Orbital Debris Assessment Report Rev C Final

0240-EX-CN-2018 Text Documents

University of Virginia

2018-09-28ELS_217012

ELVL—2018—0045207 Rev C June 20, 2018 Orbital Debris Assessment for The CubeSats on the CRS OA—10/ELaNa—21 Mission per NASA—STD 8719.14A

Signature Page Yus son, Analyst, a.i. solutions, AIS2 Scott Higginbotham, MissidAManager, NASA KSC VA—C

National Aeronauties and Space Administration John F. Kennedy Space Center, Florida Kennedy Space Center, FL 32899 ELVL—2018—0045207 Rev B Reply to Attnof: VA—H1 June 20, 2018 TO: Scott Higginbotham, LSP Mission Manager, NASA/KSC/VA—C FROM: Yusef Johnson, a.i. solutions/KSC/AIS2 SUBJECT: Orbital Debris Assessment Report (ODAR) for the ELaNa—21 Mission REFERENCES: A. NASA Procedural Requirements for Limiting Orbital Debris Generation, NPR 8715.6A, 5 February 2008 B. Process for Limiting Orbital Debris, NASA—STD—8719.14A, 25 May 2012 C. International Space Station Reference Trajectory, delivered May 2017 D. McKissock, Barbara, Patricia Loyselle, and Elisa Vogel. Guidelines on Lithium— ion Battery Use in Space Applications. Tech. no. RP—08—75. NASA Glenn Research Center Cleveland, Ohio UL Standardfor Safety for Lithium Batteries, UL 1642. UL Standard. 4th ed. Northbrook, IL, Underwriters Laboratories, 2007 Kwas, Robert. Thermal Analysis of ELaNa—4 CubeSat Batteries, ELVL—2012— 78 0043254; Nov 2012 Range Safety User Requirements Manual Volume 3— Launch Vehicles, Q Payloads, and Ground Support Systems Requirements, AFSCM 91—710 V3. HQ OSMA Policy Memo/Email to 8719.14: CubeSat Battery Non—Passivation, t Suzanne Aleman to Justin Treptow, 10, March 2014 HQ OSMA Email:6U CubeSat Battery Non Passivation Suzanne Aleman to i Justin Treptow, 8 August 2017 TechEdSat—8 Orbital Debris Assessment Report (ODAR), T8MP—06—XS001 Rev 0, NASA Ames Research Center The intent of this report is to satisfy the orbital debris requirements listed in ref. (a) for the ELaNa—21 auxiliary mission launching on the CRS OA—10 vehicle. It serves as the final submittal in support of the spacecraft Safety and Mission Success Review (SMSR). Sections 1 through 8 of ref. (b) are addressed in this document; sections 9 through 14 fall under the requirements levied on the primary mission and are not presented here.

RECORD OF REVISIONS REV DESCRIPTION DATE 0 ODAR Submission for TJREVERB and VCC March 2018 CubeSats A Combined original submission with full ELaNa—21 May 2018 complement B Updated mass properties for CySat June 2018 C Incorporated changes to VCC CubeSats CONOPS June 2018

The following table summarizes the compliance status of the ELaNa—21 payload mission to be flown on the OA—10 vehicle. The 13 CubeSats comprising the ELaNa—21 mission are fully compliant with all applicable requirements. Table 1: Orbital Debris Requirement Com liance Matrix Requirement Compliance Assessment Comments 4.3—la Not applicable No planned debris release 4.3—1b Not applicable No planned debris release 4.3—2 Not applicable No planned debris release 4.4—1 Compliant On board energy source (batteries) incapable of debris—producing failure 4.4—2 . Compliant On board energy source (batteries) incapable of debris—producing failure 44—3 Not applicable No planned breakups 4,4—4 Not applicable No planned breakups 4.5—1 Compliant 4.5—2 Not applicable 4.6—1(a) Compliant Worst case lifetime 3.9 yrs 4.6—1(b) Not applicable 4.6—1(c) Not applicable 4.6—2 Not applicable 4.6—3 Not applicable 4.6—4 Not applicable Passive disposal 4.6—5 Compliant 4.7—1 Compliant Non—credible risk of human casualty 4.8—1 Compliant No planned tether release under ELaNa—21 mission

Section 1: Program Management and Mission Overview The ELaNa—21 mission is sponsored by the Human Exploration and Operations Mission Directorate at NASA Headquarters. The Program Executive is Jason Crusan. Responsible program/project manager and senior scientific and management personnel are as follows: CapSat: McKale Berg, Project Manager, University of Illinois CySat 1: Rami Shoukih, Project Manager, Iowa State University KickSat—2: BJ Jaroux, Project Manager, NASA Ames Research Center HARP: Dr. J. Vanderlei Martins, Principal Investigator OPAL: Dr. Charles Swenson, Principal Investigator, Utah State Phoenix: Sarah Rogers, Project Manager, Arizona State University SPACE HAUC: Supriya Chakrabarti, Principal Investigator, University of Massachusetts—Lowell TechEdSat 8: Marcus Murbach, Project Manager, Ames Research Center TJREVERB: Michael Piccione, Principal Investigator, Thomas Jefferson High School UNITE: Glen Kissel, Principal Investigator, University of Southern Indiana Virginia CubeSat Consortium (Aeternitas, Ceres, Libertas): Mary Sandy, Principal Investigator, Virginia Space Grant Consortium

Program Milestone Schedule Task Date CubeSat Selection September 15, 2017 CubeSat Delivery to NanoRacks August 20th, 2018 Launch November 175%, 2018 Figure 1: Program Milestone Schedule The ELaNa—21 CubeSat complement will be launched as payloads on the OA—10 Antares launch vehicle to the International Space Station. The ELaNa—21 mission will deploy 13 pico—satellites (or CubeSats) from the International Space Station, using the NanoRacks CubeSat dispenser. Each CubeSat is identified in Table 2: ELaNa—21 CubeSats. The ELaNa—21 manifest includes: CapSat, CySat, HARP, KickSat—2, OPAL, Phoenix, SPACE HAUC, TechEdSat 8, TJREVERB, UNITE, and the three Virginia CubeSat Consortium CubeSats (Aeternitas, Ceres, and Libertas). The current launch date is projected to be November 17"", 2018. The CubeSats on this mission range in size from a 10 cm cube to 60 cm x 10 cm x 10 cm, with masses from about 1.2 kg to 3.5 kg, with a total mass of roughly 20 kg being manifested on this mission. The CubeSats have been designed and universities and government agencies and each have their own mission goals.

Section 2: Spacecraft Description There are 13 CubeSats flying on the ELaNa—21 Mission. Table 2: ELaNa—21 CubeSats outlines their generic attributes. Table 2: ELaNa—21 CubeSats CubeSat CubeSat CubeSat Names § CubeSat size (mm°) Masses Quantity (kg) CapSat 1 300 x 100 x 100 2.8 CySat 1 340 x 100 x 100 2.7 * HARP 1 368 x 100 x 100 4.1 * KickSat—2 1 300 x 100 x 100 ol * OPAL 1 368 x 100 x 100 5.0 Phoenix 1 325 x 100 x 100 3.2 SPACE HAUC 1 340 x 100 x 100 2.9 *TechEdSat 8 1 600 x 100 x 100 7.9 TJREVERB 1 227 x 100 x 100 2.6 UNITE 1 340 x 108 x 108 3.5 Virginia CC — Aeternitas 1 113 x : 100 x 78 1.2 Virginia Ceres CC — 1 113 x 106 x 106 1.2 Virginia CC — Libertas 1 118 x 105 x 106 1.4 *The following pages describe the CubeSats flying on the ELaNa—21 mission, with the omissions noted below. ODARs for these CubeSats were previously submitted to the Agency as follows: HARP: ELaNa—22 Rev A ODAR 5/2017 KickSat—2: KickSat—2 9/2015 OPAL: ELaNA—22 ODAR 10/2016 TechEdSat—8‘s ODAR was drafted by NASA Ames (Document No. T8SMP—06— XS001 Rev 0)

CAPSat — University of Illinois — 3U Figure 2: CAPSat Expanded View Overview The Cooling, Pointing, Annealing Satellite (CAPSat) is a CubeSat under development by the University of Illinois and Bradley University. The mission, which is expected to last approximately one year, encompasses three technology demonstrations, each advancing the technology readiness level of NASA roadmap technologies. The experiments are: strain—actuated deployable panels for improved pointing control and jitter reduction, an active thermal control system, and single—photon avalanche detectors (SPADs) to test methods of mitigating space radiation damage. CONOPS Thirty minutes after all three separations switches register deployment, the power board will set a flag to initiate full boot. The C&DH will be brought online, and attempt to fire the thermal knives to release the antenna. After three attempts, it will begin a Bdot detumbling algorithm to attempt to reduce all angular motion. Beaconing will begin after the antenna has attempted deployment. Once we are able to uplink its TLE and a date stamp update, the ADCS algorithm will switch to controlling the satellite such that service plate is the ram. After a few weeks of commissioning and testing the payloads, science operations will begin. Data will be transmitted down on the NanoCom AX100 radio to our ground station. Science will continue until the satellite re—enters.

Materials Satellite structure is made from AL60601T6, while the solar panels are Carbon fiber with an aluminum backing. The Cooling Payload consists of many small stainless steel components and its deployable panel is made of carbon fiber. The annealing payload is comprised of two circuit boards. The Pointing Payload is mostly circuit boards with an iron vibration motor and its deployable panel is made of a thing sheet of stainless steel. Hazards Regarding the restriction on pressure vessels for this launch, one of the CAPSat payloads contains a fluid loop containing 50/50 glycol/water, which under normal atmospheric conditions would not be considered a pressure vessel. The system has an operating pressure limit of 29.4 PSIA and a safety margin will be placed on the operating pressure of the system. The payload will undergo thorough leak and pressure testing in addition to standard vacuum, thermal, and vibration testing. There are no other hazards or exotic materials. Power System/Batteries The electrical power storage system consists of common lithium—ion batteries with over— charge/current protection circuitry. The charging system incorporates an MPPT logic. The lithium batteries carry the UL—listing number MH12210. 10

CySat — lowa State University — 3U Radio APSS = *——— Payioad o _Side Panet Motherboard No Bottom Panel Structure Monopole Artenna Payload Side Pane! Board Stricture /® Side Panel Dipole Antennas Figure 3: CySat Expanded View Overview CySat will operate in a Low Earth Orbit (LEO) environment to test out a state—of—the—art radiometer payload based off a Software Defined Radio (SDR) to observe the Earth and measure the soil moisture. CONOPS Once CySat is deployed power will begin flowing and the countdown timers for the deployable antenna and the communication subsystem will initiate. After 45 minutes have passed, the antenna will deploy. The ground station, will then attempt to pick up CySat‘s beacon and establish contact. The satellite will be in a passive mode at this point, and will stay in this mode for roughly the first 24 hours of operations. This involves an ASCII message containing minimal system status information and a welcome message for radio amateurs. A command will then be sent to CySat to ensure health and housekeeping data is gathered. This will continue for no more than a week. Once functions are determined to be nominal, CySat will be transitioned for primary operations and all primary payload routines will be active at this time. Payload activities are desired to continue for at least one year. Materials The CySat structure is made of Aluminum 6061—T6. It contains standard commercial off the shelf (COTS) materials, electrical components, PCBs and solar cells. 11

Hazards There are no pressure vessels, hazardous or exotic materials. Batteries The electrical power storage system consists of Lithium ion batteries with cell overcurrent — charge, cell overcurrent — discharge, cell voltage and cell under — voltage protection circuits on each cell as well as on the entire battery assembly. Additional over — current bus protection and battery under — voltage protection is also provided by the electric power system (EPS). The UL — listing number for the batteries is: UL 1642. 12

Phoenix, Arizona State University — 3U UHF Antenina Batteries $ all Electronics Stack Mount —¥ Panel Chassis Roil Solar Parel EPS $—Band Transmitter $—Band Antenna ©BC UHF Transcelver Motherboard interface Board GPS GPS Antenna Camera Mounts Sun Sensor FLIR Tau 2 640 Camera —ZFace Figure 4: Phoenix Expanded View Overview Phoenix is a 3U CubeSat designed to study the Urban Heat Island Effect over several US cities. The payload is the Tau2 640 infrared camera, which is a commercially available, uncooled microbolometer produced by FLIR Technologies. CcoONOPS After the satellite is deployed from the ISS, it will initiate power to it components and start a countdown timer. After 30 minutes, the UHF antenna will deploy. After 45 minutes, the UHF beacon will be activated to communicate satellite health. Phoenix will undergo a week of checkout operations, where mission operators will monitor the health of the satellite, capture calibration images, and solidify the satellite‘s trajectory before beginning the mission objective. Mission operations are expected to last up to two years, and yield a total of 8,000 thermal IR images before the satellite re—enters. Materials Phoenix is comprised of COTS hardware. Therefore, all electrical components, PCBs, and solar cells are rated for the environment of space. The chassis is made of Aluminum 7075—16. Stainless—steel bolts will be used to assemble the chassis and all cabling will be comprised of copper alloy material. 13

Hazards There are no pressure vessels, hazardous, or exotic materials. Batteries The electrical power storage system consists of Lithium ion batteries with overcharge/current protection circuitry. The UL listing number for the batteries is: UL1642. 14

SPACE HAUC — University of Massachusetts, Lowell — 3U Coomim. Sytein ©CAIDLT Solar Puneln JAAA __..2 wtil [ Thecraadi Inmifation * ftrncfie Commrerc Figure 5: SPACE HAUC Expanded View Overview SPACE HAUC will demonstrate that high data transmission rates can be achieved by using a X—Band Phased—Array antenna with an electronically steered beam on a CubeSat. CONOPS Immediately upon deployment, SPACE HAUC will power up anddetermine if it is spin stabilized. If not, the Attitude Determination and Control System will stabilize the spin. It will then determine if it is sun pointed, if not the Attitude Determination and Control System will point SPACE HAUC at the sun. SPACE HAUC will then wait for a beacon signal from the ground, upon receipt of the beacon, SPACE HAUC will take pictures of the sun and transmit them down. The process of waiting for the beacon signal will be repeated whenever the beacon signal is lost. Materials The CubeSat structure is made of Aluminum 7075—T6. It contains all standard commercial off the shelf (COTS) materials, electrical components, PCBs and solar cells except for the RF front end board and patch antennas which are custom designed. The high—speed radio uses a ceramic patch antenna. 15

Hazards There are no pressure vessels, hazardous, or exotic materials.. Batteries The lithium—ion battery is charged with all the available power from the photo—voltaic inputs that is not drained by the loads on the external power busses. The battery is protected against voltage being too high or too low. The software high voltage protection implements a constant voltage charge scheme that will keep the battery at its maximum voltage. The full mode regulation works by lowering voltage on the solar panel inputs, thereby only taking in the power needed. The software low voltage protection is a four state system. Should the battery voltage drop below 7.2 V, the battery hardware will switch to a ‘safe mode‘ configuration, which allows for the switching off of all essential systems and leaves only a simple power beacon running. Should the battery drop below 6.5 V, the software will switch off all user outputs. 16

