ODAR

0698-EX-CN-2018 Text Documents

Astro Digital US, Inc.

2018-09-12ELS_215719

DocuSign Envelope ID: D457D7AC-EFA7-48D8-BEBE-91A3FFBAA62B Momentus X1 ODAR – Version 1.0 Astro Digital Momentus X1 Orbital Debris Assessment Report (ODAR) MOMENTUS X1-ODAR-1.0 This report is presented as compliance with NASA-STD-8719.14, APPENDIX A. Report Version: 1.0, 11/12/2015 Astro Digital US, Inc. 3171 Jay St. Santa Clara, CA 95054 Document Data is Not Restricted. This document contains no proprietary, ITAR, or export controlled information. 1

DocuSign Envelope ID: D457D7AC-EFA7-48D8-BEBE-91A3FFBAA62B Momentus X1 ODAR – Version 1.0 DAS Software Version Used In Analysis: v2.0.2 Astro Digital Momentus X1 Orbital Debris Assessment Report MOMENTUS X1-ODAR-1.0 APPROVAL: Chris Biddy CEO, Astro Digital ______________________________ Jan A. King CTO, Astro Digital _______________________________________ Brian Cooper Mission Manager Momentus X1 Program ________________________________________ 2

DocuSign Envelope ID: D457D7AC-EFA7-48D8-BEBE-91A3FFBAA62B Momentus X1 ODAR – Version 1.0 Revision Record Revision: Date: Affected Pages: Changes: Author(s): 1.0 9/11/2018 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 Momentus X1 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: D457D7AC-EFA7-48D8-BEBE-91A3FFBAA62B Momentus X1 ODAR – Version 1.0 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 Momentus X1 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 Momentus X1 satellite. Sections 9 through 14 apply to the launch vehicle ODAR and are not covered here. Momentus X1 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: D457D7AC-EFA7-48D8-BEBE-91A3FFBAA62B Momentus X1 ODAR – Version 1.0 Summary of NASA’s responsibility under the governing agreement(s): N/A Schedule of upcoming mission milestones: • Shipment of spacecraft: Q4 2018 • First Launch: Q1 2019 Mission Overview: Momentus X1 is a technology demonstration spacecraft built to the 16U CubeSat standard. It includes the VigorideTM thruster which energizes water steam into a plasma using microwaves and ejects the super-heated propellant to efficiently produce thrust. The spacecraft will be launched aboard a Soyuz rocket in a 16U CubeSat Deployer designed and built by ECM Technologies of Berlin. The spacecraft bus is the Corvus-16 design. The common 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: Soyuz, Baikonur Cosmodrome, Kazakhstan. Proposed Initial Launch Date: Q1 2019 Mission Duration: The anticipated mission duration is 9 months (nominal). 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 Momentus X1 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 Momentus X1 spacecraft will deploy a UHF antenna once deployed from the ECM deployer. The spacecraft will decay naturally from operational orbits within the following ranges: Average Orbital Altitude: 583 km to 587 km Eccentricity: 0.0000 to 0.0033 Inclination: 97.4° to 98.6° The Momentus X1 propulsion system has a very small total impulse, and as such is not being considered as a viable deorbit method. The spacecraft will be launched into an orbit that will result in a natural orbital decay in less than 25 years. 5

DocuSign Envelope ID: D457D7AC-EFA7-48D8-BEBE-91A3FFBAA62B Momentus X1 ODAR – Version 1.0 ODAR Section 2: Spacecraft Description: Physical description of the spacecraft: Momentus X1 uses the standard Corvus-16 bus, which is based on the 16U CubeSat form factor. Basic physical dimensions are 454.0 mm x 246.3 mm x 246.3 mm with a mass of no more than 24 kg. The superstructure is comprised of 5 outer panels and one internal panel separating the telescope from the bus electronics. There are L rails along each of the 454 mm edges. These accommodate the deployment of the satellite from the deployer. The bus electronics provide additional internal stability to the structure. The Momentus X1 Vigoride thruster nozzle is located on the +Z face of the spacecraft. The spacecraft bus includes a spring-loaded UHF antenna which is deployed after jettison from the deployer by a burn wire controlled by a software timer 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 is depicted in Figure 1. Figure 1: Momentus X1 Spacecraft 6

