Initial ODAR

0064-EX-CN-2016 Text Documents

Cornell University

2016-09-16ELS_182095

            Cislunar  Explorers  Formal  Orbital   Debris  Assessment  Report  (ODAR)     Report  Version:  1.1   September  21st  2016      

The  Cislunar  Explorers  Formal  Orbital  Debris  Assessment  Report  (ODAR)  has  been  prepared  for   compliance  with  NASA-­‐STD  8719.14  and  NPR  8715.6A  for  submittal  as  part  of  the  Ground   Tournament  3  competition  of  the  CubeQuest  Challenge.       Submitted  by:       ____________________________________     ________________         Amil  Vira             Date         Cornell  University         Document  Preparer                                 ____________________________________     ________________         Kyle  Doyle             Date         Cornell  University         Lead  Engineer                                 ____________________________________     ________________         Dr.  Mason  Peck           Date         Cornell  University         Team  Leader       Concurrence  by:       ____________________________________     ________________         ?               Date       Approval  and  Risk  accepted  by:     ____________________________________     ________________         ?               Date        

DOCUMENT  HISTORY  LOG   _________________________________________________________________     Document  Revision   Effective  Date   Description     -­‐   09/21/2016   Initial  Release                                            

  Orbital  Debris  Assessment  Report  Self  Evaluation   Launch  Vehicle   Spacecraft   Requirement   Standard   Not   Not   Comments   #   Compliant   Incomplete   Non-­‐ Compliant   Incomplete   Compliant   Compliant   Compliant   4.3-­‐1.a   o   o   x   o   x   o   o   N/A   4.3-­‐1.b   o   o   x   o   x   o   o   N/A   4.3-­‐2   o   o   x   o   x   o   o   N/A   4.4-­‐1   o   o   x   o   x   o   o   See  Section  4   4.4-­‐2   o   o   x   o   x   o   o   See  Section  4   4.4-­‐3   o   o   x   o   x   o   o   N/A   4.4-­‐4   o   o   x   o   x   o   o   N/A   4.5-­‐1   o   o   x   o   x   o   o   N/A   4.5-­‐2           x   o   o   See  Section  6   4.6-­‐1(a)   o   o   x   o   x   o   o   N/A   4.6-­‐1(b)   o   o   x   o   x   o   o   N/A   4.6-­‐1(c)   o   o   x   o   x   o   o   N/A     4.6-­‐2   o   o   x   o   x   o   o   N/A   4.6-­‐3   o   o   x   o   x   o   o   N/A   4.6-­‐4   o   o   x   o   x   o   o   N/A   4.6-­‐5   o   o   x   o   x   o   o   N/A   4.7-­‐1   o   o   x   o   x   o   o   N/A   4.8-­‐1           x   o   o   N/A        

TABLE  OF  CONTENTS   ____________________________________________________________________________   1   INTRODUCTION   7   1.1   Purpose   7   1.2   Scope   7   1.3   Software  and  Models  Used   7   2   PROGRAM  MANAGEMENT  AND  MISSION  OVERVIEW   7   2.1   Personnel  and  Management   7   2.2   Mission  Milestones   8   2.3   Mission  Description   9   3   SPACECRAFT  DESCRIPTION   11   3.1   Physical  Description   11   3.2   Release  Mechanism   12   3.3   Propulsion  and  Attitude  Control   15   3.3.1   Spin  Stabilization   17   3.3.2   RCS   17   3.4   Power  Management  System   18   3.4.1   Solar  Panels   18   3.4.2   Batteries   19   3.4.3   Controller   19   3.5   Other   20   4   ASSESSMENT  OF  SPACECRAFT  DEBRIS  RELEASED  DURING  NORMAL   OPERATIONS   21   5   ASSESSMENT  OF  SPACECRAFT  INTENTIONAL  BREAKUPS  AND  POTENTIAL  FOR   EXPLOSIONS   21   6   ASSESSMENT  OF  SPACECRAFT  POTENTIAL  FOR  ON-­‐ORBIT  COLLISIONS   22  

7   ASSESSMENT  OF  SPACECRAFT  POSTMISSION  DISPOSAL  PLANS  AND   PROCEDURES   24   8   ASSESSMENT  OF  SPACECRAFT  REENTRY  HAZARDS  AND  HAZARDOUS   MATERIALS   24   9   ASSESSMENT  FOR  TETHER  MISSIONS   24   10   LAUNCH  VEHICLE  DESCRIPTION  AND  ASSESSMENT   24      

