ODAR

0742-EX-CN-2017 Text Documents

Montana State University

2017-11-21ELS_201436

ELVL-2016-0044542 August 23, 17 Orbital Debris Assessment for The CubeSats on the CRS 0A-9/ELaNa-23 Mission per NASA-STD 8719.14A

Signature Page , Analyst, a.i. solutions, AIS2 Scott Higginbotham, MV5ion Manager, NASA KSC VA-C 2

National Aeronautics and Space Administration John F. Kennedy Space Center, Florida Kennedy Space Center, FL 32899 ELVL-XXX-XXXX Reply to Attn of: VA-H1 August 23, 17 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-23 Mission (DRAFT) 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 E. UL Standard for Safety for Lithium Batteries, UL 1642. UL Standard. 4th ed. Northbrook, IL, Underwriters Laboratories, 2007 F. Kwas, Robert. Thermal Analysis of ELaNa-4 CubeSat Batteries, ELVL-2012- 0043254; Nov 2012 G. Range Safety User Requirements Manual Volume 3- Launch Vehicles, Payloads, and Ground Support Systems Requirements, AFSCM 91-710 V3. H. HQ OSMA Policy Memo/Email to 8719.14: CubeSat Battery Non-Passivation, Suzanne Aleman to Justin Treptow, 10, March 2014 I. HQ OSMA Email:6U CubeSat Battery Non Passivation Suzanne Aleman to Justin Treptow, 8 August 2017 The intent of this report is to satisfy the orbital debris requirements listed in ref. (a) for the ELaNa-23 auxiliary mission launching on the CRS OA-9 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. 3

The following table summarizes the compliance status of the ELaNa-23 payload mission to be flown using a to-be-determined CRS vehicle. The 8 CubeSats comprising the ELaNa-23 mission are fully compliant with all applicable requirements. Table 1: Orbital Debris Requirement Compliance Matrix Requirement Compliance Assessment Comments 4.3-la Not applicable No planned debris release 4.3-lb 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 4.4-3 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 4.3 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-23 mission 4

Section 1: Program Management and Mission Overview The ELaNa-23 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: CaNOP: Kevin Crosby, Principal Investigator, Carthage College CubeRRT: Joel T. Johnson, Principal Investigator, The Ohio State University EQUISat: Rick Fleeter, Advisor, Brown University MemSat: Sangho Shin, Principal Investigator, Rowan University RadSat: Brock LaMeres, Technical POC, Montana State University RainCube: Eva Peral, Principal Investigator, The Jet Propulsion Laboratory SORTIE: Geoffrey Crowley, Principal Investigator, University of Colorado Tempest-D: Steven Reising, Principal Investigator, Colorado State University 5

Program Milestone Schedule Task Date CubeSat Selection August 16, 2016 CubeSat Delivery to NanoRacks February 1, 2018 Launch March14, 2018 (U/R) Figure 1: Program Milestone Schedule The ELaNa-23 CubeSat complement will be launched as payloads on an upcoming CRS launch vehicle to the International Space Station. The ELaNa-23 mission will deploy 8 pico-satellites (orCubeSats) from the International Space Station, using the NanoRacks CubeSat dispenser. Each CubeSat is identified in Table 2: ELaNa-23 CubeSalsTable4.: ELaNa-2-3-Gu_.,-.,-Siim . The ELaNa-23 manifest includes: CaNOP, CubeRRT, EQUISat, MemSat, RadSat, RainCube, SORTIE, and Tempest-D. The current launch date is projected to be no earlier than April 30, 2018. The CubeSats on this mission range in size from a 10 cm cube to 24 cm x 36 cm x 10 cm, with masses from about .71 kg to 11.4 kg, with a total mass of roughly 49 kg being manifested on this mission. The CubeSats have been designed and universities and government agencies and each have their own mission goals. 6

Section 2: Spacecraft Description There are 8 CubeSats flying on the ELaNa-23 Mission. Table 2: ELaNa-23 CtibeSatsTabk,---2+-.gLiOhi-17-GF:.:°-gats, outlines their generic attributes. Table 2: ELaNa-23 CubeSats CubeSat CubeSat CubeSat CubeSat size (mm3) Masses Names Quantity (kg) CaNOP 1 3U (340 x 100 x 100) 2.75 CubeRRT 1 6U (365 x 226.3 x 100) 114 EQUIS at 1 1U(100 x 113 x100) 1.20 MemSat 1 1U (100 x 113 x 100) .71 RadSat 1 3U (98.5 x 326.5 x 98.5) 1.6 RainCube 1 6U (116 x 366 x 239) 11.9 SORTIE 1 6U (239.2 x 365 x 115.4) 7.94 Tempest-D 1 6U (226.3 x 365 x 100) 11.09 The following subsections contain descriptions of these 8 CubeSats. 7

