Updated ODAR

0021-EX-CM-2016 Text Documents

Astro Digital, Incorporated

2017-10-09ELS_199461

DocuSign Envelope ID: 48DB92AB-1488-468D-A7CD-5152E6733269


                                                  Corvus-BC ODAR – Version 1.5




                     Astro Digital Corvus-BC Orbital Debris Assessment
                                       Report (ODAR)

                                                 CorvusBC-ODAR-1.5




                 This report is presented as compliance with NASA-STD-8719.14, APPENDIX A.
                 Report Version: 1.0, 11/12/2015




                 Astro Digital US, Inc.

                 NASA Ames Research Park
                 Building 503
                 M/S-19-46L, P.O. Box 1
                 Moffett Field, CA 94034-0001




                 Document Data is Not Restricted. This document contains no proprietary, ITAR, or
                 export controlled information.



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                                                  Corvus-BC ODAR – Version 1.5



                 DAS Software Version Used In Analysis: v2.0.2


                 Astro Digital Corvus-BC Orbital Debris Assessment Report
                 CorvusBC-ODAR-1.0



                                                             APPROVAL:


                                                          Chris Biddy
                                                        CEO, Astro Digital


                                              ______________________________


                                                           Jan A. King
                                                         CTO, Astro Digital


                                               _______________________________________

                                                        Brian Cooper
                                                       Mission Manager
                                                    Landmapper-BC Program



                                              ________________________________________




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                                                                    Corvus-BC ODAR – Version 1.5




                                                                            Revision Record
                     Revision:                  Date:                      Affected Pages:                          Changes:                       Author(s):
                       1.0                   5/5/2017                      All –Initial                     DAS Software Results                 B. Cooper
                                                                                                            Orbit Lifetime
                                                                                                            Analysis




                                                                              Table of Contents

                 Self-assessment and OSMA assessment of the ODAR using the format in Appendix
                 A.2 of NASA-STD-8719.14: ............................................................................................................. 3
                 Comments .............................................................................................................................................. 4
                 Assessment Report Format: ........................................................................................................... 5
                 Landmapper Description: .............................................................................................................. 5
                 ODAR Section 1: Program Management and Mission Overview ....................................... 5
                 ODAR Section 2: Spacecraft Description ...................................................................................... 6
                 ODAR Section 3: Assessment of Spacecraft Debris Released during Normal
                 Operations ................................................................................................................................................ 10
                 ODAR Section 4: Assessment of Spacecraft Intentional Breakups and Potential for
                 Explosions. ………………………………………………………………………………………………………. 11
                 ODAR Section 5: Assessment of Spacecraft Potential for On-Orbit Collisions ............ 16
                 ODAR Section 6: Assessment of Spacecraft Post-mission Disposal Plans and
                 Procedures ............................................................................................................................................... 17
                 ODAR Section 7: Assessment of Spacecraft Reentry Hazards ........................................... 19
                 ODAR Section 8: Assessment for Tether Missions................................................................... 20
                 Raw DAS 2.0.2 Output ……………………………………………………………………………………… 21
                 Appendix A: Acronyms ....................................................................................................................... 30




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                 Self-assessment of the ODAR using the format in Appendix A.2 of NASA-STD-
                 8719.14:

                 A self assessment is provided below in accordance with the assessment format
                 provided in Appendix A.2 of NASA-STD-8719.14.




                 Note 1: The primary payloads for all launch missions belong to other organizations. This is not a
                 primary mission of Astro Digital. All other portions of the launch composite are not the
                 responsibility of Astro Digital and the Landmapper Program is not the lead launch organization.

                 Assessment Report Format:

                 ODAR Technical Sections Format Requirements:

                 Astro Digital US, Inc is a US company. This ODAR follows the format in NASA-STD-
                 8719.14, Appendix A.1 and includes the content indicated as a minimum, in each of
                 sections 2 through 8 below for the Corvus-BC satellites. Sections 9 through 14 apply
                 to the launch vehicle ODAR and are not covered here.


