ODAR RANGE RevC

0311-EX-PL-2016 Text Documents

Georgia Institute of Technology

2016-10-03ELS_182776

RANGE (Ranging and Nanosatellite Guidance Experiment) Orbital Debris Assessment Report (ODAR) Revision C 3 October 2016 Approved by: Dr. Brian Gunter (Principal Investigator) Prepared by: Michael Herman (Mission Design Lead) Daniel Groesbeck Byron Davis (Program Manager) Prepared in compliance with NASA-STD-8719.14 by the Georgia Institute of Technology. NASA Debris Analysis Software (DAS) version 2.02 was used in preparing this report.

Record of Revisions Revision Date Affected Pages Description of Authors Change C 3 October 2016 Section 5 Additional text Brian Gunter regarding collision risk/debris B 2 September 2016 Section 1 Updated schedule Brian Gunter Section 7 Updated DAS Daniel Groesbeck output A 12 February 2016 All Initial Revisions Michael Herman Daniel Groesbeck

Table of Contents RECORD OF REVISIONS   2   TABLE OF CONTENTS   3   SELF-ASSESSMENT AND OSMA ASSESSMENT OF THE ODAR   4   ASSESSMENT REPORT FORMAT   5   ODAR SECTION 1: PROGRAM MANAGEMENT AND MISSION OVERVIEW   5   MISSION DESCRIPTION   5   SECTION 2: SPACECRAFT DESCRIPTION   9   ODAR SECTION 3: ASSESSMENT OF SPACECRAFT DEBRIS RELEASED DURING NORMAL OPERATIONS   11   ODAR SECTION 4: ASSESSMENT OF SPACECRAFT INTENTIONAL BREAKUPS AND POTENTIAL FOR EXPLOSIONS   11   ODAR SECTION 5: ASSESSMENT OF SPACECRAFT POTENTIAL FOR ON-ORBIT COLLISIONS   15   ODAR SECTION 6: ASSESSMENT OF SPACECRAFT POSTMISSION DISPOSAL PLANS AND PROCEDURES   18   ODAR SECTION 7: ASSESSMENT OF SPACECRAFT REENTRY HAZARDS   20   ODAR SECTION 7A: ASSESSMENT OF SPACECRAFT HAZARDOUS MATERIALS   30   ODAR SECTION 8: ASSESSMENT FOR TETHER MISSIONS   30  

Self-Assessment and OSMA Assessment of the ODAR A self-assessment is provided below in accordance with the assessment format provided in Appendix A.2 of NASA-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: RANGE Mission Requirement Launch Spacecraft Comments # Vehicle Compliant Not Incomplete Standard Compliant Not Incomplete Compliant Non- Compliant Compliant 4.3-1.a X X No Debris Released in LEO. 4.3-1.b X X No Debris Released in LEO. 4.3-2 X X No Debris Released in LEO. 4.4-1 X X 4.4-2 X X 4.4-3 X X No Planned Breakups. 4.4-4 X X No Planned Breakups. 4.5-1 X X 4.5-2 X X 4.6-1(a) X X 4.6-1(b) X X 4.6-1(c) X X 4.6-2 X X 4.6-3 X X 4.6-4 X X 4.6-5 X X 4.7-1 X X 4.8-1 X X No Tethers Used.

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 RANGE satellite. Sections 9 through 14 apply to the launch vehicle ODAR and are omitted. ODAR Section 1: Program Management and Mission Overview Mission Description The RANGE mission will demonstrate improved absolute positioning of two 1.5U nanosatellites through a low-cost inter-satellite ranging instrument. The positioning will be validated through ground laser measurements. The satellites will be propulsion-less and rely on differential drag for control. The satellite will launch from Vandenburg AFB and deploy from the Minotaur-C launch vehicle. It will be inserted into an orbit at 500 km perigee and apogee altitude on an inclination from the equator at 97.4065 degrees. Atmospheric drag will slow the satellite and reduce the altitude of the orbit, until de- orbiting occurs approximately 7.84 years after launch and will conclude the mission. Launch vehicle and launch site: Minotaur-C from Vandenberg AFB Proposed launch date: January 2017 Mission duration: 1 year Launch and deployment profile, including all parking, transfer, and operational orbits with apogee, perigee, and inclination: The RANGE orbital elements are defined as follows: Apogee: 500 km Perigee: 500 km Inclination: 97.4065 deg RANGE has no propulsion and therefore does not actively change orbits. There is no parking or transfer orbit. At this time, we know of no potential interaction or physical interference between RANGE and any other operational spacecraft. Further analysis is planned on proving beyond reasonable doubt that ISS interference is avoidable through a probabilistic differential drag collision avoidance maneuver study.