CaNOP — Carthage College — 3U Solar Panels Comm. Board ....r .../ Bat: plli s / Camera .0/ Microcontroller Skeleton GPS Receiver "........N.N... "■................. Camera il t 1 In I ;lasecamera ■•■,............ OBC vt ADCS 's ir S-Band / Antenna Antenna Hatch LAIIIIIIL Solar Panels Figure 1: CaNOP Expanded View. Overview The CaNOP CubeSat mission objective is to provide an educational experience in aerospace engineering to undergraduate students through designing, building, testing, and operating a 3U CubeSat designed to replicate early Landsat remote sensing capabilities. To achieve this objective, a multispectral linescan imaging sensor and associated optics capable of resolving a ground sample distance of 60 m and a scene size of 120 km will be utilized. CONOPS Upon deployment from the ISS, CaNOP will power up and attempt a status test of internal systems. After 45 minutes, the ADCS proceeds to de-tumble CaNOP . When CaNOP is approaching an imaging target, the camera will be activated. The camera will acquire spectrally resolved images along the orbital path. When CaNOP CubeSat is near a Near Earth Network (NEN) ground station with S-band capabilities, the communication system will be activated and begin transmitting handshake signals until downlink is established. Once downlink is established, compressed image data and telemetry will be broadcast to the closest ground station within range. When CaNOP is low on power the EPS will stop powering all unnecessary components and allow CaNOP to orbit until the 30Whr battery has been sufficiently recharged. 8

Materials The CaNOP CubeSat components are commercial-off-the-shelf (COTS) materials, electrical components, and solar cells. The primary provider for COTS technology used in CaNOP is ClydeSpace, Inc.. The primary structural component is the 3U frame provided by ClydeSpace, Inc. The ClydeSpace 3U structure is machined from anodized 6061 aluminum. Ii.z zards There are no pressure vessels, hazardous or exotic materials. Power. Systeni/Batteries The electrical power storage system consists of common lithium ion polymer (LiPo) cell batteries with Overcharge, Over-discharge, Overcurrent, Overvoltage, and Under- temperature protection. The LiPo cells are qualified to NASA standards EP-Wi-032. 9

CubeRRT - The Ohio State University/Goddard Spaceflight Center/The Jet Propulsion Laboratory/Blue Canyon Technologies - 6U CubeSat Solar Arrays Instrument Electronics `may, Avionics Module Machined Aluminum Housing HDRM and UHF antenna Antenna Assembly Figure 2: CubeRRT Deployed View/CubeRRT Electronics Overview The CubeSat Radiometer Radio. Frequency Interference Technology Validation (CubeRRT) mission will observe, detect, and mitigate radio frequency interference (RFI) for microwave radiometers. RFI is a growing concern for Earth science observations due to its negative impact on science measurements with microwave radiometers. CubeRRT is being developed through collaboration between OSU, NASA's Goddard Space Flight Center, and NASA Jet Propulsion Laboratory (JPL). Goddard and JPL will build the front-end instrument and the backend electronics for the spacecraft, respectively, while OSU will implement CubeRRT's antenna subsystem and manage the project. Blue Canyon Technologies (BCT) will integrate the CubeRRT payload with the 6U spacecraft bus and perform environmental testing of the complete spacecraft. The spacecraft will be operated from BCT's Mission Operations Center located in Boulder, Colorado. CONOPS Once CubeRRT is ejected from the dispenser and separation is confirmed, the CubeSat begins an autonomous initialization sequence which begins with the CubeSat entering a sun pointing mode and begins recording vehicle state of health (SOH) data. The spacecraft's guidance then puts the vehicle into a Sun Safe attitude. 30 minutes after separation, the solar arrays and UHF antenna are deployed, and the mirror is unlocked. At 45 minutes after separation, the spacecraft begins beaconing of SOH data, using the GlobalStar system. At 55 minutes after separation, GPS beings cycling in order to 10

determine a position solution. Once the ground analyzes the data and finds it satisfactory, calibration of CubeRRT's instruments begins. Once nominal subsystem performance is confirmed, the ground will command deployment of the payload antenna. Materials The majority of the bus is machined 6061-T6 aluminum structure, with carbon-fiber composite solar array panels, fused silica star tracker optics, and various commercial electronics parts. Hazards There are no pressure vessels, hazardous or exotic materials. Batteries The power system consists of a peak power tracker that charges 6x lithium-ion batteries (2 strings of 3 cells). The battery design incorporates a cell balancing circuit and protection for over-charge, over-discharge and short-circuits. The UL-listing number is BBCV2.MH19896. 11

