ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
35 (0.5”) thickness, 100-mm (4”) spherical radius and 45 ° interface flats drove the resulting geometry. Since  the resulting fittings had a diameter of 83 mm (3.25”), the fittings had to be installed with a single fastener  centered on the fitting so that positioning was mainta ined, and the CTE mismatch between the fittings and  the near-zero CTE backplane structure woul d not result in unacceptably high loads. 

Figure 8.  Primary Mirror Hinge and Latch Sets      The preload device for the latches uses a 3/8-24 UNJF-3A screw driven by  a stepper gear-motor  that engages and screws into a floating nut in the mating latch (Figure 10). The two halves of the device are pulled toward each other but  never bottom out so the preload device never  provides a stiff load path. It is critical to have the nut properly aligned to the drive screw before it  starts turning to prevent cross threading. The  floating nut is spring loaded in a loose housing that prevents nut rotation during tightening but  allows a small amount of angular motion for nut  to drive screw alignment during engagement. A feature was also added to the tip of the drive  screw to guide the nut onto the screw and  prevents the nut from rotating the mating threads out of alignment. The screw material is  A286 CRES chosen for its high strength, and  Nitronic 60 was chosen for the mating nut for its anti-galling characteristics. To minimize the  torque resistance to the gear-motor, the drive  screw has MoS 2 lubrication and a thrust bearing was used between the base of the drive screw and its  housing. The thrust bearing also provides a de sired compliant element to the preload device.   Drive Hinge  Passive Hin ge1541 mm      60”  Preload  Device with  Cryo motors  (4K Op temp.)  •The Quasi-Kinematic I/F reduces stresses due to initial adjustment  mis-alignments and thermal deformations from cool-down • Cryo motors allow for final     latching after cool-down   • The wing deployment  repeatability is within  .075 mm- V1 & .1 mm V2 & V3 (At the Latch Interface) Y X A1-DOR  A1-DOR  A1-DORA1-DOR A1-DOR 1-DOR2-DOR 3-DORZ1027 mm      40”  Figure 9.  Latch Design Details  Wet Install Fasteners All latch mating surfaces are Titanium with a Tiodize 203 mm 8.0”  203 mm   8.0”  Adjustable Pin for over  constrained direction Drive Screw Latch    provides 600 lbs preloadCup / Ball Fittings  12.7 ± 1.27 mm  A single fastener  removes thermal  mismatch problem
36   Figure 10.  Wing Latch Engagement and Preload 

Secondary Mirror Support Structure    The SMSS design is a deployed tripod as shown in Figure 11. A tripod design was selected because of  its superior dynamic stability and mass efficiency. Th e deployment system is a four-bar linkage driven by  a stepper gear-motor. Since wavefront error (WFE) is ve ry strongly affected by axial despace (defocus) of  the secondary mirror assembly (SMA),  the design uses a very low CTE composite material system in the  SMSS tubes. This produces a very small despace contri bution to wave front error over the worst-case  hot-to-cold conditions.  

