ODAR v5

0880-EX-CN-2018 Text Documents

Tyvak Nano Satellite Systems

2018-11-08ELS_219521

PREPARED BY NUMBER CAGE CODE 6J8J8 Kelly Cheng Systems Engineer APPROVALS TYPE Analysis Procedure Approved By: Richard Prasad DATE Lead Systems Engineer Final release date Concurrence: Austin Williams SUPERSEDES SPEC DATED NA Chief Technology Officer REV 05 TITLE: TYVAK-0085 Orbital Debris Assessment Report (ODAR) / End of Mission Plan (EOMP) All future revisions to this document shall be approved by the controlling organization prior to release. Page 1 of 41

REVISION SUMMARY RELEASE EFFECTIVE REV NO. DATE BRIEF DESCRIPTION/REASON FOR CHANGE PAGES 01 10/23/18 Initial draft for review All 2 10/23/18 TYVAK-0085/-0086/-0087 ODAR results consolidation All 3 10/24/18 Removed proprietary markings 1 4 11/1/18 Cleaned up artifact propulsion text 13, 17 Page 2 of 41

TABLE OF CONTENTS ORBITAL DEBRIS SELF-ASSESSMENT: TYVAK-0085/-0086/-0087 MISSION ...............6 1.0 PROGRAM MANAGEMENT AND MISSION OVERVIEW ........................................7 1.1 Program Management .......................................................................................................7 1.2 Mission Overview .............................................................................................................7 1.2.1 Mission Design and Development Milestones ........................................................7 1.2.2 Mission Overview ....................................................................................................7 2.0 SPACECRAFT DESCRIPTION .........................................................................................9 2.1 Physical Description of Spacecraft ...................................................................................9 2.1.1 Description of Propulsion Systems ........................................................................10 2.1.2 Description of attitude control system ...................................................................10 2.1.3 Description of normal attitude of the spacecraft with respect to the velocity vector ................................................................................................................................10 2.1.4 Description of any range safety or other pyrotechnic devices ...............................11 2.1.5 Description of the electrical generation and storage system ..................................11 3.0 ASSESSMENT OF SPACECRAFT DEBRIS RELEASED DURING NORMAL OPERATIONS ....................................................................................................................12 4.0 ASESSMENT OF SPACECRAFT POTENTIAL FOR EXPLOSIONS AND INTENTIONAL BREAKUPS ...........................................................................................13 4.1 Potential causes of spacecraft breakup during deployment and mission operations ......13 4.2 Summary of failure modes and effects analysis of all credible failure modes ...............13 4.3 Detailed plan for any designed spacecraft breakup ........................................................13 4.4 List of components which shall be passivated at End of Mission (EOM)......................13 4.5 Rational for all items which are required to be passivated, but cannot be due to their design ..............................................................................................................................13 4.6 Assessment of spacecraft compliance with Requirements 4.4-1 through 4.4-4 .............13 5.0 ASSESSMENT OF SPACECRAFT POTENTIAL FOR ON-ORBIT COLLISIONS 16 5.1 Assessment of spacecraft compliance with Requirements 4.5-1 and 4.5-2: ...................16 6.0 ASSESSMENT OF SPACECRAFT POSTMISSION DISPOSAL PLANS AND PROCEDURES ...................................................................................................................17 6.1 Description of spacecraft disposal option selected .........................................................17 6.2 Plan for any spacecraft maneuvers required to accomplish postmission disposal: ........17 6.3 Calculation of area-to-mass ratio after postmission disposal: ........................................17 6.4 Assessment of spacecraft compliance with Requirements 4.6-1 through 4.6-5: ............17 6.5 Detailed plan for passivating (depleting all energy sources) of the spacecraft: .............20 7.0 ASSESSMENT OF SPACECRAFT REENTERY HAZARDS ......................................21 7.1 Assessment of spacecraft compliance with Requirement 4.7-1: ....................................21 8.0 ASSESSMENT FOR TETHER MISSIONS ....................................................................22 APPENDIX A – FMEA DETAILS AND SUPPORTING RATIONALE ..............................23 Battery Explosion Failure: .......................................................................................................23 APPENDIX B - REQUIREMENT 4.5-1 DAS 2.0.1 LOG ........................................................26 APPENDIX C - REQUIREMENT 4.6 DAS 2.0.1 LOG ...........................................................28 APPENDIX D - REQUIREMENT 4.7-1 DAS 2.0.1 LOG ........................................................31 Page 3 of 41