EquiSat — Brown University — 2U Solar Panels Vi 4.11 14. Main Control Board A s,f it Chassis Li-ion Batteries Radio Solar Panels 4mw Antenna LIFePO Batteries Flash Panel Figure 3: EquiSat Expanded View Overview EquiSat will be visible to observers on the ground, within earshot of amateur radio users, and will be testing experimental lithium iron phosphate battery technology. The assembly will prove the feasibility of low cost spaceflight through the use of student designed and manufactured components.. CONOPS Upon deployment from the ISS, EquiSat will power up and start counting down timers. At 45 minutes, the antenna will be deployed and our radio will start transmitting. Once system checkout has been complete, the spacecraft will progress to its primary operating mode and begin payload tests. Operations will continue until loss of contact with spacecraft. :Materials The CubeS at structure is milled out of a single block of 6061 Aluminum. It contains standard off the shelf materials, electrical components, PCBs and solar cells. I Lizards 12

There are no hazardous systems on board. There are no pressure vessels nor thrusters nor any chemical reactants. Batteries The electrical power storage system consists of 2 lithium-ion (primary) and 4 lithium iron phosphate batteries (payload) with over-charge/current protection circuitry. The payload batteries carry the UL-listing number MH61931 and the primary batteries are of type 18650. 13

MemSat —Rowan University — 1U Figure 4: MemSat Expanded View Overview MemSat's mission is to evaluate the behavior of novel memristive memory elements against standard, silicon-based memory technologies to determine potential advantages— and/or disadvantages—of memristors for space applications. The key figure of merit to be quantified between technologies for this mission experiment is comparison of single event upset (SEU) rate. CONOPS Once MemSat has been deployed and the mandated 45 minute radio silence period has passed, the antenna will be deployed and begin to run satellite diagnostics. Once the satellite is operational, it will attempt to communicate with its ground station. After a preliminary status update about the satellite has been sent, the MemSat will begin the Minimum Viable Experiment. The primary experiment will run for 30 days, with upkeep and diagnostics running in the background. After these 30 days, the mission operations team will begin to upload new experimental profiles to the MemSat to be performed. Material The CubeSat structure is a Pumpkin 1U Structure made from Al 6061 Stainless Steel. It contains all standard commercial off the shelf (COTS) materials, electrical components, PCBs and solar cells. Attitude stabilization will be obtained utilizing neodymium magnets. 14

Hazards There are no pressure vessels, hazardous, or exotic materials. Power Systems/Hazards The electrical power storage system consists of common lithium-ion batteries with over- charge/current protection circuitry. The lithium batteries carry the UL-listing number ICR18650. 15

RadSat-g — Montana State University — 3U Structure Side Panels I Side Panel Antenna I I I dat Li-ion Battery Pack Side Panel I •••• .••• Solid State Radiation Sensor •-• 00 0. Data Logging Board ." -,'r • FPGA Board • • Payload Power Board • Passive Magnetic Stabilization Board Payload Interface Board Radio Board Power Board CLEM Board Figure 5: RadSat-g Expanded View O erview RadSat-g will demonstrate a radiation tolerant computer technology in Low Earth Orbit. The technology is based on commercial off the shelf (COTS), Field Programmable Gate Arrays. RadSat-g deploys a novel single event effect (SEE) mitigation strategy based on spatial avoidance of faults using an array of redundant processors with selective activation and background repair through partial reconfiguration. ...7 0NOPS Upon deployment from the ISS, RadSat-g will power up and start a timer system. At 45 minutes RadSat-g will turn on and the flight computer will engage the antenna deployment sequence. The system will then begin UHF beacon transmissions at a 30 second cadence. For the first week of operation the ground station operators will attempt communications to perform checkouts of the spacecraft. Approximately 7 days from launch, data from the computer technology will begin to be downloaded and will continue for approximately 1 year. 16

Materials The RadSat-g structural components are made of the following aluminum alloys: 5052- H32 and 7075-T6. It contains all standard commercial off the shelf (COTS) materials, electrical components, PCBs and solar cells. The UHF/UHF antennas are made of spring steel. Hazards There are no pressure vessels, hazardous or exotic materials. Batteries The electrical power storage system consists of common lithium-ion batteries with over- charge/current protection circuitry. The lithium batteries carry the UL-listing number MH48285. 17