TJREVERB — Thomas Jefferson High School for Science and Technology 2U +X Solar Panel $—Band Patch +¥ Solar Panel . tridium SBD Modem / Interface bqard l +Z mounting pl P s[ . --:‘\ h Raguarok Flight Computer and 8 h e S@ APRS Transc, EPS/HeatSink 3X > a q L evite t 2 ( GPS C Iridiuon ISIS L j Antenna Magnotorquer N «y Solar Panet GPS Parch Antenna Figure 6: TJREVERB Expanded View Overview TJREVERB (Thomas Jefferson High School for Science and Technology Research and Education Vehicle for Evaluating Radio Broadcasts) will be a 2U CubeSat with magnetic torque control. It will be using a VHF APRS transceiver on 145.825MHz for command and control. It will also have a 2.2—2.3 GHz transceiver and a 1.616—1.6265 GHz short burst data (SBD) modem to test the ability to send and receive data packets and compare the usage of the Near Earth Network and the Tridium satellite network. The SBD modem will also provide secondary command and control. CONOPS Thirty mins after deployment, the spacecraft will deploy its antenna and start to detumble. After 45 mins the spacecraft establishes communications link, establish GPS link, clock synch, orbit determination daily, transmit AMSAT APRS signals, and perform operations modes (Charging, Comms check, and update) and science modes. Science modes consist of running various transmission activities while orbiting in various attitude orientations modes such as spin—stabilized and 3—axis regulation. Materials TJREVERB‘s chassis is made of Aluminum 606. It contains standard commercial off— the—shelf (COTS) materials, electrical components, PCBs and solar cells. 17

Hazards TJREVERB does not include any hazardous systems or pressure vessels. Batteries The Orbtronics 18650B cell is a modified standard Panasonic 18650B NCR cell with UL listing MH12210 with flight heritage aboard past CubeSats such as GeneSat, SporeSaat, OREOS, and Pharmasat. Each cell is 65 mm in length and 18.6 mm in diameter. The Graphite/LiNiCoAIO2 (NCA) chemistry provides for maximum capacity of 3400 mAhr at a full charge. A total of 40 Whr battery capacity is provided via 2 packs of 2 battery cells in series, @S2P, each at 20Whr. Each cell contains a Positive Temperature Coefficient (PTC) device, Current Interrupt Device (CID), and an exhaust gas hole built into each battery cell to prevent cell rupture. The cell builds on the safety features of the 18650 cell by including a Seiko Protection Integrated Circuit (IC) that provides over voltage protections (OVP) at 4.35V, over— discharge (UVLO) protection (OCD) at 10—12A, and over—heating protection. 18

UNITE— University of Southern Indiana — 3U Mu—metal Rods a««= Ballast Mass ( Auxiliary Batteries o z_ Langmuir Probe i Magnetometer > <@ Command Board Optical Bench ==———» Magnets & Magnet Holder Duplex A..ntenna t§ “ Langmuir Probe Board \\\\\\/ A & TEeMPETTUTE mxmnumup 4 Sensor (x8 _ (x81 Horizon Sensy _ RBF Pin Simplex Antenna / * 8 \U Deployment Switch (x4) Diagnostie Port (x2) GPS Antenna Figure 7; UNITE Expanded View Overview The Undergraduate Nano Ionospheric Temperature Explorer (UNITE) CubeSat is a 3U nanosatellite that will explore Low Earth Orbit until re—entry into the atmosphere around 90 km. The mission of UNITE is to conduct space weather measurements with a Langmuir plasma probe, measure interior and exterior temperature of the craft, and model the craft‘s orbit in the final hours of re—entry. The lower ionosphere is a relatively unexplored region of space and the scientific data collected and transmitted by UNITE will contribute to the understanding of the region. CONOPS Once deployed, the satellite‘s inhibit switches will be released. However, the satellite will not power on until the solar panels receive light. This is due to the "solar enable" feature of the EPS purchased from NearSpace Launch that acts as a third inhibit mechanism to the satellite powering on. Once powered on, no transmissions will be made for the first 45 minutes. Once this initial deployment period has passed, the satellite will begin collecting data and transmitting to the Globalstar satellite constellation. All transmission from UNITE will be to the Globalstar constellation as no ground station is used for the UNITE mission. The software of UNITE will change the rates of data collection and transmission based on the altitude. The satellite will continue to collect data and transmit until it burns up during re—entry. Materials The structure of UNITE is a 3U chassis made of anodized 6061 aluminum. External to the chassis are solar panels, consisting of PCB and glass covered solar cells, and ceramic patch antennae. The internal components of the satellite are commercial off the shelf 19

(COTS) materials, two 1/8" thick aluminum plates (optical benches in exploded view), copper ballast masses, electrical components, PCBs, and batteries. Hazards There are no pressure vessels, hazardous or exotic materials. Ratteries There are four 2—cell lithium—polymer battery packs on UNITE, bringing the maximum total stored energy to 64 watt—hours. Each battery pack contains over—charge/current protection circuitry. The NearSpace Launch EPS that interfaces with the batteries also contains over—current and over—voltage protection. The UL listing number for the batteries is 30156—1. 20

Aeternitas — Old Dominion University (Virginia CubeSat Constellation) Sole Posel {4) Trag Break Potais fiatu Bowe Antorion losd Plate~ Yendloem Antenen Cover Pxte — .0 Rewiled Rods (4} Figure 8: Aeternitas Expanded View Overview The Virginia CubeSat Constellation (VCC) mission is a joint operation between teams at Old Dominion University, University of Virginia, Virginia Tech, and Hampton University. ODU, UVA, and VT are each building 1U CubeSats (Aeternitas, Libertas, and Ceres, respectively) that will fly in low earth orbit. The universities are working collaboratively as each university builds its own 1U CubeSat and the data collected is to be shared between the three universities. The mission objectives are to provide undergraduate students with a hands—on flight project experience, to obtain data on atmospheric density and variability in LEO. A Hampton University student team will perform analysis of spacecraft attitude, location, and orbital data to measure variations in atmospheric density in low earth orbit. Differing from Libertas and Ceres, Aeternitas will deploy a petal—like drag brake (similar to a deployable solar panel array) and will deorbit at an accelerated rate for the purposes of providing additional atmospheric drag data. CONOPS After deployment from the NanoRacks deployer and remaining off for the required 3O0min, the antenna will deploy. Once enough power has been stored and the attitude has been determined, detumbling via magnetorquers will commence in short bursts. Once the 21

desired attitude stabilization is reached, Aeternitas will proceed with normal operations in which attitude and GPS data is recorded. The scientific data, and health updates will be downlinked to the ODU ground station during overflights. After initial data has been collected and downlinked, Aeternitas will deploy four drag brake petals that will remain connected to the satellite during de—orbit. Materials Aeternitas‘ chassis is made of Aluminum 6061—16. It contains standard commercial off— the—shelf (COTS) materials, electrical components, PCBs and solar cells. The Aeternitas‘ payload includes a ceramic patch antenna and the cover plate for the antenna assembly will be printed from Windform. Hazards There are no pressure vessels, hazardous, or exotic materials. Batteries Aeternitas is using the GOMspace NanoPower P31u EPS which controls the charging and discharging of two 1—cell lithium—ion batteries. The EPS features under—voltage and over—voltage protection as well as over—current protection via power distribution switches. 22

Libertas —University of Virginia (Virginia CubeSat Constellation) — 1U Custom Solar Panel Version B.. "Top Mounted Solar Panel ho , {EnduroSat Integrated) ADCS Permanent Magnet Mounting Assembly "\ mR EnduroSal UHF Antenna Radio Board ~, S 1 ~, [Depioyed State) Motherboard and Pluggable \ Cover Plate with 3 Inhibit Switches Processor Module ~*~.__ ® € ~ ~—Solar Panel Clips (8) & & s d PiPatch GPS _ f Antenna ~~~~~ é& mt ~~ Custom Solar Panel Version A {3} Figure 9: Libertas Expanded View Overview The Virginia CubeSat Constellation (VCC) mission is a joint operation between teams at Old Dominion University, University of Virginia, Virginia Tech, and Hampton University. ODU, UVA, and VT are each building 1U CubeSats (Aeternitas, Libertas, and Ceres, respectively) that will fly in low earth orbit. The universities are working collaboratively as each university builds its own 1U CubeSat and the data collected is to be shared between the three universities. The mission objectives are to provide undergraduate students with a hands—on flight project experience, to obtain data on _ atmospheric density and variability in LEO. A Hampton University student team will perform analysis of spacecraft attitude, location, and orbital data to measure variations in atmospheric density in low earth orbit. CONOPS Upon deployment from NanoRacks, Libertas will initiate a 30 minute countdown timer before powering up, as required by the NanoRacks deployer ICD. The satellite will enter a commissioning period in which the satellite has its initial power—up, deploys the UHF antenna if there is sufficient battery power available, and performs a system health check. The CubeSat will detumble using a passive magnetic attitude control system. Once the desired attitude stabilization is reached, Libertas will proceed with normal operations in 23

which attitude and GPS data is recorded. This data and health updates will be downlinked to the UVA ground station during overflights. Material The Pumpkin CubeSat Kit 1U chassis is constructed primarily from Aluminum 5052. Internal components are either commercial—off—the—shelf or fabricated from common materials such as custom PCBs and aluminum brackets inside the spacecraft for securing magnets used for PMAC and separation switches. Hazards There are no pressure vessels, hazardous, or exotic materials. Power Systems/Hazards The electrical power storage system will consist of a Clyde Space 3rd Generation. EPS and battery system that uses lithium—ion polymer cells with over—charge/current protection circuitry. 24

Ceres — Virginia Tech (Virginia CubeSat Constellation) — 3U Clyde Space Solar Panet EPS/Batftery Board Motherboard Basepiate — s Clyde Space Rail \—Custom Side Panel ‘UHF Antenna w _ '\\:f/ Top Sotar Panel Figure 10: Ceres Expanded View Overview The Virginia CubeSat Constellation (VCC) mission is a joint operation between teams at Old Dominion University, University of Virginia, Virginia Tech, and Hampton University. ODU, UVA, and VT are each building 1U CubeSats (Aeternitas, Libertas, and Ceres, respectively) that will fly in low earth orbit. The universities are working collaboratively as each university builds its own 1U CubeSat and the data collected is to be shared between the three universities. The mission objectives are to provide— undergraduate students with a hands—on flight project experience, to obtain data on atmospheric density and variability in LEO. A Hampton University student team will perform analysis of spacecraft attitude, location, and orbital data to measure variations in atmospheric density in low earth orbit. ; 25

CONOPS Following deployment, Ceres will power up and start a countdown timer. After thirty minutes have passed, a UHF turnstile antenna will be deployed. For the first few passes the ground station operators will attempt communications to perform checkouts of the spacecraft. Following successful checkout, the primary science mission will begin and continue for at least 3 months. This includes recording attitude and GPS data. The scientific data, and health updates will be downlinked to the VT ground station during overflights. Materials The CubeSat rail structure and skeleton is made of Aluminum 5052—H32. Non—critical parts of the chassis are made of a 3D printed Ultem 1010 derivative with added carbon nanotubes, similar to GSC31264. It contains all standard commercial off the shelf (COTS) materials, electrical components, PCBs and solar cells. Hazards There are no pressure vessels, hazardous or exotic materials. Batteries The electrical power storage system consists of a Clyde Space 3rd Generation EPS and battery system that uses lithium—ion polymer cells with over—charge/current protection circuitry.. 26

Section 3: Assessment of Spacecraft Debris Released during Normal Operations The assessment of spacecraft debris requires the identification of any object (>1 mm) expected to be released from the spacecraft any time after launch, including object dimensions, mass, and material. The section 3 requires rationale/necessity for release of each object, time of release of each object, relative to launch time, release velocity of each object with respect to spacecraft, expected orbital parameters (apogee, perigee, and inclination) of each object after release, calculated orbital lifetime of each object, including time spent in Low Earth Orbit (LEO), and an assessment of spacecraft compliance with Requirements 4.3—1 and 4.3—2. No releases are planned on the ELaNa—21 CubeSat mission therefore this section is not applicable. 27

Section 4: Assessment of Spacecraft Intentional Breakups and Potential for Explosions. There are NO plans for designed spacecraft breakups, explosions, or intentional collisions on the ELaNa—21 mission. The probability of battery explosion is very low, and, due to the very small mass of the satellites and their short orbital lifetimes the effect of an explosion on the far—term LEO environment is negligible (ref (h)). . The CubeSats batteries still meet Req. 56450 (4.4—2) by virtue of the HQ OSMA policy regarding CubeSat battery disconnect stating; "CubeSats as a satellite class need not disconnect their batteries if flown in LEO with orbital lifetimes less than 25 years." (ref. (h)) Limitations in space and mass prevent the inclusion of the necessary resources to disconnect the battery or the solar arrays at EOM. However, the low charges and small battery cells on the CubeSat‘s power system prevents a catastrophic failure, so that passivation at EOM is not necessary to prevent an explosion or deflagration large enough to release orbital debris. The 6U CubeSat in this complement satisfy Requirements 4.4—1 and 4.4—2 if their batteries at equipped with protection circuitry, and they meet International Space Station (ISS) safety requirements for secondary payloads. Additionally, these CubeSats are being deployed from a very low altitude (ISS orbits at approximately 400 km), meaning any accidental explosions during mission operations or post—mission will have negligible long—term effects to the space environment. Assessment of spacecraft compliance with Requirements 4.4—1 through 4.4—4 shows that with a maximum CubeSat lifetime of 3.9 years maximum, the ELaNa—21 CubeSats are compliant. 28

Section 5: Assessment of Spacecraft Potential for On—Orbit Collisions Calculation of spacecraft probability of collision with space objects larger than 10 cm in diameter during the orbital lifetime of the spacecraft takes into account both the mean cross sectional area and orbital lifetime. The largest mean cross sectional area (CSA) among the 13 CubeSats is that of the SPACE HAUC CubeSat with solar arrays deployed. Comm. System Solar Paneis 7 uid § wca0h | z* T Thermal ho Insuilation Structure Camera Figure 10:; SPACE HAUC Expanded View (with solar panels deployed) [2 * (w * 1 :4 * (w * A)] Mean CSA = ZSurfl:ce Area _ Equation 1: Mean Cross Sectional Area for Convex Objects A +Aq + A Mean CSA = (ma—ng Equation 2: Mean Cross Sectional Area for Complex Objects All CubeSats evaluated for this ODAR are stowed in a convex configuration, indicating thereare no elements of the CubeSats obscuring another element of the same CubeSats from view. Thus, the mean CSA for all stowed CubeSats was calculated using Equation 1. This configuration renders the longest orbital life times for all CubeSats. Once a CubeSat has been ejected from the NanoRacks dispenser and deployables have been extended, Equation 2 is utilized to determine the mean CSA. Amax is identified as the view that yields the maximum cross—sectional area. A1 and A» are the two cross— sectional areas orthogonal to Amax. Refer to Appendix A for component dimensions used in these calculations The SPACE HAUC (2.9 kg) orbit at deployment is 408 km apogee altitude by 400 km perigee altitude, with an inclination of 51.6 degrees. With an area to mass ratio of 0.00398 m*/kg, DAS yields 3.9 years for orbit lifetime for its stowed state, which in turn 29

is used to obtain the collision probability. Even with the variation in CubeSat design and orbital lifetime ELaNa—21 CubeSats see an average of 0.0 probability of collision. All CubeSats on ELaNa—21 were calculated to have a probability of collision of 0.0. Table 3 below provides complete results. There will be no post—mission disposal operation. As such the identification of all systems and components required to accomplish post—mission disposal operation, including passivation and maneuvering, is not applicable. 30

Table 3: CubeSat Orbital Lifetime & Collision Probability CubeSat CapSat CySat Phoenix SPACE HAUC TechEdSat 8 Mass (kg) 2.8 2.7 3.2 2.9 7.9 Mean C/S Area {m"2) 0.0371 0.027 0.015 0.0116 0.104 "O $ Area—to Mass (m*2/kg) 0.0133 0.014 0.012 0.004 0.0132 g Orbital Lifetime (yrs) 2.5 2.3 3.7 3.9 2.2 Probability of collision (10"X) 0.00000 0.00000 0.00000 0.00000 0.00000 & Mean C/S Area (m"~2) 0.1263 0.0709 0.564 § Area—to Mass (m*2/kg) 0.0454 0.0243 0.0714 G Orbital Lifetime (yrs) 0.62 1.09 0.482 3 Probability of collision (10"X) 0.00000 0.00000 0.00000 Solar Flux Table Dated 8/14/2017 **Note: Blacked out areas represent CubeSats which do not have deployables or have deployable antennae with negligible areas with respect to on—orbit dwell time calculation. Data for TechEdSat—8 taken from Ames—submitted ODAR report.