List of Figures Figure 2-1: Spacecraft Vehicle Layout ..................................................................................... 10 Figure 6-1: TYVAK-0085 Deorbit Lifetime ..............................Error! Bookmark not defined. Page 4 of 41

List of Tables Table 1-1: Summary of Program Management Personnel ....................................................... 7 Table 1-2: Summary of TYVAK-0085 Mission Parameters ......................................................... 8 Table 1-3: Summary of TYVAK-0086 Mission Parameters ......................................................... 8 Table 1-4: Summary of TYVAK-0087 Mission Parameters ......................................................... 8 Table 2-1: Summary of Spacecraft Parameters ........................................................................... 10 Table 3-1: Summary of Spacecraft Debris Released During Normal Operations ....................... 12 Page 5 of 41

ORBITAL DEBRIS SELF-ASSESSMENT: TYVAK-0085/-0086/-0087 MISSION Requirement Launch Vehicle Spacecraft Comments Standard Non Compliant Not Compliant Not Compliant Incomplete Incomplete Compliant Compliant 4.3-1.a X X No debris released in LEO 4.3-1.b X X No debris released in LEO 4.3-2 X X No debris released in GEO 4.4-1 X X Less than 0.001 probability 4.4-2 X X Design to electrical power system and reaction wheels 4.4-3 X X No planned breakups 4.4-4 X X No planned breakups 4.5-1 X X Probability 0.00000 (requirement < 0.001) 4.5-2 X X Probability 0.00000 (requirement < 0.01) 4.6-1(a) X X Predicted orbital lifetime 3.565 years 4.6-1(b) X X N/A – using atmospheric entry 4.6-1(c) X X N/A – using atmospheric entry 4.6-2 X X N/A – Not GEO 4.6-3 X X N/A – Not between LEO and GEO 4.6-4 X X Expected probability < 0.001 4.7-1 X X No pieces survive reentry 4.8-1 X No tethers used TYVAK-0085, TYVAK-0086 and TYVAK-0087 are currently manifested to fly as a secondary “rideshare” payload. Compliance with requirements levied by NASA-STD 9719.14A on the launch vehicle are not applicable to this document and the responsibility of the launch provider. Page 6 of 41

1.0 PROGRAM MANAGEMENT AND MISSION OVERVIEW 1.1 Program Management Parameter Value Mission Directorate N/A Program Executive Tom Yunck (GeoOptics) / Marco Villa (Tyvak) Program/project Manager Marco Villa (Tyvak) Senior Scientist Tom Yunck (GeoOptics) Senior Management N/A Foreign government or space agency N/A participation Summary of NASA’s responsibility under N/A the governing agreement(s) Table 1-1: Summary of Program Management Personnel 1.2 Mission Overview 1.2.1 Mission Design and Development Milestones The schedule of mission design and development milestones is provided in Table 1.2. TYVAK-0085 Launch January 12, 2018 TYVAK-0086 Launch November 11, 2018 TYVAK-0087 Launch November 26, 2018 Table 1.2 – Summary of Mission Design and Development Milestones 1.2.2 Mission Overview The goal of the TYVAK-0085/-0086/-0087 Missions are to perform GPS Radio Occultation (RO) measurement demonstration utilizing a single 6U CubeSat. The collection of RO data will be used to validate the TYVAK-0085/-0086/-0087 system and quality of data collected. Parameter Value Launch vehicle and launch site Sriharikota, India Launch date Q1 2018 Mission duration 2+ year Launch and deployment profile The PSLV launch vehicle will launch the primary mission satellite. After which, it will deploy the TYVAK-0085 satellite into their final mission orbit (504km, circular, sun-synchronous orbit 97.56° inclination). There is no parking or transfer orbit. Page 7 of 41

The TYVAK-0085 satellite will decay naturally for debris mitigation and will re-enter within 25 years after completion of mission. Table 1-2: Summary of TYVAK-0085 Mission Parameters Parameter Value Launch vehicle and launch site Mahia, New Zealand Launch date November 2018 Mission duration 2+ year Launch and deployment profile The Electron launch vehicle will launch and deploy the TYVAK-0086 satellite into its final mission orbit (500km, circular orbit, 85° inclination). There is no parking or transfer orbit. The TYVAK-0086 satellite will decay naturally for debris mitigation and will re-enter within 25 years after completion of mission. Table 1-3: Summary of TYVAK-0086 Mission Parameters Parameter Value Launch vehicle and launch site Sriharikota, India Launch date November 2018 Mission duration 2+ year Launch and deployment profile The PSLV launch vehicle will launch the primary mission satellite. After which, it will deploy the TYVAK-0087 satellite into their final mission orbit (505km, circular, sun-synchronous orbit 97.4° inclination). There is no parking or transfer orbit. The TYVAK-0087 satellite will decay naturally for debris mitigation and will re-enter within 25 years after completion of mission. Table 1-4: Summary of TYVAK-0087 Mission Parameters Page 8 of 41