                 Landmapper-BC constellation (Corvus-BC) Space Mission Program:

                 ODAR Section 1: Program Management and Mission Overview

                 Program Mission Manager: Brian Cooper

                 Senior Management: Chris Biddy

                 Foreign government or space agency participation: None.


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                                                  Corvus-BC ODAR – Version 1.5



                 Summary of NASA’s responsibility under the governing agreement(s): N/A


                 Schedule of upcoming mission milestones:

                     •   Shipment of spacecraft scheduled October 27, 2017
                     •   Launch: scheduled for November 28, 2017

                 Mission Overview: Landmapper-BC is a remote sensing satellite constellation
                 consisting of 10 satellites built from Astro Digital’s Corvus-BC satellite bus. These
                 satellites will be launched into a sun synchronous orbits (SSO) with an average
                 altitude of 550 +/- 50 km. The satellites are designed to the CubeSat standard: 6U XL
                 for Corvus-BC bus. The satellite will be contained in a Quadpack 6U Deployer. These
                 deployers are to be included on-board a variety of launch vehicles, including the
                 SpaceX Falcon 9, Rocket Lab Electron, Glavkosmos Soyuz, Antrix PSLV, and
                 European Space Agency Vega.

                 Each Corvus-BC spacecraft carries three separate cameras to gather 22 meter
                 resolution imagery in the Red, Green, and Near-Infrared spectral bands. This
                 imagery is processed on-board and then downlinked over a miniaturized high-speed
                 Ka-band transmitter. The satellite bus uses reaction wheels, magnetic torque coils,
                 star trackers, magnetometers, sun sensors, and gyroscopes to enable precision 3-
                 axis pointing without the use of propellant.

                 Launch Vehicles and Launch Sites: Falcon 9, Vandenberg AFB, United States.
                 Electron, Mahia, New Zealand. Soyuz, Baikonur Cosmodrome, Kazakhstan. PSLV,
                 Satish Dhawan Space Centre, India. Vega, Centre Spatial Guyanais, French Guyana.

                 Proposed Initial Launch Date: November 28, 2017

                 Mission Duration: The anticipated lifetime of the spacecraft (pl.) is ≥ 5 year in LEO.

                 Launch and deployment profile, including all parking, transfer, and
                 operational orbits with apogee, perigee, and inclination: The selected launch
                 vehicle will transport multiple mission payloads to orbit. The Corvus-BC spacecraft
                 will be deployed into a sun synchronous low Earth orbit. Once the final stage has
                 burned out, the primary payloads will be dispensed. After the primary payloads are
                 clear, the secondary payload will separate. The Corvus-BC spacecraft will deploy a
                 UHF antenna and two solar panels once deployed from the deployer. The
                 spacecraft will decay naturally from operational orbits within the following ranges:

                 Average Orbital Altitude: 550 km +/- 50 km

                 Eccentricity: 0.0000 to 0.0033



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                 Inclination: 97.6° to 97.9°

                 Corvus-BC satellites do not currently include propulsion and as such will not
                 conduct end of life deorbit maneuvers. All Corvus-BC satellites will be launched into
                 low enough orbits such that they will decay and reenter naturally within 25 years
                 after end of life.

                 ODAR Section 2: Spacecraft Description:

                 Physical description of the spacecraft:
                 Corvus-BC is based on the 6U XL CubeSat form factor. Basic physical dimensions are
                 366 mm x 239 mm x 113 mm with a mass of approximately 11.5 kg. The
                 superstructure is comprised of six rectangular plates forming the sides of the
                 structure with interior stiffening members. There are L rails along each of the 366
                 mm corner edges. These accommodate the deployment of the satellite from the
                 deployer. Additional stiffness is provided by various major module components
                 mounted within the spacecraft structure. These include the Imaging Payload, the
                 Ka-Band transmitter, the Attitude Control Module, and the Data and Power Module.

                 This spacecraft designs include a spring-loaded UHF antenna and two solar panels
                 that are deployed after jettison from the deployer by two independent burn wires
                 controlled by software timers via the flight computer. Power is locked away from all
                 spacecraft platform and payload components by means of redundant series
                 separation switches. These switches cannot be activated until the spacecraft
                 separates from the deployer structure. The spacecraft are depicted in Figure 1.