Table 4: CubeSat Orbital Lifetime & Collision Probability CubeSat HBEVERD UNITE Aeternitas Ceres Libertas Mass (kg) 2.6 3.5 1.2 1.2 1.4 Mean C/S Area {mA2) 0.085 0.079 0.0163 0.0176 0.0182 G Area—to Mass (m~2/kg) 0.0327 0.0225 0.0136 0.0147 0.0130 3 Orbital Lifetime (yrs) . 0.77 1.2 L 2.3 . 2.0 2.4 2 42 — > P'°b-ab"(';z,‘\’;)°°"'s'°" 0.00000 0.00000 0.00000 0.00000 0.00000 § Mean C/S Area {(m42) * 3 Area—to Mass (m‘2/kg) 8 _Orbital Lifetime (yrs) g} Probability of collision (104%) Table 3: CubeSat Orbital Lifetime & Collision Probability (cont.)

The probability of any ELaNa—21 spacecraft collision with debris and meteoroids greater than 10 cm in diameter and capable of preventing post—mission disposal is less than 0.00000, for any configuration. This satisfies the 0.001 maximum probability requirement 4.5—1. The VCC CubeSat Aeternitas will deploy a petal—like drag brake, for the purpose of providing data regarding drag effects upon its orbit. This feature does not increase the probability of on—orbit collision. The ELaNa—21 CubeSats have no capability or plan for end—of—mission disposal, therefore requirement 4.5—2 is not applicable. In summary, assessment of spacecraft compliance with Requirements 4.5—1 shows ©ELaNa—21 to be compliant. Requirement 4.5—2 is not applicable to this mission. Section 6: Assessment of Spacecraft Post Mission Disposal Plans and Procedures All ELaNa—21 spacecraft will naturally decay from orbit within 25 years after end of the mission, satisfying requirement 4.6—1a detailing the spacecraft disposal option. Planning for spacecraft maneuvers to accomplish post—mission disposal is not applicable. Disposal is achieved via passive atmospheric reentry. Calculating the area—to—mass ratio for the worst—case (smallest Area—to—Mass) post— mission disposal among the CubeSats finds SPACE HAUC in its stowed configuration as the worst case. The area—to—mass is calculated for is as follows: Mean C/SArea (m") a m m W = Area — to — Mass (E) Equation 3: Area to Mass 0.0116 m _ 0004m2 29kg _ kg SPACE HAUC has the smallest Area—to—Mass ratio and as a result will have the longest orbital lifetime. The assessment of the spacecraft illustrates they are compliant with Requirements 4.6—1 through 4.6—5. DAS 2.1.1 Orbital Lifetime Calculations: DAS inputs are: 408 km maximum apogee 400 km maximum perigee altitudes with an inclination of 51.6° at deployment no earlier than April 2018. An area to mass ratio of ~0.004 m*/kg for the SPACE HAUC CubeSat was used. DAS 2:1.1 yields a 3.9 years orbit lifetime for SPACE HAUC in its stowed state. ~ This meets requirement 4.6—1. For the complete list of CubeSat orbital lifetimes reference Table 4: CubeSat Orbital Lifetime & Collision Probability. Assessment results show compliance. 33

Section 7: Assessment of Spacecraft Reentry Hazards A detailed assessment of the components to be flown on ELaNa—21 was performed. (Data provided for TechEdSat—8 in their submitted ODAR report was reviewed as well). The assessment used DAS 2.1.1, a conservative tool used by the NASA Orbital Debris Office to verify Requirement 4.7—1. The analysis is intended to provide a bounding analysis for characterizing the survivability of a CubeSat‘s component during re—entry. For example, when DAS shows a component surviving reentry it is not taking into account the material ablating away or charring due to oxidative heating. Both physical effects are experienced upon reentry and will decrease the mass and size of the real—life components as the reenter the atmosphere, reducing the risk they pose still further. An assessment of the components flown on TechEdSat—8 is contained in Reference J. The following steps are used to identify and evaluate a components potential reentry risk relative to the 4.7—1 requirement of having less than 15 J of kinetic energy and a 1:10,000 probability of a human casualty in the event the survive reentry. 1. Low melting temperature (less than 1000 °C) components are identified as materials that would never survive reentry and pose no risk to human casualty. This is confirmed through DAS analysis that showed materials with melting temperatures equal to or below that of copper (1080 °C) will always demise upon reentry for any size component up to the dimensions of a 1U CubeSat. 2. The remaining high temperature materials are shown to pose negligible risk to human casualty through a bounding DAS analysis of the highest temperature components, stainless steel (1500°C). If a component is of similar dimensions and has a melting temperature between 1000 °C and 1500°C, it can be expected to possess the same negligible risk as stainless steel components. Table 4: ELaNa—21 High Melting Temperature Material Analysis CAPSat Antenriae Stainless Steel .0176 0 0 CAPSat Pointing Panel 301 Stainless Steel .0382 0 10 CAPSat Face Seal Edge Connector 316 Stainless Steel .0093 77.5 o CAPSat Gear Pump 316 Stainless Steel .110 68.6 o CAPSat Bellows Accumulator 316 Stainless Steel .218 63.8 0 CAPSat Pressure Sensors 316 Stainless Steel .079 70.3 0 CAPSat Radiator Panel Hinge Unfinished Steel .0068 76.5 0 CAPSat napste! 9t Standoffs 18—8 Stainless Steel 0055 73.8 o CAPSat Pipe Fittings sunilessSiee] (generic) various 75.2 0 CySat Rods AanlessSite {generic} .080 0 o CySat Standoffs ie (generic) .084 72.8 o 34

CySat Fasteners ssibles es! {generic) .040 — 77.0 0 X Stainless Steel h . k f CySat Separation Switches flsr 028| 0 ‘0_ CySat RBF Pin esc {generic) .017 74.7 o CySat Separation Springs est (generic} .0002 77.3 , > 0 CySat Reaction Wheel Brass .060 _ Jaiye 0 _ CySat Magnetometer _ Aluminum .005 77.3 26 c Deployable ns 1 Ts CySat egherameter Brass .002 77.8 0 Phoenix Screws stainiess .StEEI 6.94 77.7 0 (generic) Phoenix Nuts Steinlges Stes! {generic) 3.92 77.6 o Phoenix Electronics Stack Rod seinless .Steel 4,29 76.8 0 (generic) Phoenix Separation Springs Stainless _Steel 0.072 77.9 0 C m _ {generic) _ Stainless Steel TJREVERB Standoff screws % .020 77.7 0 (generic) TiRrevers & 6 mm c screws stoier—Sice! (generic) .064 717.5 o SRX Virginia CC: Acternitas Antenna Blades Steel/copper plate 0005 0 0 Vifeihls 55: Separation Switches Beryllium Copper 003 0 0 Ceres Virginia CC: Solar Pane} Retaining Stainless Steel .oo1 o o Ceres Clips Virginia CC: Magnet Mounting. Aluminum — .050 o o Ceres Plates Virginia CC: & — : Libertas Separation Switches Beryllium Copper .003 0 0 The majority of stainless steel components demise upon reentry and all CubeSats comply with the 1:10,000 probability of Human Casualty Requirement 4.7—1. A breakdown of the determined probabilities follows: 35

Table 5: Requirement 4.7—1 Compliance by CubeSat CapSat Compliant 1:0 CySat Compliant 1:0 SPACE HAUC Compliant 1:0 TechEdSat—8 Compliant 1:0 TJREVERB Compliant 1:0 UNITE Compliant 1:0 Virginia & . CC:Aeternitas Compliaht Fo Virginia CC: Compliant 1:0 Ceres Virginia CC: . . Libertas Compliant 1:0 *Requirement 4.7—1 Probability of Human Casualty > 1:10,000 If a component survives to the ground but has less than 15 Joules of kinetic energy, it is not included in the Debris Casualty Area that inputs into the Probability of Human Casualty calculation. This is why all of the ELaNa—21 CubeSats have a 1:0 probability as none of their components have more than 15J of energy. All CubeSats launching under the ELaNa—21 mission are shown to be in compliance with Requirement 4.7—1 of NASA—STD—8719.14A. 36

Section 8: Assessment for Tether Missions ELaNa—21 CubeSats will not be deploying any tethers. ELaNa—21 CubeSats satisfy Section 8‘s requirement 4.8—1. 37

Section 9—14 ODAR sections 9 through 14 pertain to the launch vehicle, and are not covered here. Launch vehicle sections of the ODAR are the responsibility of the CRS provider. If you have any questions, please contact the undersigned at 321—867—2098. /original signed by/ Yusef A. Johnson Flight Design Analyst a.i. solutions/KSC/AIS2 ce: VA—H/Mr. Carney VA—H1/Mr. Beaver VA—H1/Mr. Haddox VA—C/Mr. Higginbotham VA—C/Mrs. Nufer VA—G2/Mr. Treptow SA—D2/Mr. Frattin SA—D2/Mr. Hale SA—D2/Mr. Henry Analex—3/Mr. Davis Analex—22/Ms. Ramos 38

Appendix Index: Appendix A. ELaNa—21 Component List by CubeSat: CAPSat Appendix B. ELaNa—21 Component List by CubeSat: CySat Appendix C. ELaNa—21 Component List by CubeSat: Phoenix Appendix D. ELaNa—21 Component List by CubeSat: SPACE HAUC Appendix E. ELaNa—21 Component List by CubeSat: TJREVERB Appendix F. ELaNa—21 Component List by CubeSat: UNITE Appendix G. ELaNa—21 Component List by CubeSat: Virginia CC: Aeternitas Appendix H. ELaNa—21 Component List by CubeSat: Virginia CC: Ceres Appendix I. ELaNa—21 Component List by CubeSat: Virginia CC: Libertas Appendix J. ELaNa—21 TechEdSat—8 ODAR (produced by NASA—Ames) 39

Appendix A. ELaNa—21 Component List by CubeSat: CapSat 1 Aluminum 6061 Rail 2 Bottom Plate 1 Aluminum 6061 Plate 110.6 100 100 100 No — Demise 3 Antennae 4 Stainless Steel Strip 4.4 6.7 22 100 Yes 2642° Okm 4 Short Radiation Shielding 2 Carbon Fiber Plate 38.7 80 284.3 100 No — Demise 5 Tall Radiation Shielding 1 Carbon Fiber Plate 44.5 80 327 76 No — Demise 6 ooiaic® *n 1 Carbon Fiber Plate 442 so 327 20 No — Demise 7 Solar Cell — 25 Solar Cell Panel 3.2 40 70 N/D No — Demise 8 Short Flex Cable 2 Kapton and PCB riy Cable 4.4 54 18.34 ND No — Demise ~ @ Components 9 Tall Flex Cable 2 Kaptonand PCB Components riuCable 4.5 54 18.34 ND No — Demise 10 Top Plate 1 Alurainum 6061 Plate 76.3 100 100 90.17 No — Demise 11 Middle Plate 1 Aluminum 6061 Plate 65.3 93.65 93.65 46 No — Demise 12 Batteries 4 Lithium—on BACCY chemistry Cyinger 46. 184 65 90 No — Demise 13 Battery Support Plate 1 Aluminum 6061 Plate 25.2 86 86 33.02 No — Demise 14 Magnetometer 4 Circuit Board Board 3.7 40 30 90.17 No — Demise 15 Daughter Card 1 Circuit Board Plate 10.1 60 30 90 No — Demise 16 Power Board 1 Circuit Board Board 56.5 94 90 70 No — Demise 17 C&DH Board (was CPU) 1 Circuit Board Board 40 90 90 90 No — Demise 18 Torque Coil 6 FR—4 Plate 52.218 86 76.2 90.17 No — Demise 19 Torque Coil Plate 1 Aluminum 6061 Plate 17.8 74 74 87 No — Demise 20 GOMSPace Radio 2 Aluminum 6061 and Circuit Board Plate 24.5 65 40 31.75 — No — Demise 21 Radio Carry Board 2 Circuit Board Board 19.1 60 90 40 No — Demise 22 Pointing Panel 1 301 Stainless Steel Panel 128 80 326.5 14.7 Yes 2800 ° 0 km 23 Strain gauge 8 Silicon Board 2 7.5 10.8 6.5 No — Demise 40

Piezoelectric Ceramic 24 PL140 Bending Acutator 4 (PIC 252) Board 2.1 11.2 45.5 15.5 No — Demise 25 L—bracket 2 Aluminium 6061 Board 4.67 17 39454 51.64 No — Demise 26 Shaft 1 Aluminium 6061 Board 2.54 7.071 44 ND No — Demise 27 Q—614 Rotary Actuator 2 Piezoelectric Ceramic Board 9 17.91 17.5 ND No — Demise 28 Circuit Board 3 Circuit Board Board 20 46.5 91.5 N/D No — Demise 29 Standoffs 12 Aluminium 6061 Hexagonal Cylinder | g37 5.2 ND No — Demise 30 Vibration Motor 1 Cast Iron Board 88 32 32 8.26 No — Demise Carbon Fiber Radiator Panel Carbon Fib 31 with —Thermal Coati Control 1 arvon FLDOCT Composite Plate 62.6 80 327 1.5 No — Demise oating 32 Face—Seal Edge Connector 2| 316 Stainless Steel "**Z&MREU®T 466 9.53 1210 19 Yes 2500° Demise 33 Gear Pump 1 316 Stainless Steel Cylinder 110.00 31.90 60.67 32 No 2500° Demise 34 Water Block Heat Exchanger 1 Aluminum 6061 Re‘gfi;fi‘:lar 17.00 25.30 25.30 1.2 No — Demise 35 Bellows—Accumulator 1 316 Stainless Steel 5‘;}1’1‘32 218.00 43.94 29.85 5.00 No 2500° Demise 36 Pressure Sensors 2 316 Stainless Steel Cylinder 39.70 18.90 31.00 N/A No 2500° Demise 37 Radiator Control Board 1 Circuit Board Board 49.89 93.65 85.25 12.40 No — Demise 38 Kapton Heater 1 Kapton Sheet 0.23 25.30 25.30 N/A No — Demise 39 Radiator Panel Hinge 2 Unfinished Steel Plate 3.44 32.00 18.40 N/A No 2500° Demise 40 Cooling — Top Plate 1 Aluminum 6061 Plate 61.71 93.65 93.65 1.60 No — Demise 41 Cooling — Bottom Plate 1 Aluminum 6061 Plate 75.26 93.65 93.65 0.25 No — Demise 42 Cooling — Back Plate 1 Aluminum 6061 Plate 32.79 47.50 93.65 0.60 No — Demise 43 Cooling — Side Plate 1 1 Aluminum 6061 Plate 24.62 47.50 91.15 2.50 No — Demise 44 Cooling — Side Plate 2 1 Aluminum 6061 Plate 23.06 47.50 91.15 2.50 No — Demise Bent Sheet 45 Loop Clamp 2 Aluminum 6061 Metal (in 0.75 16.60 6.50 2.50 No — Demise half—circle) Bent Sheet 46 Hex Clamp 2 Aluminum 6061 Mi‘;}f‘“ 0.60 9.00 20.77 2.50 No — Demise hexagon) 41