RainCube — The Jet Propulsion Laboratory - 6U Deploy Solar Array UHF Transceiver Battery Modules IRM S-Band TX GPS Antenna RainCube Payload Ka-Antenna Canister Ka-Antenna Figure 6: RainCube View Overview The RainCube (Radar in a Cubesat) mission is a technology demonstration mission to enable Ka-band precipitation radar technologies on a low-cost, quick turnaround platform. RainCube is demonstrating two new technologies: a novel radar architecture called miniKaAR-C (miniature Ka-band Atmospheric Radar for CubeSats) and an ultra compact deployable Ka-band antenna called KaRPDA (Ka-band Radar Parabolic Deployable Antenna). The mission utilizes Tyvak's Endeavor suite architecture for power generation and management, telemetry and commanding (TC), Command and Data Handling (CDH), thermal management, and Guidance Navigation and control (GNC). The Tyvak system utilizes two lithium ion battery modules and deployable solar arrays for power generation and management. The TC consists of a UHF transceiver for bus telemetry downlinks and command uplinks; as well as an S-Band transmitter for payload data downlink The CDH and GNC are packaged into the Inertial Reference Module (IRM) which consists of a CDH & GNC processor, an IMU, three magnetorquers, and three reaction wheel assemblies, and two star trackers. The payload, comprised of miniKaAR-C and KaRPDA, is designed and manufactured by JPL. CONOPS The RainCube vehicle powers on following release from its deployer through the release of three independent enable switches (two high-side, one low-side), and begins operations to de-tumble the vehicle and enter a sun-pointing orientation. Following a 45 minute timer expiration, the UHF antennas are automatically deployed and beacons occur over designated ground stations via UHF. Approximately 90 minutes after deployment, the solar arrays are deployed and the vehicle begins bus commissioning operations, which are anticipated to take several days. Following bus commissioning, the Ka-band antenna is deployed and the vehicle begins payload commissioning operations, which are 18

anticipated to take several days as well. Following payload commissioning, the vehicle being nominal operations to collect science data. Nominal operations will switch between Ka-band measurements with the antenna boresight facing Nadir and sunpointing with the deploy arrays normal to the sun vector. Materials The primary structure for RainCube is composed of Aluminum 7075-T7. The spacecraft is largely composed of components manufactured by Tyvak, which consist of electrical components, PCBs or FR4, and solar cells. Both the S-Band transmit antenna and the GPS receive antenna are ceramic patches. Hazards There are no pressure vessels, hazardous or exotic materials. Batteries The electrical power storage system consists of common lithium-ion batteries with over- charge/current protection circuitry. The lithium batteries used are LG 18650 and have passed IEC/UN38.3. The theoretically peak energy storage is 120Whrs, though circuits limit the peak charge to approximately 109.2 Whrs. 19

SORTIE — The Astra Corporation — 6U rnIVNI Payload Antenna Deployment CSD SEP Connector System Right SEP Switch Sun Sensor 1 ode, Right Solar Panel state Electrical Power Battery Pack I System --- ARM Cortex M4F Payload Controller Left SEP Switch Left Solar Panel Routing Beard MSP4SS CADH L3 Cadet Radio Moflwbaard CSC. Rail oPLP Payload BCT %ACT ADACS — Battery Pack 2 Battery Pack 3 ID, Charge. ARM/Safe Connectors Figure 8: SORTIE Expanded/Deployed View Over iew Scintillation Observations and Response of The Ionosphere to Electrodynamics (SORTIE)'s mission is to collect data over the course of 6 months, which will allow scientists to describe the distribution of wave-like structures in the plasma density of the ionospheric F-region and to connect these variations to wave sources in the troposphere and in the high latitude thermosphere. SORTIE is a 6U CubeSat that will carry two instruments for determining the structure of equatorial plasma: a miniature Ion Velocity Meter (mini-IVM) built by The University of Texas at Dallas, and a micro Planar Langmuir Probe (tPLP) built by the Air Force Reseirch Laboratory (AFRL) in Albuquerque. CONOPS After deployment from the ISS, SORTIE will turn on in a low-power mode. This mode will perform initial system tests, battery tests and start a timer for 45min. After timeout the antenna will be deployed, and SORTIE will wait to hear from the ground station. Upon initial contact with the ground station, SORTIE will download system health telemetry, and payload testing telemetry. This will continue through subsequent orbits until operations team send command to change mode to normal operations. It is expected that a move to normal operations will take more than a week to insure the spacecraft is functioning as expected. After entering normal operations mode, the spacecraft will collect and download payload and system telemetry until the end of life. Materials 20

The SORTIE structure is made of Aluminum 6061-T6. It contains all standard commercial off the shelf (COTS) materials, electrical components, PCBs and GaAs solar cells. Hazards There are no pressure vessels, hazardous or exotic materials. Batteries The batteries chosen are lithium-ion from Panasonic ACA4000PE3. They have acquired UL1642. The power system has current protection and overcharge protection. TEMPEST-D — Colorado State University/The Jet Propulsion Lab— 1U 21