Figure 1: RANGE 4-view for undeployed stowed stack configuration

Figure 2: RANGE 4-view separated undeployed configuration

Figure 3: RANGE 4-view separated deployed configuration Interaction or potential physical interference with other operational spacecraft: • At this time, we know of no potential interaction or physical interference between RANGE and any other operational spacecraft.  

Project Management: • Principal Investigator: Dr. Brian Gunter • Project Manager: Byron Davis Key engineering personnel: • ADCS Lead: Rohan Deshmukh • COMM Lead: Michael Lucchi • GNC Lead: Michael Herman • Laser Ranging Lead: Zachary Levine • OBC Lead: John Ridderhof • Structures Lead: William O'Donoghue and Ariana Keeling • Thermal Control System Lead: • EPS Lead: Austin Claybrook Foreign government or space agency participation: • No foreign agency is participating in this mission. All personnel are United States citizens. Summary of NASA’s responsibility under the governing agreement(s): • Not applicable. Schedule of mission design and development milestones from mission selection through proposed launch date, including spacecraft PDR and CDR (or equivalent) dates*: Date Milestone May 2016 Design and Fabrication July 2016 Integration August 2016 Testing December 2016 Shipment January 2017 Launch January 2017 Deployment and Operations Section 2: Spacecraft Description Physical description of the spacecraft: The RANGE satellites are two separated 1.5U nanosatellites with dimensions of 15 cm X 10 cm X 10 cm and a total mass of about 2.125 kg each with un-deployed solar arrays. The un-separated configuration has dimensions of 30 cm X 10 cm X 10 cm with un-deployed solar arrays and a total mass of 4.25 kg. The satellites have 1U symmetric single sided deployable solar arrays that increase the separated deployed configuration to 15 cm X 30 cm X 10 cm. Each RANGE satellite will contain the following systems: • GomSpace ANT430 antenna • GomSpace NanoPower p31us EPS board with Lithium Ion battery pack • GomSpace NanoCom AX100 UHF/VHF transceiver

• GomSpace NanoHub • GomSpace NanoMind 3200 onboard computer • Solar MEMS Nano SSOC D60 fine sun sensor • Four in-house built coarse sun sensors • Novatel OEM628 GPS receiver • ANTCOM GPS antenna • CubeSense single axis reaction wheel • In-house built three axis magnetorquer system • In-house built laser ranging system • Micrometer • ThorLabs laser diode • Princeton Lightwave Geiger-mode avalanche photodiode • Microsemi GPS-2750 10 MHz CSAC-based Disciplined Oscillator with QUANTUM SA.45s Chip Scale Atomic Clock • Pumpkin solar cells. Total satellite mass at launch, including all propellants and fluids: • 4.25 kg Dry mass of satellite at launch, excluding solid rocket motor propellants: • 4.25 kg Description of all propulsion systems (cold gas, mono-propellant, bi-propellant, electric, nuclear): • There will be no propulsion systems on RANGE. 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 as there will be no fluids or gasses on board. Fluids in Pressurized Batteries: • None. RANGE uses unpressurized standard COTS Lithium-Ion battery cells. Description of attitude control system and indication of the normal attitude of the spacecraft with respect to the velocity vector: • RANGE uses magnetorquer and reaction wheel control systems. The magnetorquers will provide three degree of freedom control with redundancy. The reaction wheel will only provide single y- axis control for differential drag ratio modifications. Description of any range safety or other pyrotechnic devices: • RANGE will use a burn wire for deployment of the single-sided single-deploy 1U solar arrays. The wire will burn at a resistor temperature of 200 C and a voltage between 12.5 and 16.8 V. Description of the electrical generation and storage system: • The power will be generated using solar panels and Lithium-Ion batteries. The batteries used will be the GomSpace NanoPower P31us with 6-8.4 V onboard lithium ion battery pack. The solar