2.0 SPACECRAFT DESCRIPTION 2.1 Physical Description of Spacecraft The TYVAK-0085/-0086/-0087 vehicles have been designed to support a 2+ year mission in LEO, and it is compatible with the P-POD launch environments and designed to the requirements in the CubeSat Design Specification (CDS). The TYVAK-0085/-0086/-0087 vehicles are 6U CubeSats with the vehicles being 30cm x 20cm x 10cm with a mass of 10.226 kg. The TYVAK-0085/-0086/-0087 vehicle design uses subsystem modules built from printed circuit boards (PCB) or miniature enclosures mounted to the open frame primary structure. The open structure permits the vehicle to be built incrementally with open access for securing interconnects. The body mounted side panels attach directly to the primary structure and are used for thermal management and can be easily removed to get access to the interior of the vehicle. The vehicle is primarily constructed out of aluminum and PCB materials. The TYVAK-0085/-0086/-0087 payload utilizes a GPS array antenna mounted to the Minus- Y panel to receive GPS RO signals. The signals are then captured by the CION Payload board which outputs the data files to the TYVAK-0085/-0086/-0087 CDH for storage and later transmission to the ground via X-Band of UHF transmission. Page 9 of 41

Figure 2-1: Spacecraft Vehicle Layout Parameter Value Total satellite mass at launch, including all propellants and fluids 10.226 kg Dry Mass of satellite at launch, excluding solid rocket motor propellants 10.226 kg Identification, including mass and pressure, of all fluids NONE. TYVAK-0085/-0086/-0087 have no propulsion Fluids in Pressurized batteries NONE. TYVAK-0085/-0086/-0087 use unpressurized standard COTS Li-ion battery cells Identification of any other sources of stored energy NONE Identification of any radioactive materials on board NONE Table 2-1: Summary of Spacecraft Parameters 2.1.1 Description of Propulsion Systems None. 2.1.2 Description of attitude control system The TYVAK-0085/-0086/-0087 attitude determination and control system consists of a processor, Inertial Reference Module (IRM), nano-Reaction Wheel Array (nRWA), GPS receiver, Sun sensors, magnetometers, and integrated torque coils. 2.1.3 Description of normal attitude of the spacecraft with respect to the velocity vector The nominal attitude of the TYVAK-0085/-0086/-0087 vehicles are in an LVLH orientation with the long axis (z-axis) facing Nadir towards the Earth and the Y-axis aligned with the velocity vector. The vehicles will rotate about the x-axis for pointing the boresight of the vehicle GPS array at the Earth limb during RO collections and will slew for sun-pointing periodically. Page 10 of 41

2.1.4 Description of any range safety or other pyrotechnic devices None. 2.1.5 Description of the electrical generation and storage system Energy generation is accomplished using the primary solar panels. Energy storage is accomplished using standard COTS Li-ion battery cells. The cells are recharged by the solar cells mounted on the deployable and body panels. The power management and distribution is provided by the electrical power system and battery protection circuitry. Page 11 of 41

3.0 ASSESSMENT OF SPACECRAFT DEBRIS RELEASED DURING NORMAL OPERATIONS No intentional release of any object > 1mm is expected. Parameter Value Identification of any object (>1mm) expected to be released from the None spacecraft at any time after launch Rationale/necessity for release of 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 of each object after release N/A Calculated orbital lifetime of each object N/A Compliance 4.3-1 Mission related debris passing through GEO COMPLIANT Compliance 4.3-2 Mission related debris passing through LEO COMPLIANT Table 3-1: Summary of Spacecraft Debris Released During Normal Operations Page 12 of 41

4.0 ASESSMENT OF SPACECRAFT POTENTIAL FOR EXPLOSIONS AND INTENTIONAL BREAKUPS 4.1 Potential causes of spacecraft breakup during deployment and mission operations There is no credible scenario that would result in spacecraft breakup during normal deployment and operations. 4.2 Summary of failure modes and effects analysis of all credible failure modes In-mission 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 (Appendix A, see requirement 4.4-1) describe the combined faults that must occur for any of seven (7) independent, mutually exclusive failure modes to lead to explosion. 4.3 Detailed plan for any designed spacecraft breakup There are no planned breakups. 4.4 List of components which shall be passivated at End of Mission (EOM) The nRWA will be passivated at EOM through a series of commands to reduce wheel momentum to a minimum level and then to transition the vehicle to free drift mode. The batteries will be passivated by discharging the cells to a minimum state and then disconnecting them from the solar panels and charging circuitry. 4.5 Rational for all items which are required to be passivated, but cannot be due to their design None. 4.6 Assessment of spacecraft compliance with Requirements 4.4-1 through 4.4-4 Page 13 of 41