Machined Aluminum Housing Solar Arrays Machined Globalstar Aluminum Antenna Side Wall Instrument Coarse Sun Sensor PSC Tabs or NLAS Corners Avionics Module UHF Interconnects HDRM Figure 9: TEMPEST-D Expanded View Overview The objective of the TEMPEST-D mission is to reduce the risk, cost and development time for a future TEMPEST mission. A future five-satellite TEMPEST mission would provide the first temporal observations of cloud and precipitation processes on a global scale. These observations are important to understanding the linkages in and between Earth's water and energy balance and to the understanding of cloud model microphysical processes that are vital to climate change prediction. This understanding will also have significant impact on agricultural forecasting, forest management and disaster preparedness. Blue Canyon Technologies (BCT) will deliver the 6U spacecraft to the project, led by Colorado State University (CSU). TEMPEST-D is supported by NASA's Science Mission Directorate, Earth Science Division and is managed by NASA's Earth Science Technology Office. The NASA/CalTech Jet Propulsion Laboratory (JPL) will provide the five-channel millimeter-wave radiometer instrument. CONOPS After deployment, they rapidly boot-up and image the launch vehicle. Out of view of the upper stage, each will characterize the other's tumble using the cameras. At 30 mins, the cameras will be used to observe each PIC's own antenna deployment. At 45 mins, the system will downlink primary mission data, including telemetry used for evaluating system performance. Materials 22

The majority of the bus is machined 6061-T6 aluminum structure, with carbon-fiber composite solar array panels, titanium thermal bench standoffs, fused silica star tracker optics, and various commercial electronics parts. Hazards There are no pressure vessels, hazardous or exotic materials. Batteries The power system consists of a peak power tracker that charges 6x lithium-ion batteries (2 strings of 3 cells). The battery design incorporates a cell balancing circuit and protection for over-charge, over-discharge and short-circuits. The UL-listing number is BBCV2.MH19896. 23

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-23 CubeSat mission therefore this section is not applicable. 24

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-23 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 CubeSats 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 4.3 years maximum, the ELaNa-23 CubeSats are compliant. 25

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 8 CubeSats is that of the Raincube CubeSat with solar arrays and Ka antenna deployed. Deploy Solar Array UHF Transceiver Battery Modules IRM S-Band TX GPS Antenna RainCube Payload Ka-Antenna Canister Ka-Antenna Figure 6: Raincube Deployed Configuration E Surface Area [2 * (w * + 4 * (w * Mean CSA = 4 4 Equation 1: Mean Cross Sectional Area for Convex Objects l + A1) Mean CSA = (Amax + A 2 Equation 2: Mean Cross Sectional Area for Complex Objects All CubeSats evaluated for this ODAR are stowed in a convex configuration, indicating there are no elements of the CubeSats obscuring another element of the same CubeSats from view. Thus, mean CSA for all stowed CubeSats was calculated using Eiwtion 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, : is utilized to determine the mean CSA. Amax is identified as the view that yields the maximum cross-sectional area. Ai and A2 are the two cross-sectional areas orthogonal to Amax. Refer to Appendix A for component dimensions used in these calculations The RainCube (11.9 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 26

0.0041m2/kg, DAS yields 3.9 years for orbit lifetime for its stowed state, which in turn is used to obtain the collision probability. Even with the variation in CubeS at design and orbital lifetime ELaNa-23 CubeSats see an average of 0.0 probability of collision. All CubeSats on ELaNa-23 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. 27

0 fD -o m ep rr C C ta 3 a) DJ fD I- NJ N N.) O O O O NJ O (."" O O al O O O O O O O fD 03 O N O O 3 cn cr, fD Stowed O O O 00 O NJ (JJ (sJA) aLugam !elm° O O !AI O O O O O O O O O O O O O O O O Probabilityofcollision(10^X) O O O O O O O O O O MeanC/SArea(m^2) 0.040 -0 cu Area-toMass(m^2/kg) 0.003 OrbitalLifetime(yrs) 4.3 o Probabilityof collision(10^X) 0.00000 pa4 e0 awlxrqd Jelos seam no paveig :aloN** Table 3: CubeSat Orbital Lifetime &Collision Probability

▪ Deployed Stowed a) 03 4 3 C Table 3: CubeSatOrbitalLifetime & CollisionProbability(cont.) O a) A U) CD sse —s ag 3 3 w) NJ 3 NJ O O 0 0 O b O 0 0 W o b O W 0 0 -I=. O 0 03 ui .P. CO (A) O CO 0 O O O O O O (JJ O O O O O O O O G-1S3d1A131 O O O O O . b ix) O • O O O • (..0 00 O O 1-k u-1 O 0 .11b UJ I—h O 0 O Cn 0