panels will be Pumpkin single-sided single-deploy solar array configuration with two cells on the front body mounted face, two cells on the back body mounted face, and six cells on the deployable solar arrays. The body mounted side, top, and bottom faces will be clear of solar cells. Identification of any other sources of stored energy not noted above: • None. Identification of any radioactive materials on board: • None. ODAR Section 3: Assessment of Spacecraft Debris Released during Normal Operations Identification of any object (>1 mm) expected to be released from the spacecraft any time after launch, including object dimensions, mass, and material: • There are no intentional releases. Rationale/necessity for release of each object: • Not applicable. Time of release of each object, relative to launch time: • Not applicable. Release velocity of each object with respect to spacecraft: • Not applicable. Expected orbital parameters (apogee, perigee, and inclination) of each object after release: • Not applicable. Calculated orbital lifetime of each object, including time spent in Low Earth Orbit (LEO): • Not applicable. Assessment of spacecraft compliance with Requirements 4.3-1 and 4.3-2 (per DAS v2.0): • 4.3-1, Mission Related Debris Passing Through LEO: COMPLIANT • 4.3-2, Mission Related Debris Passing Near GEO: COMPLIANT ODAR Section 4: Assessment of Spacecraft Intentional Breakups and Potential for Explosions There are no intentional breakups scheduled during on orbit operation. We are aware of no known potential causes of spacecraft breakup during deployment and mission operations. 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. A controlled separation of the 3U nanosatellite package into two separate 1.5U nanosatellites will be executed following detumble. Summary of failure modes and effects analyses of all credible failure modes which may lead to an accidental 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 independent, mutually exclusive failure modes that could lead to a battery venting. If the LiIon batteries fail, they are expected to vent gas rather than explode. Detailed plan for any designed spacecraft breakup, including explosions and intentional collisions: • There are no planned intentional breakups by explosion, collision, nor by any other means. List of components which shall be passivated at End of Mission (EOM) including method of passivation and amount which cannot be passivated: • RANGE contains no components which are passivated at EOM. The satellite will breakup in atmospheric reentry. There is no plan to passivate the batteries, however in the case of mechanical damage or short-circuit they will not explode. Rationale for all items which are required to be passivated, but cannot be due to their design: • It was deemed unnecessary to passivate the lithium ion batteries for EOM, as the satellite will break 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: For each spacecraft and launch vehicle orbital stage employed for a mission, the program or project shall demonstrate, via failure mode and effects analyses or equivalent analyses, that the integrated probability of explosion for all credible failure modes of each spacecraft and launch vehicle is less than 0.001 (excluding small particle impacts) (Requirement 56449). Compliance statement: Required Probability: 0.001. Expected probability: 0.000. Supporting Rationale and FMEA details: Battery explosion: Effect:

• All failure modes below might result in battery explosion with the possibility of orbital debris generation. However, in the unlikely event that a battery cell does explosively rupture, the small size, mass, and potential energy, of these small 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. Probability: • Extremely Low. It is believed to be less than 0.01% given that multiple independent (not common mode) faults must occur for each failure mode to cause the ultimate effect (explosion). Failure mode 1: Internal short circuit. Mitigation 1: Qualification and acceptance 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. Combined faults required for realized failure: Environmental testing AND functional charge/discharge tests must both be ineffective in discovery of the failure mode. Failure Mode 2: Internal thermal rise due to high load discharge rate. Mitigation 2: Cells were tested in lab for high load discharge rates in a variety of flight like configurations to determine if the feasibility of an out of control thermal rise in the cell. Cells were also tested in a hot environment to test the upper limit of the cells capability. No failures were seen. Combined faults required for realized failure: Spacecraft thermal design must be incorrect AND external over current detection and disconnect function must fail to enable this failure mode. Failure Mode 3: Overcharging and excessive charge rate. 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. Combined faults required for realized failure: 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 1.5G conditions (in space, perpendicular to the sun) is 124 mA. The maximum charge rate our battery can accept is 3 A. The battery is a proto-qualified Molicell from the JSC ISS program, and has two 18650 cells. The battery itself has one string of 2 cells connected in series. 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 the single string from the battery was accepting charge. 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) 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). Combined faults required for realized failure: An external load must fail/short-circuit AND external over-current detection and disconnect function must all occur to enable this failure mode. Failure Mode 5: Inoperable vents. Mitigation 5: Battery vents are not inhibited by the battery holder design or the spacecraft. Combined effects required for realized failure: The manufacturer fails to install proper venting. Failure Mode 6: Crushing. Mitigation 6: This mode is negated by spacecraft design. There are no moving parts in the proximity of the batteries. Combined faults required for realized failure: A catastrophic failure must occur in an external system AND the failure must cause a collision sufficient to crush the batteries leading to an internal short circuit AND the satellite must be in a naturally sustained orbit at the time the crushing occurs. Failure Mode 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. Combined faults required for realized failure: Abrasion or piercing failure of circuit board coating or wire insulators AND dislocation of battery packs AND failure of battery terminal insulators AND failure to detect such failures in environmental tests must occur to result in this failure mode. Failure Mode 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. Combined faults required for realized failure: Thermal analysis AND thermal design AND mission simulations in thermal-vacuum chamber testing AND the PTC device must fail AND over- current

monitoring and control must all fail for this failure mode to occur. Failure Mode 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 safemode processor and radio receiver are enabled and charging the battery. 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. Combined faults required for realized failure: 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. Requirement 4.4-2: Design for passivation after completion of mission operations while in orbit about Earth or the Moon: Design of all spacecraft and launch vehicle orbital stages shall include the ability to deplete all onboard sources of stored energy and disconnect all energy generation sources when they are no longer required for mission operations or postmission disposal or control to a level which can not cause an explosion or deflagration large enough to release orbital debris or break up the spacecraft (Requirement 56450). Compliance statement: • SkyCube's battery charge circuits include overcharge protection and to limit the risk of battery failure. 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. Requirement 4.4-3. Limiting the long-term risk to other space systems from planned breakups: Compliance statement: • This requirement is not applicable. There are no planned breakups. Requirement 4.4-4: Limiting the short-term risk to other space systems from planned breakups: Compliance statement: • This requirement is not applicable. There are no planned breakups. ODAR Section 5: Assessment of Spacecraft Potential for On-Orbit Collisions Assessment  of  spacecraft  compliance  with  Requirements  4.5-­‐1  and  4.5-­‐2  (per  DAS  v2.02,   and  calculation  methods  provided  in  NASA-­‐STD-­‐8719.14,  section  4.5.4):  

• Requirement 4.5-1. Limiting debris generated by collisions with large objects when operating in Earth orbit: For each spacecraft and launch vehicle orbital stage in or passing through LEO, the program or project shall demonstrate that, during the orbital lifetime of each spacecraft and orbital stage, the probability of accidental collision with space objects larger than 10 cm in diameter is less than 0.001 (Requirement 56506). Large Object Impact and Debris Generation Probability: 0.000000; COMPLIANT. • Requirement 4.5-2. Limiting debris generated by collisions with small objects when operating in Earth or lunar orbit: For each spacecraft, the program or project shall demonstrate that, during the mission of the spacecraft, the probability of accidental collision with orbital debris and meteoroids sufficient to prevent compliance with the applicable postmission disposal requirements is less than 0.01 (Requirement 56507). Small Object Impact and Debris Generation Probability: 0.000000; COMPLIANT Figures 4 and 5 show the DAS derived ISS avoidance maneuver time analyses. Figure 4 illustrates how switching between high and low drag modes will yield an avoidance maneuver time based on the given radial keepout distance. Figure 4 uses a radial keepout distance of ±1km. Figure 5 shows how the avoidance maneuver time varies with the radial keepout distance. The most realistic case is a radial keepout distance of ±10km. This keepout distance yields an avoidance maneuver time of approximately 40 days and is defined as the nominal avoidance scenario.