Bent Sheet 47 Bellows Clamp 4 Aluminum 6061 Metal 0.52 5.00 46.39 2.50 No — Demise 48 Bellows Ring Clamp 1 Aluminum 6061 Cylinder 1.49 22.00 2.66 0.80 No — Demise 49 Radiator Board Standoffs 4 18—8 Stainless Steel Hgfifiggfi‘l 1.38 4.50 12.00 0.80 No — Demise 50 Pipe Fitting316 — MCB—1016— 4 316 Stainless Steel HEX@800A4! Cylinder 437 7.92 15.37 0.80 No 2500° Demise 51 Pipe Fitting — MBV—1010 1 316 Stainless Steel R“;;‘;fn"l‘“ 1043 1334 17.90 N/A No 2500° Demise 52 Pipe Fitting — MTS—1010 1 316 Stainless Steet RTAMEUST 759 7.94 15.88 N/A No 2500° Demise 53 Pipe Fitting — SMCBT—1016 2 316 Stainless Steet R*TAMETA 925 7.92 32.77 N/A No 2500° Demise 54 Pipe Fitting — MPFA—1810 2 316 Stainless Steel Cylinder 13.52 13.97 14.48 7.92 No 2500° Demise 55 Pipe Fitting — $S—100—1—2RT 2 316 Stainless Steel Hg’y‘fifig‘;’;‘l 1237 1111 26.24 8.00 No 2500° Demise 56 Pipe Fitting — $S—100—1—1 2 316 Stainless Steel Hgffifig‘e‘fl 8.39 787 23.88 792 No 2500° Demise 57 Pipe Fitting — MF—1010 2 316 Stainless Steel }g;‘fifiggfl 3.07 7.94 9.12 N/A No 2500° Demise 58 Pipe Fitting — MH—1031—316 2 316 Stainless Steel Hg;‘fifigg:‘ 2.52 7.94 8.66 N/A No 2500° Demise 59 Pipe Fitting — $S$—100—3 1 316 Stainless Steel Recli‘i':f:laf 20.48 22.56 35.56 N/A No 2500° Demise 60 Pipe Fitting — CF—316—05— 316—E 1 316 Stainless Steel HEX@8ONAl Cylinder ;7;7 7.94 10.80 N/A No 2500° Demise 61 Metal Tubing — Segment 1 1 316 Stainless Steel Bent Tubing 0.64 1.59 46.30 N/A No 2500° Demise 62 Metal Tubing — Segment 2 1 316 Stainless Steel Bent Tubing 0.32 1.59 23.21 9.53 No 2500° Demise 63 Metal Tubing — Segment 3 1 316 Stainless Steel Bent Tubing 0.20 1.59 14.77 N/A No 2500° Demise 64 Metal Tubing — Segment 4 1 316 Stainless Steel Bent Tubing 0.49 1.59 35.27 N/A No 2500° Demise 65 Metal Tubing — Segment 5 1 316 Stainless Steel Bent Tubing 0.53 1.59 38.22 N/A No 2500° Demise 66 Metal Tubing — Segment 6 1 316 Stainless Steel Bent Tubing 0.53 1.59 38.22 N/A No 2500° Demise 67 Metal Tubing — Segment 7 1 316 Stainless Steel Bent Tubing 0.90 1.59 65.14 N/A No 2500° Demise 68 Metal Tubing — Segment 8 1 316 Stainless Steel Bent Tubing 0.67 1.59 48.38 N/A No 2500° Demise 69 Flexible Tubing — Segment 1 2 Teflon Bent Tubing 0.40 1.59 121.25 N/A No — Demise 70 Multimode Optic Fiber 2 Fiberglass Bent Tubing 0.019 0.125 150 N/A No — Demise 71 Photodiode 4 Silicon Cylinder 0.018 1.02 2.41 N/A No — Demise 42

Appendix B. ELaNa—21 Component List by CubeSat: CySat 1 1° Aluminum 6061 Box 10 340.5 No — 2 Rods 4 18—8 Stainless Steel Cylinder 20 3 340.5 Yes 2650° 0 km 3 Standoffs 28 18—8 Staintess Steet LOHO® Cylinder 3 6 15 No 2650° Demise 4 Fasteners (2M) 50 18—8 Stainless Steel Screw 0.8 2 5 ——— No 2650° Demise 5 Separation Switches A “"m"gfis"“““’ Box 7 3.378 20 12268 No 2550° 0 km 6 RBF Pin 1 Stainless Steel Pin 17 4.67 22.6 ——— No 2550° Demise 7 Separation Springs 2 316 Stainless Steel Cylinder 0.1 2.843 5.258 ——— No 2650° Demise Brass (flywheel), 8 Reaction Wheel 1 Alucoat 650 coated Box 60 28 28 26.2 No 1724° Demise Aluminum (housing) 9 Magnetometer — Deployment 1 Alucoat 6§0 coated Box 5 83.3 16.8 6.5 No _ Demise & Shell Aluminum 10 Depl°yableBbg§i“et°mfl"” ~ 1 Brass Box 2 83.3 16.8 6.5 No 1724° Demise 11 Course Sun Sensor 6 FR4 — ];14});,;“er AS Box 1 3.7 10.8 1.5 No — Demise Supra 50 (Core), Alucoat 650 Cylinder 12 CubeTorquer (magnetorquer) 2 Aluminum (caps), with box 28 13.5 60 18 No — Demise Enamel coated caps copper (windings) Alucaot 650 . Aluminum (body), . 13 Cubecoil (magnetorquer) 1 Enamel coated Plate 46 90 96 8 No — Demise copper (windings) Cube Sense (fine sun sensor FR4 — Elpemer AS { 14 & nadir sensor, PCB) 1 2467 Plate 80 90 96 35 No — Demise 15 CubeComputer (ADCS 1 FR4 — Elpemer AS Plate 56 90 96 10 No _ Demise computer) 2467 16 Cubecontrol 1 Pra —Piponer AS 60 90 96 30 Yes — Demise 17 Battery 1 Lithium—Ion Polymer Box 111.15 90.17 95.89 14 Yes = Demise 43

Electronic Power System 18 (BPS) 1 PCB Material (FR4) Box 126.9 90.17 95.39 124 No Demise 19 Solar Panels PCB Material (FR4) Rectangle approx varies varies 1.6 No Demise 20 XTJ Prime Solar Cells 1 ~ Box 2.27 6.91 3.97 0.225 No Demise 21 *. AaigM 0C — Rectangle 0.75 26.3 10 .01 No Demise 22 Primary Radio 6 Aluminum 6082 Rectangle 90 90.18 95.89 14.56 No Demise 23 PCB 28 FR—4 Substrate Rectangle 54 96.52 90.17 1.6 No Demise 24 g?;;‘;‘;‘;;‘fi‘;%"éé 12 Silica(Amorphoug)A Triangle *57"° 12 12 1.6 No Demise 25 Memory 1 Silica(Amorphous}A Box 0.54 8.26 5.33 2.03 No Demise 26 Linear V"A“;%lléeg"‘a“’" 1| Silica(Amorphoug)A Box 0.008. 24 29. 1.025 No Demise 27 Buffer — SN74LS125ADR 2 Silica(Amorphous)A Box 0.129 8.2 10.5 2 No Demise 28 Buffer — PCAQSL7AD118 4 Silica(Amorphous}A mgfgf‘l‘ifd 0.0745 6.2 5 1.75 No Demise copper, Rogers 4003, Integrated 29 Software—Defined Radio 1 FR4 370HR, stainless “gfii; 260 100 61.97 127 No Demise steel, etc. 30 Antennae 7 flexible polymer Intfgrat.ed 700 No Demise Circuit 31 Payload Board 4 PCB layers, ctc copper, Integrated Circuit §1g 90.17 95.89 127 No Demise 44

Appendix C. ELaNa—21 Component List by CubeSat: Phoenix 1 Phoenix 3U CubeSat 1 Aluminum 7075 Box 3.2 100 325 — — 2 +Z panel 1 Aluminum 7075 Flat Plate 49.76 100 337 100 No Demise 3 —Z panel 1 Aluminaum 7075 Flat Plate 56.71 97 96 97 No Demise 4 +X panel 1 Aluminum 7075 Flat Plate 38.75 83 64 325 No Demise 5 —X panel 1 Aluminum 7075 Flat Plate 38.75 83 34 325 No Demise 6 +Y panel 1 Aluminum 7075 Flat Plate 38.75 83 37.5 325 No Demise 7 —¥ panel 1 Aluminum 7075 Flat Plate 38.75 83 19.34 325 No Demise 3 Comer Rails 4 AMmiMrcoe, FlatPlate 89.65 18.5 0.1528 340 No Demise 9 3U Solar Panels 2 PCB FR—4/Fiberglass Flat Plate 56 83 25 322.5 No Demise 10 S—Band Patch Antenna 1 Aluminum Sphere 50 89 27.94 81.5 No Demise 11 GPS Patch Antenna 1 PCB FR—4/Fiberglass Flat Plate 50 70 1.59 70 No Demise 12 UHF T“r;‘:;fiz Antenna 1 Al‘/‘\’;“;‘;‘({;‘?gi{)ard Box 74 98 12.7 98 No Demise 13 bHS T“"Ef)’(’}: Antenna 4 SMA Cylinder 0.25 1.6 326 136.9 No Demise 14 Sun Sensors 6 PCB FR—4/Fiberglass Box 3.5 27.94 96 17.15 No Demise 15 Deployment Switches 3 Thermoplastic Box 0.016 3.37 96 23.4 No Demise 16 Battery 1 PCB FR—4/Fiberglass Box 335 95.9 93.39 90.2 No Demise 17 EPS 1 PCB FR—4/Fiberglass Flat Plate 148 95.9 87.44 90.2 No Demise 18 MAI ADCS 1 Fintinished Aluminum Box 694 97 16.26 97 No Demise 19 FLIR Tau 2 640 IR Camera 1 olinmsp Cylinder 475 82 144.51 No Demise 20 S—Band Transmitter 1 PCB FR—4/Fiberglass Box 95 96 27 4 90.2 No Demise 21 UHF Transciever 1 polimide Box 24.5 40 19.05 65 No Demise 22 GPS 1 PCB FR—4/Fiberglass Box 24 46 96 71 No Demise 23 OBC 1 polimide Box 24 40 70 65 No Demise 45

24 Motherboard 1 polimide Flat Plate 51 92 20 88.9 No — Demise 25 Interface Board 1 PCB FR—4/Fiberglass Flat Plate 150 92 33.02 88.9 No — Demise 26 FLIR Breakout Board 1 PCB FR—4/Fiberglass Flat Plate 0.88 25 100 14.48 No — Demise 27 Payload Lens Mount 1 Aluminum 7075 Box 47.54 87 152.4 87 No — Demise 28 Payload Core Mount t Aluminum 7075 Box 58.96 87 92 87 No — Demise 29 E‘“"°“i°(st:;;‘c“ Moant 1 Aluminum 7075 Box 44.1 97 144 97 No — Demise 30 Elecmnic(sms‘t;“’k No 1 Aluminum 7075 Box 47. 97 30 97 No Demise 31 G10 Washers 27 G—10 Cylinder 0.06 5 22 0 No — Demise 32 Screws 112 Stainless Steel Cylinder 0.062 2.5 6.93 8 Yes 2550° Demise 33 Nuts 112 Stainless Steel Cylinder 0.035 2.5 96 2 Yes. 2550° Demise 34 Cabling 35 Copper Alloy Cylinder 0.28 26 AWG 96 various No — Demise 35 Thermal Heat Straps 8 Copper Alloy Flat Plate 0.25 10 96 50 No Demise 36 Electronics Stack Rod 4 Stainless Steel Cylinder 429 3.18 96 117.17 Yes 2550° Demise 37 Separation Springs 2 Stainless Steel Cylinder 0.036 4 96 13 Yes 2550° Demise 46

Appendix D. ELaNa—21 Component List by CubeSat: SPACE HAUC 1 SpaceHAUC 3U CubeSat 1 N/A Box 2.92 100 340 12 — — — 2 Spacecraft Bus Side 2 Aluminum 7075—T6 Box 266.62 82.2 337 2.83 No — Demise 3 Solar Panel Frame 4 Aluminum 7075—T6 Panel 308.72 95 96 9.175 No — Demise 4 Camera Plate 1 Aluminum 7075—T6 Plate 71.69 10 64 10 No — Demise 5 Hinge Base: 4 Aluminum 7075—T6 Box 25.08 20 34 5 No — Demise 6 Hinge Rotor 4 Aluminum 7075—T6 Plate 18.12 2.02 37.5 N/A No — Demise 7 Hinge Pin 4 Aluminum 7075—T6 Pin 1.68 0.305 19.34 N/A No — Demise 8 1800 Torsion Spring 4 Alst 3‘;';{ai“‘“s Spring 0.1528 0.0382 o1s2s 0.305 No 2550° Demise 9 4—40 Screws ie Bolt o 6.35 25 ~8.5 No 2550° Demise 10 Dowel Holster 4 Aluminum 7075—T6 Box 9.72 3.175 27.94 N/A No — Demise 11 Dowel 4 Aluminum 7075—T6 Pin 2.192 4.16 1.59 N/A No — Demise 12 Dowel Hex Nut g ABISQ!Samcs Nut 14712 144 127 4.57 No — Demise 13 Compression Spring 4 AISI 3‘;‘:621‘““1"'“ Spring 0.4684 $2.6 326 23 No — Demise 14 Solar Panels 4 Commercial FR4 Panel 624 95 96 5 No — Demise 15 EPS Front Mount 1 Aluminum 7075—1T6 Plate 81.53 95 96 5 No Demise 16 EPS Back Mount 1 Aluminum 7075—T6 Plate 78.24 92.92 93.39 36.75 No — Demise 17 E'°°"°“iig;’;er Supply 1 Commercial FR4 Box 100 93.34 g744 28.71 No — Demise 18 Battery Pack 1 Glass/Polymide Box 270 6.35 16.26 1245 No — Demise 19 Deployment Switch 2 Box 4 No — Demise 20 Wires 1 Copper Wires 20 14 274 5.9 No — Demise 21 NanoSSOC—A60 Sensor Fine Sun ; Box 4 16.51 19.05 1.63 No — Demise 22 TSL2561 Coarse Sun Sensor 8 Chip 24 95 96 10 No — Demise 47