DocuSign Envelope ID: D457D7AC-EFA7-48D8-BEBE-91A3FFBAA62B Momentus X1 ODAR – Version 1.0 Total satellite mass at launch, including all propellants and fluids: Momentus X1: 22.0 kg +/- 2.0 kg Dry mass of satellites at launch: Momentus X1: 21.85 kg +/- 2.0 kg Description of all propulsion systems (cold gas, mono-propellant, bi- propellant, electric, nuclear): The Vigoride propulsion system energizes distilled water into a plasma using RF microwave energy. The plasma is expelled out of the thruster using a nozzle to produce thrust at an Isp exceeding traditional chemical propulsion systems. The expected thrust and Isp will vary with power input levels. Thrust will not exceed 20 mN maximum. The Momentus X1 propulsion system includes a nominal 150 grams of water propellant with a tolerance of - 50grams/+100grams. The thruster is not expected to be operated at peak efficiency throughout the mission, but the absolute maximum impulse possible if it were would be less than 1000 N-s. The propulsion system is pump fed, but the tank contains a diaphragm and will be pressurized with inert gaseous nitrogen (N2). N2 is commonly used as a pressurant for many fluid and propulsion systems. At spacecraft integration the tank is fueled with water and then pressurized to 1 atmosphere (14 psi, nominal ambient sea level atmospheric pressure) with the nitrogen. The tank then remains at this pressure during transportation to the launch site, launch and after launch until thruster operations. As the water propellant is pumped out from the tank, the pressure will be reduced. At launch, the system is not considered pressurized because of its 1 atmosphere specification. The fuel tank has been proof tested to greater than three times the limit load. At the end of the mission, the propellant flow valve will be fully opened which will allow all propellant and pressurant to vent into space and remove all energy from the system. The spacecraft will be oriented in such a way that any resultant thrust from the drain operation will result in a lower orbital altitude. Please note that the accompanying DAS analysis assumes the worst case end of life mass of 23.9 kg, which coincides with the worst case ballistic coefficient (highest dry mass and lowest potential propellant load of 100 grams of water). 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: A nominal 100 grams (but up to 250 grams) of benign distilled water and no less than 83 ml of N2 pressurant gas at 14 psi. These fluids will be loaded prior to integration of the spacecraft into the standard CubeSat deployer. There will be zero 7

DocuSign Envelope ID: D457D7AC-EFA7-48D8-BEBE-91A3FFBAA62B Momentus X1 ODAR – Version 1.0 gauge pressure at the time of loading through launch. The atmospheric pressure present will result in the tank pressurizing to 14 psi during launch. The pressure vessel has been tested to 3x the proof pressure and is qualified for transportation by the DOT, as it is not a pressure vessel while in the atmosphere. Fluids in Pressurized Batteries: None The Corvus-16 satellite design uses eight 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 96 W-Hrs. Description of attitude control system and indication of the normal attitude of the spacecraft with respect to the velocity vector: The Momentus X1 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 panels 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 thrust axis to be pointed in any direction in inertial space. 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 its antenna using a burn wire system. The mechanical pressurant system will use solenoids and electric motors to reduce the volume of the pressurant tank. System power is locked off during launch by two series and two parallel deployment switches but, the ECM deployer prevents any form of premature deployment, 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 42 watts at the end-of-life of the mission. Typical bus operations consume 8 watts of power on average. The thruster can consume up to 150 watts in short bursts. 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 8

DocuSign Envelope ID: D457D7AC-EFA7-48D8-BEBE-91A3FFBAA62B Momentus X1 ODAR – Version 1.0 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. Potential causes of spacecraft breakup during deployment and mission operations: There are two potential scenarios that could potentially lead to a breakup of the satellite. In order of credibility: 1) Rupture of the propellant tank (H20, N2) 2) Lithium-ion battery cell failure 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. 9

DocuSign Envelope ID: D457D7AC-EFA7-48D8-BEBE-91A3FFBAA62B Momentus X1 ODAR – Version 1.0 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: Eight (8) 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, Momentus X1: 0.0000 Supporting Rationale and FMEA details: Pressure Tank Explosion: Effect: A rupture of a the propellant tank would release water and nitrogen. Due to the low pressure (14 psia), the penetrating energy of any debris would be relatively low. The tank is enclosed in the solid aluminum structural panels of the spacecraft. These aluminum walls would contain any released debris within the body of the spacecraft. Probability: Very low. A structural failure of the tank would need to occur, and the mechanisms by which these failures occur are very well understood. CubeSats are typically volume-limited as opposed to mass-limited. This means that it is very easy to add mass to a given structure to protect against failure, and structural strength margins can be very high. This is employed in the design of the pressure vessels for Momentus X1. Whereas typical aerospace components would have a margin of 10

DocuSign Envelope ID: D457D7AC-EFA7-48D8-BEBE-91A3FFBAA62B Momentus X1 ODAR – Version 1.0 safety under 2, all structures on the Corvus satellite designs have strength to failure margins of 3 or greater. 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 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 both spacecraft, thermal vacuum cycling of both 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). 11