1 INTRODUCTION   1.1 Purpose   This  is  the  Orbital  Debris  Assessment  Report  (ODAR)  for  Cislunar  Explorers,  a  CubeQuest  payload.   The  purpose  of  this  report  is  to  assess  the  debris  generation  potential  and  the  mitigation  options.   This  ODAR  follows  the  format  in  NASA-­‐STD-­‐8719.14,  Appendix  A.1  and  includes  the  content   indicated  at  a  minimum  in  Sections  2  through  8  below  for  BioSentinel.  Sections  9  through  14  apply   to  the  launch  vehicle  ODAR  and  are  not  covered  here.  This  report  will  be  updated  as  necessary  in   accordance  with  NPR  8715.6A.  A  summary  of  the  requirements  and  the  compliances  is  located  in   a  table  in  section  11.   1.2 Scope   This  document  shows  compliance  of  Cislunar  Explorers  with  the  requirements  of  NPR  8715.6A,   “NASA  Procedural  Requirements  for  Limiting  Orbital  Debris”.  The  orbital  debris  assessment  covers   the  following  five  aspects  according  to  NASA-­‐STD  8719.14A:   ● Debris  generated  during  normal  operations   ● Debris  generated  by  explosion  or  intentional  breakup   ● Debris  generated  by  on-­‐orbit  collisions  during  and  after  mission  operations   ● Reliable  disposal  of  spacecraft  and  launch  vehicle  orbital  stages  after  mission  completion   ● Structural  components  impacting  the  Earth  following  post-­‐mission  disposal  by  atmospheric   reentry   1.3 Software  and  Models  Used   No  specific  debris  assessment  software  was  used  for  this  assessment,  due  to  the  unusual  orbits.   2 PROGRAM  MANAGEMENT  AND  MISSION  OVERVIEW   2.1 Personnel  and  Management   Headquarters  Mission  Directorate:  Space  Technology  Mission  Directorate   Program  Executive:  Dr.  Mason  Peck   Lead  Engineer:  Kyle  Doyle   Planetary  Protection/Orbital  Debris  Engineer:  Amil  Vira   Cube  Quest  Program  Manager:  James  Cockrell     Secondary  Payload  Integration  Manager:  Kevin  Sykes     Foreign  Government  or  Space  Agency  Participation:  None     Summary  of  NASA’s  responsibility  under  the  governing  agreement(s):  N/A      

2.2 Mission  Milestones   • September  2015:  CubeQuest  Ground  Tournament  1   • October  2015:  Phase  0  Safety  Review   • February  2016:  CubeQuest  Ground  Tournament  2   • May  2016:  Phase  I  Safety  Review   • June  2016:  Final  preparations  for  internal  CDR.   • Life  cycle  testing  of  propulsion  subsystem.   • Upgrades  to  ground  station.   • Optical  navigation  system  simulated  mission.   • “software  flatsat”  debugging.   • July  1st  2016:  Internal  critical  design  review  prior  to  fabrication  of  EDU   • July-­‐August  2016:  Engineering  Development  Unit  fabrication.   • “August  5th”  2016:  Submittals  for  CubeQuest  Ground  Tournament  3  due   • Several  sources  have  said  this  is  delayed  until  at  least  September.   • “September  7th”  2016:  CubeQuest  Ground  Tournament  3  Face  to  Face   • Would  be  delayed  until  at  least  October.   • August-­‐October  2016:  EDU  testing.   • October  2016:  Testing  of  EDU  completed.   • Fabrication  of  flight  units  commences.   • October  2016:  Phase  II  Safety  Review.   • December  2016:  Fabrication  of  flight  units  complete.     • February  3rd,  2017:  Submittals  for  CubeQuest  Ground  Tournament  4  due   • March  1st,  2017:  CubeQuest  Ground  Tournament  4  Face  to  Face   • March-­‐July  2017:  Development  of  mission  products.   • July  2017-­‐Launch:  Mission  rehearsals.   • 2017:  Integration  with  dispenser.   • NLT  30  days  prior  to  next  level  integration:  Phase  III  Safety  Review   • CubeQuest  in  space  portion  begins.  Present  final  safety  analysis  with  all  verification   methods  and  status.   • Obtain  final  panel  endorsement.   • February  1st,  2018:  Integrated  payload-­‐dispenser  delivery  to  KSC   • February  2017-­‐Launch:  Integration  of  stack  at  KSC,  storage,  pre-­‐launch.   • Fall  2018:  EM-­‐1  Launch   • CubeQuest  in  space  portion  begins.   • T+1  year:  Competition  ends      