23 Magnetorquer Board 1 Aluminum 7075—T6 Plate 98.22 8.5 70 N/A No — Demise 24 Magnetorquer Rods 3 Copper Cylinder 90 6 20 6.5 No — Demise 25 Magnetorquer Rod Collar 4 Aluminum 7075—T6 Box 5.2 22.86 33.02 2.36 No — Demise 26 gMz(;Eefi)‘:n"g:f 1 Commercial FR4 Chip 2.8 62 100 11.73 No — Demise 27 ADRV9361 1 Board 60 76.2 1524 6.65 No — Demise 28 ADRV9361 Breakout Board 1 Board 80 54 92 1.57 No — Demise 29 Auxilary Mounting Board 1 Aluminum 7075—T6 Plate 26.94 88 144 1.57 No — Demise 30 Base Board 1 Aluminum 7075—T6 Plate 64.09 30 30 41.2 No Demise 31 Camera 1 Cylinder 21 2.5 22 N/A No — Demise 32 Standoff_Camera 4 Aluminum 7075—T6 Cylinder 0.448 6 6.93 20 No — Demise 33 Standoff_ADRV 8 Cylinder 84 95 96 8 No — Demise 34 Tape Antenna Base 1 Aluminum 7075—T6 Plate 114.54 95 96 10 No — Demise 35 Antenna Mounting Brace 1 Aluminum 7075—T6 Plate 123.67 95 96 1.57 No — Demise 36 Back End Board 1 RO4000 Board 35 95 96 1.57 No — Demise 37 Daughter Board 1 RO4000 Board 35 95 96 1.57 No — Demise 38 Patch Antenna 1 RO4000 Board 14 16 23.6 4.74 No — Demise 39 Tope CageMonopole 2 Aluminum 7075—T6 Plate 1.5 22 9 N/A No — Demise 40 Pulley_Monopole Antenna 2 Alurainum 7075—T6 Cylinder 10.76 4.5 5 4.5 No — Demise 41 Spacer_RF Boards 8 AISI 32‘;21‘3“’1"“ Cylinder 4.312 Yes 2550° Demise 42 Wires 1 Copper Wires 20 100 340 0.2 No — Demise 43 Multi—Layer Insulation 1 Insulation Panel 40 No — Demise 44 Radiators 2 Aluminum 7075—T6 Panel 150 N/A N/A N/A No — Demise 45 Paints 1 AZ—93 White Paint Paint 5 100 340 12 No — Demise 48

Appendix E. ELaNa—21 Component List by CubeSat: TJREVERB * -|_ 3 Material _ 2U CubeSat Structural . Aluminum 5052—H32 Chasis 3 GaAs, G10 . 2 SIDE Solar Panel 4 Fiberslass Panel 100 100 226* 2.5 No Demise 3 pos Z Mounting Plate 1 Aluminum 5052 ls,gx 14.935 100 100 1 No Demise a —Z Mounting Plate 1 Aluminum 5052 ggfi; 14.935 100 100 1 No Demise EPS Block, Ragnarok Flight Circuit Boards (FR—4 Plate—like 5 Computer, Aluminum Heat 1 Fiberglass), block 250 96 90 45 No Demise Sink Aluminum Heat Sink 6 12850 “‘I‘é‘efia““y Dust 2 Lithium polymer cylinder 256 96 91 21 No Demise ISIS 3—axis Magnetorquer PCB FR—4 Fiberglass, . 7 Board 1 Aluminum, Copper Board 196 90.1 95.9 17 No Demise Iridium Radio (Iridium 9603— I daughterboard on FR—4 Fiberglass, . 8 motherboard from NAL 1 Aluminum Heat Sink box 130 47 80 10 No Demise Research) GomSpace S—band Radio FR—4 Fiberglass, . 9 (TR&00) 1 Aluminum box 200 92.682 88.875 19.531 No Demise FR—4 Fiberglass, 10 APRS Radio (SATT4) 1 Aluminum, Stainless box 150 95.885 86.17 9.087 No Demise Steel 11 S—Band Patch Antenna 1 Aluminum 8062 box 50 76 — 4 No Demise 12 Patch Antenna(GPS) 1 Aluminum, Ceramic Box 50 25 25 8 No Demise 13 Patch Antenna Near Earth Network 1 — Aluminum, m — Ceramic Box 10 17 17 9 No is Demise 14 S—Band Heat Sink Block 2 Aluminum Box 60 97* 97* 10 No Demise 98 ISIS Antenna Depolyer « 5 Square 98 7 : 15 System (Turnstile) 1 Aluminum 6061 plate 100 (stowed) (sto)wed (stowed) No Demise 16 Interface. Board GPS/fridium 1 FR—4 Fiberglass, Aluminum Square 50 96 92 11.7 No Demise 17 Circuit board standoffs 20 Aluminum 5052* cylinder 1 3 — various No Demise 49

Molex PicoBlade 4 Pin 18 Connector Female 51021 8 Stainless Steel connector 0.3376 — variable No — Demise Series Molex PicoBlade 12 Pin 19 Connector Female 51021 4 Stainless Steel connector 0.4256 — variable No — Demise Series 20 2P Sh"mrfi“[f;d as an RBF , Stainless Steel pin 10 5.08 6.5 2.54 No — Demise 21 M3, 8mm Screw A(standoff 20 Stainless Steel Screws 1 3* — 8 Yes 2500° Demise screws) »» MolexPicoBlade 4 Pin $ Snss Snd ncA ; 1f I $F 245 N Demi Connector Male 53047—0210 giutices Sice conteer j ' ° ~ mise 23 M2.5, 6mm screw 64 Stainless Steel Screws 1 2.5* — 6 Yes 2500° Demise . Acrylic 24 Kapton Tape — Tape Adhesive 22.5 — — — No — Demise (Coating) 50

Appendix F. Elana—23 Component List by CubeSat: UNITE UNITE CubeSat 1 AnodizedMQMINUM Boy 108.15 10815 No — Demise 2 NSL 3U Bus 2 AmodizedAIOMINUM piang 835 99.95 31622 99.95 No — Demise 3 End Plates g AnodizedMOMIAUM plangr 275 1212 9995 9995 Yes — Demise 4 Side Panels 3 Ceramic, PCB FR—4 Planar 204 1.59 98.95 82.95 No — Demise 5 Patch Antenna 1 Ceramic, PCB FR—4 Planar 45 1.75 35.1 35.1 No — Demise 6 Duplex Antenna 3 PCB FR—4 Planar 21.3 9.66 48.41 48.41 No — Demise 7 8—Cell Solar Panels 1 PCB FR—4 Planar 291.9 82.95 316.25 1.59 Yes — Demise 8 6—Cell Solar Panel 30 GaAs Planar 64.8 82.59 240.18 1.59 No — Demise 9 Solar Cells 4 PCB FR—4 Planar 93 39.7 69.11 0.2 No — Demise 10 PCB Fins 1 _ Cylindrical 16.4 0.79 80 9.75 No — ‘Demise 11 Horizon Sensor 38 Nylon Cylindrical 1.1 8.03 20.2 N/A No — Demise 12 Spacers 146 Stainless Steel Cylindrical 2.28 3.18 N/A 3.18 No — Demise 13 Fasteners 4 _ Box 29.2 2.18 N/A N/A No 2500° Demise 14 Deployment Switches 2 _ Box 13 6.35 6.5 20 No — Demise 15 Diagnostic Port 1 _ Cylindrical 10 7.17 13.29 5.38 No — Demise 16 RBFPN i o> 200 10 40 swas _|._"" _ Demise 17 Batteries 4 Lithium Polymer Box 13.8 31.75 31.75 9.59 No ~ Demise 18 Magnetometer Board 1 PCB FR—4 Planar 35 1.59 57.14 45.2 No ~ PDemise 19 Langmuir Probe Board 1 PCB FR—4 Planar sn 37 90 471 No — Demise 20 NSL EPS/Simplex Board 1 PCB FR—4 Box 125 61 1187 21.59 No ~ Demise 21 NSL Duplex Board 1 PCB FR—4 Planar 27 70 47.5 6.5 No > Demise 22 GPS Board 1 PCB FR—4 Planar a go 80 7p2 No — Besise 23 C&DH Board 1 PCB FR—4 Planar 10 71.5 sois 635 No — Demise 51

24 Magnet Holder 1 Lexan Planar 47 12.7 67.7 6.35 No — Demise 25 Mu — Metal Rod 2 HyMu—80 Cylindrical 420 80 310 3.175 No 2650° Demise 26 Optical Bench 2 Aluminum Planar 3 2.8 12.31 2.8 No — Demise 27 Magnets 3 Neodymium Cylindrical 12.2 10.45 1.64 11.9 No — Demise 28 oBogo_ 2 PCB FR—4 Planar 19.2 218 N/A N/A No — Demise 29 Internal Fasteners 96 Stainless Steel Cylindrical 12 2.18 N/A 6.35 No 2500° Demise 30 Spacers 48 Aluminum Cylindrical 50 N/A N/A N/A No — Demise 31 Cabling — C‘;figfif;gg:fi Linear 377 45 75 5 No — Demise 32 Ballast Mass 2 Copper Alloy Planar 50 N/A N/A N/A No — Demise 33 Silicon — — = 10 N/A N/A N/A No — Demise 34 Epoxy — — — 200 70 40 37.18 No — Demise 52

Appendix G. Elana—23 Component List by CubeSat: Virginia CC — Aeternitas 1 Aeternitas ODU 1U Chassis 1 — ©Box ~ — — — — — — 2 CubeSat Structure +x—Axis — Rails / 1 Aluminum 6061 RS@DEUA r Sheet 57pyo 100 113.5 69 No — Demise 3 CubeSat Structure . x—Axis — Rails // — 1 Aluminum 6061 ReHAPEUA r Sheet 57595 100 113.5 6.9 No — Demise 4 CubeSat Structure +y—Axis — Span // 1 Aluminum 6061 R°RPEUA r Box 15959 19 78.35 415 No — Demise 5 CubeSat Structure y—Axis — Span // — 1 Aluminum 6061 Rec®ne*s r Box 13.99j 15 78.35 415 No — Demise 6 CubeSat Squclure — Bolts and Fasteners 30 Steel Alloy CYHPCMCAl Rods pgyrg 2625 11375 2625 No 2500° Demise 7 Antenna — Cover Plate 1 Windform Box 12443 98 98 1 No — Demise 8 Antenna — Base Plate 1 Aluminum 6061 Box 66.18 96.8 96.8 19.9 No — Demise 9 Anterna—Antcnna Swig 2 Windform L—shaped 1.944 25 50 73 No — Demise 10 Antenna — Antenna Blades 4 Steel/copper plate Sheet 0.5 6 187.4 0.4 Yes 2500° 0 km 11 Antenna — GPS/Tridium Patch Antenna — Toaglas 1 Ceramic Box < 64 25 25 4 No — Demise 12 Drag Brake — Hinge — Top 4 Aluminum 6061 B"(;“’ifiayll‘“ 1 6.921 30 6 No — Demise 13 Drag Brake — Hinge — Bottom 4 Aluminum 6061 Bo‘;(r/i%glllln 1 30 5 1 No — Demise 14 Drag Brake — Petals — petal 1 1 Lexan Box 8 65.4 70.8 1.6 No — Demise s Drag Brake — ie‘als — petal 2— 3 Lexan Box 9 654 70.8 16 No — Demise 16 Drag Brake — Springs 4 Alloy Steel Cylindrical 0.578 4.7244 4.7244 0.5334 Yes 2500° Demise Solar Panels with . Fectanouln 17 Maganetorquers/CSS — 4 Germanium 8 57 No — Demise § r Sheet GOMSpace 18 pNAYy uP3L1 Labs — SkyBox 1 FR4, Metal Alloy Rectangula r Box ,, g4 35 12 No — Demise 19 Lithium Radio — Astro Dev 1 FR4, Aluminum Rei‘;‘;fi“h 48 62 32 1112 No — Demise 53

20 EPS/Battery — GOMSpace 1 Lithium Ion, FR4 R r Boiula 220 96 90 No — Demise 21 Radio Board 1 FR4 5{};::? 40 96 90 2 No — Demise 22 GPS Board 1 rR4 Tae 45 96 90 2 No — Demise 23 Processor Board 1 FR4 S]c>|11;?;c 40 96 90 2 No — Demise Mounting Hardware (4 Stainless Steel Cylindrical 24 Threaded Rods,12 Spacers, 1 * Alumin * Rod, 20 — ~ — Yes 2500° Demise 12 Nuts) o rnine Toroid Pre—evacuated : enamel copper Rectangula _ . 25 Z—axis magnetorquer 1 wire,Space grade rBox 7.5 50 50 4.3 No Demise epoxy 3M 26 Iridium Radio 1 —— Re‘;‘g‘(‘)fi“"‘ 114 29.6 31.5 8.1 No — Demise 27 ©Cables/Connectors —— Copper alloy, —— — — —— —— No — Demise Insulator 28 IMU — MPU—9250 1 Ceramic, X7R Square 1 3 3 1 No — Demise Intersat Radio — HopeRF . 5 29 RFM6OHCW 1 Ceramic, FR4 Square 1 16 16 1.8 No ~ Demise 54

Appendix H. ELaNa—23 Component List by CubeSat: Virginia CC — Ceres 1 Ceres 1U CubeSat — Box — 106.7 106.7 113.5 — — — Ultem 1010 CubeSat Structure (Side Substrate with . 2 Walls) Carbon Nano Tube Plate 8 1 83 95 No — Demise Matrix 3 CubeSat Plate) Structure (Top Aluminum H32 5032— Plate 35 100 100 11.58 No — Demise CubeSat Structure (Bottom Aluminum 5052— ie . 4 Plate) H32 Plate 35 100 100 38.3 No — Demise 5 CubeSat Structure (Rails and Aluminum 5052— Plate 5 §.5 §.5 113.5 No _ Demise Feet) H32 Mother Board; TI . 6 MSP430FR5994 FR4 Plate 88 96 90 1.6 No — Demise 7 Clyde Space 3rd Gen. EPS FR4 Plate 86 95.89 90.17 23.24 No — Demise Processing Module; TI . . 8 MSPA30FS438A FR4 Plate 11 54.6 534 1.6 No — Demise Batteries; ClydeSpace Lithium Ion : 9 c20WHr Polymer, FR4 Plate 246 95.89 90.17 21.4 No — Demise 10 ClydeSpace Solar Panels FR4 Plate 46 83 97 1.6 No — Demise 11 EnduroSat Solar Panel FR4—Tg170 Plate 48 98 98 3.1 No — Demise 12 piNAV GPS FR4, Metal Alloy Box 47 84 35 12 No — Demise 13 Skyfox Lake Silatch OPS ntenna FR4, GPS L1 Patch Plate 50 98 98 5.5 No — Demise EnduroSat UHF Antenna Hard Annodized s 14 Assembly Aluminum, FR4 Plate 85 98 98 5.6 No — Demise 15 Radio Board FR4 Plate 24 96 90 1.6 No Demise 16 GPS and IMU Board FR4 Plate 25 96 90 1.6 No — Demise 17 Astro Dev Radio Li—1 FR4, Aluminum Box 30 62 32 11.12 No — Demise 18 Separation Switches onsue Beryllium Copper Box 3 12.3 20 3.38 Yes 2349° 0 km 19 Separation Spring ASTM A228 Cylinder 1 3 — 10 No — Demise 55

20 Bondable Terminals 2 Plate <1 2.7 1.65 0.6 No — Demise 21 Strain Gauge 2 encapsulated K—alloy Plate <l1 3.18 6.35 0.6 No — Demise 22 Mounting Hardware (Solar Panel Retaining Clips) 8 ts Stainless Steel Plate 1 o20 3 20 0.5 Yes 2500° 0 km 23 IMU; Invensense MPU9250 1 Ceramic, X7R Box 1 3 3 1 No — Demise Intersatellite Radio; HopeRF ; : 24 RFM6OHCW 1 Ceramic, FR4 Box 1 16 16 1.8 No — Demise Mounting Hardware (4 . . Cylindrical 25 ThreadediRods/12 Spacers, 1 iorng. uminum Rod,g 20 2 p C Yes 2500° 0 km 12 Nuts) Toroid 26 Separating Switch Mounts 4 Aluminum Plate 4 12.3 20 20 No — Demise Mounting Hardware (Nuts . Toroid, o . 27 and Bolts Pairs) 30 Stainless Steel Cylindrical 2 3 — 7 Yes 2750 Demise hy . Copper alloy, Flexible . 28 Cabling (Electrical) 1 Insulator Cable 15 2 300 — No — Demise : L Copper alloy, Flexible ; 29 Cabling (Co—ax) 1 Insulator Cable 15 3 400 — No — Demise 56