Figure 4: RANGE ISS Avoidance Maneuver Time Analysis from DAS Output Figure 5: RANGE ISS Avoidance Maneuver Time versus Radial Keepout Distance from DAS Output Regarding  the  risk  of  the  two  satellites  colliding  after  separation  and  generating  debris,  this   is  considered  extremely  unlikely.    Because  the  satellites  have  no  active  propulsion  system,   the  formation  flying  will  be  managed  through  differential  drag.    By  changing  the  orientation   of  one  of  the  satellites,  a  change  in  drag  area  occurs,  inducing  a  relative  velocity  between   the  spacecraft.      This  allows  the  relative  distance  between  the  spacecraft  to  be  controlled.     Initially,   the   RANGE   satellites   will   remain   within   a   few   km   from   each   other,   but   the   intention  is  to  extend  this  distance  later  in  the  mission  to  test  the  limits  of  the  intersatellite   ranging   techniques.     As   such,   the   probability   of   collision   is   extremely   low.     First,   the   differential   drag   technique   is   not   nearly   precise   enough   to   bring   the   satellites   back   to   within   centimeters   of   each   other,   especially   after   extended   operations.     Second,   the   two   spacecraft   will   be   physically   attached   upon   deployment   using   a   frangibolt   (manufactured   by   TiNi   Aerospace,   Inc.),   and   will   only   separate   upon   command,   with   spring   mechanisms   (located   on   the   structure   feet)   used   to   ensure   separation.     The   frangibolt   itself   is   a   notched   titanium  screw  that  is  broken  with  a  thermal  actuator.    These  frangibolts  are  designed  for   space  applications,  have  extensive  flight  heritage,  and  do  not  generate  debris.    The  use  of   the   frangibolt   allows   for   a   very   controlled   separation,   further   reducing   the   chance   for   an  

accidental   collision.     Lastly,   the   relative   velocities   between   the   spacecraft   will   only   reach   the  cm/s  level,  so  even  in  the  unlikely  event  of  a  collision,  the  velocities  are  not  high  enough   to  damage  the  spacecraft  or  generate  debris.   ODAR Section 6: Assessment of Spacecraft Postmission Disposal Plans and Procedures 6.1  Description  of  spacecraft  disposal  option  selected:   • The  satellite  will  de-­‐orbit  naturally  by  atmospheric  re-­‐entry.  There  is  no  propulsion   system.       6.2  Plan  for  any  spacecraft  maneuvers  required  to  accomplish  postmission  disposal:   • None.       6.3  Calculation  of  area-­‐to-­‐mass  ratio  after  postmission  disposal,  if  the  controlled  reentry   option  is  not  selected:     • Spacecraft Mass: 2.125 kg • Cross-sectional Area: 0.015 m^2 (Calculated by DAS 2.02 for the configuration in Figure 3). • Area to mass ratio: 0.015/2.125 = 0.007059 m^2/kg 6.4  Assessment  of  spacecraft  compliance  with  Requirements  4.6-­‐1  through  4.6-­‐5  (per  DAS  v   2.0  and  NASA-­‐STD-­‐8719.14  section):     • Requirement  4.6-­‐1.  Disposal  for  space  structures  passing  through  LEO:  A  spacecraft  or   orbital  stage  with  a  perigee  altitude  below  2000  km  shall  be  disposed  of  by  one  of  three   methods:  (Requirement  56557)       a. Atmospheric reentry option: • Leave the space structure in an orbit in which natural forces will lead to atmospheric reentry within 25 years after the completion of mission but no more than 30 years after launch; or • Maneuver the space structure into a controlled de-orbit trajectory as soon as practical after completion of mission. b. Storage orbit option: • Maneuver the space structure into an orbit with perigee altitude greater than 2000 km and apogee less than GEO - 500 km. c. Direct retrieval: • Retrieve the space structure and remove it from orbit within 10 years after completion of mission. Analysis: The RANGE satellite reentry is COMPLIANT using Method “a”. RANGE will re-enter approximately 7.84 years after launch with orbit history as shown in Figure 6 (analysis assumes a nadir pointing configuration).