2.3 Mission  Description   Cislunar  Explorers  is  competing  in  the  CubeQuest  Challenge,  which  is  sponsored  by  NASA’s  Space   Technology  Mission  Directorate  as  a  part  of  the  Centennial  Challenge  Program.  The  team  plans  to   compete  in  the  Lunar  Derby  in  pursuit  of  the  Lunar  Propulsion  and  Spacecraft  Longevity  prizes.   For  the  Lunar  Propulsion  Prize,  the  Cislunar  Explorers  spacecraft  must  achieve  a  verifiable  lunar   orbit.  The  Spacecraft  Longevity  Prize  will  be  judged  based  on  the  number  of  elapsed  days   between  the  first  and  last  confirmed  reception  of  a  1024-­‐bit  data  block  from  the  spacecraft.     Cislunar  Explorers  also  aims  to  demonstrate  the  viability  of  electrolysis  propulsion  for  spacecraft   with  a  special  emphasis  on  application  in  nanosatellites.  The  electrolysis  propulsion  system  will   separate  water  into  hydrogen  and  oxygen  gas,  which  will  then  be  combusted  resulting  in  a  Δv   between  650  and  800  m/s.  Other  goals  include  the  demonstration  of  passive  spin  stabilization,   optical  navigation,  and  a  3D  printed  nozzle  for  Technology  Readiness  Level  advancement.  In  order   to  operate,  the  two  halves  of  the  spacecraft  will  spin  about  their  major  axes,  and  the  water  in  the   propellant  tank  will  provide  a  viscous  damping  effect,  passively  stabilize  the  satellites’  spins.  The   satellites  will  be  able  to  determine  its  position  optically  by  taking  photographs  of  the  Sun,  Earth,   and  Moon.  Flight  software  will  analyze  the  appearance  of  these  bodies  in  the  photographs  and   use  the  data  in  conjunction  with  data  regarding  their  instantaneous  positions  to  triangulate  the   position  of  the  satellite.       The  Cislunar  Explorers  spacecraft  will  be  launched  as  one  of  several  secondary  payloads  on  the   Space  Launch  System  (SLS)  Block  I  from  Kennedy  Space  Center.  The  spacecraft  will  be  launched   during  the  SLS  Exploration  Mission  1,  which  is  scheduled  to  take  place  in  2018.  The  mission  to   achieve  lunar  orbit  is  just  over  one  month  in  duration,  with  an  extended  mission  in  lunar  orbit   lasting  no  longer  than  one  year  before  controlled  impact  into  the  lunar  surface.       Following  SLS  launch,  the  upper  stage  performs  a  Trans  Lunar  Injection  (TLI)  burn  placing  the   upper  stage  on  a  Trans  Lunar  trajectory.    The  Multi-­‐Purpose  Crewed  Vehicle  (MPCV)  then   separates  from  Interim  Cryogenic  Propulsion  Stage  (ICPS)  to  continue  its  lunar  flyby.    Once  the   MPCV  is  clear  of  the  ICPS,  the  ICPS  will  perform  a  disposal  maneuver.    At  this  point,  the  Secondary   Payload  Deployer  System  (SPDS)  sequencer  system  is  activated  and  will  deploy  Cislunar  Explorers   from  the  Dispenser  at  the  deployment  interval  negotiated  ahead  of  time.    Once  Cislunar  Explorers   is  clear  of  ICPS,  it  will  begin  a  preprogramed  activation  and  deployment  sequence  of  its  onboard   systems.     Cislunar  Explorers  will  tweak  its  trajectory  to  perform  a  gravitational  swingby  of  the  Moon,  with   the  intent  to  facilitate  a  second  lunar  encounter.  While  beyond  the  orbit  of  the  Moon,  Cislunar   Explorers  will  perform  additional  course  corrections  followed  by  a  lunar  orbit  injection.  The   spacecraft  will  eventually  circularize  to  a  lunar  orbit  of  no  greater  than  10,000  km  apogee.  The   precise  orbit  is  not  important  to  the  mission  as  the  goal  is  to  achieve  lunar  orbit  within  10,000  km;   there  is  no  scientific  component  of  the  mission.    

  Figure  1:  Deployment  Overview.     There  is  concern  over  potential  inadvertent  interaction  with  the  SLS  second  stage  or  other   secondary  payloads  immediately  after  deployment.  For  this  reason,  the  spacecraft  will  inhibit  its   boot  up  for  a  short  time  after  deployment  from  the  secondary  payload  dispenser.  Cislunar   Explorers  team  has  submitted  a  Safety  Data  Package  to  and  is  preparing  for  a  Phase  II  Safety   Review  with  the  SLS  Payload  Safety  Review  Panel,  to  assure  that  there  will  be  no  potential  for   inadvertent,  hazardous  interactions  with  SLS  or  other  secondary  payloads.     Figure  2:  Post-­‐Deployment  Trajectory  

3 SPACECRAFT  DESCRIPTION   3.1 Physical  Description   The  Cislunar  Explorers  spacecraft  is  a  6U  cubesat  which  splits  into  two  L  shaped  3U  spacecraft   after  separation  from  the  launch  vehicle.  The  spacecraft  are  called  Cislunar  Explorer  1  and   Cislunar  Explorer  2  (CE-­‐1  and  CE-­‐2).  The  total  mass  of  spacecraft  at  launch  is  14  and  the  total  dry   mass  of  the  spacecraft  at  launch  is  11  kg.  The  mass  of  propellant  on  each  3U  spacecraft  is  1.5  kg.   CE-­‐1  weighs  7  kg  and  CE-­‐2  weighs  7  kg.       Each  spacecraft  has  a  full  set  of  all  subsystems  and  operates  independently  of  the  other  after   splitting.  The  splitting  mechanism  consists  of  a  release  mechanism  held  in  unstable  equilibrium  by   a  burn  wire.  Once  the  release  mechanism  is  triggered,  CE-­‐1  and  CE-­‐2  will  be  separated  by  springs.   The  Spacecraft  will  deploy  radio  antennas  after  splitting.  The  spacecraft  will  have  solar  panels  on   approximately  80  percent  of  their  surfaces.  Figures  3  –  5  show  up  to  date  models  of  the  Cislunar   Explorers  Spacecraft.       Figure  3:  6U  Storage  Configuration     m     er   nis am b ch a  Ch Me ion se   u st lea mb Re C o