                                                 Figure 1: Corvus-BC Spacecraft.




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                 Total satellite mass at launch, including all propellants and fluids:
                 Corvus-BC: 11.0 kg +/- 1.0 kg

                 Dry mass of satellites at launch:
                 Corvus-BC: 10.5 kg +/- 1.0 kg

                 Description of all propulsion systems (cold gas, mono-propellant, bi-
                 propellant, electric, nuclear): None.

                 Identification, including mass and pressure, of all fluids (liquids and gases)
                 planned to be on board and a description of the fluid loading plan or
                 strategies, excluding fluids in sealed heat pipes: None

                 Fluids in Pressurized Batteries: None

                 The Corvus-BC satellite design uses four unpressurized standard COTS Lithium-Ion
                 battery cells in each spacecraft. The energy capacity of each battery is 12 W-Hrs.
                 The total capacity energy capacity per spacecraft is 48 W-Hrs.

                 Description of attitude control system and indication of the normal attitude of
                 the spacecraft with respect to the velocity vector: All Crovus-BC spacecraft will
                 be initially controlled by magnetic torque coils embedded in the fixed solar panels of
                 the spacecraft. These will be used to detumble the spacecraft to a low enough rate
                 such that the reaction wheels can take over and provide precision 3-axis attitude
                 control.

                     •    A sun pointing mode that is optimized for solar power generation from the
                          satellite. The spacecraft’s large fixed panel and deployable panel will be
                          oriented towards the sun. This mode will make use of magnetometers, sun
                          sensors, reaction wheels, and magnetic torquers to orient the spacecraft
                          correctly.
                     •    A targeted tracking mode, which will allow the Imager or Ka-Band antenna to
                          be directed at any location on the Earth’s surface. This mode is used for
                          taking multi-spectral imagery and for downlinking payload data to a Ka-
                          band ground station. This mode will make use of reaction wheels and a star
                          tracker to orient the spacecraft.

                 Description of any range safety or other pyrotechnic devices: None. The
                 spacecraft deploy their antennae and solar panels using a burn wire system. System
                 power is locked off during launch by two series and two parallel deployment
                 switches but, the Quadpack deployer prevents any form of premature deployment,




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                 in any case. The antenna and panel spring constants are very low and can be held in
                 place by hand.

                 Description of the electrical generation and storage system: Standard COTS
                 Lithium-Ion battery cells are charged before payload integration and provide
                 electrical energy during the eclipse portion of the satellites’ orbit. The batteries are
                 operated in an “all-parallel” arrangement that results in increased safety thanks to
                 natural voltage balancing between cells. A series of Triple Junction Solar Cells
                 generate a maximum on-orbit power of approximately 34 watts at the end-of-life of
                 the mission (5 years for calculation purposes). Typical operational mode for
                 Corvus-BC consumes 17 watts of power on average. The charge/discharge cycle is
                 managed by a power management system overseen by the Flight Computer.

                 Identification of any other sources of stored energy not noted above: None

                 Identification of any radioactive materials on board: None


                 ODAR Section 3: Assessment of Spacecraft Debris Released during Normal
                 Operations:

                 Identification of any object (>1 mm) expected to be released from the
                 spacecraft any time after launch, including object dimensions, mass, and
                 material: None.

                 Rationale/necessity for release of each object: N/A.

                 Time of release of each object, relative to launch time: N/A.

                 Release velocity of each object with respect to spacecraft: N/A.
                 Expected orbital parameters (apogee, perigee, and inclination) of each object
                 after release: N/A.

                 Calculated orbital lifetime of each object, including time spent in Low Earth
                 Orbit (LEO): N/A.

                 Assessment of spacecraft compliance with Requirements 4.3-1 and 4.3-2 (per
                 DAS v2.0.2)
                 4.3-1, Mission Related Debris Passing Through LEO: COMPLIANT
                 4.3-2, Mission Related Debris Passing Near GEO: COMPLIANT


                 ODAR Section 4: Assessment of Spacecraft Intentional Breakups and Potential
                 for Explosions.