Figure 6: RANGE orbit history for low-drag configuration Figure 7: RANGE orbit history for high-drag configuration Requirement 4.6-2. Disposal for space structures near GEO. Analysis: Not applicable. RANGE orbit is LEO. Requirement 4.6-3. Disposal for space structures between LEO and GEO. Analysis: Not applicable. RANGE orbit is LEO. Requirement 4.6-4. Reliability of Postmission Disposal Operations

Analysis: RANGE de-orbiting does not rely on de-orbiting devices. Deployment from launch vehicle will result in de-orbiting in approximately 7.84 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 of human casualty: The potential for human casualty is assumed for any object with an impacting kinetic energy in excess of 15 joules: a. For uncontrolled reentry, the risk of human casualty from surviving debris shall not exceed 0.0001 (1:10,000) (Requirement 56626). Summary Analysis Results: DAS v2.0 reports that RANGE is compliant with the requirement. It predicts that no components reach the ground. As seen in the analysis outputs below, the impact kinetic energies are 0.000000 Joules and impact casualty areas are all 0.000000 square meters. 08 20 2016; 20:52:22PM DAS Application Started 08 20 2016; 20:52:25PM Opened Project C:\Program Files (x86)\NASA\DAS 2.0\project\Range\ 08 20 2016; 20:52:35PM Processing Requirement 4.3-1: Return Status : Not Run ===================== No Project Data Available ===================== =============== End of Requirement 4.3-1 =============== 08 20 2016; 20:52:37PM Processing Requirement 4.3-2: Return Status : Passed ===================== No Project Data Available ===================== =============== End of Requirement 4.3-2 =============== 08 20 2016; 20:52:39PM Requirement 4.4-3: Compliant =============== End of Requirement 4.4-3 =============== 08 20 2016; 20:52:45PM Processing Requirement 4.5-1: Return Status : Passed ============== Run Data ============== **INPUT** Space Structure Name = Range 1 Space Structure Type = Payload

Perigee Altitude = 500.000000 (km) Apogee Altitude = 500.000000 (km) Inclination = 97.406500 (deg) RAAN = 0.000000 (deg) Argument of Perigee = 0.000000 (deg) Mean Anomaly = 0.000000 (deg) Final Area-To-Mass Ratio = 0.007059 (m^2/kg) Start Year = 2016.830000 (yr) Initial Mass = 2.250000 (kg) Final Mass = 2.250000 (kg) Duration = 6.642000 (yr) Station-Kept = False Abandoned = True PMD Perigee Altitude = -1.000000 (km) PMD Apogee Altitude = -1.000000 (km) PMD Inclination = 0.000000 (deg) PMD RAAN = 0.000000 (deg) PMD Argument of Perigee = 0.000000 (deg) PMD Mean Anomaly = 0.000000 (deg) **OUTPUT** Collision Probability = 0.000000 Returned Error Message: Normal Processing Date Range Error Message: Normal Date Range Status = Pass ============== =============== End of Requirement 4.5-1 =============== 08 20 2016; 20:52:51PM Requirement 4.5-2: Compliant 08 20 2016; 20:52:52PM Processing Requirement 4.6 Return Status : Passed ============== Project Data ============== **INPUT** Space Structure Name = Range 1 Space Structure Type = Payload Perigee Altitude = 500.000000 (km) Apogee Altitude = 500.000000 (km) Inclination = 97.406500 (deg) RAAN = 0.000000 (deg) Argument of Perigee = 0.000000 (deg) Mean Anomaly = 0.000000 (deg) Area-To-Mass Ratio = 0.007059 (m^2/kg) Start Year = 2016.830000 (yr) Initial Mass = 2.250000 (kg) Final Mass = 2.250000 (kg) Duration = 6.642000 (yr) Station Kept = False