le   ozz r  N u ste Thr ste r    G as   e   an i old alv s  C C id  V a eno ld  G So l Co nk   l  Ta Fue 3 1u   e  p S pac G om   ies ter Ba t Pi   ry   ber R a sp   Figure  4:  3U  Internal  Layout  of  Components       Figure  5:  3U  Spacecraft  Dimensions   3.2 Release  Mechanism   Once  ejected  from  the  dispenser  the  satellite  shall  not  deploy  any  mechanisms  for  a  minimum  of   30  minutes.  After  this  time  the  single  6U  CubeSat  shall  split  into  two  3U  CubeSats  that  each  have   the  ability  to  complete  the  mission.    The  driving  force  behind  this  deployment  is  a  set  of  four   conical  springs  in  compression  located  between  the  two  3U  satellites.    A  release  mechanism  holds  

the  satellites  together  during  storage,  launch,  and  deployment.  This  mechanism  then  releases  the   two  satellites  when  a  command  is  received  from  the  flight  computer.         Figure  6:  Mechanism  Locked     Figure  6  shows  the  release  mechanism  in  the  locked  position.    A  single  length  of  high-­‐strength   cord  holds  the  arms  shown.  A  burn  wire  circuit  shall  be  wrapped  around  this  length  of  cord  to   sever  it  when  prompted  by  the  flight  computer.    To  avoid  a  single  point  of  failure  scenario,   multiple  burn  wire  circuits  shall  be  attached  to  the  length  of  cord  to  ensure  it  is  severed.    The   individual  burn  wire  circuits  shall  be  designed  to  run  using  the  batteries  on  either  of  the   satellites  in  the  event  that  one  of  the  batteries  fails.  The  cord  represents  a  single  point  of  failure   for  this  system,  which  is  why  it  and  all  other  components  of  this  mechanism  have  been  designed   with  a  factor  of  safety  of  at  least  3  for  the  expected  conditions;  this  is  more  than  double  the   factor  of  safety  of  1.4  required  in  the  Secondary  Payload  User’s  guide.         Figure  7:  Unlocking     Once  the  burn  wire  holding  the  release  mechanism  in  the  locked  position  is  severed,  the   mechanism  shall  begin  to  rotate  to  the  unlocked  position.  The  force  causing  this  rotation  is   provided  by  the  springs  that  are  also  responsible  for  the  separation.          

  Figure  8:  Separation     After  the  satellites  are  no  longer  attached  by  the  release  mechanism,  they  shall  pivot  about  each   other  at  the  end  of  the  satellites  furthest  away  from  the  release  mechanism  as  shown  in  Figure   8  This  allows  the  satellites  to  both  separate  as  well  as  spin-­‐up  to  the  desired  angular  velocity   required  for  the  spin-­‐stabilization  of  the  satellite.  By  using  the  energy  stored  in  the  springs,  the   satellite  is  able  to  spin-­‐up  without  using  propellant,  therefore  saving  it  for  use  later  on  in  the   mission.     The    reliability    and    safety    of    the    release    mechanism    has    been    evaluated    using    the    finite   element  method.  The  results  from  the  analysis  are  shown  below  in  Figures  9  and  10.    The   release  mechanism  exceeds  the  desired  factor  of  safety  of  3.       Figure  9:  Stresses  on  Release  Mechanism    

  Figure  10:  Stresses  on  Release  Mechanism   3.3 Propulsion  and  Attitude  Control   3.3.1 Propulsion  System   The  Cislunar  Explorers  spacecraft  will  make  use  of  an  electrolysis  propulsion  system  to  provide   the  necessary  Δv.  The  propulsion  system  will  produce  thrust  by  separating  water  into  a   combustible  mixture  of  hydrogen  and  oxygen  gas  and  then  combusting  the  gaseous  mixture.   Both  spacecraft  include  a  propellant  tank  that  holds  1.5  kg  of  inert  liquid  water  which  is  at  1  atm   at  launch  and  deployment.  This  much  propellant  is  expected  to  produce  a  Δv  of  650  m/s  and   only  417  m/s  is  required  to  achieve  lunar  orbit.  The  tank  is  made  of  two  Ti-­‐6Al-­‐4V  halves  welded   together  with  the  electrolyzers  inside.  It  is  designed  to  hold  a  maximum  pressure  of  150  psi  with   a  factor  of  safety  of  2.17.  We  consider  the  potential  hazard  of  propellant  leakage  mitigated  by   the  design  in  which  the  propellant  is  stored  inertly  and  at  low  pressure  until  after  deployment.     The  combustion  chamber  is  3D  printed  titanium  and  can  hold  a  maximum  pressure  of  1000  psi   with  a  factor  of  safety  of  2.06.  The  propulsion  system  also  includes  two  electrolyzes,  two   pressure  transducers,  a  solenoid  valve,  a  detonation  flame  arrestor,  and  a  3D  printed  titanium   nozzle.  The  fluid  loading  plan  for  this  system  consists  solely  of  filling  the  propellant  tank  with   liquid  water  prior  to  the  commencement  of  the  mission.     Attitude  control  is  done  primarily  by  the  Reaction  Control  System.  Undesirable  nutation  of  the   spin  axis  is  cause  by  the  imperfect  alignment  of  the  main  thruster  firing  axis  and  the  center  of   mass.  This  problem  is  addressed  by  the  water  in  the  propellant  tank,  which  provides  passive   spin-­‐stabilization  by  damping  theis  nutation.      