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                 Potential causes of spacecraft breakup during deployment and mission operations:
                 A Lithium-ion battery cell failure is the only potential scenario that could potentially
                 lead to a breakup of the satellite.

                 Summary of failure modes and effects analyses of all credible failure modes
                 which may lead to an accidental explosion: The in-orbit failure of a battery cell
                 protection circuit could lead to a short circuit resulting in overheating and a very
                 remote possibility of battery cell explosion. The battery safety systems discussed in
                 the FMEA (see requirement 4.4-1 below) describe the combined faults that must
                 occur for any of seven (7) independent, mutually exclusive failure modes to lead to
                 such an explosion.

                 Detailed plan for any designed spacecraft breakup, including explosions and
                 intentional collisions: There are no planned breakups.

                 List of components which shall be passivated at End of Mission (EOM)
                 including method of passivation and amount which cannot be passivated:
                 Four (4) Lithium Ion Battery Cells

                 Rationale for all items which are required to be passivated, but cannot be due
                 to their design: None

                 Assessment of spacecraft compliance with Requirements 4.4-1 through 4.4-4:

                 Requirement 4.4-1: Limiting the risk to other space systems from accidental
                 explosions during deployment and mission operations while in orbit about
                 Earth or the Moon: “For each spacecraft and launch vehicle orbital stage employed
                 for a mission, the program or project shall demonstrate, via failure mode and effects
                 analyses or equivalent analyses, that the integrated probability of explosion for all
                 credible failure modes of each spacecraft and launch vehicle is less than 0.001
                 (excluding small particle impacts) (Requirement 56449).”

                 Compliance statement:

                 Required Probability: 0.001

                 Expected probability, Corvus-BC: 0.000

                 Supporting Rationale and FMEA details:

                 Battery explosion:

                 Effect: All failure modes below might result in battery explosion with the possibility
                 of orbital debris generation. However, in the unlikely event that a battery cell does
                 explosively rupture, the small size, mass, and potential energy, of these small


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                 batteries is such that while the spacecraft could be expected to vent gases, most
                 debris from the battery rupture should be contained within the spacecraft due to
                 the lack of penetration energy to the multiple enclosures surrounding the batteries.

                 Probability: Extremely Low. It is believed to be less than 0.01% given that multiple
                 independent (not common mode) faults must occur for each failure mode to cause
                 the ultimate effect (explosion).

                 Failure mode 1: Internal short circuit.

                 Mitigation 1: Protoflight level sine burst, sine and random vibration in three axes of
                 the spacecraft, thermal cycling of the spacecraft and extensive functional testing
                 followed by maximum system rate-limited charge and discharge cycles were
                 performed to prove that no internal short circuit sensitivity exists. Additional
                 environmental and functional testing of the batteries at the power subsystem
                 vendor facilities were also conducted on the batteries at the component level.

                 Combined faults required for realized failure: Environmental testing AND functional
                 charge/discharge tests must both be ineffective in discovery of the failure mode.

                 Failure Mode 2: Internal thermal rise due to high load discharge rate.

                 Mitigation 2: Battery cells were tested in lab for high load discharge rates in a
                 variety of flight-like configurations to determine if the feasibility of an out-of-control
                 thermal rise in the cell. Cells were also tested in a hot, thermal vacuum environment
                 (5 cycles at 50° C, then to -20°C) in order to test the upper limit of the cells
                 capability. No failures were observed or identified via satellite telemetry or via
                 external monitoring circuitry.

                 Combined faults required for realized failure: Spacecraft thermal design must be
                 incorrect AND external over-current detection and disconnect function must fail to
                 enable this failure mode.

                 Failure Mode 3: Excessive discharge rate or short-circuit due to external device
                 failure or terminal contact with conductors not at battery voltage levels (due to
                 abrasion or inadequate proximity separation).

                 Mitigation 3: This failure mode is negated by:

                 a) qualification tested short circuit protection on each external circuit,

                 b) design of battery packs and insulators such that no contact with nearby board
                 traces is possible without being caused by some other mechanical failure,

                 c) observation of such other mechanical failures by protoflight level environmental
                 tests (sine burst, random vibration, thermal cycling, and thermal-vacuum tests).