Abandoned = True PMD Perigee Altitude = -1.000000 (km) PMD Apogee Altitude = -1.000000 (km) PMD Inclination = 0.000000 (deg) PMD RAAN = 0.000000 (deg) PMD Argument of Perigee = 0.000000 (deg) PMD Mean Anomaly = 0.000000 (deg) **OUTPUT** Suggested Perigee Altitude = 500.000000 (km) Suggested Apogee Altitude = 500.000000 (km) Returned Error Message = Reentry during mission (no PMD req.). Released Year = 2023 (yr) Requirement = 61 Compliance Status = Pass ============== =============== End of Requirement 4.6 =============== 08 20 2016; 20:53:02PM *********Processing Requirement 4.7-1 Return Status : Passed ***********INPUT**** Item Number = 1 name = Range 1 quantity = 1 parent = 0 materialID = 5 type = Box Aero Mass = 2.250000 Thermal Mass = 2.250000 Diameter/Width = 0.100000 Length = 0.170000 Height = 0.100000 name = CSAC quantity = 1 parent = 1 materialID = 23 type = Box Aero Mass = 0.035000 Thermal Mass = 0.035000 Diameter/Width = 0.065000 Length = 0.090150 Height = 0.015000 name = CubeSense Camera quantity = 1 parent = 1 materialID = 23 type = Box

Aero Mass = 0.080000 Thermal Mass = 0.080000 Diameter/Width = 0.090000 Length = 0.096000 Height = 0.010000 name = Mid Panel for NanoPower quantity = 1 parent = 1 materialID = 5 type = Flat Plate Aero Mass = 0.100000 Thermal Mass = 0.100000 Diameter/Width = 0.100000 Length = 0.100000 name = Lidar Housing quantity = 1 parent = 1 materialID = 23 type = Box Aero Mass = 0.022000 Thermal Mass = 0.022000 Diameter/Width = 0.037000 Length = 0.048000 Height = 0.020000 name = L-Plate quantity = 2 parent = 1 materialID = 5 type = Box Aero Mass = 0.093213 Thermal Mass = 0.093213 Diameter/Width = 0.100000 Length = 0.125000 Height = 0.100000 name = Side PCB quantity = 4 parent = 1 materialID = 5 type = Box Aero Mass = 0.004500 Thermal Mass = 0.004500 Diameter/Width = 0.080000 Length = 0.095000 Height = 0.002000 name = GPS Antenna quantity = 1 parent = 1 materialID = 23 type = Cylinder

Aero Mass = 0.001500 Thermal Mass = 0.001500 Diameter/Width = 0.024000 Length = 0.089000 name = Reaction Wheel Mount quantity = 1 parent = 1 materialID = 5 type = Box Aero Mass = 0.001323 Thermal Mass = 0.001323 Diameter/Width = 0.028000 Length = 0.031000 Height = 0.001000 name = Reaction Wheel quantity = 1 parent = 1 materialID = 5 type = Box Aero Mass = 0.058590 Thermal Mass = 0.058590 Diameter/Width = 0.031000 Length = 0.031000 Height = 0.028000 name = Sun Sensor quantity = 1 parent = 1 materialID = 24 type = Box Aero Mass = 0.006500 Thermal Mass = 0.006500 Diameter/Width = 0.014000 Length = 0.043000 Height = 0.006000 name = Solar Panel quantity = 4 parent = 1 materialID = 24 type = Flat Plate Aero Mass = 0.302682 Thermal Mass = 0.302682 Diameter/Width = 0.083000 Length = 0.098000 name = Bottom Panel quantity = 1 parent = 1 materialID = 5 type = Box Aero Mass = 0.094997