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: 0.000 COMPLIANT 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 postmission 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: The batteries will be passivated by discharging the cells to a minimum state and then disconnecting them from the solar panels and charging circuit. In the unlikely event that a battery cell does explosively rupture, the small size, mass, and potential energy of these batteries is such that while the spacecraft could be expected to vent gases, most debris from the battery rupture would be contained within the vehicle due to lack of penetration energy and also because the cells are housed in a substantial aluminum bracket. The nRWA will be passivated at EOM through a series of commands to reduce wheel momentum to a minimum level and then to transition the vehicle to free drift mode. Requirement 4.4-3. Limiting the long-term risk to other space systems from planned breakups: Page 14 of 41

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. Page 15 of 41

5.0 ASSESSMENT OF SPACECRAFT POTENTIAL FOR ON-ORBIT COLLISIONS 5.1 Assessment of spacecraft compliance with Requirements 4.5-1 and 4.5-2: 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 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). Compliance statement: (Large Object Impact and Debris Generation Probability) Required Probability: 0.001 Expected probability: 0.000001 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 postmission disposal requirements is less than 0.01 (Requirement 56507). Compliance statement: (Small Object Impact and Debris Generation Probability) Required Probability: 0.01 Expected probability: 0.00000 COMPLIANT Page 16 of 41

6.0 ASSESSMENT OF SPACECRAFT POSTMISSION DISPOSAL PLANS AND PROCEDURES 6.1 Description of spacecraft disposal option selected The satellite will de-orbit naturally by atmospheric re-entry. 6.2 Plan for any spacecraft maneuvers required to accomplish postmission disposal: None. 6.3 Calculation of area-to-mass ratio after postmission disposal: Spacecraft Mass: ~10.226 kg (dry mass) Cross-sectional Area: 0.117516 m^2 Area to mass ratio: (0.117516 m^2)/(10.226kg) = 0.01149 m^2/kg 6.4 Assessment of spacecraft compliance with Requirements 4.6-1 through 4.6-5: 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) 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 Compliance statement: The orbit used for disposal of structure analysis is 504 km (TYVAK-0085), 500km (TYVAK- 0086) and 505km (TYVAK-0087). The worst-case orbital lifetime is predicted to be 3.565 years; COMPLIANT Page 17 of 41

Figure 6-1: TYVAK-0085 Deorbit Lifetime Figure 6-2: TYVAK-0086 Deorbit Lifetime Page 18 of 41

Figure 6-3: TYVAK-0087 Deorbit Lifetime Requirement 4.6-2. Disposal for space structures near GEO. Compliance statement: Not applicable. TYVAK-0085/-0086/-0087 mission orbit is LEO. Requirement 4.6-3. Disposal for space structures between LEO and GEO. Compliance statement: Not applicable. TYVAK-0085/-0086/-0087 mission orbit is LEO. Requirement 4.6-4. Reliability of Postmission Disposal Operations Page 19 of 41

Compliance statement: Not applicable. The satellite will reenter passively without the need for post mission disposal operations within the allowable timeframe. 6.5 Detailed plan for passivating (depleting all energy sources) of the spacecraft: The nRWA will be passivated at EOM through a series of commands to reduce wheel momentum to a minimum level and then to transition the vehicle to free drift mode. The free drift mode does not utilize any attitude control actuators, specifically the nRWA. The power service to the nRWA will also be deactivated so that no inadvertent switch to another attitude control mode can actuate the nRWA. The batteries will be passivated by permanently disconnecting solar array power from the battery module and discharging the cells to a minimum state under load of the spacecraft bus. Page 20 of 41

7.0 ASSESSMENT OF SPACECRAFT REENTERY HAZARDS 7.1 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). Compliance statement: DAS v2.0.2 reports that TYVAK-0085/-0086/-0087 is COMPLIANT with the requirement. The vehicle is primarily composed of Aluminum and PCB (Fiberglass) material and none of the components is expected to survive re-entry. The predicted Total Debris Casualty Area is 0.00. Appendix D located in the back of this report contains the DAS 2.0.2 modeling input and results. Requirement 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). Compliance statement: Not applicable. No controlled reentry planned. Requirement 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). Compliance statement: Not applicable. No controlled reentry planned. Page 21 of 41