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                 Combined faults required for realized failure: An external load must fail/short-circuit
                 AND external over-current detection and disconnect function must all occur to
                 enable this failure mode.



                 Failure Mode 4: Inoperable vents.

                 Mitigation 4: Battery venting is not inhibited by the battery holder design or the
                 spacecraft design. The battery can vent gases to the external environment.

                 Combined effects required for realized failure: The cell manufacturer OR the satellite
                 integrator fails to install proper venting.

                 Failure Mode 5: Crushing

                 Mitigation 5: This mode is negated by spacecraft design. There are no moving parts
                 in the proximity of the batteries.

                 Combined faults required for realized failure: A catastrophic failure must occur in an
                 external system AND the failure must cause a collision sufficient to crush the
                 batteries leading to an internal short circuit AND the satellite must be in a naturally
                 sustained orbit at the time the crushing occurs.

                 Failure Mode 6: Low level current leakage or short-circuit through battery pack
                 case or due to moisture-based degradation of insulators.

                 Mitigation 6: These modes are negated by:

                     a) battery holder/case design made of non-conductive plastic, and

                     b) operation in vacuum such that no moisture can affect insulators.


                 Combined faults required for realized failure: Abrasion or piercing failure of circuit
                 board coating or wire insulators AND dislocation of battery packs AND failure of
                 battery terminal insulators AND failure to detect such failures in environmental
                 tests must occur to result in this failure mode.

                 Failure Mode 7: Excess temperatures due to orbital environment and high
                 discharge combined.

                 Mitigation 7: The spacecraft thermal design will negate this possibility. Thermal rise
                 has been analyzed in combination with space environment temperatures showing
                 that batteries do not exceed normal allowable operating temperatures under a


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                 variety of modeled cases, including worst case orbital scenarios. Analysis shows
                 these temperatures to be well below temperatures of concern for explosions.

                 Combined faults required for realized failure: Thermal analysis AND thermal design
                 AND mission simulations in thermal-vacuum chamber testing AND over-current
                 monitoring and control must all fail for this failure mode to occur.

                 Requirement 4.4-2: Design for passivation after completion of mission
                 operations while in orbit about Earth or the Moon:

                 ‘Design of all spacecraft and launch vehicle orbital stages shall include the ability to
                 deplete all onboard sources of stored energy and disconnect all energy generation
                 sources when they are no longer required for mission operations or post-mission
                 disposal or control to a level which can not cause an explosion or deflagration large
                 enough to release orbital debris or break up the spacecraft (Requirement 56450).”

                 Compliance statement: Corvus-BC includes the ability to fully disconnect the
                 Lithium Ion cells from the charging current of the solar arrays. At End-Of-Life, this
                 feature can be used to completely passivate the batteries by removing all energy
                 from them. In the unlikely event that a battery cell does explosively rupture, the
                 small size, mass, and potential energy, of these small batteries is such that while the
                 spacecraft could be expected to vent gases, most debris from the battery rupture
                 should be contained within the spacecraft due to the lack of penetration energy to
                 the multiple enclosures surrounding the batteries.

                 Requirement 4.4-3. Limiting the long-term risk to other space systems from
                 planned breakups: Compliance statement: This requirement is not applicable.
                 There are no planned breakups.

                 Requirement 4.4-4: Limiting the short-term risk to other space systems from
                 planned breakups: Compliance statement: This requirement is not applicable.
                 There are no planned breakups.


                 ODAR Section 5: Assessment of Spacecraft Potential for On-Orbit Collisions


                 Assessment of spacecraft compliance with Requirements 4.5-1 and 4.5-2 (per
                 DAS v2.0.2, and calculation methods provided in NASA-STD-8719.14, section
                 4.5.4):

                 Requirement 4.5-1. Limiting debris generated by collisions with large objects
                 when operating in Earth orbit:

                 “For each spacecraft and launch vehicle orbital stage in or passing through LEO, the
                 program or project shall demonstrate that, during the orbital lifetime of each


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                 spacecraft and orbital stage, the probability of accidental collision with space objects
                 larger than 10 cm in diameter is less than 0.001 (Requirement 56506).”