Thermal Mass = 0.094997 Diameter/Width = 0.100000 Length = 0.100000 Height = 0.015000 name = Mid Panel quantity = 1 parent = 1 materialID = 5 type = Flat Plate Aero Mass = 0.011389 Thermal Mass = 0.011389 Diameter/Width = 0.100000 Length = 0.100000 name = End Cap quantity = 1 parent = 1 materialID = 5 type = Box Aero Mass = 0.189559 Thermal Mass = 0.189559 Diameter/Width = 0.100000 Length = 0.100000 Height = 0.008000 name = NanoHub Internal quantity = 1 parent = 1 materialID = 23 type = Box Aero Mass = 0.045000 Thermal Mass = 0.045000 Diameter/Width = 0.089000 Length = 0.096000 Height = 0.018000 name = Novatell GPS Receiver quantity = 1 parent = 1 materialID = 23 type = Box Aero Mass = 0.037000 Thermal Mass = 0.037000 Diameter/Width = 0.060000 Length = 0.100000 Height = 0.009000 name = Nanocom-ant430 quantity = 1 parent = 1 materialID = 5 type = Box Aero Mass = 0.024000

Thermal Mass = 0.024000 Diameter/Width = 0.113000 Length = 0.113000 Height = 0.098000 name = MotherBoard quantity = 1 parent = 1 materialID = 23 type = Box Aero Mass = 0.138404 Thermal Mass = 0.138404 Diameter/Width = 0.089000 Length = 0.092000 Height = 0.019000 name = BP4 Case quantity = 1 parent = 1 materialID = 23 type = Box Aero Mass = 0.075400 Thermal Mass = 0.075400 Diameter/Width = 0.087000 Length = 0.093000 Height = 0.029000 name = Placeholder Side Panel quantity = 2 parent = 1 materialID = 5 type = Flat Plate Aero Mass = 0.042383 Thermal Mass = 0.042383 Diameter/Width = 0.079000 Length = 0.145000 name = NanoPower-P31us quantity = 1 parent = 1 materialID = 5 type = Box Aero Mass = 0.200000 Thermal Mass = 0.200000 Diameter/Width = 0.089000 Length = 0.093000 Height = 0.012000 name = BP4 Batteries quantity = 4 parent = 1 materialID = 46 type = Cylinder Aero Mass = 0.048600

Thermal Mass = 0.048600 Diameter/Width = 0.018500 Length = 0.065000 **************OUTPUT**** Item Number = 1 name = Range 1 Demise Altitude = 77.993918 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = CSAC Demise Altitude = 77.412707 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = CubeSense Camera Demise Altitude = 76.841738 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = Mid Panel for NanoPower Demise Altitude = 75.631840 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = Lidar Housing Demise Altitude = 77.240004 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = L-Plate Demise Altitude = 77.231433 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = Side PCB Demise Altitude = 77.876629 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = GPS Antenna Demise Altitude = 77.944535 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000

************************************* name = Reaction Wheel Mount Demise Altitude = 77.784285 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = Reaction Wheel Demise Altitude = 74.003472 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = Sun Sensor Demise Altitude = 77.726293 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = Solar Panel Demise Altitude = 75.790379 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = Bottom Panel Demise Altitude = 76.103363 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = Mid Panel Demise Altitude = 77.723676 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = End Cap Demise Altitude = 73.958019 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = NanoHub Internal Demise Altitude = 77.400199 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = Novatell GPS Receiver Demise Altitude = 77.354097 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000

************************************* name = Nanocom-ant430 Demise Altitude = 77.785988 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = MotherBoard Demise Altitude = 76.141238 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = BP4 Case Demise Altitude = 77.081277 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = Placeholder Side Panel Demise Altitude = 77.190613 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = NanoPower-P31us Demise Altitude = 73.489878 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* name = BP4 Batteries Demise Altitude = 71.910441 Debris Casualty Area = 0.000000 Impact Kinetic Energy = 0.000000 ************************************* =============== End of Requirement 4.7-1 =============== Requirements 4.7-1b and 4.7-1c below are non-applicable requirements because RANGE 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). 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 7A: Assessment of Spacecraft Hazardous Materials Not Applicable. There are no hazardous materials contained on the spacecraft. ODAR Section 8: Assessment for Tether Missions Not applicable. There are no tethers in the RANGE mission. END of ODAR for RANGE.


Document Created: 2260-04-24 00:00:00
Document Modified: 2260-04-24 00:00:00
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