SMSS Deployment Drive   The SMSS is driven by a single stepper gear-motor locat ed at the inboard single strut hinge as shown in  Figure 12. This hinge was selected as the drive loca tion since it provides t he best mechanical advantage  for the system. The gear-motor pinion drives a spur gear mounted on the inboard single strut providing an  additional 4/1 gear ratio over the 22.5 Nm (200 in-l b) torque capability of the gear-motor. This provides  roughly 90 Nm (800 in-lb) of drive torque capability.   The SMSS, like the PM wings, is configured to minimi ze stresses in the system while latching. Although  all five SMSS hinges are latched, the mid-hinge is th e only one to be preloaded into hard stops. Because      Figure 11. Secondary Mirror Support Structure    Fully  Deployed  Wing   Drive Screw Clears Housing   (.020 Capture range) Drive Screw Engages Nut   .5 lb spring and nut alignment  feature locates nut  Nut is pushed BackHinges  bring  latches to  within thei r  capture  range  Drive Screw Latches   Latching and preload adjustments  occur at cryo temperatures  Latches   Engage   Hinges are  now out of  the  stiffness  path  Inboard Single  Strut is restrainedto ISIM Enclosure Dual Struts are restrained to the upper center BP   Inboard Hinge  drives deploymentMid-hinge latches first into hard stops creating a rigid singlestrut  Repeatable within 3mm @ SM 123 Heaters  control when  latches are  engaged  4 The dual, inboard  and outboard hinge  latches, which can latch in any  position, latch last  to minimize distortions into the  SMS S structure
37  200 In-lb  stepper gear-motor meets  4x flight & 2x  ground testing  torque reqmnt1.00 Dia. Pinion 4” Dia. Deployment Gear   bolted to the Inboard single-strut  Thermal clamp  latch pin   Figure 12: SMSS Deployment Drive   the four remaining latches take advantage of the  thermal shrinkage of their aluminum components  to clamp them in place during cooldown to operating temperature, t hey can latch at any  angle of rotation.  SMSS Mid-hinge

The mid-hinge is deployed through a 168-degree rotation and bottoms out on hard stops at the end of its travel. There are two leaf spring latches  (one on each side of the hinge) that engage at  this point capturing catch fittings on the mating  hinge half. The catch and latch base fittings are  machined titanium and the leaf-spring latch is 7075 Aluminum. The latches are centered on the  hinge (Figure 13), so that as the temperature drops fr om RT down to ~30K, the aluminum latches shrink  and preload the M55J composite hinge at its two adjustable hard stops and two hinge pin to bushing  interfaces. The stops provide non-co nforming sphere-on-flat interfaces and the pin bushings are notched  to provide non-conforming line contact interfaces. 

Figure 13.  Mid-hinge Latches and Hard Stops 

A spring feature was added to the latch to help compens ate for any creep that coul d occur over the life of  the mission and help provide a well-def ined stiffness path. The stack of Bellville washers, shown in Figure  14, is roughly half the stiffness of  the aluminum latch and has a 1-mm (0.040”) stroke. Two stops were  required to meet the natural frequency requirement of the hinge, and the process for adjusting them to get  an equal load distribution has been demonstrated on the DOTA hardware.    SMSS End-hinge

The end hinge design is common to the dual stru t hinges and the inboard and outboard single strut  hinges. It has a rotation pin that provides precision deployment location plus a clamping pin that preloads  the lug and clevis bushings together once the hinges ar e in their final deployed locations. The latches will  90º Notched Bushing in clevis  Notched Bushing in lug  The Bushings are notched to give  “Line Contact” at 4 locations for   reduced jitter and improved repeatabilityPin Design   Notches provide an anti- jitter feature Adjustable Stop Design   Stops give high deployment repeatability  and balanced thermal stabilit y  Stops Bear on Hinge Fittings and  Strut Tubes  Fasteners hold stops in place Threaded  Stop with  Jam Nut  (4.25”) (8.50”)The Thermal Spring latches provide 2224N (500 lb) min.  to preload both the pin and  adjustable stops  Stops Bear on Hinge Fittings and Strut Tubes 108 mm 216 mm
38 latch in any position without forcing the hinge to a preset location and provides a 10x load carrying  capability over operational loading. The end hinge design details are shown in Figure 15. 

Figure 14.  Mid-hinge Preload Mechanism 

   Figure 15.  End Hinge Latch Design   Deployment Repeatability    The deployment repeatability of the SMSS must be such that once the latches are engaged and  preloaded, the secondary mirror is within the wave front capture range of the wavefront sensing and  control system (WFS&C). The tripod design is a geom etrically robust design for locating the SMA within  the allowable capture range and is driven primarily by  the size of the gap between the end hinge pins and  bushings. The repeatability requirement and analysis result are shown in Table 5.  

