and signal circuits of three and four times respectively. This is in addition to the electrical redundancy in the gimbal itself.    Component-Level Test Issue During component thermal testing, the ECRA exhibited higher than expected noise in the lines. While the  ECRA exhibited acceptable noise performance at temperatures greater than 0oC, the noise levels on  several circuits increased considerably at sub-zero temperatures. The primary cause was determined to  be water contamination of the brush-groove lubricant. It was determined that the nitrogen environment in  the thermal chamber was insufficient to purge the assembly of all water vapor. The primary solution was  to perform the thermal testing in vacuum. Additional steps that were taken included a vacuum bakeout  without the external housing installed before final assembly, a run-in before testing, and efforts to improve  420