ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
Conclu sion 

The m echani sms presented in this pa per were d esigned to conform to the  idiosyncra sies of optics in  the  far an d extre me ultraviolet. Ho w the se mech anisms were  to be used on the ground was just as im portant  in their d esign a s how they we re to  be u sed in space. Anal ysis a nd te st wa s used extensively to  determi ne if the me chanisms would meet their req uirements. In  certain cases the se analyses prove d  benefi cial fo r improving an already  satisfacto ry de sign or findi ng fla ws bef ore they could affect th e  missi on. NRL is contin uing to ben efit fro m the effort put forth into t hese me chanisms a s their designs are  incorporated into future ins truments .     Ackn owledgements 

The Air Forc e Spac e Tes t Program  funded the ARGOS s pacecraft. The LORAAS mec hanisms described  in this pape r were co ntracted out to the followin g instit utions:  Swal es Aerospa ce, Inc. for the analysi s and  construction of the DDM, and Research Systems In c. for the de sign and construction of the USA.   

Referen ces 

1. McCoy, Rob ert P. et al; “Speci al Sen sor Ultravio let Li mb Image r: An Iono spheric and Neutral Density  Profiler for the Defense M eteorol ogical Satellite Program Satellites.” Optical En ginee ring, Vol. 33 No.  2 (Feb ruary 1994) 4 23-429. 

2. Samso n, Jam es A., Ederer,  David L. Vacuum Ultraviolet Spectro scopy.  Academi c Press (2000) 

3. Whitely, M.J. , Pearson, J.F., Fraser, G.W ., and Barstow, M .A., “The Stability of CsI-Coated  Microchan nel Plate X-Ray Detecto rs” Nuclear Instrum ents and Method s in Physi cs Re search , Vol  224 (1 984) 287-297 

13
The STERE O mission h as a pri me observin g period of two years. Du ring thi s time, each coron agraph  telescop e will  requi re 600,000 motio ns of the HCM. The typical demand s of spaceflight instrumentatio n  requi red the HCM  design to be light weight, low power, an d relatively low cost. Becau se of its fun ction  within th e optical p ath of the tele scop e, it is required to have rigorous me chanical tole rances, highly  repeatable m otion, and to  be extremel y clean, a s particulates ca n be a signifi cant sou rce of stray light  scatter in a coron agraph. 

The HCM was designed, built a nd tested by collaboratio n of LMSAL with  H. Ma gneti cs [1]. He re we  descri be the  developm ent and testi ng of proto-qualificati on mech anism that wa s bu ilt to flight  specification s and su bjected to full lifetime simul ation that includ ed vibration, thermal cycli ng, operation  unde r vacu um, and a full battery of pre - and po st-life  func tional te sting. Specifi cally, we di scuss a method   develop ed to characte rize the repeatab ility of the HCM . This test  verified that the mech anism perform s  within spe cification s over i ts proje cted li fe but also re vealed the li mitation our p articul ar design. 

Finally, we p resent th e results of te sting of th e fou r flight me chanisms that have bee n delivered to the   corona graph  teams and the impa ct of our d evelopment effort on de signs for similar m otors on th e  upcoming Helioseismic Magnetic Imag er (HMI) on the Solar Dyn amics Observato ry (SDO ). 

Hollo w-Core Motor Requi rements and  Design  

The SECCHI HCM i s shown in Figu re 2. It is de rived fro m a similar thin -section motor, th e Michel son  Tuning  Moto r (MTM ), that is presently b eing used in the Michael son Doppler Image r (MDI) on the  Sola r  and Helio spheric Ob servatory [1]. Thre e of these mech anisms have pe rform ed mo re tha n 70 millio n  operation s during a continu ing mission o f more than e ight years. Modified versio ns of the MDI  MTMs a re  also being successfully u sed in  the guid e telescope  on the Transiti on Region and Coronal  Explore r which  has b een in o rbit for more than 5 years. 