Appendix I: ELaNa—23 Component List by CubeSat: Virginia CC — Libertas Libertas UVA 1U CubeSat 1 — Box — 106.75 105.66 1186 No — Demise Pumpkin CubeSat Kit . 2 Structure (Side Walls and 1 Al"““’}’g; 5052— Box 104 100 100 113.5 No — Demise Feet) Pumpkin CubeSat Kit Aluminum 5052— Square . 3 Stracture (Top Plate) 1 32 ce 45 100 100 11.58 No — Demise Pumpkin CubeSat Kit Aluminum 5052— Square . 4 Structure (Bottom Plate) 1 H32 ——| _— Plate — 58 160 100 38.3 No ~ Demise Pumpkin CubeSat Kit Square a 5 (CHEiMothsiboard 1 FR4 K# 88 96 90 1.6 No — Demise 6 Clyde Space EPS 1 FR4 Sgl‘;i‘;e 86 95.89 90.17 23.24 No — Demise Pumpkin CSK Pluggable Rectangula ; 7 Processing Moouie 1 FR4 yora 11 54.6 534 1.6 No — Demise Clyde Space 20 WHr Battery Lithium Ion Square _ . 8 Pack (integrated with EPS) 1 Polymer, FR4 Plate 246 95.89 9017 214 No Demise 9 Clyde Space Solar Panels 3 FR4 R"'rci,a]’;tg;“a 46 83 97 1.6 No — Demise Clyde Space Solar Panel Rectangula : 10 RBF Cutout) 1 FR4 1 Plate 46 81.74 111 3.58 No — Demise 11 EnduroSat Solar Panel 1 FR4—Tg170 Rcf“g’;i“la 48 98 98 3.1 No — Demise 12 Skyfox Labs piNAV GPS 1 FR4, Metal Alloy Ref‘g‘(‘)fi“"’ 47 84 35 12 No « Demise Skyfox Labs PiPatch GPS FR4, GPS L1 Square : 13 Antenna 1 $s Pige 50 98 98 5.5 No — Demise EnduroSat UHF Antenna Hard Anodized Square . 14 Assembly 1 Aluminum, FR4 Plate 85 98 98 56 No ~ Demise 15 Radio Board 1 FR4 Slfl;it‘? 24 96 90 1.6 No — Demise 16 GPS and IMU Board 1 FR4 Sg;‘;‘f 25 96 90 1.6 No — Demise 17 Astro Dev Lithium Radio 1 FR4, Aluminum Ref‘g‘;fi“‘a 30 62 32 11.12 No — Demise 18 Magnets for PMAC 20 Al Ni Co Cylindrical 1 3.175 4.953 Yes 2651° Demise 19 Separation Switches 3 nonmeneilhnn eryllium Copper r Box 3 123 20 3.38 Yes 2349° 0 km 57

Separation/deployment Spring . 20 Springs (in CSK cover plate) 1 ASTM A228 Coil 1 3 10 No ~ Demise Mounting Hardware (Solar . o . 21 Panel Retaining Clips) 8 Stainless Steel Bent Plate 1 20 20 0.5 Yes 2500 Demise 22 Invensense MPU9250 IMU 1 Ceramic X7R Box 1 3 3 +1 No — Demise 23 nepeh" REMESTEY Intersatellite radio 1 Cermaic/FR4 Box 1 16 16 1.8 No — Demise 24 Separation Switch Mounts 4 Aluminum Bent Plate 4 12.3 20 20 No — Demise 25 Magnet Mounting Hardware 4 Aluminum Cylindrical 4 4.175 ~ 25 No — Demise Mounting Hardware (Nuts . Toroid, _ . 26 and Bolts Pairs) 45 Stainless Steel Cylindrical 2 3 — 7 Yes Demise Mounting Hardware (4 Stainless Steel Cylindrical 27 Threaded Rods,12 Spacers, 1 ainiess »icel, Rod, 20 ~ — = Yes 2500° Demise \ Aluminum R 12 Nuts) Toroid 28 Cabling(Electrical) 1 Copper alloy, Insulator Flexible Cable 15 2 600 — No — Demise : Copper alloy, Flexible _ . 29 Cabling (RG178 Coax) 1 Insulator Cable 15 3 400 — No — Demise 58

Appendix J: ELaNa—21 TechEdSat—8 ODAR 59

TechEdSat—7 T7MP—06 «* * 3 Orbital Debris Assessment Report (ODAR) XS001 Rev 2 . z TechEdSat—7 Orbital Debris Assessment Report (ODAR) & End of Mission Plan(EOMP) In accordance with NPR 8715.6A, this report is presented as compliance with the required reporting format per NASA—STD—8719.14, APPENDIX A. Report Version: 2 (11/09/2017) DAS Software Used in This Analysis: DAS v2.1.1 Page 1 of 31

| | _ Tech TechEdSat—7 a Orbital Debris Assessment Report (ODAR) TIMP—06— C * XSO01 Rev 2 M VERSION APPROVAL and/or FINAL APPROVAL*: Michael J. Wright Marcus S. Murbach ESM Program Manager TechEdSat 7 Project Manager NASA Ames Research Center NASA Ames Research Center Richard Morrison Michel Liu Safety and Mission Assurance Office Director of Safety and Mission Assurance NASA Ames Research Center NASA Ames Research Center Prepared By: Meredith Campbell Ali Guarneros Luna TechEdSat ME TechEdSat S&MA NASA Ames Research Center NASA Ames Research Center Page 2 of 31

TechEdSat—7 ¥‘ c‘ . . T7ZMP—06— % Orbital Debris Assessment Report (ODAR) Xe001 Rey 7 &n Record of Revisions Rev Date *ricug‘ AGES Description or Chance | Author (s) 0 10/24/2017 All Initial Draft Meredith Campbell 1 10/27/2017 All New mass an launch date Meredith Campbell Update information on 2 11/9/2017 All Revision of ODAR Report Ali Guarneros Luna 8719.14B Page 3 of 31

TechEdSat—7 TIMP—06 4# Orbital Debris Assessment Report (ODAR) XS5001 Rev 2 &.\, Table of Contents Self—assessment and OSMA assessment of the ODAR using the format in Appendix A.2 of NASA—STD—8719.14; Assessment Report Format Mission Description ODAR Section 1: Program Management and Mission Overview ODAR Section 2: Spacecraft Description ODAR Section 3: Assessment of Spacecraft DebrisReleased during Normal Operations ODAR Section 4: Assessment of Spacecraft Intentional Breakups and Potential for Explosions. ODAR Section 5: Assessment of Spacecraft Potential for On—Orbit Collisions ODAR Section 6: Assessment ofSpacecraft Postmission Disposal Plans and Procedures ODAR Section 7: Assessment of Spacecraft Reentry Hazards ODAR Section 8: Assessment for Tether Missions Appendix A: Acronyms Appendix B: Battery Data Sheet Appendix C: Wiring Schematics Page 4 of 31

TechEdSat—7 *# #¥ + Orbital Debris Assessment Report (ODAR) x:;mpl;:—z + Self—assessment and OSMA 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 ofNASA—STD—8719.14. In the final ODAR document, this assessment will reflect any inputs received from OSMA as well. Orbital Debris Self—Assessment Report Evaluation: TechEdSat—7 Mission 1. . aunch Vehicle Spacecraft Comments For all incompletes, includerisk Poneo nt" Compliant (low, Medium, or of non—compliance& Project Risk Tracking # 4.3—1.a Debris Released in LEO. 25 vears note 1. 4.3—1.b Debris Released inLEO. <100 object year limit note 1. 4,3—2 Debris Released in GEO. GEO +/— 200K m note 1. 4.4—1 . . is no explosive hazard. <0.001 6 sion Risk is no explosive hazard. Passive I 4.4—3 o planned breakups. Limit Lerm Risk 4.4—4 Limit BU Short term Risk planned breakups. 1. 4.5—1 1 10cm Risk note 1. 4.5—2 Postmission Risk 4.6—1(a) note 1. )‘!ven‘ Bnoroy 1(b) note 1. is Orhit 0) note 1. t Refricval 4.6—2 note 1. GBO isal 4.6—3 te 1 note 1. MEO Disposal 4.6—4 Disposal Reliability note 1. 4.6—5 note 1. of DeOrbit 4.7—1 Risk note 1. Ground 4.8—1 tethers used. I sR Notes: 1. All of the other portions of the launch stack are non—NASA and TechEdSat—7 is not the lead. Page 5 of 31

TechEdSat—7 r7Mp.—06 24 0t C _ Orbital Debris Assessment Report (ODAR QDl * # port ( ) XSO01 Rev 2 Pre—Launch EOMP Evaluation: TechEdSat Mission Spacecraft EOMP Regm‘t # Comments Compliant N/A Not Compliané N/A en K O 0 O 4.3—1.b <100 object year limit 3 D D D 4.3—2 GEO +/200Km g D D D 4.4—1 0.00L Exslosion Risk D D D Passive RF systems. No 4.4—2 planned breakups. Very low P anairtaalal M C [)] D probability ofbreakup or sepiMe SHCLRY SPLitee debris generation due to explosion. Bd bonsflls 44—3 uis O O C No pianned breakups. t 4.4—4 Limit BW Shom terin Risk g D D D 4.5—1 ©0.001 IGoin Imnact Risk g D D D 4.5—2 Postmission Disaosal R) «k IX D D D 4.6—1(a) Atmosphere Inergy E D D D Ontion 4.6—1(b) N/A. No ability to maneuver Storage Orbit m D D D to higher orbit. 4.6—1 . ies %(,2&‘“1 IZ D D D N/A. Atmospheric re—entry. op ?;gi?;w f [Fe) C O DJ NA NotingEo 4.6—3 N/A. Orbit not between MFEO Disposal D D D LEO and GEO. No operation is required to 4.6—4 > — Disposal Retiability D D D :fi;(;ute atmospheric re— 4.6—5 Summary of DeQrbit E D D D 4.7—1 Ground Population Risk g D D D 4.8—1 Tethers Risk g D D D Page 6 of 31

TechEdSat—7 TIMP—06 + * _ Orbital Debris Assessment Report (ODAR) XS001 Rev 2 M ”b\ Assessment Report Format: ODAR Technical Sections Format Requirements: This ODAR follows the format in NASA—STD—8719.14, Appendix A.1 and includes the content indicated at a minimum in each section 2 through 8 below for the TechEdSat—7 satellite. Sections 9 through 14 apply to the launch vehicle ODAR and are not covered here. ODAR Section 1: Program Management and Mission Overview Mission Directorate: ARC Code R Office Engineer Director: David Korsmeyer, ARC Mission Design Division, Division Chief: Charles Richey, ARC Project Manager/Senior Scientist: Marcus Murbach Schedule of mission design and development milestones from NASA mission selection through proposed launch date, including spacecraft PDR and CDR (or equivalent) dates*: Mission Selection: August 2017 Mission Preliminary Design Review: August 2017 Mission Critical Design Review: January 2018 Launch: April 2018 Begin Operation: April 2018 Mission Overview: The Technical Education Satellite 7 (TechEdSat—7) satellite will be integrated onto Virgin Orbit‘s LauncherOne. TechEdSat—7 will test and validate two different technologies in Low Earth Orbit (LEO): demonstration of the Exo—Brake and demonstration ofthe viability of the Iridium 9602 communication module. The satellite will be inserted into orbit at an apogee of approximately 500 km, perigee of 500 km, with an inclination of 90 degrees. Transmission of data will begin 1 minute after deployment from the launch vehicle. The Exo—Brake will deorbit the satellite approximately 26 weeks after deployment concluding the mission. TechEdSat—7 will fly on the Virgin Orbit CRS—13 mission, and will utilize the Xtenti FANTM RAiL separation system. There are no propellants. Launch vehicle and launch site: Virgin Orbit LauncherOne, Mojave Air & Space Port (MHV) with air launch over the Pacific Ocean Proposed launch date: April, 2018 Page 7 of 31

TechEdSat-7 TIMP—06 +4 * a f Orbital Debris Assessment Report (ODAR); XS001 Rev 2 _ ~ M* Mission duration: 26 weeks Launch and deployment profile, including all parking, transfer, and operational orbits with apogee, perigee, and inclination: TechEdSat—7 will be launched on a Virgin Orbit LauncherOne launch vehicle using the Xtenti FANTM RAiL separation system. The TechEdSat—7 orbit is defined as follows: Apogee: 500 km Perigee: 500 km Inclination: 90 degrees. TechEdSat—7 has no propulsion and therefore does not actively change orbits. TechEdSat—7 will deploy the Exo—Brake, slow down, lose altitude, and then disintegrate upon atmospheric re—entry approximately 26 weeks after deployment. If the Exo—Brake fails to deploy the satellite will re— enter in 256 weeks. Interaction or potential physical interference with other operational Spacecraft: The main risks of this satellite are the Canon BP—930 battery used by the spacecraft (flown in and certified by the ISS program) and the possibility of the TechEdSat—7 impacting another deployment. Since the TechEdSat—7 is a 3.5U CubeSat being launched from the system, and NanoRacks has shown that the likelihood of any CubeSat impacting the ISS is very minimal (validated by the ISS Program Office). Page 8 of 31

TechEdSat—7 17Mp—06 ¥ fi Orbital Debris Assessment Report {(ODAR) 5001 Rew 2 fi z ODAR Section 2: Spacecraft Description Physical description of the spacecraft: TechEdSat—7 is a 2U nanosatellite with dimensions of 10 cm x 10 cm x 21.7 cm and a total mass approximately equal to 2.5 kg. TechEdSat—7‘s payload carries a deployable Exo—Brake as a technology demonstration. The deployed Exo—Brake has a cross—sectional area of 1.25 m*2. TechEdSat—7 will contain the following systems: one power board, one CUBIT RFID tag, one Crayfish board, one Iridium 9602 modem, one OEM 615 GPS, two Canon BP—930 batteries, two patch antennas, and one helical antenna. e The Iridium 9602 modem will have one patch antenna. e The OEM615 GPS shares a dual patch antenna with the Iridium 9602 modem. e The power board will control the deployment ofthe Exo—Brake. Figure 1: TechEdSat—7 Fully Deployed View Page 9 of 31

TechEdSat—7 T7MP—06 + * 9 Orbital Debris Assessment Report (ODAR) XS001 Rev 2 fii Total satellite mass at launch, including all propellants and fluids: 2.5 kg Dry mass of satellite at launch, excluding solid rocket motor propellants: 2.5 kg Description of all propulsion systems (cold gas, mono—propellant, bi—propellant, electric, nuclear): > There will be no propulsion systems on TechEdSat—7. 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. ' Not applicable, there will be no fluids or gasses on board. Fluids in Pressurized Batteries: None. TechEdSat—7 uses unpressurized standard COTS Lithium TIon battery cells. Description of attitude control system and indication of the normal attitude of the spacecraft with respect to the velocity vector: TechEdSat—7 does not have any attitude control system, but it does include an IMU to determine the orientation of the satellite (any attitude control comes from the aerodynamics of the Exo— Brake). Description of any range safety or other pyrotechnic devices: None. The TechEdSat—7 satellite will be launched powered off and a Remove—Before—Flight (RBF) pin is used to prevent accidental activation. Description of the electrical generation and storage system: The power will be generated by solar panels and stored in two Lithium Ion batteries. The batteries that will be used are Canon BP—930 (supplied by the ISS Program Office). See attached data sheet (Appendix B). This battery is approved by the ISS for flight. The dimensions of the battery are 4 x 7 x 3.8 cm and the weight is 0.18 kg. Identification of any other sources of stored energy not noted above: None. Identification of any radioactive materials on board: None. Page 10 of 31