Lug Hinge Bushing  (INVAR) Compression Tube (GFRP)  Provides load path for strut  clamped bushings Heaters will be located  on the Compression TubeClevis Hinge Bushing (INVAR)  Rotation Pin - Provides  precision axis of rotation  (INVAR with MoS2)  Clamping Pin - Clamps  mating hinge bushings  together producing torque carrying capability  (Aluminum pin fits inside INVAR rotation pin) Cryo Operation:  Set initial gap to   0.25mm ± 0.05mm  (.010 ± .002”)  Room Temperature Operation: Torque the Pin Retainer to 7.57 ±.23 N-m (67 ±2 ft-1bs) Dual  Strut Outboard Sin gle Will latch in any positionwithout forcing the hinge to a pre-set location  The latch relies on friction  but is designed to a 10x capability to expected load ratio  Inboard Single  Strut Latch Adjustment &  Preload and RT Latching Bolt   Belleville Washers used to reduce preload sensitivity to  initial adjustment and  temperature changes Thermal Spring Latch(Aluminum) Latch  Guard (Titanium) Latch Catch  (Titanium)  Bearing Double r (M55J) Bearing Double r (M55J) Bearing Surface Heater Located on  GFRP Hinge fittings Latch Base(Titanium)Bearing Surface Photograph of  test hardware
39 Table 5. SMSS Deployment Repeatability 

 X Despace  (mm) Y Decenter  (mm) Z Decenter  (mm) Theta Z  (arcmin)  Requirement 3 3 3 1  Analysis Result < 1 < .5 < .5 < .5 

Development Testing 

This section covers the DOTA testing and associ ated results including; deployment capability and  measurements of the repeatability of the PM wing  mechanisms, characterization of the micro-dynamic  stability of the wing latching system , functional latching of the SMSS hinge latches, and measurement of  the load carrying capability of the SMSS end hinge. Deta ils of the test metrol ogy and procedures and  additional thermal stability results can be found in [2] .  Figure 16.  Cryogenic Development Optical Telescope Assembly (DOTA) 

In order to understand the behavior and reduce th e risk associated with very large, deployed optics  support structures, NGST developed and tested the DOTA PM wing and SMSS hinges. The DOTA wing replicates a full-scale portion of the JWST PM support structure and includes a hinge-line with a full set of  latches so the overall stability of the system c an be characterized. The DOTA SMSS hinges and latches  replicate full-scale end and mid-hinges. The DOTA hardware is shown in Figure 16.    The DOTA wing hinge/latch testing was broken into tw o parts. The first part had the hinges and latches  integrated into an aluminum test fixture that simulate d the flight structure. They were then tested as a unit  at room temperature (RT) for deployment repeatabilit y and micro-dynamic stability. Once this testing was  complete, the mechanisms were removed from the test fixture and integrated onto the DOTA wing  structure. Then, as an integrated  structure, the DOTA wing was te sted at cryogenic temperatures for  thermal and micro-dynamic stability.  Hinge/latch repeatability

The fixture that the wing hinges and latches were tested on is shown in Figure 17. The hinge-line was scaled down in the Z direction to facilitate testi ng. The deployment repeatabilit y requirements are derived  from the JWST need to have the Primary Mirror wi thin the capture range of the WFS&C system.  Deployment repeatability was measured in X, Y, Z a nd Theta Z directions. For each test, the hinge was  driven 103 degrees from the stowed to the deployed position using a flight-like stepper gear-motor. The  latches were then engaged using a torque wrench to accurately assess the torque being applied. The test  was repeated 10 times for both a 1.13 Nm (10 in-lb) la tching torque and the 11.25 Nm (100 in-lb) latching Theta Z Test DOTA Mid-Hinge DOTA End-Hinge DOTA Wing
40 torque. Repeatability measurement results  are given in Table 6. The latch repeatability  tests demonstrate that the DOTA wing  latches easily meet  the repeatability  requirements for wavefront capture. The  repeatability tended to degrade with higher  preload and we believe this was due to the lack of stiffness in the aluminum test fixture.  Although the test data met all the  requirements with margin, even smaller repeatability errors are expected with the  stiffer flight backplane. 