The HCM  has an overall fairly sim ple design. A 48- mm-dia meter ape rture i s surroun ded by a thin-se ction  beari ng an d a bru shless DC mot or. The motor consists of a three-p hase, wy e windi ng on a 72-p ole iro n  stator paired with a 48-ma gnet rotor. This de sign results in 144 discrete  dete nts with a static detent torq ue  of approxima tely 64 mN•m. A space i ndustry stan dard Ti mken  thin-sectio n beari ng with Teflon toroi d  separators was selected for this design. The bearing was deemed a low-risk choice because of its long  heritag e in space-based applicatio ns and  successful use in previou s LMSAL  mechanisms. In addition, th e  mech anism’s mode st life requi rement coul d easily be accommo dated with this type of be aring. Two   beari ngs were use d in a b ack-to -back hard pre-loaded condition to  approxim ately 100 N.   

Figures 2a & 2b. SECCHI Hollo w-Core Motor 

16
this instability. As can be seen, som e of the detent  positions are very repeatable, but the aggregate  repeatability is dominated by thos e target detents that are not. 

The po sition of this unstable   transitio n re gion vari es slightly   betwe en positions, a s shown in   Figure 8, but there is a wid e range  of stabl e del ay values co mmon to   all mot or positions. Th e mo st  repeatable  moveme nts o ccur  approximatel y at the middle of the  stable delay range. The high delay  setting s tha t cause consi stent  overshootin g of the  target  position  are al so stable, but are m uch less  repeatable. I n this case, o f course,  the me chanism al so stop s  approximatel y 2.5° from its  intende d position. 05010015020025010356085110135 -3-2-10123Position Offset,  degrees (from  2.5° sp acing) Delay SettingDetent

Figure 7.  Stopping Positions for 12 0° Rota tions to  Six HCM  Targe t Detents While Var ying Dela y Setting   

-3-2-10123 0 50 100 150 200 250 Delay  SettingPosition Offset  (degrees fro m referen ce point) Most Repeatabl e Delay  Setting and Final Pos ition Under shoot F inal PositionOvershoot F inal Position 10-11 Detent Transition 9-10 Detent Transition

Figure 8.  An nota ted Plot of Stopping Positions for 120° Rota tions to Targ et Detent 1 0. Dela y  Setting Varie d by 2 from 0 to 254. 

Mechanism Qualification Testing 

The SECCHI HCM was qualified a nd its performan ce characte rized throug h lifetime sim ulation testin g a  single me chanism. T his life-test consisted of construction and testing of the me chanism to flight  specification s, vibration testing, therm al functional  testing at am bient pressu re to the me chanism’s  operating te mperature ex treme s, operation in vacu um to  four times the e stimated missio n life, post-te st  function al testing, and  disassembly and inspectio n. Figure 9 shows the sequence an d requirements for a   single sequence.   

Vibration – GEV S  2x Du ration Thermal C ycling  -10° to 5 0°C Vacuum Ope ration  > 2 Million O perationsFunctional Testing /  Disassembly  Assembly  / Brea k-in /  Functional Testing  Figure 9.  HCM Qualific ation Ba sic Sequen ce  20
Two life -test sequences, both discu ssed here, were  required to qualify the m echanism. During  the first   sequence, co ntamination  in the motor beari ng re sulted in a stall in the vacuum operation portion  of the  test. In the secon d sequence, the me chanism p erformed well th rougho ut the test. 

HCM T hermal Function al Testing 

After a stan dard vib ration test se quence, a therm al function al test was completed to characteri ze th e  perfo rman ce of the mecha nism ove r the extreme s of its operatin g tempe ratures. Althoug h the nomin al  predi cted t emperature  range fo r the HCM  is 20 to 30°C, the  mechani sm was op erated ove r five  temperature cycle s from –10 to 50°C while perfo rming movem ents simila r to those expe cted o n-orbit.  Periodi c automated  funct ional te sts that me asured the motor frictio n and stea dy stat e spin-sp eed  provide d information on the  chan ges in t he mechani sm at different temperature s.  