The probability of any ELaNa-23 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. Since the CubeSats have no capability or plan for end-of-mission disposal, requirement 4.5:2 is not applicable. Assessment of spacecraft compliance with Requirements 4.5-1 shows ELaNa-23 to be compliant. Requirement 4.5-2 is not applicable to this mission. Section 6: Assessment of Spacecraft Postmission Disposal Plans and Procedures All ELaNa-23 spacecraft will naturally decay from orbit within 25 years after end of the mission, satisfying requirement 4.6-la 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 RainCube in its stowed configuration as the worst case. The area-to-mass is calculated for is as follows: Mean %Area (m2) m2 = Area — to — Mass (— ) Mass (kg) kg Equation 3: Area to Mass 0.04088 m2 m2 11.5 k g — 0.00355 g — k Tempest 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 0.0041 m2/kg for the RainCube CubeSat was used. DAS 2.1.1 yields a 4.1 years orbit lifetime for RainCube in its deployed state. This meets requirement 4.6-1. For the complete list of CubeSat orbital lifetimes reference Table 3: CubeSat Orbital :Li l'etitne & Collision ProhabilityTable-3: Cubt.)&144)-thital


Section 7: Assessment of Spacecraft Reentry Hazards A detailed assessment of the components to be flown on ELaNa-23 was performed. 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. 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-23 High Melting Temperature Material Analysis Demise Kenetic CubeSat Nam Material Mass (kg) Alt (km) Energy (.1) ADCS momentum Callop Tungsten 0.09 0 55 wheel ADCS magnetic Callop Iron 0.018 0 3 torquers CubeRRT Hinge retainer clipl Titanium (generic) 0.002 0 0 CubeRRT Hinge retainer clip 2 Titanium (generic) 0.0015 0 0 CubeRRT Hinge retainer clip 3 Titanium (generic) 0.0014 0 0 CubeRRT Hinge retainer clip 4 Titanium (generic) 0.0025 0 0 Stainless Steel CubeRRT Screws 0.0006 76.5 0 (generic) Stainless Steel CubeRRT Washers 0.0009 78.0 0 (generic) EQUISat Hysteresis rods Iron 0.0004 0 0 Stainless Steel EQUISat 3/8 screws 0.0058 0 0 (generic) Stainless Steel EQUISat Nuts 0.0228 78 0 (generic) Stainless Steel EQUISat .6 studs .0085 0 0 (generic) Stainless Steel EQUISat 1.0 studs .147 0 0 (generic) 32

Stainless Steel EQUISat .1875 slotted screws 004 78.0 0 (generic) Stainless Steel EQUISat 1,75 slotted screws .004 0 0 (generic) Stainless Steel EQUISat 2-56 nuts .004 0 0 (generic) Stainless Steel EQUISat i6 screws .005 0 0 (generic) Stainless Steel EQUISat '/4 screws .006 0 0 (generic) MemSat Antennae Steel (AISI 410) .066 0 1 Stainless Steel MemSat Separation springs .031 77.8 0 (generic) MemSat Magnets Neodymium .0035 78.0 0 RadSat Antenna Steel (AISI 410) .008 0 0 RadSat Magnet Neodymium .005 77.8 0 Stainless Steel RainCube Ka band antenna 1.1 0 0 (generic) RainCube fasteners Steel (AISI 316) .0005 76.5 • 0 SORTIE Antennae Steel (AISI 410) .150 0 85 SORTIE Separation switches Steel (AISI 316) .0019 0 0 SORTIE Fasteners Steel A-286 .0006 77.2 0 Tempest-D Hinge retainer clip 1 Titanium (generic) .002 0 0 Tem pest-D Hinge retainer clip 2 Titanium (generic) .0015 0 1 Tempest-D Hinge retainer clip 3 Titanium (generic) .0014 0 0 Tempest-D Hinge retainer clip 4 Titanium (generic) .0025 0 0 Tempest-D RHS structural standoff Titanium (generic) .397 0 157 Motor mount Tempest-D Titanium {generic) .339 0 275 structural standof 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: Table 5: Requirement 4.7-1 Compliance by CubeSat Risk of Human Name Status Casualty Callop Compliant 1:145200 CubeRRT Compliant 1:0 EQUISat Compliant 1:0 MemSat Compliant 1:0 RadSat Compliant 1:0 RainCube Compliant 1:0 SORTIE Compliant 1:136600 Tempest-D Compliant 1:60600 *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 33

not included in the Debris Casualty Area that inputs into the Probability of Human Casualty calculation. This is why CubeSats that have surviving components like CubeRRT, EQUISat, and RainCube have a 1:0 probability as none of their components have more than 15J of energy. TEMPEST-D has 2 components with greater than 15J of energy but the CubeSat's probability of risk of human casualty still exceeds the NASA requirement on an order of magnitude. This issue is the same in the case of SORTIE, and Callop. All CubeSats launching under the ELaNa-23 mission are shown to be in compliance with Requirement 4.7-1 of NASA-STD-8719.14A. 34

