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. 1

DocuSign Envelope ID: 48DB92AB-1488-468D-A7CD-5152E6733269 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 ________________________________________ 2

DocuSign Envelope ID: 48DB92AB-1488-468D-A7CD-5152E6733269 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 3

DocuSign Envelope ID: 48DB92AB-1488-468D-A7CD-5152E6733269 Corvus-BC ODAR – Version 1.5 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. 4

DocuSign Envelope ID: 48DB92AB-1488-468D-A7CD-5152E6733269 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 5

DocuSign Envelope ID: 48DB92AB-1488-468D-A7CD-5152E6733269 Corvus-BC ODAR – Version 1.5 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. 6

DocuSign Envelope ID: 48DB92AB-1488-468D-A7CD-5152E6733269 Corvus-BC ODAR – Version 1.5 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, 7

DocuSign Envelope ID: 48DB92AB-1488-468D-A7CD-5152E6733269 Corvus-BC ODAR – Version 1.5 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. 8

DocuSign Envelope ID: 48DB92AB-1488-468D-A7CD-5152E6733269 Corvus-BC ODAR – Version 1.5 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 9

DocuSign Envelope ID: 48DB92AB-1488-468D-A7CD-5152E6733269 Corvus-BC ODAR – Version 1.5 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). 10

DocuSign Envelope ID: 48DB92AB-1488-468D-A7CD-5152E6733269 Corvus-BC ODAR – Version 1.5 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 11

DocuSign Envelope ID: 48DB92AB-1488-468D-A7CD-5152E6733269 Corvus-BC ODAR – Version 1.5 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 12

DocuSign Envelope ID: 48DB92AB-1488-468D-A7CD-5152E6733269 Corvus-BC ODAR – Version 1.5 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) 13

DocuSign Envelope ID: 48DB92AB-1488-468D-A7CD-5152E6733269 Corvus-BC ODAR – Version 1.5 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. 14

DocuSign Envelope ID: 48DB92AB-1488-468D-A7CD-5152E6733269 Corvus-BC ODAR – Version 1.5 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: 15

DocuSign Envelope ID: 48DB92AB-1488-468D-A7CD-5152E6733269 Corvus-BC ODAR – Version 1.5 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 16

DocuSign Envelope ID: 48DB92AB-1488-468D-A7CD-5152E6733269 Corvus-BC ODAR – Version 1.5 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) 17

DocuSign Envelope ID: 48DB92AB-1488-468D-A7CD-5152E6733269 Corvus-BC ODAR – Version 1.5 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 18

DocuSign Envelope ID: 48DB92AB-1488-468D-A7CD-5152E6733269 Corvus-BC ODAR – Version 1.5 **************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 19

DocuSign Envelope ID: 48DB92AB-1488-468D-A7CD-5152E6733269 Corvus-BC ODAR – Version 1.5 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* 20

DocuSign Envelope ID: 48DB92AB-1488-468D-A7CD-5152E6733269 Corvus-BC ODAR – Version 1.5 * Wertz, J., Space Mission Analysis & Design, Version 4. 21


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