Figure 10 sh ows the cha nge in mech anism ag gregate friction ov er the tempe rature ra nge. In the figure,  the co mponents of fri ction have not b een separated and consist of changes in motor hy stere sis as well a s  changes in the friction prop erties of the b earin g. The result s from the thermal fun ctional test co nfirmed the   expecte d correlation b etween motor fri ction and te mperat ure. T he HCM perf ormance with  variations in  temperature  is typical of t he pe rforma nce of Bray cote grea se lubricated bearing s [4]. T he de creased  friction  at lo wer temp eratures  – with the atten dant increa se in runnin g speed – is typical to  grease  lubricants and indicate s that there is a thinner lay er of lubri cant bearing si nce the bea ring has forced   away the hig her visco sity Teflon bind er. Understan ding th is effect is criti cal to desig ning a tu ning sche me  if lower than anticip ated te mperature s are encountered duri ng on orbit ope rations. 4.04.55.05.56.06.57.0 -10 0 10 20 30 40 50 60 Temperature ( °C)Spin Speed (Rad / S)1st Thermal Cycle 2nd Therm al Cyc le 3rd Thermal Cycle 4th Th ermal Cyc le     02468101214 -10 10 30 50 Temperature (°C)Damping (oz-in / Rad / S)

Figure 10.  Mechanism S peed Depen dent Fric tion and   Nominal Voltage Spin Speed with Te mperatu re 

The extern al enco der system was u sed durin g the t hermal fun ctional test to cha racterize temperature   induced change in the repeatability of t he mechani sm and to prov ide guideline information for tuni ng the  0510152025 -10 0 10 20 30 40 50 60 Temper ature (°C)Aggregate Repeatability  (arcseconds)   020406080100120140 -10 0 10 20 30 40 50 60 Temp erature (°C)Optimum D elay Set ting Optimum De lay Min Acceptable Delay Max Acce ptable  Delay    Figure 11a & 11b.  HCM Repeatabilit y Dela y Setting with Temp eratu re  21
While un der vacuum, the mech anism p erform ed as expecte d for 2.8 million operatio ns – slightly more   than five tim es the expe cted life. As shown in Figure 14, the f ree-run speed profile of the  HCM in  the  second series was similar to that of the  first se ries and the minim um start -up voltage for the  mech anism  remai ned under th e goal of 7.5 V fo r the duration of the vacuum  run. The 1 Volt, or 21%, i ncrease in the  minimum sta rt-up voltage i ndicated a de crease in torq ue margin fro m 160% to about 115%.   

Upon compl etion of the v acuum p ortion of the second q ualification seri es, th e HCM was subjected to  function al testing for comp arison to the baseline pa rameters. Both pre - and po st-life-te st funct ional testing   was perfo rmed at 19 to 21° C.  

In most parameters, a s shown in Table 4, the  post-life me asure ments compared well with  the ea rly life  measurement s. Ho wever, the agg regate repeata bility of the mecha nism de creased by more than 200%.   The final valu e of 38 a rcseconds fell slightly outsi de of the requi rement that the  mech anism repeata bilityto-po sition be less th an 30 arcse conds. The free run speed also decreased over the lifetime  from a value   of 5.47 rad/ sec to 4.65 ra d/sec for a total  chan ge of 15 %. 

Subse quent repeata bility testing  with  the exte rnal encoder syste m elimin ated mea surement erro r as a  possibl e sou rce of the d iscrep ancy but also showed a drastic de cline in the re peata bility of the  mech anism f rom test-to-test  over the  next 60 0,000   operation s. A decrea se in  free-run speed wa s also n oted  durin g t his p eriod,  accomp anied by an incre ase  in the variability of the spinspeed over multiple  revolution s. 