8.0 ASSESSMENT FOR TETHER MISSIONS Not applicable. There are no tethers in the TYVAK-0085/-0086/-0087 missions. Page 22 of 41

APPENDIX A – FMEA DETAILS AND SUPPORTING RATIONALE Battery Explosion Failure: 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 batteries is such that while the spacecraft could be expected to vent gases, most debris from the battery rupture should be contained within the vessel due to the lack of penetration energy. The battery is housed within a substantial aluminum bracket. Probability: Very Low. It is believed to be less than 0.1% given that multiple independent (not common mode) faults must occur for each failure mode to cause the ultimate effect (explosion). Failure mode 1: Battery Internal short circuit. Mitigation 1: Qualification and acceptance tests include vibration, thermal cycling, and vacuum tests followed by maximum system rate-limited charge and discharge to prove that no internal short circuit sensitivity exists. Mitigation 2: Over/under voltage cell protection circuitry guards against stress conditions that can cause the development of internal shorts. Combined faults required for realized failure: Environmental testing AND functional charge/discharge tests must both be ineffective in discovery of infant mortality failure rate (IMFR) related faults OR protection circuitry malfunctions and fails to protect cells from stress conditions. Failure Mode 2: Internal thermal rise due to high load discharge rate. Mitigation 3: Each cell includes an internal positive temperature coefficient (PTC) variable resistance device that reduces discharge current as cell temperature increases to prevent thermal runaway. Mitigation 4: External under-voltage lockout circuitry disconnects battery when battery discharge voltage droop crosses a predefined threshold. Combined faults required for realized failure: Spacecraft thermal design must be incorrect AND internal AND external over current detection and protection must fail for this failure mode to occur. Failure Mode 3: Overcharging and excessive charge rate. Mitigation 5: The satellite bus battery charging circuit design eliminates the possibility of the batteries being overcharged if circuits function nominally. This circuit will be extensively bench-tested and be proto-qualified for survival in vibration, and thermal-vacuum environments. The charge circuit disconnects the incoming current when cell voltage indicates normal full charge at 4.2V and limits charge current within battery specification. If this circuit fails to operate, continuing or excessive charge current can cause gas generation. The batteries Page 23 of 41

include overpressure release vents that allow gas to escape, virtually eliminating any explosion hazard. Combined faults required for realized failure: 1) For overcharging: The charge control circuit must fail to limit charge voltage AND the PTC device must fail (or temperatures generated must be insufficient to cause the PTC device to modulate) AND the overpressure relief device must be inadequate to vent generated gasses at acceptable rates to avoid explosion. 2) For excessive charge rate: The charge control circuitry must fail to limit charge current AND the PTC device must fail (or temperatures generated must be insufficient to cause the PTC device to modulate) AND the overpressure relief device must be inadequate to vent generated gasses at acceptable rates to avoid explosion. Failure Mode 4: 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 6: This failure mode is negated by a) proto-qualification tested short circuit protection on each external circuit, b) design of battery packs and insulators such that no contact with nearby board traces or structure is possible without being caused by some other mechanical failure, c) obviation of such other mechanical failures by proto-qualification and acceptance environmental tests (shock, vibration, thermal cycling, and thermal-vacuum tests). Combined faults required for realized failure: The PTC must fail AND an external load must fail/short-circuit AND external over-current detection and disconnect function must fail to enable this failure mode. Failure Mode 5: Inoperable vents. Mitigation 7: Battery vents are not inhibited by the battery holder design or the spacecraft. Combined effects required for realized failure: The spacecraft design inhibits cell venting, or cell venting clearance is sensitive to environmental stress. Failure Mode 6: Crushing. Mitigation 8: This mode is negated by spacecraft design. There are no moving parts in the proximity of the batteries. Qualification and acceptance tests including vibration, thermal cycling, and vacuum tests will demonstrate cell venting clearance insensitivity to environmental stress. 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. Page 24 of 41

Failure Mode 7: Excess temperatures due to orbital environment and high discharge combined. Mitigation 9: The spacecraft thermal design will negate this possibility. Thermal rise will be analyzed in combination with space environment temperatures showing that batteries do not exceed normal allowable operating temperatures which are 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 the PTC device must fail AND over-current monitoring and control must all fail for this failure mode to occur. Failure Mode 8: Polarity Reversal Due to Over-Discharge Mitigation 10: The spacecraft battery chemistry (Li-ion) is not susceptible to polarity reversal due to over-discharge. Combined faults required for realized failure: Spacecraft battery module assembled with incorrect cell chemistry AND failure of cell protection circuitry Page 25 of 41