   Table 6.  Deployment and latchi ng repeatability results, one-sigma 

 X Despace  (mm) Y Decenter (mm) Z Decenter (mm) Theta Z (arcmin)  1.13 Nm error 0.020 0.030 0.043 0.15  11.25 Nm error 0.033 0.043 0.030 0.74  Requirement 1 0.1 0.1 1.0 

Latch Micro-dynamic Testing   Over 100 load cycles were applied during the RT  micro-dynamic testing of the DOTA latches with shear  and moment loads significantly higher than those  expected operationally. The first set of tests applied cyclical quasi-static loading at 0.1 Hz in each shear  direction. Since it’s believe d that the micro-dynamics in  a deployable structure is related to hysteretic  response, the displacements were measured in all  three directions and the hysteresis was determined by removing the linear response. Sample hysteresis  results are shown in Figure 19.    The next set of testing also applied cyclical loading to  the latch interface, while looking for evidence of  “nanolurches”. Sample results for this testing are shown in Figure 20. The 9-nm lurc h shown occurred with the latches at half of the operational preload  and with an applied load 100 times operation loading. The final set of tests run with this latch micro-dynamic test setup applied a thermo-elastic load acro ss the latched interface 10,000x greater than the  operational thermo-elastic load. The series of test s revealed no evidence of “nanolurches” when the  preload applied to the latch was consistent with its operational design of 2669 N (600 lb). Moreover, only  3 nanolurches were detected during all of the testing, with these occurring when latch preloads of 1334.5  N (300 lb) or less were applied.
41 

 Thermal Stability Testing

After the latch testing was completed, the latches were integrated onto the DOTA wing structure. The  DOTA wing was then transported to the Marshall Space Flight Center in Huntsville, Alabama for thermal  stability testing down to cryogenic temperatures in th e X-Ray Calibration Facility. Figure 22 illustrates how  distance-measuring interferometers were configured to monitor the latch interface at several locations along the hinge-line. Similar to the micro-dynamics te sting discussed previously, the interferometers were  used to look for discontinuities in the latch behavior.  

The interferometers detected frequent slips as the DOTA first achieved temperature, as depicted in  Figure 23. Three of the four latches were powered fo r the DOTA testing and these latches were released  and re-tightened (Figure 24) as planned after cool-dow n. As expected, this significantly reduced the  amplitude and frequency of the slips. Bulk temper ature testing and gradient testing was then performed  on the DOTA. The bulk temperature testing consisted of a 20K thermal excursion. This was 100 times  larger than the expected on-orbit t hermal excursion of 0.2K creating 100x the thermal loading in the latch  interface. Moreover, the gradient test ing far exceeded this multiplier compared to on-orbit expectations.  

Axis #9Axis #11 Axis #7 Axis #8Axis #10West EastLatch #4 Latch #3 (2-DoR) Latch #2Latch #1 (unpowered, 3-DoR)Laser Can and delivery optics remain  at room temperature on ASF via  heaters and proportional controllers  along with 20 layer MLIMetrology attaches to main beams on  each side of hingeline1 1213 111Latch DoR configuration Axis #9Axis #11 Axis #7 Axis #8Axis #10West EastLatch #4 Latch #3 (2-DoR) Latch #2Latch #1 (unpowered, 3-DoR)Laser Can and delivery optics remain  at room temperature on ASF via  heaters and proportional controllers  along with 20 layer MLIMetrology attaches to main beams on  each side of hingeline1 1213 111Latch DoR configuration Figure 22.  DOTA Hinge-line Metrology Beam Paths
42 Figure 26.  SMSS Mid-hinge Load vs. Temp NGST Mid Hinge Latch 1 & 2 Load Vs. Temp Run 1 -1000100200300400500600 -300 -250 -200 -150 -100 -50 0 50 100 150 Temperature in Degrees FahrenheitLoad in LbsCorrctd1 - L1 Corrctd1 - L2 Corrctd2 - L1 Corrctd2 - L2 Corrctd3 - L1 Corrctd3- L2  Under these high loads, interferometer axis 10  exhibited stick-slip behavio r in two instances, with  no other measurable discontinuities in the data.  Without the ability to re-torque this latch, it  suffered the loss in pre-load that results from cooling from room temperature to 50K. The  observed slippage at the latch with the lower  preload that is absent at the latches with operational preloads correlates with the room  temperature micro-dynamic performance  measurements. 