The HCM was di sassem bled  and inspe cted. No damag e to  the moto r or the be aring balls  and ra ces was observe d, nor  was extensi ve wear to  the  Teflon toroid ball s eparators .  The Bray cote grea se lu bricant sho wed slight wea r characterized by small patch es of viscou s and slig htly  darken ed ma terial pu shed away from t he bea ring races. No p article contami nation like th at seen i n the  previou s HCM failure was found in the lubri cant. Table 4.  HCM Life Tes t Resul ts 

Parameter (Bench T est ~20°C) Pre-LT Post- LT Chan ge  Detent To rque + Start-up Friction  (mN•m) 67.1 77.7 15.8%  Speed Depen dent Friction  (mN• m / rad/sec) 10.0 17.2 71.8%  Drag  ( mN•m)  35.5 35.2 1.0%  Nominal Average  Spin Curren t  (mA) 106.2 106.6 0.4%  Minimum Spin Voltage  (V) 45.9 53.0 15.4%  Minimum Full Motor Start -up Volta ge  (V) 52.3 55.8 6.8%  Nominal Dela y Setting 80 100 25.0%  Repeatabilit y at Nominal Dela y Setting    (arcseconds, 1- σ) 12 38 216.7% 01234567 0.0 1.0 2.0 3.0 4.0 Operatio ns (millio ns)Speed (rad/sec)           0.01.02.03.04.05.06.07.08.0 0.0 1.0 2.0 3.0 4.0 Operations (millions )Voltage

Figure 14a & 14b.  HCM Qualifica tion Testing Res ults –  Stead y State  Spin Speed and Minimum Start-up Voltage  24
The me chanism was reassembl ed with  a new set of beari ngs and subj ected to  break-in sequen ces an d functio nal  testing identi cal to the previou s two   qualification  seri es. Th e mechani sm  perfo rmed as expecte d in all mea sured  parameters a nd sho wed an aggregate  repeatability of 13 arcseconds single  stand ard dev iation a nd was tu ned for  best perform ance at  a delay setting  of  100. 

Flight Mecha nism Accep tance Te sting  

Four HCM s were co nstructed an d tested for flight at  LMSAL. The testing of these mechan isms p rovide d  more  info rmation a bout t he characteristics of  the desi gn and  the relation ship between each of the  properties. The key ea rly-life testing results for the se mech anisms are presented in T able 5. As expe cted,  the re sults fo r the flight mech anisms were quite similar to those for the HCM developm ent testing. The   HCM s with  higher fri ction had l ower ste ady-state  spin spee ds an d high er minim um start-up v oltage s and  requi red required g reater delay setting s for optimu m repeata bility. All all of t he mech anisms excee ded the  repeatabililty and torq ue m argin requirements for the mech anism. Table 5.  Post-Life-Test A dditional Repeatabilit y Testing 

Ops Spin Spe ed  (Rad / sec) Spin σ Repea tability  (arcsec 1 - σ) Dela y  Settin g  134,360 5.47 0.05 12 80  3,683,930 4.65 0.12 38 100  3,760,952 4.34 0.13 62 120  3,854,396 3.71 0.14 80 150  3,951,228 3.64 0.14 140 150  4,057,168 3.77 0.16 160 160  4,276,882 3.02 0.17 250 160  Table 5.  Flight Mech anism Pre-Shipment Tes ting Resul ts 

Parameter L20-FM L30-FM L40-FM L50-FM  Minimum Spin Voltage 5.9 6.3 5.5 5.5  Minimum Full-Motor Start- up  Voltage (V) 7.2 7.5 7.4 7.3  Damping (mN•m / rad/sec) 10.5 18.9 14.3 14.1  Drag (mN•m) 20.1 23.2 2.8 18.5  Spin Current  (mA) 97 103 92 92  Nominal Free -Run Speed ( rad / se c) 6.8 4.0 6.0 5.6  Nominal Dela y Setting (20° C) 30 130 50 90  Aggregate Rep eatability  at Nomi nal  Delay Setting 19 14 17 15  Total Ops Prio r to Shipment 254,000 281,000 284,000 324,000 020406080100120 0.0 1.0 2.0 3.0 4.0 Total Op erations (millions)Repeatability  (arcseconds, 1-sigma)Aggregat e  Repe atability  Requ irement 30 Arcsecs1x Lifetime .8M test,  .6M on-orbi t2x Lifetime .8M test,  1.2M on-orbi t    0123456 0.01.02.03.0 4.0 Total Operations (millions)Steady-S tate S pin Speed (Rad / S )1x Lifetim e .8M test,  .6M on-orbit2x Lifetim e .8M test,  1.2M on -orbit