DocuSign Envelope ID: D457D7AC-EFA7-48D8-BEBE-91A3FFBAA62B Momentus X1 ODAR – Version 1.0 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). 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. 12

DocuSign Envelope ID: D457D7AC-EFA7-48D8-BEBE-91A3FFBAA62B Momentus X1 ODAR – Version 1.0 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 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: Momentus X1 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, the 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. As discussed above in the propulsion system section, all energy will be released from the propulsion system prior to spacecraft deactivation. The spacecraft will be oriented such that any thrust generated from propellant release results in an orbit lowering maneuver. All thruster valves will be opened until all propellant and pressurant are completely released. No attempt will be made to activate the RF microwave element, which will result in a “cold gas” thruster firing. 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. 13

DocuSign Envelope ID: D457D7AC-EFA7-48D8-BEBE-91A3FFBAA62B Momentus X1 ODAR – Version 1.0 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 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).” 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 postmission disposal requirements is less than 0.01 (Requirement 56507).” Small Object Impact and Debris Generation Probability: 0.0000; 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. 14

DocuSign Envelope ID: D457D7AC-EFA7-48D8-BEBE-91A3FFBAA62B Momentus X1 ODAR – Version 1.0 6.2 Plan for any spacecraft maneuvers required to accomplish post-mission disposal: None are required to accomplish post-mission disposal 6.3 Calculation of area-to-mass ratio after post-mission disposal, if the controlled reentry option is not selected: Spacecraft Mass: 24.0 kg (selected as worst case mass) Cross-sectional Area: 0.135 m^2 (average tumbling) (Calculated by DAS 2.0.2). Area to mass ratio: 0.005625 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) 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 Momentus X1 spacecraft will follow a concept of operations to ensure a safe disposal within 25 years of the end of the mission. To demonstrate the thruster, the perigee of the orbit will be lowered first, and the apogee will be raised afterwards. This will ensure that even if the thruster fails at any point, the lifetime requirement will still be met. The final target orbit will be a perigee of 560 km and an apogee of 615 km. This results in an orbit lifetime of 22.0 years. This analysis was performed with the NASA Debris Assessment Software 2.0.2. Figure 2 and Figure 3 show the output data from this analysis. 15

DocuSign Envelope ID: D457D7AC-EFA7-48D8-BEBE-91A3FFBAA62B Momentus X1 ODAR – Version 1.0 Figure 2: Momentus X1 Orbit History (615x560 km worst case orbit) 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: Assessment of spacecraft compliance with Requirement 4.7-1: Requirement 4.7-1. Limit the risk of human casualty: 16

DocuSign Envelope ID: D457D7AC-EFA7-48D8-BEBE-91A3FFBAA62B Momentus X1 ODAR – Version 1.0 “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 Momentus X1 is COMPLIANT with the requirement. The critical values reported by the DAS software are: • Demise Altitude = 0.0 km • Debris Casualty Area = 0.68 m^2 • Impact Kinetic Energy = 4504 Joules • Risk of Human Casualty = 1:115700 This is expected to represent the absolute maximum casualty risk, as calculated with DAS's modeling capability. Requirements 4.7-1b, and 4.7-1c: These requirements are non-applicable requirements because the spacecraft does 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 Momentus X1. The satellite does not use controlled reentry. 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. The satellite does not use controlled reentry. ODAR Section 8: Assessment for Tether Missions Not applicable. There are no tethers used in Momentus X1 END of ODAR for Momentus X1 17

DocuSign Envelope ID: D457D7AC-EFA7-48D8-BEBE-91A3FFBAA62B Momentus X1 ODAR – Version 1.0 The raw DAS report as follows for Momentus X1: 08 21 2018; 14:57:59PM Processing Requirement 4.3-1: Return Status : Not Run ===================== No Project Data Available ===================== =============== End of Requirement 4.3-1 =============== 08 21 2018; 14:58:02PM Processing Requirement 4.3-2: Return Status : Passed ===================== No Project Data Available ===================== =============== End of Requirement 4.3-2 =============== 08 21 2018; 14:58:05PM Requirement 4.4-3: Compliant =============== End of Requirement 4.4-3 =============== 08 21 2018; 15:34:46PM Processing Requirement 4.5-1: Return Status : Passed ============== Run Data ============== **INPUT** Space Structure Name = Momentus X1 Space Structure Type = Payload Perigee Altitude = 560.000000 (km) Apogee Altitude = 615.000000 (km) Inclination = 98.600000 (deg) RAAN = 0.000000 (deg) Argument of Perigee = 0.000000 (deg) Mean Anomaly = 0.000000 (deg) Final Area-To-Mass Ratio = 0.005650 (m^2/kg) Start Year = 2019.000000 (yr) Initial Mass = 24.000000 (kg) Final Mass = 23.900000 (kg) Duration = 1.000000 (yr) Station-Kept = False Abandoned = True 18