DOTA SMSS Hinge / Latch Testing

To verify the key components of the SMSS, a  prototype of each type of hinge (i.e., a mid-hinge  and an end hinge) was built and tested. The purpose of the testing was to verify functional latching at  cryogenic temperatures. In addition, the end hinge was te sted to show that it has a minimum 10x margin  against operational loads. Micro-dynamic stability test ing on the hinges is also planned for early 2005.   The cold box shown in Figure 25 was used to take the hinges down to ~88K using LN 2. Flight  temperatures of ~20K could have be en reached using liquid helium; however, the added expense couldn’t  be justified given that flight prel oads were achieved at the higher LN 2 temperatures by simply adjusting  the initial latch gaps accordingly. 

   Figure 23.  Hinge-line Displacements at the  End of the Cool-down  Figure 24.  Latch Interferometer Data  During Bulk Temperature Test
44 Concluding Remarks 

The observatory architecture and design details descri bed in this paper are a “point in time” design that  was submitted as part of the Phase 2 JWST proposal in October 2001. The program had a replan  exercise in an effort to reduce program cost. This repl an effort resulted in several significant changes to  the architecture including a reduction in the PM diam eter from 7 m to 6.5 m and a change to the number  of segments in the primary mirror from 36 to 18. Th ese architecture changes have resulted in some small  impacts to the designs described in this paper, but essentially, the designs have stayed the same. As a  result of the DOTA early development testing, t he OTE’s primary deployment system designs have been  shown to satisfy their driving requirements. Consequent ly, this testing has proven to be very effective in  reducing the program risk associated with t hese optically stable deployable structures.  

References 

1. Lake, Mark S. and Hachkowski, M. Roman. “Design of Mechanisms for Deployable, Optical  Instruments: Guidelines for Reducing Hyster esis.” NASA/TM-2000-21 0089, (March 2000)  2. Atkinson, Charlie, Gilman, Larry and Reynolds , Paul. “Technology Development for Deployable  Telescope Structures and Mechanisms.” SPIE Conference Paper 4850-45 (June 2002)
The bo resight alignment b etwee n the transmit path and the re ceive path is criti cal to the pe rforman ce of  the GLAS instrument. The  alignme nt require ment be tween the t ransmit an d receive path s was <30 arc  seconds for full scie nce, and <4 2 arcsecond s for 50% (FWHM) re sponse. The 50% respon se is  considered the minimum a cceptabl e ran ge for scie nce. These alig nment re quirements had t o be met pre - launch to po st launch, an d within a thermal ran ge of 15ºC to 23 º C5     Transm it  Optical PathLaser A ssemblyFixed Fold MirrorsStar Tracker 1 meter Telescope   

 Receiver Bench  Transmit Bench   

Figure 1.  GLAS Rec eiver and Tran smit Ben ch  

Since te sting indicated that  the boresi ght alignme nt over temp eratu re may be ma rginal, it was decided to   pursue ra pid developm ent of a Bore sight Adjustment Mech anism. This effort, in parall el with  contin ued  environ menta l testing of GLAS, wa s intende d to mi tigate pote ntial margi nal alignme nt perfo rman ce  should further tests and data reduction in dicate th at the GLAS in strument was failing to me et its align ment  budg ets. Fin al testing and data redu ction later i ndicated that the boresight align ment an d sta bility budget s  could be ma intained u sing existing in strum ent re sources. At that time, it was dete rmined that the   integratio n of the BAM ont o the G LAS instrument was not require d. The BAM would have  been m ounte d  as sho wn in Figure 2 had i t been integra ted into the GLAS instrum ent. 