Section 8: Assessment for Tether Missions ELaNa-23 CubeSats will not be deploying any tethers. ELaNa-23 CubeSats satisfy Section 8's requirement 4.8-1. 35

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/KS C/AIS 2 cc: VA-H/Mr. Carney VA-H1/Mr. Beaver VA-H1/Mr. Haddox VA-C/Mr. Higginbotham VA-G2/Mr. Treptow SA-D2/Mr. Frattin SA-D2/Mr. Hale SA-D2/Mr. Henry Analex-3/Mr. Davis Analex-22/Ms. Ramos 36

Appendix Index: Appendix A. ELaNa-23 Component List by CubeSat: CaNOP Appendix B. ELaNa-23 Component List by CubeSat: CubeRRT Appendix C. ELaNa-23 Component List by CubeSat: EQUISat Appendix D. ELaNa-23 Component List by CubeSat: MemSat Appendix E. ELaNa-23 Component List by CubeSat: RadSat Appendix F. ELaNa-23 Component List by CubeSat: RainCube Appendix G. Appendix H. ELaNa-23 Component List by CubeSat: SORTIE Appendix I. ELaNa-23 Component List by CubeSat: Tempest-D 37

I C C C C C el c (.1 nnne-)nr.) CaNOP CaNOP CaNOP CaNOP R CaNOP 1 % aNO z % % % tx aNOP 0 .0 aNOP o -0 o -0 .0 o 0 .0 P 0 .0 aNOP 0 .0 aNOP 000000 o % o .0 .0 .0 .0 o o rr. -0 P-3 s•-.1 C1 C-", ■ 01 LA -ID. IAD N ..` 81 0 to --.1 23 B N.) ■ 1-., Ir r r 13 r r r N) t. .) .' N 0 ,--• ,10 41. -• 0 ELaNa-23 Component List by CubeSat: CaNOP ADCS Momentum Wheel ADCS Magnetictorquers ADCS IREHS blocks Temperature Sensors CaNOP CubeSat3U Seperation Switches CubeSat Structure SDR Transceiver 30Whr Batteries S-Band Antenna SDR Controller 2U Solar Panel 3U Solar Panel GPSReceiver ADCSBoard GPS antenna SunSensors SDR Board EPSboard c.) tii r4z rotors OBC -0 Q7 g ir.3 ID A - - ® . - - - - , - - - 4 N - - - ® - - C PCB, lithium ion Aluminum 6061 TungstenNickel alloy MT-17C Al 606 1-T651 Ferousalloy cy, 04 polymer Al6082 .-, z z o HDd PCB PCB PCB PCB PCB PCB PCB PCB PCB N/D N/D 10:el1 ..... 0 . t.' g' n. --; cA to -(-) co HollowBox 0: rr & Box .S,E. Ino Board Board Board Board Board Board Board Board o Z o to Panel OtO <,4 N/D N/D Box Box Box ci, o.. w a g o P-.3 ...t DDn A vrt .p. -. 00 p ,c, . . 1 VD p f-0 N • • 1 '--. 8 'E--. '8 --. u..., 0 a% . so so w so '": .. i ..,,a,. - ...-.,.-J . A :_,\ \ -.i ...■ \ . . . ,--. b 1 • •4 N LA so 4, ..... o os p -a Z a Z- a Z ..8 _ -ID. '01-, 10th VD ' . 1 t. .) SC, Diameter / Width 0-0• C>-ID. 0..- 43.25 95.89 imni) 01 N/D oo 1..4 t...) (A) 200 VD CA) ,p LA Z Z ..) 46 65 20 0 LA i7, 00 bc ‘.0 CrN 0 a, b " cA b. VD a a ■ •••■ ,-. U.) ,__, e■•■ ts..) 'CI- - -{D, ..••, •C,_,.• •,._,•01 1.-■00 CD0 l 20.65 L'' ll,,. . .) 0 01 LADso LA(. D • 0° it.,. /1■ ,Z;J 4 N/D 6.5 1.6 20 r 00 r 00 ,,, (,..) ,) Z Z 14 Cr, ..- a g, ... o ,.o ,...) ..-... 90.17 31.75 C., .....4, p ,c. t-,,) 4, Z I.) -4 1- "" N/D oo 4D p b %.0 N/D 100 100 VD 6 a, a, 40 70 ;,,, 1 --3 z.-1 c, ,-, -.1 t.. c> on s 8 ii 2z z g z z z z z z z z E z z z z z z z z z 2 z z ..e v. 1..) c000 e ■ 1 t 1 s 1 ■ 17)cp,_ , , 1 1 1 . 11 1 1 1 1 1 1 54 . ,-3 E 0 CD0 W Demise Demise Demise Demise Demise Demise Demise Demise Demise Demise Demise Demise Demise Demise Demise Demise Demise Demise Demise 0km 0km rv. 0o R. n Ir. =