                 Corvus-BC
                 Large Object Impact and Debris Generation Probability: 0.00001; COMPLIANT.

                 Requirement 4.5-2. Limiting debris generated by collisions with small objects
                 when operating in Earth or lunar orbit:

                 “For each spacecraft, the program or project shall demonstrate that, during the
                 mission of the spacecraft, the probability of accidental collision with orbital debris and
                 meteoroids sufficient to prevent compliance with the applicable post mission disposal
                 requirements is less than 0.01 (Requirement 56507).”

                 Corvus-BC
                 Small Object Impact and Debris Generation Probability: 0.001239; COMPLIANT

                 Identification of all systems or components required to accomplish any post-
                 mission disposal operation, including passivation and maneuvering: None


                 ODAR Section 6: Assessment of Spacecraft Post-mission Disposal Plans and
                 Procedures


                 6.1 Description of spacecraft disposal option selected: The satellite will de-orbit
                 naturally by atmospheric re-entry. There is no propulsion system.

                 6.2 Plan for any spacecraft maneuvers required to accomplish post-mission
                 disposal: None

                 6.3 Calculation of area-to-mass ratio after post-mission disposal, if the
                 controlled reentry option is not selected:

                 Spacecraft Mass: 11.5 kg

                 Cross-sectional Area: 0.0875 m^2

                 (Calculated by DAS 2.0.2). Area to mass ratio: 0.0875/11.5 = 0.0076 m^2/kg

                 6.4 Assessment of spacecraft compliance with Requirements 4.6-1 through
                 4.6-5 (per DAS v 2.0.2 and NASA-STD-8719.14 section): Requirement 4.6-1.
                 Disposal for space structures passing through LEO:

                 “A spacecraft or orbital stage with a perigee altitude below 2000 km shall be disposed
                 of by one of three methods: (Requirement 56557)


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                 a. Atmospheric reentry option: Leave the space structure in an orbit in which natural
                 forces will lead to atmospheric reentry within 25 years after the completion of mission
                 but no more than 30 years after launch; or Maneuver the space structure into a
                 controlled de-orbit trajectory as soon as practical after completion of mission.

                 b. Storage orbit option: Maneuver the space structure into an orbit with perigee
                 altitude greater than 2000 km and apogee less than GEO - 500 km.

                 c. Direct retrieval: Retrieve the space structure and remove it from orbit within 10
                 years after completion of mission.”

                 Analysis:
                 The Corvus-BC satellite method of disposal is COMPLIANT using method “a.” The
                 spacecraft will be left in the orbit it was injected into via the launch vehicle (550 +/-
                 50 km circular sun-synchronous orbit) and will passively decay within 25 years of
                 end of mission (and within 30 years after launch). 600 km altitude represents the
                 worst case from an orbital duration and debris risk standpoint, so analysis was
                 performed for this orbital altitude. From a 600 km altitude sun-synchronous orbit
                 the simulation predicts reentering in approximately 7008 days (19.2 years) after
                 launch with orbit history as shown in Figure 2 (analysis assumes an approximate
                 random tumbling behavior). If the launch vehicle injects the satellite at a lower
                 altitude the on-orbit lifetime is reduced with the profile shown in Figure 2.




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                                       Figure 2: Corvus-BC1 & Corvus-BC2 Orbit History


                 Requirement 4.6-2. Disposal for space structures near GEO:
                 Analysis is not applicable.

                 Requirement 4.6-3. Disposal for space structures between LEO and GEO:
                 Analysis is not applicable.

                 Requirement 4.6-4. Reliability of Post-mission Disposal Operations:
                 Analysis is not applicable. The satellite will reenter passively without post mission
                 disposal operations within the allowable timeframe.