  Figure  11:  Schematic  of  Propulsion  Subsystem   3.3.2 Attitude  Determination,  Control,  and  Navigation  System   The  Cislunar  Explorers’  ADCNS  system  is  composed  of  three  Raspberry  Pi  camera  modules  and  a   gyroscope  for  position  and  attitude  determination,  one  cold-­‐gas  pulse  thruster  for  reorientation   maneuvers,  and  a  single  electrolysis  engine  for  navigation.  The  image  processing  required  to   extract  apparent  sizes  and  centroids  of  the  celestial  bodies  is  performed  on  the  Raspberry  Pi,   which  also  stores  an  onboard  ephemerides  table  and  the  spacecrafts’  control  logic.  The  optical   navigation  process  is  visually  depicted  in  Figure  12,  and  a  block  diagram  of  operations  is  in   Figure  13.  The  Raspberry  Pi  flight  computer  relays  this  data  along  with  the  telemetry  via  the   communication  subsystem,  which  provides  health  and  navigation  information  to  flight   controllers  and  receives  reorientation  and  navigation  commands.       Figure  12:  Optical  Navigation  Geometry    

  Figure  13:  Attitude  Control  Block  Diagram     3.3.3 Spin  Stabilization   The  spacecraft  are  passively  spin-­‐stabilized  by  propellant  sloshing  after  deployment  from  SLS   and  separation  of  the  3U  spacecraft.  There  are  no  rotating  parts  onboard.  Instead,  due  to  the   separation  mechanism,  the  spacecraft  spins  about  its  major  axis.  Reorientation  is  achieved  with   a  single  cold  gas  thruster  that  exerts  a  torque  affecting  the  spin  axis  depending  on  when  during   a  spin  cycle  it  is  pulsed.  Nutation  caused  by  the  reorientation  (or  by  any  other  disturbance  such   as  an  electrolysis  thruster  burn)  damps  out  due  to  the  influence  of  propellant  sloshing  in  the   tank.   3.3.4 Reaction  Control  System   The  Reaction  Control  System  (RCS)  is  the  primary  means  actuating  attitude  control.  It  consists  of   a  single  cold  gas  thruster  with  38  grams  of  carbon  dioxide  in  a  COTS  cylinder  manufacture  by   Leland  Ltd.  38g  of  carbon  dioxide  can  provide  up  to  2200  degrees  of  reorientation,  a  margin  of   5.1  times  the  360  degrees  we  require.  Details  are  provided  below  and  the  system  is  pictured  in   Figure  14.  The  MDP  is  955  psi  at  the  greatest  anticipated  stowed  temperatures  (over  140°C)     ·∙ Leland  Limited  86121z  co2  gas  cylinder  contains  38.0g  of  co2   • Pressure  vessel  at  850  psi  at  21°C   • Burst  pressure  of  7840  psi  -­‐  factor  of  safety  of  8.2  MDP.     • Certified  mil-­‐i-­‐45208a   ·∙ Lee  IEPA1221141H  valve,  factor  of  safety  1.67  MDP  proof,  2.51  MDP  ultimate   • Failure  mode  is  leakage  through  seal  after  elastomer  extrudes  through  seal,  not   burst   ·∙ Stainless  steel  tubing  with  a  maximum  pressure  of  3900  psi  for  a  factor  of  safety  of  4.08  MD   ·∙ Puncture  device  with  a  hydrostatic  minimum  test  of  7850  psi  for  a  factor  of  safety  of  8.22   MDP   ·∙ Well  tested  prototype,  see  Section  4.4.7  of  the  CubeQuest  Design  Document.   ·∙ Flight  heritage  expected  before  EM-­‐1   ·∙ Flight  heritage  expected  before  EM-­‐1  