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CUSESAT Tempest-D Tempest-D Tempest-D Tempest-D Tempest-D Tempest-D Tempest-D Tempest-D Tempest-D Tempest-D Tempest-D Tempest-D Tempest-D Tempest-D Tempest-D CD CD CD CD CD CD CD 5 3 5 5 5 5 5 5 .5 N us D sn TS CD us ,--, TS CD co, N=I CD / qa ‘4' 'r 6 6 6 6 6 6 6 6 Number I n t...) ,..., — — — — — 23 22 40. LA) t...) 1' 14 12 15 10 1•-• 0 SO CO ---.1 al L4 P■1 ‘.0 00 --41 ON LA j ELaNa-23 ComponentListbyCubeSat: Tempest-D Mechanism, Release, Linear UHF Monopole Antenna Hinge,Retainer Clip,-4 Hinge,Retainer Clip,-3 Hinge, Retainer Clip,-2 Hinge,Retainer Clip,-1 Solar Array,Horizontal Antenna, Globalstar Solar Array,-Canted c) XB1, Avionics z Panel, Access CSS, Bathtub CSS, Bathtub Rail, PSC,-3 Rail, PSC,-2 Col Harnessing 5 Side Panel Structure Chassis C.) .54 cc/a Z g tin + 1-o X in CD C.) Z t*".1 g. ,-.. • ,-, 1-, ■• ..... ,--. ■-■ Ss) 41. J.-, D-■ DJ NJ o-, •-■ or o-. t...) I..., 1 I-, i--, 1-, I-■ SO q 1 Aluminum 6061 Titani um 6AL-4V Titanium 6AL-4V Titanium 6AL-4V Titanium 6AL-4V Aluminum 7075 Aluminum 6061 Aluminum6061 Aluminum 6061 Aluminum7075 Aluminum 7075 Aluminum 6061 Various Various Various Various Various Various < < < < < . . . . .1-7. 0 0 0 0 0 0 0 0 0 0 4, cc Vi 4, (0 E. 0:i :0 CO ‘• 0XtO 0 Box Box xog Box Box Box Box Box Box Box Box CO 0:J 00 x x 0 x 0 X X 0X s. p, 0X 0X 0X 0X •/-:„%.• CD M ■ -■ 4D.. ts.) LA '1. t.11 La VI isj C.--i- Vsi N ,-. t. N3 ,, t....., ,c° o ;_,. 75 c, N L.r. '-' .4. ."- (..n N co 9° vo '') . LA 90 LA Ln . ■ -■ c ,22 22 •...2 VD C\ a . cr, co t, cs, 9" CP, , t.,, ,_, 0 ,? . '-- c., 4, ,... ° IN4 i•.4 IP ■0 4::, Li' • P 41' ■ . 0 •= L — — 0 —q — — 100.0 N u, ....., ..0 --.1 ,.. — -.2— 41 27.9 ...., 9.4 5.7 5.7 8.5 os. ,-. ,., ,_, CD - i 19 LA P :-.) b. LA ;:rs im • Lo . co .-, La •so • ‘c, • so gl ev Eh p b o g O'OL I 2263 245.0 NI, CAD l...) ts.) ^ L..) iIQ ■ -• 1... •-■ '-' Ls., ■-• s-. 1-• ; 10.2 10.2 9'6 1 ND „j a\ ,..,., • .....1 is) •%C:s • .4 P co ...1 00 L,,,, Vi ,(:. Vs .4. 4 0 o, L, LA '0 to to Lis ■ .o Ps.) PN 3 a ,r4 "--3 --a :-..] ....-1 0 ,.•J .... = o.is I 246.6 287.5 365.0 594.8 S...) -.--• mii 187.6 1,4 t.> VI LA Cr, 27.9 s GI LA N) •4:, 00 -• 1 00 144 ■1 :s3 C44 ta4 •4:7 0•-■ (44 C., L.' 2. 4, . L.4 0', LA * 70 O b -, .o . .. b b., bo La la U b b z... ' .1.- O o C 0 2 2 2', ,,-,D, 0 0 0 0 2 2 2, o o o o o o 0 0 ' 3 /z .t, o- Lw) 1.4 p--• ,--• ,..., .....3 2642° 1.,.) La 0 0 0 0, tsa N.) "1 LI1 VI Z2 2 gro ... -it -it -it 4t so cot a tv Demise Demise Demise Demise c, Demise Demise Demise Demise Demise Demise 0 C Y cY W W . 0 0km 0km 0km , CD CD ,.< ... ..G c,CD,L. . ,. 8 CD I 8 I I 8

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Document Created: 2017-08-23 14:18:36
Document Modified: 2017-08-23 14:18:36
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