APPENDIX B - REQUIREMENT 4.5-1 DAS 2.0.1 LOG 09 26 2018; 14:24:12PM Processing Requirement 4.5-1: Return Status : Passed ============== Run Data ============== **INPUT** Space Structure Name = TYVAK-0085 Space Structure Type = Payload Perigee Altitude = 493.000000 (km) Apogee Altitude = 504.000000 (km) Inclination = 97.560000 (deg) RAAN = 0.000000 (deg) Argument of Perigee = 0.000000 (deg) Mean Anomaly = 0.000000 (deg) Final Area-To-Mass Ratio = 0.011490 (m^2/kg) Start Year = 2018.000000 (yr) Initial Mass = 10.226000 (kg) Final Mass = 10.226000 (kg) Duration = 25.000000 (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 = 0.000001 Returned Error Message: Normal Processing Date Range Error Message: Normal Date Range Status = Pass ============== =============== End of Requirement 4.5-1 =============== 10 22 2018; 18:26:12PM Processing Requirement 4.5-1: Return Status : Passed ============== Run Data ============== **INPUT** Space Structure Name = TYVAK-0086 Space Structure Type = Payload Perigee Altitude = 500.000000 (km) Apogee Altitude = 500.000000 (km) Inclination = 85.000000 (deg) RAAN = 0.000000 (deg) Argument of Perigee = 0.000000 (deg) Mean Anomaly = 0.000000 (deg) Final Area-To-Mass Ratio = 0.011490 (m^2/kg) Start Year = 2018.000000 (yr) Initial Mass = 10.226000 (kg) Final Mass = 10.226000 (kg) Page 26 of 41

Duration = 25.000000 (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 = 0.000001 Returned Error Message: Normal Processing Date Range Error Message: Normal Date Range Status = Pass ============== =============== End of Requirement 4.5-1 =============== 10 22 2018; 18:28:52PM Processing Requirement 4.5-1: Return Status : Passed ============== Run Data ============== **INPUT** Space Structure Name = TYVAK-0087 Space Structure Type = Payload Perigee Altitude = 505.000000 (km) Apogee Altitude = 505.000000 (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 = 0.011490 (m^2/kg) Start Year = 2018.000000 (yr) Initial Mass = 10.226000 (kg) Final Mass = 10.226000 (kg) Duration = 25.000000 (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 = 0.000001 Returned Error Message: Normal Processing Date Range Error Message: Normal Date Range Status = Pass ============== =============== End of Requirement 4.5-1 =============== Page 27 of 41

APPENDIX C - REQUIREMENT 4.6 DAS 2.0.1 LOG 09 26 2018; 14:29:43PM Processing Requirement 4.6 Return Status : Passed ============== Project Data ============== **INPUT** Space Structure Name = TYVAK-0085 Space Structure Type = Payload Perigee Altitude = 493.000000 (km) Apogee Altitude = 504.000000 (km) Inclination = 97.560000 (deg) RAAN = 0.000000 (deg) Argument of Perigee = 0.000000 (deg) Mean Anomaly = 0.000000 (deg) Area-To-Mass Ratio = 0.011490 (m^2/kg) Start Year = 2018.000000 (yr) Initial Mass = 10.226000 (kg) Final Mass = 10.226000 (kg) Duration = 25.000000 (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** Suggested Perigee Altitude = 493.000000 (km) Suggested Apogee Altitude = 504.000000 (km) Returned Error Message = Reentry during mission (no PMD req.). Released Year = 2022 (yr) Requirement = 61 Compliance Status = Pass ============== =============== End of Requirement 4.6 =============== 10 22 2018; 18:41:46PM Processing Requirement 4.6 Return Status : Passed ============== Project Data ============== **INPUT** Space Structure Name = TYVAK-0086 Space Structure Type = Payload Perigee Altitude = 500.000000 (km) Apogee Altitude = 500.000000 (km) Inclination = 85.000000 (deg) RAAN = 0.000000 (deg) Argument of Perigee = 0.000000 (deg) Page 28 of 41