TechEdSat—7 T7MP—06 * * > Orbital Debris Assessment Report (ODAR) XS001 Rev 2 M ODAR Section 3: Assessment of Spacecraft Debris Released durin 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. There are no intentional releases. 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 v2.1) 4.3—1, Mission Related Debris Passing Through LEO: COMPLIANT. No debris released >1mm, while passing through LEO. 4.3—2, Mission Related Debris Passing Near GEO: COMPLIANT. No debris released will transverse GEO. Page 11 of 31

TechEdSat—7 . + Orbital Debris Assessment Report (ODAR) P0o + XS0O01 Rev 2 £A ODAR Section 4: Assessment of Spacecraft Intentional Breakups an Potential for Explosions. 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. Summary of failure modes and effects analyses of all credible failure modes, which may lead to an accidental explosion: 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 (see requirement 4.4—1 below) describe the combined faults that must occur for any of nine (9) independent, mutually exclusive failure modes that could lead to a battery explosion. ' Detailed plan for any designed spacecraft breakup, including explosions and intentional collisions: There are no planned breakups other than during atmospheric entry for disposal. List of components which shall be passivated at End of Mission (EOM) including method of passivation and amount which cannot be passivated: None. Rationale for all items which are required to be passivated, but cannot be due to their design: TechEdSat—7 will be in orbit for 26 weeks with successful deployment of the Exo—Brake based on the DAS analysis shown in this report. Ifthe Exo—Brake fails to deploy, TechEdSat—7 will be in orbit for 256 weeks based on the DAS analysis shown in this.report. Therefore, no post— mission passivation will be performed, as the satellite will burn up on re—entry at the end of the mission. 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, or Mars, or in the vicinity of Sun—Earth or Earth—Moon Lagrange Points: 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 does not exceed 0.001 (excluding small particle impacts) (Requirement 56449).° Page 12 of 31

TechEdSat—7 T7MP—06 4 * * Orbital Debris Assessment Report (ODAR) XS001 Rev 2 M Compliance statement: Required Probability: 0.001. Expected Probability: 0.000. Supporting Rationale and FMEA details: Payload Pressure Vessel Failure: TechEdSat—7 is vented per ISS safety standards. It is not a sealed container. 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 vessel due to the lack of penetration energy. Note also that this same battery combination has been tested extensively, and now flown several times with no noted anomaly. 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: Complete proto—qualification and environmental acceptance tests of the Canon BP—930 battery by JSC ISS program. The acceptance tests are shock, vibration, thermal cycling, and vacuum tests followed by maximum system rate—limited charge and discharge to prove that no internal short circuit sensitivity exists. Combinedfaults requiredfor realizedfailure: 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: Each cell includes a positive temperature coefficient (PTC) variable resistance device that ensures high rate discharge is limited to acceptable levels if thermal rise occurs in the battery. Combinedfaults requiredfor realizedfailure: The PTC must fail AND spacecraft thermal design must be incorrect AND external over current detection and protection must fail for this failure mode to occur. Failure Mode 3: Overcharging and excessive charge rate. Page 13 of 31

TechEdSat—7 r7mp.06 5 * * Orbital Debris Assessment Report (ODAR) + XS001 Rev 2 & C ~ M * Mitigation 3: The satellite bus battery charging circuit design eliminates the possibility of the batteries being overcharged if circuits function nominally. This circuit has been proto— qualification tested for survival in shock, vibration, and thermal—vacuum environments. The charge circuit disconnects the incoming current when battery voltage indicates normal full charge at 8.4 V. If this circuit fails to operate, continuing charge can cause gas generation. The batteries include overpressure release vents that allow gas to escape, virtually eliminating any explosion hazard. Combinedfaults requiredfor realizedfailure: 1) For overcharging: The charge control circuit must fail to function 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 maximum charging rate from a single solar panel when in AM 1.5 G conditions (on Earth, perpendicular to the sun) is 200 mA. The maximum charge rate our battery can accept is 3 A. The battery is a proto—qualified Canon BP—930 from the JSC ISS program, and has four US18650S cells. The battery itself has two parallel strings of 2 cells connected in series, and thus having 4 cells. Due to solar panel current limits and their direction—facing arrangement on the satellite, there is no physical means of exceeding charging rate limits, even if only a single string from the battery was accepting charge. For this failure mode to become active one string must fail to accept a charge AND the charge control circuit on the remaining string fails. The overpressure relief vent keeps the battery cells from rupturing, and is thus limited to worst—case effects of overcharging. 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 4: 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 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). Combinedfaults requiredfor realizedfailure: 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 5: Battery vents are not inhibited by the battery holder design or the spacecraft. Combined effects requiredfor realizedfailure: The manufacturer fails to install proper venting and ISS environmental stress screening fails to detect failed vents. Page 14 of 31

TechEdSat—7 ‘ T7MP—06 + + * Orbital Debris Assessment Report (ODAR) XS5001 Rev 2 M Failure Mode 6: Crushing. Mitigation 6: This mode is negated by spacecraft design. There are no moving parts in the proximity of the batteries. Combinedfaults requiredfor realizedfailure: 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 7: Low level current leakage or short—circuit through battery pack case or due to moisture—based degradation of insulators. Mitigation 7: 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. Combinedfaults requiredfor realizedfailure: 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 8: Excess temperatures due to orbital environment and high discharge combined. Mitigation 8: 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, which are well below temperatures of concern for explosions. Combinedfaults requiredfor realizedfailure: 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 9: Polarity reversal due to over—discharge caused by continuous load during periods of negative power generation vs. consumption. Mitigation 9: In nominal operations, the spacecraft EPS design negates this mode because the processor will stop when voltage drops too low, below 7 V. This disables ALL connected loads, creating a guaranteed power—positive charging scenario. The spacecraft will not restart or connect any loads until battery voltage is above the acceptable threshold. At this point, only the safe mode processor is enabled and charging the battery commences. Once the battery reaches 90% of the peak voltage (around 7.5 V), it will switch to nominal mode and will be able to receive ground commands for continuing mission functions. Combinedfaults requiredfor realizedfailure: The microcontroller must stop executing code AND significant loads must be commanded/stuck "on" AND power margin analysis must be wrong AND the charge control circuit must fail for this failure mode to occur. Page 15 of 31

TechEdsat—7 e jfl Orbital Debris Assessment Report (ODAR) | yegoy rey 3 M S¢ Failure Mode 10: Excess battery temperatures due to post mission orbital environment and constant solar panel overcharge while satellite is powered off. Mitigation 10: Selection of the ISS—approved Canon BP—930 battery packs (GSE from the NASA/Johnson Space Center). These battery packs have battery protection circuits, which prevent over—charge and over—heating. They are lot—tested and supplied as GSE (Government Furnished Equipment) from the NASA/Johnson Space Center. In terms of the orbit environment, the previous TechEdSat—1, TechEdSat—3, and TechEdSat—4, TechEdSat—5 (using the same packaging and battery pack) showed no signs of overeating from environmental heating. 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 cannot cause an explosion or deflagration large enough to release orbital debris or break up the spacecraft. The design of depletion burns and ventings should minimize the probability of accidental collision with tracked objects in space (Requirement 56450). Compliance statement: TechEdSat—7 will be in orbit for 24 weeks with successful deployment of the Exo—Brake. If the Exo—Brake fails to deploy, TechEdSat—7 will be in orbit for approximately 253 weeks based on the DAS analysis shown in this report. Therefore, no post—mission passivation will be performed, as the satellite will burn up on re—entry at the ‘end ofthe mission. Therefore, the TechEdSat—7 battery will meet the above requirement. Requirement 4.4—3. Limiting the long—term risk to other space systems from planned breakups for Earth, lunar, Mars, Sun—Earth Lagrange Point, and Earth—Moon Lagrange Point missions: Planned explosions or intentional collisions shall: a. For LEO—crossing missions, be conducted at an altitude such that for orbital debris fragments larger than 10 cm the object—time product does not exceed 100 object—years. For example, if the debris fragments greater than 10cm decay in the maximum allowed 1 year, a maximum of 100 such fragments can be generated by the breakup. b. Not generate debris larger than 1 mm that remains in Earth, lunar, or Mars orbits or in the vicinity of Sun—Earth or Earth—Moon Lagrange points longer than one year Compliance statement: This requirement is not applicable. There are no planned breakups. Page 16 of 31

TechEdSat—7 TIMP—06 + * * Orbital Debris Assessment Report (ODAR) XS001 Rev 2 M Requirement 4.4—4: Limiting the short—term risk to other space systems from planned breakups for Earth, lunar, Mars, Sun—Earth Lagrange Point, and Earth—Moon Lagrange Point missions: Immediately before a planned explosion or intentional collision, the probability of debris, orbital or ballistic, larger than 1 mm colliding with any operating spacecraft within 24 hours of the breakup shall be verified to not exceed 10—6. 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 v2.1, 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 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 probabilityof accidental collision with space objects larger than 10 cm in diameter does not exceed—0.001. For spacecraft and orbital stages near GEO, the time—integrated probability —when they are in the GEO protection zone —of accidental collision with space objects larger than 10 cm in diameter shall not exceed 0.001 (Requirement 56506). Large Object Impact and Debris Generation Probability: 0.000000; COMPLIANT. Requirement 4.5—2. Limiting debris generated by collisions with small objects when operating in Earth: 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 does not exceed 0.01 (Requirement 56507). Small Object Impact and Debris Generation Probability: 0.000000; COMPLIANT ODAR Section 6: Assessment of Spacecraft Postmission Disposal Plans and Procedures 6.1 Description of spacecraft disposal option selected: Two cases will be considered for this section. The first case is called "Nominal Deployment" in which the Exo—Brake successfully Page 17 of 31

TechEdSat—7 T7MP—06 4 * * Orbital Debris Assessment Report (ODAR) XS001 Rev 2 M deploys and de—orbits the satellite. The second case is called "No Deployment" in which the Exo—Brake fails to deploy and the satellite de—orbits naturally due to atmospheric friction. Case 1: Nominal Deployment The satellite will de—orbit due to the deployed Exo—Brake. There is no propulsion system and burn at re—entry. Case 2: Failed Deployment The satellite will de—orbit naturally by atmospheric re—entry. There is no propulsion system and burn at re—entry. 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: Case 1: Nominal Deployment Spacecraft Mass: 2.5 kg Cross—sectional Area: 1.25 m~2 Area to mass ratio: 1.25/2.5 = 0.5 m~2/kg Case 2: Failed Deployment Spacecraft Mass: 2.5 kg Cross—sectional Area: 0.0217 m*~2 Area to mass ratio: 0.0217/2.5 = 0.00868 m*A2/kg 6.4 Assessment of spacecraft compliance with Requirements 4.6—1 through 4.6—4 (per DAS v 2.1 and NASA—STD—8719.14 section): Requirement 4.6—1. Disposalfor 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: e Leave the space structure in an orbit in which natural forces will lead to atmospheric reentry within 25 yearsafter the completion of mission but no more than 35 years after launch; or e Maneuver the space structure into a controlled de—orbit trajectory as soon as practical after completion ofmission. 8 b. Storage orbit option: Maneuver the space structure into an orbit with perigee altitude greater than 2000 km and ensure its apogee will be bellow GEO altitude — 200 km for 100 years. c. Direct retrieval: Retrieve the space structure and remove it from orbit within 10 years after completion of mission. ' Page 18 of 31

TechEdSat—7 TIMP—06 4 Orbital Debris Assessment Report (ODAR) XS001 Rev 2 fl!,\ Analysis: Case 1: Nominal Deployment * TechEdSat—7 satellite reentry is COMPLIANT using Method "a." TechEdSat—7 will re—enter in 0.504 years (approximately 26 weeks) after launch with orbit history shown in Figure 2. Case 2: Failed Deployment TechEdSat—7 satellite reentry is COMPLIANT using Method "a." TechEdSat—7 will re—enter in 4.928 years (approximately 256 weeks) after launch with orbit history as shown in Figure 3. Altitude 1 > §z m« > — z 5 (:".3 r Wogee:’l—"engee Altitude History for a t:’:'uven (}@] B01 Apoges Alf= 500.00 km Ko2 Perigee Alt = 500.00 km BM3§ |z5cemmmramtnaames 4§2,g—] 0 0000——2 a% f.—..... e se * sc S > 192 g.| [Stort Yyear = 201830 vr SRAAA lc 424. Inclination = 90.00 deg \‘\B‘ CC R&AAN= 0.00 deg . ‘\ 4 . |arg Peri= 0.00 deg ol ® 351—7~) |mean Anomaly = 0.00 deg ud TPTtn td Area—To—Mass = 0.500 m"Z2ikg 301.§¢} «sns se i 2512 — x 4 . — I o ko ow .. BM MJ — «+s kess ssess 180.Johss es e kess ho nc n es enaseeesh sns sn c cesRyraks es |seass> 100.5 sns ds en a en es mes+>} 0.0— § j f f f > 2018.3 2018.4 2018.5 2018.6 2018.6 2018.7 Year — Figure 2: TechEdSat—7 Orbit History for Case 1: Nominal Deployment Page 19 of 31

TechEdSat—7 PMb Gs / * C i Debris Orbital ri Assessment Report eport ((ODAR) o2 ySOD! Rev > * 6 Alt’it:fle | [Apogee/Perigee Altitude History for a Given Orbit] N01 Apoges Alt= 500.00 km (km) i WO2 Perigee Alt = 500.00 km §00.0 }— cce w_ 450.0—h—~>=~~——~~<jr s« us wa «n ate s m N4 . » TaadAo “~\ ............... Start Year.= 2018.30 yr x Inclination = 90.00 deg N 400:0—}— — |RAAN = 0.00 deg i Arg Peri= 9.00 deg '-"s Mean Anomaly = 0.00 deg \ 350.0 Area—To—Mass = 0.009 m"Zikg ——~ 300.0—}— s oo ofamepmoen | 222 250.0 . | 150.0— 100.0—| — > ~— 50.0 0.0— f X : ; F 2018.3 2019.1 2019.9 2020.8 2021.6 2022.4 Yeal Figure 3: TechEdSat—7 Orbit History for Case 2: Failed Deployment Requirement 4.6—2. Disposalfor space structures near GEO.A spacecraft or orbital stage in an orbit near GEO shall be maneuvered at EOM to a disposal orbit above GEO with a predicted minimum perigee of GEO +200 km (35,986 km) or below GEO with a predicted maximum apogee of GEO —200 km (35,586 km) for a period of at least 100 years after disposal. Analysis: Not applicable. TechEdSat—7 orbit is in LEO. Requirement 4.6—3. Disposalfor space structures between LEO and GEO. a. A spacecraft or orbital stage shall be left in an orbit with a perigee greater than 2000 km above the Earth‘s surface and apogee below GEO altitude —200 km for 100 years. b. A spacecraft or orbital stage shall not use nearly circular disposal orbits near regions of high value operational space structures, such as the Global Navigation Satellite Systems near the semi—synchronous altitudes — Page 20 of 31