    Figure  14:  Cold  Gas  Thruster  Assembly     3.4 Power  Management  System   Power  will  be  supplied  by  two  main  components,  Emcore  ZTJ  Photovoltaic  Cells  and  a  commercial   battery  pack  of  18650  batteries.  The  batteries  will  be  managed  using  a  GomSpace  p31u.  For  the   vast  majority  of  the  mission  life,  the  net  power  will  remains  positive  and  will  keep  the  battery  fully   charged.   3.4.1 Solar  Panels   Each  3U  CubeSat  will  have  578  cubic  centimeters  of  solar  cell  coverage  with  cells  on  each   surface.    The  two  terminal  triple  junction  GaAs  cells  are  almost  twice  as  efficient  (29.5  percent)   as  silicon  cells.  They  are  also  capable  of  delivering  4  times  the  voltage  when  compared  to  silicon   cells.    The  solar  cells  also  offer  an  extremely  low  solar  cell  density  of  84  mg/square  cm.  They  are   arranged  with  blocking  and  bypass  diodes  as  shown  in  Figure  15.  Characteristics  of  solar  cell   performance  are  provided  in  Figure  16.       Figure  15:  Solar  Cell  and  Diode  Arrangement    

  Figure  16:  Solar  Cell  Performance   3.4.2 Batteries   18650  cells  are  used  as  the  battery  source  for  each  3U  CubeSat.  Each  18650  cell  is  rated  for  3.7V   with  a  capacity  of  2600  mAh.  The  batteries  have  a  heritage  on  several  CubeSat  space  missions   speaking  to  their  ability  as  space  rated  batteries.  They  are  configured  as  a  7.4V,  2600  mAh  stack,   and  have  built  in  protection  against  over-­‐temperature,  over-­‐current  draw,  and  over-­‐charge.  The   batteries  are  stored  open  to  the  CubeSat  environment,  on  the  controller  shown  in  Figure  4-­‐21.  It   is  grounded  and  bonded  to  the  CubeSat  structure,  specifically,  to  one  side  of  the  water   propellant  tank  for  the  dual  purpose  of  acting  as  a  heat  sink  for  the  power  system  and  helping   keep  the  water  liquid  during  the  mission.     A  potential  hazard  would  be  overcharging  or  overheating  of  the  batteries  while  onboard   SLS/ICPS.  This  is  mitigated  by  using  the  recommended  18650  cells,  which  can  survive  the  storage   and  pre-­‐deployment  environments  described  in  the  SPUG  and  have  internal  protection  against   overcharging  and  overheating  due  to  charging.  The  NASA-­‐provided  trickle  charging  will  be   carefully  monitored  for  charge  and  thermal  status,  including  the  use  of  a  thermistor  and  a  diode   on  the  positive  circuit  leg.    Only  one  of  the  two  3U  spacecraft  is  to  be  trickle  charged,  using  the   provided  trickle  charging  apparatus.     The  batteries  have  been  qualified  by  NASA,  ESA,  and  JAXA  for  the  ISS.  Qualification  included   abuse  testing  as  well  as  destructive  testing.  The  batteries  have  internal  PTC  rings,  CID,  and   pressure  relief  disks.  In  the  event  of  overpressure,  the  CID  interrupts  the  battery  current  flow   and  causes  the  venting  disk  to  open.  Vented  products  are  primarily  carbon  dioxide.  Overcharge,   overdischarge  (cell  reversal),  overheating,  and  overcurrent  are  prevented  by  BPS  circuitry.   3.4.3 Controller   The  batteries  interface  with  a  GOM  Space  P31u  power  board.  Battery  power  is  fed  through  two   buck-­‐converters  that  supply  a  3.3V  at  5A  and  5V  at  4A  output  bus.  Both  the  battery  and  power   board  are  from  GOMspace  and  as  such  will  not  have  any  interface  issues.  The  board  contains  3   photo-­‐voltaic  inputs  that  allow  for  conversion  of  GaAs  solar  cell  power  of  up  to  30W.  Low  and   high  voltage  protection  is  embedded  to  protect  the  battery  as  it  charges.    The  power  board  can   also  operate  up  to  6  configurable  output  switches  and  has  interfaces  for  a  remove-­‐before-­‐flight-­‐ pin  and  separation-­‐switch.  It  includes  heaters  to  keep  the  batteries  within  operational   temperature  ranges.  The  system  is  inhibited  from  activating  during  ground  loading  and  flight  by   the  aforementioned  separation  switches  as  well  as  a  pre-­‐flight  activation  switch.  The  controller   is  interfaced  using  I2C  to  an  onboard  microcontroller.  It  provides  onboard  housekeeping  

measurements  such  as  temperature,  battery  voltage,  and  current  draw.  A  functional  block   diagram  and  a  physical  description  can  be  found  below  in  Figure  17.       Figure  17:  Physical  Description  and  Block  Diagram   3.5 Other   Other  than  the  power  system,  the  main  source  of  stored  energy  is  potential  energy  stored  in  the   springs  used  by  the  splitting  mechanism.  This  energy  will  be  expended  after  the  splitting   mechanism  is  triggered.  Kinetic  or  potential  kinetic  energy  is  not  stored  anywhere  else  on  the   spacecraft  as  there  are  no  reaction  wheels.     There  are  no  range  safety  or  pyrotechnic  devices.  There  are  no  radioactive  materials  on  board   either  CE-­‐1  or  CE-­‐2.      