Mean Anomaly = 0.000000 (deg) Area-To-Mass Ratio = 0.011490 (m^2/kg) Start Year = 2018.000000 (yr) Initial Mass = 10.226000 (kg) Final Mass = 10.226000 (kg) Duration = 25.000000 (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** Suggested Perigee Altitude = 500.000000 (km) Suggested Apogee Altitude = 500.000000 (km) Returned Error Message = Reentry during mission (no PMD req.). Released Year = 2022 (yr) Requirement = 61 Compliance Status = Pass ============== =============== End of Requirement 4.6 =============== 10 22 2018; 18:31:37PM Processing Requirement 4.6 Return Status : Passed ============== Project Data ============== **INPUT** Space Structure Name = TYVAK-0087 Space Structure Type = Payload Perigee Altitude = 505.000000 (km) Apogee Altitude = 505.000000 (km) Inclination = 98.000000 (deg) RAAN = 0.000000 (deg) Argument of Perigee = 0.000000 (deg) Mean Anomaly = 0.000000 (deg) Area-To-Mass Ratio = 0.011490 (m^2/kg) Start Year = 2018.000000 (yr) Initial Mass = 10.226000 (kg) Final Mass = 10.226000 (kg) Duration = 25.000000 (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** Suggested Perigee Altitude = 505.000000 (km) Page 29 of 41

Suggested Apogee Altitude = 505.000000 (km) Returned Error Message = Reentry during mission (no PMD req.). Released Year = 2022 (yr) Requirement = 61 Compliance Status = Pass ============== =============== End of Requirement 4.6 =============== Page 30 of 41

APPENDIX D - REQUIREMENT 4.7-1 DAS 2.0.1 LOG Note: Only worst-case TYVAK-0085 case shown. TYVAK-0086/-0087 results at higher altitudes with lower casualty risk from reentry debris. 09 26 2018; 14:34:17PM *********Processing Requirement 4.7-1 Return Status : Passed ***********INPUT**** Item Number = 1 name = TYVAK-0085 quantity = 1 parent = 0 materialID = 9 type = Box Aero Mass = 10.226000 Thermal Mass = 10.226000 Diameter/Width = 0.366000 Length = 0.678000 Height = 0.100000 name = PX Panel quantity = 1 parent = 1 materialID = 9 type = Box Aero Mass = 0.362000 Thermal Mass = 0.362000 Diameter/Width = 0.082000 Length = 0.335000 Height = 0.010000 name = MX Panel quantity = 1 parent = 1 materialID = 9 type = Box Aero Mass = 0.700000 Thermal Mass = 0.700000 Diameter/Width = 0.082000 Page 31 of 41

Length = 0.335000 Height = 0.011500 name = PZ Panel quantity = 1 parent = 1 materialID = 9 type = Box Aero Mass = 0.272000 Thermal Mass = 0.272000 Diameter/Width = 0.100000 Length = 0.209000 Height = 0.020000 name = MZ Panel quantity = 1 parent = 1 materialID = 9 type = Box Aero Mass = 0.248000 Thermal Mass = 0.248000 Diameter/Width = 0.082000 Length = 0.205000 Height = 0.020000 name = PY Panel quantity = 1 parent = 1 materialID = 9 type = Box Aero Mass = 1.800000 Thermal Mass = 1.800000 Diameter/Width = 0.226000 Length = 0.330000 Height = 0.011500 name = MY Panel quantity = 1 parent = 1 materialID = 9 Page 32 of 41

type = Box Aero Mass = 2.100000 Thermal Mass = 2.100000 Diameter/Width = 0.226000 Length = 0.330000 Height = 0.011500 name = IRM quantity = 1 parent = 1 materialID = 23 type = Box Aero Mass = 0.225000 Thermal Mass = 0.125000 Diameter/Width = 0.100000 Length = 0.100000 Height = 0.050000 name = IMU quantity = 1 parent = 8 materialID = 9 type = Box Aero Mass = 0.018000 Thermal Mass = 0.018000 Diameter/Width = 0.024500 Length = 0.038000 Height = 0.011100 name = Star Camera quantity = 2 parent = 8 materialID = 9 type = Cylinder Aero Mass = 0.041000 Thermal Mass = 0.041000 Diameter/Width = 0.030000 Length = 0.047000 name = RWA_Motor_Flywheel Page 33 of 41

quantity = 3 parent = 1 materialID = 9 type = Cylinder Aero Mass = 0.040000 Thermal Mass = 0.020000 Diameter/Width = 0.040000 Length = 0.013000 name = RWA_Motor quantity = 3 parent = 11 materialID = 9 type = Cylinder Aero Mass = 0.010000 Thermal Mass = 0.010000 Diameter/Width = 0.040000 Length = 0.013000 name = RWA_Brackets quantity = 3 parent = 11 materialID = 9 type = Box Aero Mass = 0.010000 Thermal Mass = 0.010000 Diameter/Width = 0.040000 Length = 0.040000 Height = 0.013000 name = Battery Module_1 quantity = 2 parent = 1 materialID = 9 type = Box Aero Mass = 0.493000 Thermal Mass = 0.352000 Diameter/Width = 0.042000 Length = 0.086000 Height = 0.042000 Page 34 of 41