                 ODAR Section 7: Assessment of Spacecraft Reentry Hazards:




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                 Assessment of spacecraft compliance with Requirement 4.7-1: Requirement
                 4.7-1. Limit the risk of human casualty:

                 “The potential for human casualty is assumed for any object with an impacting kinetic
                 energy in excess of 15 joules:
                 a) For uncontrolled reentry, the risk of human casualty from surviving debris shall not
                 exceed 0.0001 (1:10,000) (Requirement 56626).”
                 Summary Analysis Results: DAS v2.0.2 reports that the Corvus-BC satellites are
                 COMPLIANT with the requirement. The critical values reported by the DAS software
                 are:

                     •    Demise Altitude = 54.5 km
                     •    Debris Casualty Area = 0.000000
                     •    Impact Kinetic Energy = 0.000000

                 This is expected to represent the absolute maximum casualty risk, as calculated
                 with DAS's limited modeling capability. The DAS Output Summary Follows:

                 -------------------------------------------------------------------------------------------------
                 05 04 2015; 18:18:35PM    Processing Requirement 4.3-1:
                      Return Status : Not Run

                 =====================
                 No Project Data Available
                 =====================

                 =============== End of Requirement 4.3-1 ===============
                 05 04 2015; 18:18:37PM     Processing Requirement 4.3-2: Return
                 Status : Passed

                 =====================
                 No Project Data Available
                 =====================

                 =============== End of Requirement 4.3-2 ===============
                 05 04 2015; 18:18:39PM     Requirement 4.4-3: Compliant

                 =============== End of Requirement 4.4-3 ===============
                 05 04 2015; 18:18:45PM     Processing Requirement 4.5-1:
                      Return Status : Passed

                 ==============
                 Run Data
                 ==============

                 **INPUT**

                          Space Structure Name = Corvus-BC
                          Space Structure Type = Payload


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                         Perigee Altitude = 600.000000 (km)
                         Apogee Altitude = 600.000000 (km)
                         Inclination = 98.000000 (deg)
                         RAAN = 0.000000 (deg)
                         Argument of Perigee = 0.000000 (deg)
                         Mean Anomaly = 0.000000 (deg)
                         Final Area-To-Mass Ratio = 0.007600 (m^2/kg)
                         Start Year = 2016.000000 (yr)
                         Initial Mass = 11.500000 (kg)
                         Final Mass = 11.500000 (kg)
                         Duration = 5.000000 (yr)
                         Station-Kept = False
                         Abandoned = True
                         PMD Perigee Altitude = -1.000000 (km)
                         PMD Apogee Altitude = -1.000000 (km)
                         PMD Inclination = 0.000000 (deg)
                         PMD RAAN = 0.000000 (deg)
                         PMD Argument of Perigee = 0.000000 (deg)
                         PMD Mean Anomaly = 0.000000 (deg)

                 **OUTPUT**

                         Collision Probability = 0.000005
                         Returned Error Message: Normal Processing
                         Date Range Error Message: Normal Date Range
                         Status = Pass

                 ==============

                 =============== End of Requirement 4.5-1 ===============
                 05 04 2015; 18:18:47PM     Requirement 4.5-2: Compliant
                 05 04 2015; 18:18:49PM     Processing Requirement 4.6 Return
                 Status : Passed

                 ==============
                 Project Data
                 ==============

                 **INPUT**

                         Space Structure Name = Corvus-BC
                         Space Structure Type = Payload

                         Perigee Altitude = 600.000000 (km)
                         Apogee Altitude = 600.000000 (km)
                         Inclination = 98.000000 (deg)
                         RAAN = 0.000000 (deg)
                         Argument of Perigee = 0.000000 (deg)
                         Mean Anomaly = 0.000000 (deg)
                         Area-To-Mass Ratio = 0.007600 (m^2/kg)
                         Start Year = 2016.000000 (yr)
                         Initial Mass = 11.500000 (kg)


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                         Final Mass = 11.500000 (kg)
                         Duration = 5.000000 (yr)
                         Station Kept = False
                         Abandoned = True
                         PMD Perigee Altitude = 596.566149 (km)
                         PMD Apogee Altitude = 596.566149 (km)
                         PMD Inclination = 97.920712 (deg)
                         PMD RAAN = 40.107088 (deg)
                         PMD Argument of Perigee = 333.302300 (deg)
                         PMD Mean Anomaly = 0.000000 (deg)

                 **OUTPUT**

                         Suggested Perigee Altitude = 596.566149 (km)
                         Suggested Apogee Altitude = 596.566149 (km)
                         Returned Error Message = Passes LEO reentry orbit criteria.