2 t C TechEdSat—7 Orbital Debris Assessment Report (ODAR) TPME06— + XS001 Rev 2 4 Analysis: Not applicable. TechEdSat—7 orbit is in LEO. Requirement 4.6—4. Reliability ofPost mission Disposal Operations in Earth Orbit: NASA space programs and projects shall ensure that all post mission disposal operations to meet Requirements 4.6—1, 4.6—2, and/or 4.6—3 are designed for a probability of success as follows: a. Be no less than 0.90 at EOM. b. For controlled reentry, the probability of success at the time of reentry burn must be sufficiently high so as not to cause a violation of Requirement 4.7—1 pertaining to limiting the risk of human casualty. Analysis: Case 1: Nominal Deployment TechEdSat—7 de—orbiting relies on the Exo—Brake de—orbiting device. Release of the Exo— Brake will result in de—orbiting in approximately 26 weeks with no disposal or de—orbiting actions required. Case 2: Failed Deployment TechEdSat—7 de—orbiting does not rely on de—orbiting devices. Release with a downward, retrograde vector will result in de—orbiting in approximately 5 years with no disposal or de— orbiting actions required. ODAR Section 7: Assessment of Spacecraft Reentry Hazards Assessment of spacecraft compliance with Requirement 4.7—1: Requirement 4.7—1. Limit the risk ofhuman 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 v2.1 reports that TechEdSat—7 is compliant with the requirement. It predicts that no components on board has more than 15 joules of impact kinetic energy. The majority of TechEdSat—7 including its components and the Exo—Brake will burn up on re—entry. As seen in the analysis outputs below, the highest impact kinetic energies is 0 Joules. Also, there are no titanium components that will be used on TechEdSat— 7. 10 27 2017; 16:50:56PM ****k*x***Processing Requirement 4.7—1 Return Status : Passed Page 21 of 31

TechEdSat—7 TZMP—06— Q Orbital Debris Assessment Report (ODAR) XS001 Rev 2 i8 ***********INPUT**** Item Number = 1 name = TES7 quantity = 1 parent = 0 materialID = 8 type = Box Aero Mass = 2.500000 Thermal Mass = 2.500000 Diameter/Width = 0.100000 Length = 0.217000 Height = 0.100000 name = Door Assembly quantity = 1 parent = 1 materialID = 8 type = Flat Plate Aero Mass = 0.083000 Thermal Mass = 0.083000 Diameter/Width = 0.100000 Length = 0.100000 name = Ejection Plate Assembly quantity = 1 parent = 1 materialID = 77 type = Box Aero Mass = 0.145000 Thermal Mass = 0.145000 Diameter/Width = 0.100000 Length = 0.100000 Height = 0.044000 name = Lower Stack Assembly quantity = 1 parent = 1 materialID = 8 type = Box Aero Mass = 1.512000 Thermal Mass = 1.140000 Diameter/Width = 0.100000 Length = 0.100000 Height = 0.050000 name = Battery quantity = 2 parent = 4 materialID = 70 Page 22 of 31

TechEdSat—7 TZMP—06— Orbital Debris Assessment Report (ODAR) XS001 Rev 2 type = Box Aero Mass = 0.186000 Thermal Mass = 0.186000 Diameter/Width = 0.040000 Length = 0.075000 Height = 0.040000 name = Upper Stack Assembly quantity = 1 parent = 1 materialID = 58 type = Box Aero Mass = 0.142500 Thermal Mass = 0.031500 Diameter/Width = 0.030000 Length = 0.030000 Height = 0.015000 name = Crayfish quantity = 1 parent = 6 materialID = 23 type = Flat Plate Aero Mass = 0.035000 Thermal Mass = 0.035000 Diameter/Width = 0.100000 Length = 0.100000 name = Adam 12 Power Board quantity = 1 parent = 6 materialID = 4 type = Flat Plate Aero Mass = 0.076000 Thermal Mass = 0.076000 Diameter/Width = 0.300000 Length = 0.400000 name = Solar Panel quantity = 4 parent = 1 materialID = 23 type = Flat Plate Aero Mass = 0.053000 Thermal Mass = 0.053000 Diameter/Width = 0.100000 Length = 0.157500 name = Structure quantity = 1 Page 23 of 31

TechEdSat—7 TZMP—O6— Orbital Debris Assessment Report (ODAR) XS0O01 Rev 2 parent = 1 materialID = 8 type = Box Aero Mass = 0.370000 Thermal Mass = 0.370000 Diameter/Width = 0.100000 Length = 0.217000 Height = 0.100000 name = ExoBrake quantity = 1 parent = 1 materialID = 44 type = Flat Plate Aero Mass = 0.135500 Thermal Mass = 0.135500 Diameter/Width = 1.250000 Length = 1.250000 k*************OUTPUT**** Item Number = 1 name = TES7 Demise Altitudée = 77.995979 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 t o ohe ohe ol se whe se ol nle ol se e ol t se ohe ie the whe ol o whe ol n ole e nhe whe le sle ole nhe e o se oke name = Door Assembly Demise Altitude = 76.112137 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 e mhe ohe ce ode e se ohe ce se ohe ue ce ohe ohe nhe ol ohe ue o ce ohe e ohe ce e ohe o tie ce ode e e ohe oke e t name = Ejection Plate Assembly Demise Altitude = 77.550102 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 te se ohe e ohe ol se e se oke ue ce ce ohe ol se nle ohe e se ohe t ole se se nhe se ce ce e e ce ohe o t o oke name = Lower Stack Assembly Demise Altitude = 65.798409 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000600 se se ve e h oe uk o oke whe ole ce ohe ce ce ohe whe e se ohe ohe o ohe o ole ohe se ode ohe ce c ohe ohe ol e se oke name = Battery Demise Altitude = 64.908737 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 Page 24 of 31

TechEdSat—7 Orbital Debris Assessment Report (ODAR) TZMP—06— XS001 Rev 2 we se nhe se ie e whe se w ohe e se ohe ote ce whe ol ohe ole e se ohe ol e se nle se e ce wle ol ohe ob ie o ole ie name = Upper Stack Assembly Demise Altitude = 73.529701 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 t we ohe se se e w ohe ohe ole ohe se ohe ce t w d ohe ie ce se nde ce se wl ole nde se ob ohe wl ohe ob ie w ob ote name = Crayfish Demise Altitude = 73.024727 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 te d e e se oke ohe w nhe ohe ce e nhe ole ce w whe ce ue ohe e nde ie e ce ohe n se ohe ohe se ohe ue e e se e name = Adam 12 Power Board Demise Altitude = 73.255264 Debris Casualty Area = 0.000800 Impact Kinetic Energy = 0.000000 te se w ote ce ohe t w she ie se ce whe e ie ohe ohe nle ohe se ie w e e se ohe e ol w ole ohe ole oke se uh e e name = Solar Panel Demise Altitude = 77.472542 Debris Casualty Area = 0.0006000 Impact Kinetic Energy = 0.000000 S e t uie se se se ohe t e ie se t ole ohe ce t ohe se ie se ue ol nle ohe t h ohe ohe o se ohe ol ie ue h ue name = Structure Demise Altitude = 76.177994 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 te ohe e se ohe ue e ohe ole ce e ce se ol ce obe w wle ole ce ce ole ce ohe ohe ohe se ote ue ce ol se ol e whe o e name = ExoBrake Demise Altitude = 77.988823 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 t ohe w se se se ohe whe ole se ol ohe wle she ce ole se ode ohe c oke nhe ol ce ole se se ohe ole e ol ote ote ie whe ohe oke =============== End of Requirement 4.7—1 ====s=s==ses==e==== Requirements 4.7—1b and 4.7—1¢ below are non—applicable requirements because TechEdSat— 7 does not use controlled reentry. 4.7—1, b) NOT APPLICABLE. 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). Page 25 of 31

. + | TechEdSat—7 Orbital Debris Assessment Report (ODAR) T‘MP—96° + 3 XS001 Rev 2 — 5e * 4.7—1 c) NOT APPLICABLE. 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). ODAR Section 8: Assessment for Tether Missions Requirement 4.8—1. Mitigate the collision hazards ofspace tethers in protected regions of space: Intact and remnants of severed tether systems in Earth, lunar, or Mars orbit, in the Sun— Earth Lagrange Points, or in the Earth—Moon Lagrange Points shall limit the generation of orbital debris from on—orbit collisions with other operational spacecraft. Not applicable. There are no tethers in the TechEdSat—7 mission. ODAR Sections 9—14: Launch Vehicle Since the TechEdSat—7 launch vehicle is managed by Virgin Orbit the orbital debris assessment for the launch vehicle will be performed by Virgin Orbit. The following note from NPR 8715.6A, Paragraph P.2.2, is applied, "Note: It is recognized that NASA has no involvement or control in the design or operation ofFederal Aviation Administration (FAA)—licensed launches orforeign or Department ofDefense (DoD)—furnished launch services, and, therefore, these are not subject to the requirements in this NPRfor the launch portion." END of ODAR for TechEdSat—7. Page 26 of 31

TechEdSat—7 TZMP—06— Orbital Debris Assessment Report (ODAR) XS0O01 Rev 2 Appendix A: Acronyms AFRL _ |Air Force Research Lab ARC |Ames Research Center |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 ESMD Exploration Systems Mission Directorate FRR Flight Readiness Review GEO Geosynchronous Earth Orbit ITAR International Traffic In Arms Regulations kilogram m kilometer LEO [Low Earth Orbit Li—lon [Lithium Ion m*~2 Meters squared ml milliliter mm millimeter N/A —« [Not Applicable. ODAR Orbital Debris Assessment Report [TechEdSat—7 Technical Education Satellite—6 ORR Operations Readiness Review OSMA Office of Safety and Mission Assurance PDR Preliminary Design Review PL Payload P—POD |Poly Picosatellite Orbital Deployer PSIa [Pounds Per Square Inch, absolute SRR. [Pre—Ship Readiness Review RAAN Right Ascension of the Ascending Node SESLO Space environment survivability of live organisms (payload) SMA Safety and Mission Assurance Ti Titanium USAF United States Air Force [UTJ UItra Triple Junction L ear Page 27 of 31

TechEdSat—7 Orbital Debris Assessment Report (ODaR) T‘MP—96 XS001 Rev 2 Appendix B: Battery Data Sheet MATERIAL SAFETY DATA SHEET Page 1 of 2 MSDS#—BA0035—01—090218 SECTION 1 IDENTIFICATION OF THE SUBSTANCE/MIXTURE AND OF THE COMPANY/UNDERTAKING Product Name: Lithiumfon Battery omm mm omm mm on _ Product Code: BP—930 Company Name: _CanonIna. c s e Address; _302, Mmmfiho@m—h;flyy 1464;5(& Japan__ Emss e Use ofthe Product: Battery for Video camera _ Supplier: Address: Phone cumber: With regard to airtransport, the International Civil Aviation Organization (ICAO) Packing Instruction 965 Part 1 complies with the Recommendation as is, further, the International Air Transport Association (LATA) adopts ICAO PackingInstraction 965 Part 1. In addition, the regulations ofthe US DepartmentofTransportation forland, sea and airtransportation are based on the UN Recommendations. SECTION 2 MATERIALS AND INGREDIENTS INFORMATION IMPORT ANTNOTE: The battery pack uses four US 186508 lithium—ion rechargeable cells and control cireuit on the PWB. The cells are connected in 2 parallel strings of2 cells in series. The battery pack should not be opened or burned since the following ingredients contained within the cells could be harmful under some circumstance ifexposed or misused. The cells contain neither metallic lithium nor lithium alloy. Cathode: Lithium—Cobalt Dioxides {active material) Polyvinyldiene Fluoride (binder) Graphite (conductive material) Anode: Graphite (active material) Polyvinyldiene Fluoride (binder) Electrolyte: Organic Solvent (non—aqueous liquid) Lithium Salt Others: Heavy metals such as Mercury, Cadmium, Lead, and Chromium are not used in the cells. Enclosure: Plastic (PC) SECTION 3 FRE HAZARD DATA In ease offire, use CO; or dry chemical extinguishers. Date of Issue: September 8, 2009 Revised Date: — Ver. 2009/6/01 Page 28 of 31

TechEdSat—7 17Mp—06 a Orbital Debris Assessment Report (ODAR) SOOL Rev 2 # MATERIAL SAFETY DATA SHEET Page 2 of 2 MSDS#:BA0035—01—090218 SECTION4 HEALTH HAZARD DATA Under normal condition ofuse, these chemicals are contained in sealed can. Risk ofexposure occurs only ifthe cells are mechanically abused. Inhalation: Contents ofan opened cell can cause respiratory itritation. Remove to fresh air immediately and call a doctor. Skin Contact: Contents ofan opened cell can cause skin irritation. ‘Wash skin with soap and water. Eye Contact: Contents ofan opened cell can causeeyeirtitation. Immediately flush eyes thoroughly with water for at least 15 minutes. Seek medical attention. SECTION 5 PRECAUTIONS FOR SAFE HANDLING AND USE Storage: Store within the rec ded limit of—20 degrees C to 45 degrees C (—4 degrees F to 113 degrees F), well—ventilated area Do not expose to high temperature (60 degrees C/140 degrees F). Since short circuit can cause burn hazard or safety vent to open, do notstore with metal jewelry, metal covered tables, or metal belt. Handling: Do not disassemble, remodel, or solder. Do not short + and — terminals with a metal, Do not open the battery pack. Charging: Charge within the limits of0 degrees C to 40 degrees C (32 degrees F to 104 degrees F) temp Chargewith specified charger designed forthis battery pack. Discharging: Discharge within the limits of—10 degrees C to 50 degrees C (14 degrees F to 122 degrees F) temperature. Disposa): Dispose in accordance with applicable federal, state and local regulation. Caution: Attach the cover to the battery pack to prevent short circuits. Do not disassemble. Do not incinerate. Do not expose to temperature above 140 degrecs F. SECTION 6 SPECIAL PROTECTION INFORMATION Respiratory Protection: Not necessary under normal use. Ventilation: Not necessary under normal use. Eye Protection: Not necessary under normal use. Protective Gloves: Not necessary under normal use. Date of Issue: September 8, 2009 Revised Date: — Ver. 2009/6/01 Page 29 of 31

TechEdSat—7 Orbital Debris Assessment Report (ODAR) TZIMP—06— XS0O01 Rev 2 Appendix C: Wiring Schematics TechEdSat—7 Wiring Diagram Iridium + GPS 1 Crayfish Dual Antenna Solar Panel 7@ C | Side 1 Cricker * Solar ricke Power Pame Side 2 . e Board | 2X BP—930 | Ali All t Solar Batteries | Switch Switch Panel _ iufusttsescccucccecdl Inhibit Inhibit Side 3 Ali __—— RBF Switch Inhibit CUBIT IS,Z:E Inhibit l . Side 4 JST SH 2 Pin Connectors are used for Eigfiila connections on the Power Board. MCX Connections are used for the GPS antenna, while MMCX Connections are used for the Iridium and U.FL Connections are used for the CUBIT. Wire—to—wire connections use Molex connectors. Page 30 of 31

TechEdSat—7 TZMP—06— Orbital Debris Assessment Report (ODAR) XS001 Rev 2 Second Cricket and CUBIT Wiring Diagram F—é Inhibit Group — gooemiommmmnennonmmmsnsene | Coin Cell | (CR2032) _Era;F;h bemmmenes -)))‘ Board CUBIT [ Ground\ * )))Statlon Page 31 of 31


Document Created: 2018-06-21 09:19:29
Document Modified: 2018-06-21 09:19:29
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