name = Batteries quantity = 6 parent = 14 materialID = 54 type = Cylinder Aero Mass = 0.047000 Thermal Mass = 0.047000 Diameter/Width = 0.018000 Length = 0.065000 name = POD Antenna quantity = 1 parent = 1 materialID = 9 type = Cylinder Aero Mass = 0.075000 Thermal Mass = 0.075000 Diameter/Width = 0.060000 Length = 0.030000 name = CION Payload quantity = 1 parent = 1 materialID = 23 type = Box Aero Mass = 0.700000 Thermal Mass = 0.100000 Diameter/Width = 0.200000 Length = 0.300000 Height = 0.005000 name = CION Cover quantity = 1 parent = 17 materialID = 9 type = Box Aero Mass = 0.600000 Thermal Mass = 0.600000 Diameter/Width = 0.200000 Page 35 of 41

Length = 0.300000 Height = 0.005000 name = KL_BPF quantity = 1 parent = 1 materialID = 9 type = Box Aero Mass = 0.010000 Thermal Mass = 0.010000 Diameter/Width = 0.015000 Length = 0.030000 Height = 0.010000 name = Miteq_LNA quantity = 1 parent = 1 materialID = 9 type = Box Aero Mass = 0.010000 Thermal Mass = 0.010000 Diameter/Width = 0.015000 Length = 0.030000 Height = 0.010000 name = Backplane quantity = 1 parent = 1 materialID = 23 type = Box Aero Mass = 0.200000 Thermal Mass = 0.200000 Diameter/Width = 0.200000 Length = 0.300000 Height = 0.004000 name = UHF_assembly quantity = 1 parent = 1 materialID = 23 Page 36 of 41

type = Box Aero Mass = 0.070000 Thermal Mass = 0.040000 Diameter/Width = 0.050000 Length = 0.100000 Height = 0.050000 name = UHF_radio quantity = 1 parent = 22 materialID = 9 type = Box Aero Mass = 0.030000 Thermal Mass = 0.030000 Diameter/Width = 0.036000 Length = 0.083000 Height = 0.007000 name = OEM628 quantity = 1 parent = 1 materialID = 23 type = Box Aero Mass = 0.537000 Thermal Mass = 0.037000 Diameter/Width = 0.060000 Length = 0.100000 Height = 0.009000 name = GPS Housing quantity = 1 parent = 24 materialID = 9 type = Box Aero Mass = 0.500000 Thermal Mass = 0.500000 Diameter/Width = 0.060000 Length = 0.150000 Height = 0.030000 Page 37 of 41

**************OUTPUT**** Item Number = 1 name = TYVAK-0085 Demise Altitude = 77.999543 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = PX Panel Demise Altitude = 74.253027 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = MX Panel Demise Altitude = 71.191738 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = PZ Panel Demise Altitude = 74.929558 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = MZ Panel Demise Altitude = 74.869714 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = PY Panel Demise Altitude = 68.447316 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = MY Panel Page 38 of 41

Demise Altitude = 67.023792 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = IRM Demise Altitude = 76.521191 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = IMU Demise Altitude = 74.379300 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = Star Camera Demise Altitude = 73.346589 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = RWA_Motor_Flywheel Demise Altitude = 75.771191 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = RWA_Motor Demise Altitude = 74.585988 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = RWA_Brackets Demise Altitude = 74.927160 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 Page 39 of 41

************************************* name = Battery Module_1 Demise Altitude = 70.537644 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = Batteries Demise Altitude = 65.407839 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = POD Antenna Demise Altitude = 74.320652 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = CION Payload Demise Altitude = 77.563715 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = CION Cover Demise Altitude = 73.338964 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = KL_BPF Demise Altitude = 76.137347 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = Miteq_LNA Demise Altitude = 76.137347 Debris Casualty Area = 0.000000 Page 40 of 41

Impact Kinetic Energy = 0.000000 ************************************* name = Backplane Demise Altitude = 76.976152 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = UHF_assembly Demise Altitude = 77.428566 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = UHF_radio Demise Altitude = 75.732074 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = OEM628 Demise Altitude = 77.179410 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = GPS Housing Demise Altitude = 69.849331 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* =============== End of Requirement 4.7-1 =============== Page 41 of 41


Document Created: 2018-11-08 14:06:26
Document Modified: 2018-11-08 14:06:26
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