                         Released Year = 2035 (yr)
                         Requirement = 61
                         Compliance Status = Pass

                 ==============

                 =============== End of Requirement 4.6 ===============
                 05 04 2015; 18:18:54PM     *********Processing Requirement 4.7-1
                      Return Status : Passed

                 ***********INPUT****
                  Item Number = 1

                 name = Corvus-BC
                 quantity = 1
                 parent = 0
                 materialID = 5
                 type = Box
                 Aero Mass = 11.500000
                 Thermal Mass = 11.500000
                 Diameter/Width = 0.220000
                 Length = 0.366000
                 Height = 0.110000

                 name = Corvus-BC
                 quantity = 1
                 parent = 1
                 materialID = 5
                 type = Box
                 Aero Mass = 11.500000
                 Thermal Mass = 11.500000
                 Diameter/Width = 0.220000
                 Length = 0.366000
                 Height = 0.110000



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                 **************OUTPUT****
                 Item Number = 1

                 name =     Corvus-BC
                 Demise     Altitude = 77.994629
                 Debris     Casualty Area = 0.000000
                 Impact     Kinetic Energy = 0.000000

                 *************************************
                 name = Corvus-BC
                 Demise Altitude = 54.545510
                 Debris Casualty Area = 0.000000
                 Impact Kinetic Energy = 0.000000

                 *************************************

                 =============== End of Requirement 4.7-1 ===============

                 *************************************


                 Requirements 4.7-1b, and 4.7-1c:
                 These requirements are non-applicable requirements because Corvus-BC satellites
                 do not use controlled reentry.

                 4.7-1, b): “For controlled reentry, the selected trajectory shall ensure that no
                 surviving debris impact with a kinetic energy greater than 15 joules is closer than 370
                 km from foreign landmasses, or is within 50 km from the continental U.S., territories of
                 the U.S., and the permanent ice pack of Antarctica (Requirement 56627).”

                 Not applicable to Corvus-BC satellites. They do not use controlled reentry and no
                 debris is expected to survive.

                 4.7-1 c): “For controlled reentries, the product of the probability of failure of the
                 reentry burn (from Requirement 4.6-4.b) and the risk of human casualty assuming
                 uncontrolled reentry shall not exceed 0.0001 (1:10,000) (Requirement 56628).”
                 Not applicable to Corvus-BC1 & Corvus-BC2. They do not use controlled reentry and
                 no debris is expected to survive.


                 ODAR Section 8: Assessment for Tether Missions
                 Not applicable. There are no tethers used in the Corvus-BC satellites.

                 END of ODAR for Corvus-BC satellite

                 --------------------------------------------------------------

                 Appendix A: Acronyms


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                 Arg peri        Argument of Perigee
                 CDR             Critical Design Review
                 Cm              centimeter
                 COTS            Commercial Off-The-Shelf (items)
                 DAS             Debris Assessment Software
                 EOM             End Of Mission
                 FRR             Flight Readiness Review
                 GEO             Geosynchronous Earth Orbit
                 ITAR            International Traffic In Arms Regulations
                 Kg              kilogram
                 Km              kilometer
                 LEO             Low Earth Orbit
                 Li-Ion          Lithium Ion
                 m^2             Meters squared
                 ml              milliliter
                 mm              millimeter
                 N/A             Not Applicable.
                 NET             Not Earlier Than
                 ODAR            Orbital Debris Assessment Report
                 OSMA            Office of Safety and Mission Assurance
                 PDR             Preliminary Design Review
                 PL              Payload
                 ISIPOD          ISIS CubeSat Deployer
                 PSIa            Pounds Per Square Inch, absolute
                 RAAN            Right Ascension of the Ascending Node
                 SMA             Safety and Mission Assurance
                 Ti              Titanium
                 Yr              year




                 Appendix B: Orbit Lifetime Analysis Using Space Mission Analysis and Design*




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         * Wertz, J., Space Mission Analysis & Design, Version 4.




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Document Created: 2017-10-09 09:20:46
Document Modified: 2017-10-09 09:20:46

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