  4 ASSESSMENT  OF  SPACECRAFT  DEBRIS  RELEASED  DURING  NORMAL   OPERATIONS   The  Cislunar  Explorers  spacecraft  will  not  release  any  debris  larger  than  1  mm  during  normal   operations.  They  will  split  apart  from  a  single  6U  unit  into  two  3U  spacecraft.  However,  we  do  not   consider  this  a  debris  release  as  both  are  functional  spacecraft  and  no  other  components  separate.   If  one  spacecraft  is  inactive  it  will  not  be  separated  but  retained  in  a  6U  configuration;  the  other  will   continue  to  operate  with  the  defunct  3U  unit  attached.  Requirements  4.3-­‐1  and  4.3-­‐2  in  NASA-­‐STD-­‐ 8719.14A    do  not  apply  because  the  spacecraft  will  not  enter  a  Low  Earth  Orbit  or  a   Geosynchronous  Earth  Orbit  during  the  mission.   5 ASSESSMENT  OF  SPACECRAFT  INTENTIONAL  BREAKUPS  AND   POTENTIAL  FOR  EXPLOSIONS   The  only  intentional  break  up  designed  is  splitting  of  the  6U  unit  into  two  3U  spacecraft.  This  event   will  occur  30  minutes  after  separation  from  the  launch  vehicle.  The  release  mechanism,  how  it   functions,  and  its  factor  of  safety  are  all  described  in  section  3.2  of  this  document.  This  break  up   will  not  produce  any  debris.     Failure  Mode  1:  Explosion  of  Pressurized  Vessels   The  cold  gas  thruster  system  contains  38  g  of  CO2  stored  at  a  designed-­‐and-­‐tested  margin  of   safety  of  >2.5  against  maximum  expected  pressure,  and  is  thus  a  low  risk  for  explosion.  The   electrolysis  propulsion  system  never  contains  more  than  a  small  amount  of  combustible  propellant   at  any  time.  The  propellant  is  stored  as  inert,  liquid  water.  Small  amounts  up  to  1  g  are  electrolyzed   at  any  time,  up  to  a  pressure  of  150  psi  with  a  factor  of  safety  greater  than  2.  We  therefore   consider  this  to  have  a  very  low  risk  of  any  explosion  and  a  low  amount  of  energy  for  a  potential   explosion  in  any  case.       Failure  Mode  2:  Failure  of  Splitting  Mechanism   The  splitting  mechanism  springs  do  not  store  energy  after  deployment  and  prior  to  activation  the   mechanism  stores  at  a  factor  of  safety  greater  than  3.2  over  maximum  design  stress.     Failure  Mode  3:  Batteries   Because  of  the  above  points,  we  consider  the  batteries  to  pose  the  most  significant  risk  of   explosion.  This  could  be  due  to  overcharge,  overheating,  or  short-­‐circuit.  The  risk  of  this  is   considered  minimal  because  the  batteries  have  internal  protection  against  overheating,  pressure   relief  disks,  and  current  interruption  devices.  Additionally,  danger  from  overheating,  cell  reversal,   overcurrent  and  overcharge/discharge  are  prevented  by  the  power  system  circuitry.  This  system  is   described  in  section  3.4  of  this  document.     Because  the  Cislunar  Explorers  are  a  secondary  payload,  NASA  SLS  will  be  responsible  for   calculating  the  integrated  probability  of  explosion  for  the  launch  vehicle.  The  Cislunar  Explorers’   probability  of  explosion  has  been  assessed  to  be  very  low.      