DocuSign Envelope ID: D457D7AC-EFA7-48D8-BEBE-91A3FFBAA62B Momentus X1 ODAR – Version 1.0 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 = 0.000009 Returned Error Message: Normal Processing Date Range Error Message: Normal Date Range Status = Pass ============== =============== End of Requirement 4.5-1 =============== 08 21 2018; 15:35:02PM Requirement 4.5-2: Compliant 08 21 2018; 15:35:04PM Processing Requirement 4.6 Return Status : Passed ============== Project Data ============== **INPUT** Space Structure Name = Momentus X1 Space Structure Type = Payload Perigee Altitude = 560.000000 (km) Apogee Altitude = 615.000000 (km) Inclination = 98.600000 (deg) RAAN = 0.000000 (deg) Argument of Perigee = 0.000000 (deg) Mean Anomaly = 0.000000 (deg) Area-To-Mass Ratio = 0.005650 (m^2/kg) Start Year = 2019.000000 (yr) Initial Mass = 24.000000 (kg) Final Mass = 23.900000 (kg) Duration = 1.000000 (yr) Station Kept = False Abandoned = True PMD Perigee Altitude = 565.717223 (km) PMD Apogee Altitude = 608.933833 (km) PMD Inclination = 98.587346 (deg) 19

DocuSign Envelope ID: D457D7AC-EFA7-48D8-BEBE-91A3FFBAA62B Momentus X1 ODAR – Version 1.0 PMD RAAN = 38.340690 (deg) PMD Argument of Perigee = 231.768575 (deg) PMD Mean Anomaly = 0.000000 (deg) **OUTPUT** Suggested Perigee Altitude = 565.717223 (km) Suggested Apogee Altitude = 608.933833 (km) Returned Error Message = Passes LEO reentry orbit criteria. Released Year = 2041 (yr) Requirement = 61 Compliance Status = Pass ============== =============== End of Requirement 4.6 =============== 08 21 2018; 15:50:46PM *********Processing Requirement 4.7-1 Return Status : Passed ***********INPUT**** Item Number = 1 name = Momentus X1 quantity = 1 parent = 0 materialID = 5 type = Box Aero Mass = 23.900000 Thermal Mass = 23.900000 Diameter/Width = 0.246000 Length = 0.454000 Height = 0.246000 name = Chamber quantity = 1 parent = 1 materialID = 54 type = Box Aero Mass = 0.008000 Thermal Mass = 0.008000 Diameter/Width = 0.006000 Length = 0.031000 Height = 0.006000 20

DocuSign Envelope ID: D457D7AC-EFA7-48D8-BEBE-91A3FFBAA62B Momentus X1 ODAR – Version 1.0 name = Tank quantity = 1 parent = 1 materialID = 54 type = Box Aero Mass = 5.000000 Thermal Mass = 5.000000 Diameter/Width = 0.200000 Length = 0.250000 Height = 0.200000 name = Window quantity = 1 parent = 1 materialID = -1 type = Box Aero Mass = 0.014000 Thermal Mass = 0.014000 Diameter/Width = 0.030000 Length = 0.038000 Height = 0.030000 name = Chamber2 quantity = 1 parent = 1 materialID = 33 type = Box Aero Mass = 0.020000 Thermal Mass = 0.020000 Diameter/Width = 0.010000 Length = 0.056000 Height = 0.010000 **************OUTPUT**** Item Number = 1 name = Momentus X1 Demise Altitude = 77.995811 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = Chamber Demise Altitude = 75.317726 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 21

DocuSign Envelope ID: D457D7AC-EFA7-48D8-BEBE-91A3FFBAA62B Momentus X1 ODAR – Version 1.0 ************************************* name = Tank Demise Altitude = 0.000000 Debris Casualty Area = 0.678328 Impact Kinetic Energy = 4503.674316 ************************************* name = Window Demise Altitude = 0.000000 Debris Casualty Area = 0.401657 Impact Kinetic Energy = 1.551344 ************************************* name = Chamber2 Demise Altitude = 75.390045 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* =============== End of Requirement 4.7-1 =============== 22

DocuSign Envelope ID: D457D7AC-EFA7-48D8-BEBE-91A3FFBAA62B Momentus X1 ODAR – Version 1.0 Appendix A: Acronyms 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 23


Document Created: 2018-09-11 12:06:26
Document Modified: 2018-09-11 12:06:26
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