Figure 15a & 15b.  Lifetime Repeatabilit y Measurements ( with Requir ements)   and Mech anism Speed  25
29 Design and Development of the Primar y and Secondary Mirror Deployment  Systems for the Cryogenic JWST 

Paul Reynolds*, Charlie Atkinson* and Larry Gliman* 

Abstract 

With a 7-meter primary mirror (PM) aperture, the James Webb Space Telescope (Figure 1) will require  structures that remain stable to levels on the order of 10 nanometers out of plane under dynamic and  thermal loading while operating at cryogenic temperat ures. Moreover, the JWST will be the first telescope  in space to deploy primary and secondary mirrors. Th e resulting primary mirror (PM) aperture will not only  be segmented, but will have hinge-lin es and associated latc hes. The secondary mi rror will be deployed  with folding booms that latch to support it approximately 7 m away from the PM. This paper describes the  design of the JWST Optical Telescope Element (OTE) structures and mechanisms, focusing primarily on  the primary and secondary mirror deployment system s. It discusses the driving design requirements, how  the resulting designs satisfy those requirements, and how the risk associated with these very large,  stable, deployed structures was reduced through development and testing of the Development Optical  Telescope Assembly (DOTA).  

Introduction 

The James Webb Space Telescope (JWST) is a 7-m cryogenic telescope with near and mid-infrared  instruments for imaging and spectroscopy. It will be  used to help understand the shape and chemical  composition of the universe, and the evolution of galaxies, stars and planets. The program held its  Systems SRR in December of 2003, with PDR sc heduled for 3/06, CDR in 3/07 and launch in 2011.  NASA’s Goddard Space Flight  Center runs the JWST project out  of Greenbelt, Mary land and is also  responsible for delivering the Integr ated Science Instrument Module (ISI M) to the observatory. Northrop  Grumman Space Technologies (NGST) is the prime c ontractor and is teamed with: Ball Aerospace, who  will provide the telescopes optics and wavefront se nsing and control system, Kodak, who will integrate  the optics onto the telescope struct ure assembly, and Alliant Techsystems, who will design and build the  OTE’s precision backplane structure and Secondary Mi rror Support Structure (SMSS) struts. As the prime  contractor, NGST has overall program responsibility and is responsible for the design and fabrication of                                                    * Northrop Grumman Space Technology, Redondo Beach, CA 

Proceedings of the 37th Aerospace Mechanisms Symposium, Johnson Space Center, May 19-21, 2004  Figure 1.  James Webb Space Telescope
30 the OTE, spacecraft, and sunshield a nd the overall integratio n of the observatory. As part of NGST’s  responsibilities we are designing and building the deployment systems for the OTE. The OTE has 4  principal deployments: a tower deploy ment that thermally isolates t he cryogenic telescope from the warm  spacecraft and sunshield, two primary mirror wing deployments, and a secondary mirror deployment.  

This paper will focus on the development of the deploy ment systems for the primary mirror wings and the  secondary mirror. The discussion will center on some of the unique design challenges including  deployment and operation at cryogenic temperatur es, and the need for nanometer level stability over  extended observations and after slewing from one tar get to the next. It will also discuss in detail the  development and testing of the DOTA structures and mechanisms, how the DOTA designs relate to the  JWST designs, and how the DOTA test results relate to the JWST requirements.  

