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.	
  	
  
	
  



Document Created: 2560-04-24 00:00:00
Document Modified: 2560-04-24 00:00:00

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