Passivation  at  end  of  mission:     There  are  no  components  that  are  required  to  be  passivated  but  cannot  be  passivated  due  to  their   design.  The  Cislunar  Explorers  are  compliant  with  requirement  4.4-­‐1  and  4.4-­‐2.  Stored  sources  of   energy  to  be  passivated  include:     • Batteries,  to  be  passivated  by  opening  the  solar  array  switches  and  run  the  flight  computer   and  communications  until  the  batteries  are  drained.   • Cold  gas  pressure  vessel,  to  be  passivated  by  opening  the  thruster  valve  until  expended.   • Water  propellant  tank  with  any  remaining  water  propellant  and  electrolyzed  gas.  Energy  is   stored  here  in  the  form  of  low  pressure  gas  as  well  as  the  potential  combustion  of  the   electrolyzed  oxyhydrogen  mixture.  To  be  passivated  by  firing  the  electrolysis  propulsion   thruster  several  times  to  reduce  the  pressure  of  electrolyzed  gas  remaining  in  the   propellant  tank.  Only  a  small  amount  is  ever  present  at  any  one  time.  Any  remaining  water   can  be  left  as  it  does  not  pose  a  stored  energy  hazard  without  being  electrolyzed.   • There  are  no  other  sources  of  stored  energy  (e.g.  no  reaction  wheels)  onboard.     Requirement  4.4-­‐3  is  not  applicable  because  no  debris  larger  than  10  cm  will  be  created  and  no   debris  will  be  released  in  Earth  orbit.  Requirement  4.4-­‐4  is  not  applicable  because  no  debris  will  be   produced  by  the  splitting  of  the  Cislunar  Explorers  spacecraft;  they  split  into  two  separate,   functional,  independent  spacecraft.   6 ASSESSMENT  OF  SPACECRAFT  POTENTIAL  FOR  ON-­‐ORBIT   COLLISIONS   The  spacecraft  are  3U  CubeSats,  which  are  very  small.  Additionally  they  will  be  in  lunar  orbit  where   there  is  practically  no  man-­‐made  debris  presence  and  only  a  few  ongoing  missions  compared  to  the   crowded  space  in  Earth  orbit.  Hence,  the  risk  of  collision  with  a  large  object  is  extremely  small.  The   launch  vehicle  will  pass  through  LEO  very  briefly,  resulting  in  practically  no  exposure  to  orbital   debris  still  in  orbit.  Prior  to  launch  a  Collision  On  Launch  Avoidance  (COLA)  analysis  of  the  launch   trajectory  will  be  performed  by  the  SLS  EM-­‐1  mission  to  ensure  that  it  does  not  intersect  with   existing  satellites  or  debris  objects  tracked  by  the  US  Space  Surveillance  Network.     Using  the  Micrometeoroid  Engineering  Model  supplied  by  the  NASA  Micrometeoroid  Environment   Office,  it  was  determined  that  the  probability  of  a  damaging  collision  with  small  objects  is   extremely  low.  As  shown  in  Figure  18,  the  flux  of  milligram  or  greater  micrometeorites  capable  of   preventing  postmission  disposal  is  well  below  0.01  over  the  course  of  the  one  year  mission.  The   probability  of  0.1  milligram  and  larger  micrometeorites  is  shown  below  in  Figure  19  and  is  also   below  0.01.  The  flux  in  both  of  these  figures  is  per  square  meter  of  surface  area;  the  combined   surface  area  of  the  Cislunar  Explorers  is  approximately  0.3  square  meters,  further  reducing  the   probability  of  collision.    

  Figure  18:  Flux  of  milligram  and  greater  sized  micrometeoroids  on  Cislunar  Explorers       Figure  19:  Flux  of  0.1  milligram  and  greater  sized  micrometeoroids  on  Cislunar  Explorers    

Requirement  4.5-­‐1  is  not  applicable  because  the  Cislunar  Explorers  spacecraft  will  not  be  in  Lower   Earth  Orbit.  The  Cislunar  Explorers  are  compliant  with  requirement  4.5-­‐2.   7 ASSESSMENT  OF  SPACECRAFT  POSTMISSION  DISPOSAL  PLANS  AND   PROCEDURES   Requirements  in  section  4.6  of  NASA-­‐STD-­‐8719.14A  do  not  apply  to  this  mission  because  the   Cislunar  Explorers  spacecraft  will  not  be  in  Earth  orbit.  The  spacecraft  will  orbit  the  moon  until  they   have  ran  out  of  the  allotted  amount  of  propellant  and  then  will  be  disposed  upon  the  surface  of  the   Moon  through  a  controlled  collision.  If  needed,  course  corrections  will  be  made  to  avoid  any   historically  significant  sites  on  the  Moon.   8 ASSESSMENT  OF  SPACECRAFT  REENTRY  HAZARDS  AND   HAZARDOUS  MATERIALS   This  section  addresses  ODAR  sections  7  and  7A  as  outlined  in  Appendix  A  of  NASA-­‐STD-­‐8719.14A.   There  will  be  no  procedures  for  mitigating  reentry  hazards  for  the  Cislunar  Explorers  mission.  The   Cislunar  Explorers  spacecraft  will  not  be  reentering  the  Earth’s  atmosphere  at  any  point  in  its   mission  and  poses  no  human  casualty  risk.  Assessment  of  hazardous  materials  for  the  purpose  of   measuring  risk  to  humans  will  also  not  be  necessary.    Requirements  in  section  4.7  of  NASA-­‐STD-­‐ 8719.14A  do  not  apply  to  this  mission.   9 ASSESSMENT  FOR  TETHER  MISSIONS   The  Cislunar  Explorers  spacecraft  does  not  include  any  tethers  in  its  design.  Requirements  in   section  4.8  of  NASA-­‐STD-­‐8719.14A  do  not  apply  to  this  mission.     10 LAUNCH  VEHICLE  DESCRIPTION  AND  ASSESSMENT   This  section  addresses  ODAR  sections  9  through  14  as  outlined  in  Appendix  A  of  NASA-­‐STD-­‐ 8719.14A.  The  Cislunar  Explorers  spacecraft  will  be  launched  as  a  secondary  payload  to  NASA’s   Exploration  Mission  1  which  is  scheduled  to  launch  in  2018.  The  launch  vehicle  for  this  mission  is   the  SLS  Block  1  rocket.  Since  the  Cislunar  Explorers  spacecraft  is  a  secondary  payload,  NASA  SLS  will   be  responsible  for  the  launch  vehicle  orbital  debris  assessment.      


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