Requirements and Design Drivers 

The JWST observatory will orbit the second Lagrangian point,  L2, which is located 1.5 million k ilometers from earth, in line  with the earth and sun (Figure 2). The reason for this choice of orbit is the desire to passively cool the telescope to cryogenic temperatures with a deployable sunshield that  blocks light from the Sun, Earth and Moon. The telescope  needs to be at cryogenic temper atures to prevent it from  emitting more radiation than the infrared light from faint and  very distant objects.    One of the primary drivers to shape the JWST architecture  was the need to configure a large primary mirror that would  stow along with its associated secondary mirror, instrument suite, sunshield, and spacecraft, within the volume provided  by a medium Evolved Expendable Launch Vehicle (EELV). Essential to this objective was the need to keep deployments as low risk as possible and create optics supp ort structures stable to nanometer levels.    Figure 3 shows the  deployed observatory  with some of its critical dimensions. The primary  mirror is 7 m in diameter  and the secondary mirror is 7.2 m forward of the  primary mirror vertex.  The sunshield, which passively cools the  telescope, is  approximately 25 m long by 10 m wide. A  telescoping deployment  tower extends 1.5m to separate the OTE from  the spacecraft, resulting  in a total observatory height of over 10m. 

Since the EELV static envelope is 4.2 m in diameter , the primary mirror, secondary mirror, sunshield,  solar arrays, radiator shades and High Gain Antenna (HGA) needed to be folded or compressed to fit within the allowable volume. The OTE to Spacecraft Deployment Tower raises the OTE off of its launch  JWST  Figure 2.  JWST Orbit Around L2 SMSS7m Primar y Mirro r  Sunshield Solar ArrayHGA Deployment Towe rXZ Y7.2m Figure 3.  Deployed JWST  >10m
31 lock attachment to the spacecraft and also prov ides thermal separation. Figure 4 shows the stowed  observatory with the two wings and SMSS  folded to provide clearance to the Atlas V  static envelope.    The launch stiffness requirements for the  observatory are specified in the EELV users guide. Based on the requirements and  maturity of the design, the observatory is  being designed to have a stowed natural frequency of 11 Hz. It must also be  designed to survive the EELV launch loads  listed in Table 1. 

Table 1.  EELV Inertial Launch Loads 

Case Axial Lateral  Maximum Axial 2.0 ±5.5 2.0 at OTE/SV  Interface 2.6 at OTE  Top Linear variation  between  Maximum Lateral 2.0 ±2.60 3.5 at OTE/SV  Interface 4.5 at OTE  Top Linear variation  between 

Once the observatory has separated from the launch vehicle, the solar arrays  and radiator shades are  deployed so the observatory can begi n generating power. This is follow ed by the deployment of the HGA.  The sunshield is deployed next and must deploy prio r to the OTE since it cocoons the OTE when it is  stowed. Next, the tower deploys the telescope away  from the spacecraft and sunshield for thermal  isolation. The planned deployment sequence has th e SMSS deploying next, followed by the +Y and –Y  primary mirror wings, though this order is not required. 

Nominally, the tower, SMSS and wings deploy directly a fter the sunshield, while the telescope is still well  above its cryogenic operating temperature. However, the OTE must be designed for any contingency  including delays that could result in structure and mechanism temperatures as low as 30K prior to  deployment. A combination of cryogenic deployment  motors, heaters and careful selection of the  mechanism materials will be used to make deploym ent at these very low temperatures possible.   Both the Primary Mirror Segment Assemblies (PMSA’ s) and the Secondary Mirror Assembly (SMA) have  active control to correct for deployment errors or di stortions that occur during cool down. However, the  amount of allowable deployment error is limited by t he wavefront sensing and control systems ability to  capture an initial image so it can make its correcti ons. The deployment repe atability requirements are  shown below in Table 2.   Table 2.  Deployment Repeatability Requirements 

JWST Requirement X Despace  (mm) Y Decenter  (mm) Z Decenter  (mm) Theta Z  (arcmin)  Primary Mirror Wings 1 0.1 0.1 1  Secondary Mirror 3 3 3 5 

Ø 4521.2mm  Static Envelope Figure 4.  Stowed Observatory in Atlas V