The ope rating prin ciple, requireme nts, desig n, and  testing of the BAM are descri bed. T he BAM was  develop ed o ver a 4-mo nth perio d. Du ring that time a proto -flight mecha nism, flight electronics, sp are  electroni cs, and ground  support equipment  were de signed, fabri cated, and  fully f light qu alified.  Photograph s of the BAM and the BAM El ectro nics (BAME) are shown in Figu re 3. The BAM utilize s two  motori zed counter-rotatin g prisms to preci sely steer t he in comi ng laser radiat ion. Thi s configuration  is  comm only ref erred to as a risley pai r. This simpl e, yet precise, bea m steeri ng te chnique was the baseline  for the BAM. 

46
Laser A ssembly 

The Bore sight Adjustment Mech anism  Fixed Fold    Mirrors  Laser Sel ect  Mechanis ms  BAMLaser Pick -off

Figure 2.  Proposed BAM  Location on  GLAS Las er Transmit Be nch  

 Figure 3.  Photos o f the BAM and the BAME  

Operating Pri nciple (Ri sley Pair Equatio ns)  The ang ular deviation δ of a beam pa ssing throu gh a single prism is given by:  

)1(− =nαδ  

Where  α = prism a ngle a nd n = refractiv e index of pri sm  47
• Drive the 2 -Phase step per motors on the  BAM   • No re sources available fro m the GLAS instru ment  • Comm anding limited to using sp are High Level Di screte Comm and (HL DC) lin es from spa cecraft    Each HLDC i s a co mman dable solid stat e relay output      Only one HL DC can b e high at a time    Logic high = +24 to +31V  Logic low = high imped ance    Fixed 50 ms ± 5 ms pul se width with ri se and fall time >1 µsec a nd <10 µ sec    Maximum of 10 Hz com mand rate    Maximum cu rrent pe r com mand line 2 80mA  • Additional po wer avail able from sp are la ser heater supply circuit:  +30V, 28 watts peak  • No telemet ry available  • Must provide 100% ele ctrical isolatio n from motor when not ope rating  • Electronics  Box Mass :  <2 kg  • Harne ss length:  >3 m and  <10 m  • Design u sing available spare flight app roved parts fro m other GSF C programs 

Mecha nism Des cription 

The BAM a ssembly i s composed of th e followi ng compon ents:  the mou nting plate, the m ain structure   and two ide ntical p rism dri ve assembli es. The drive asse mblies are mo unted back to ba ck on either side  of the main structu re. Each  drive assem bly suppo rts one of t he risl ey pair pri sms on the inne r diameter of  the rotating shaft. The laser radiatio n travels through the center of the BAM as defined by the rotational   axes of the shafts. Figu re 5 shows a photograph and a cro ss-section of the BAM. Each d rive assembly  indep ende ntly rotates the  resp ective p rism via a spur  gear that  meshe s with the output pinion of the   stepp er motor actuators. T he mai n structure was bol ted to the mounting pl ate through a shi mmed 3 -point  mount to prov ide ce nterlin e to mounti ng plane h eight adjustment capability.  

Drive Ass embly   

Laser B eam Path Risley Prism s 

 Mountin g Pla te   Figure 5.  BAM Photo graph and Cro ss-Se ction  

The m ain structural comp onent s of the  BAM, the m ounting plate,  main hou sing, and bea ring housings  were ma de of Titanium Ti6al4v. Thi s material  provided a good CTE mat ch to the bearin gs and  compli ance flexure s cut into the mounti ng plate p rovided am ple margi n for th e CTE mism atch bet ween  the BAM pl ate an d composite optical bench. Sin ce the BAM would have  been attach ed after th e fact,  50