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Design of a Representative Leo Satellite and Hypervelocity Impact Test to Improve the Nasa Standard Breakup Model

Permanent Link: http://ufdc.ufl.edu/UFE0045626/00001

Material Information

Title: Design of a Representative Leo Satellite and Hypervelocity Impact Test to Improve the Nasa Standard Breakup Model
Physical Description: 1 online resource (93 p.)
Language: english
Creator: Clark, Sheldon C
Publisher: University of Florida
Place of Publication: Gainesville, Fla.
Publication Date: 2013

Subjects

Subjects / Keywords: airforce -- collision -- debris -- debrisat -- hypervelocity -- impact -- nasa -- orbital
Mechanical and Aerospace Engineering -- Dissertations, Academic -- UF
Genre: Aerospace Engineering thesis, M.S.
bibliography   ( marcgt )
theses   ( marcgt )
government publication (state, provincial, terriorial, dependent)   ( marcgt )
born-digital   ( sobekcm )
Electronic Thesis or Dissertation

Notes

Abstract: The growth of the man made orbital debris population poses accelerating risk to the operational safety of both manned and unmanned space assets. Several government and private entities around the world identify,track, and monitor the debris population, though current tracking technology does not allow for an exhaustive mapping of the complete debris field.Currently the Space Surveillance Network, operated by the U.S. Strategic Command, is a recognized authority for tracking space objects and has the capability of identifying orbiting objects down to approximately 10cm in diameter. SSN tracks over 20,000 objects now in orbit. However, there is an estimated 500,000 objects 1cm to 10cm in diameter and over 100 million objects smaller then 1cm. These objects consist of intact operational and non-operational spacecraft, upper stage rocket bodies, and natural materials such as meteoroids. The vast majority of concerning objects, however, are debris fragments resulting from on-orbit breakups – primarily explosions of upper stage rocket bodies and collisions with intact satellites. To account for these fragmentation objects that are generated during an on-orbit collision,orbital debris scientists employ a leading statistical population model developed by NASA called the Standard Breakup Model. The Standard Breakup Model uses a combination of on-orbit tracking data and terrestrial ground hypervelocity impact tests to predict the population and ejection velocity of collision debris fragments. The Standard Breakup Model is based in part on a series of breakup tests – the most important being a 1992 hypervelocity impact test, the Satellite Orbital-debris Characterization Impact Test (SOCIT), on an un-flown U.S. Navy Transit satellite. This impact test served as the primary input to the Standard Breakup Model when the fully operational Iridium-33 satellite accidentally collided with a defunct Russian Kosmos satellite. When the resulting debris field was tracked by the SSN and compared with the results of the breakup model, a noticeable discrepancy appeared. While the Standard Breakup Model successfully predicted the population of the Kosmos debris field, it did not successfully predict the population of the Iridium-33 debris field. This discrepancy was attributed to the fact that the satellite used in the primary hypervelocity impact test, SOCIT, was an obsolete 1960’s satellite with outdated materials, components, and construction methods. Therefore, it was determined that the Standard Breakup Model should be updated to better predict the debris field of modern satellites and that a new ground based hypervelocity impact test needed to be conducted. This thesis addresses specific design considerations for a representative satellite currently being designed to improve the Standard Breakup Model and presents hypothetical hypervelocity impact test considerations that could be used to simulate an on-orbit fragmentation.
General Note: In the series University of Florida Digital Collections.
General Note: Includes vita.
Bibliography: Includes bibliographical references.
Source of Description: Description based on online resource; title from PDF title page.
Source of Description: This bibliographic record is available under the Creative Commons CC0 public domain dedication. The University of Florida Libraries, as creator of this bibliographic record, has waived all rights to it worldwide under copyright law, including all related and neighboring rights, to the extent allowed by law.
Statement of Responsibility: by Sheldon C Clark.
Thesis: Thesis (M.S.)--University of Florida, 2013.
Local: Adviser: Fitz-Coy, Norman G.

Record Information

Source Institution: UFRGP
Rights Management: Applicable rights reserved.
Classification: lcc - LD1780 2013
System ID: UFE0045626:00001

Permanent Link: http://ufdc.ufl.edu/UFE0045626/00001

Material Information

Title: Design of a Representative Leo Satellite and Hypervelocity Impact Test to Improve the Nasa Standard Breakup Model
Physical Description: 1 online resource (93 p.)
Language: english
Creator: Clark, Sheldon C
Publisher: University of Florida
Place of Publication: Gainesville, Fla.
Publication Date: 2013

Subjects

Subjects / Keywords: airforce -- collision -- debris -- debrisat -- hypervelocity -- impact -- nasa -- orbital
Mechanical and Aerospace Engineering -- Dissertations, Academic -- UF
Genre: Aerospace Engineering thesis, M.S.
bibliography   ( marcgt )
theses   ( marcgt )
government publication (state, provincial, terriorial, dependent)   ( marcgt )
born-digital   ( sobekcm )
Electronic Thesis or Dissertation

Notes

Abstract: The growth of the man made orbital debris population poses accelerating risk to the operational safety of both manned and unmanned space assets. Several government and private entities around the world identify,track, and monitor the debris population, though current tracking technology does not allow for an exhaustive mapping of the complete debris field.Currently the Space Surveillance Network, operated by the U.S. Strategic Command, is a recognized authority for tracking space objects and has the capability of identifying orbiting objects down to approximately 10cm in diameter. SSN tracks over 20,000 objects now in orbit. However, there is an estimated 500,000 objects 1cm to 10cm in diameter and over 100 million objects smaller then 1cm. These objects consist of intact operational and non-operational spacecraft, upper stage rocket bodies, and natural materials such as meteoroids. The vast majority of concerning objects, however, are debris fragments resulting from on-orbit breakups – primarily explosions of upper stage rocket bodies and collisions with intact satellites. To account for these fragmentation objects that are generated during an on-orbit collision,orbital debris scientists employ a leading statistical population model developed by NASA called the Standard Breakup Model. The Standard Breakup Model uses a combination of on-orbit tracking data and terrestrial ground hypervelocity impact tests to predict the population and ejection velocity of collision debris fragments. The Standard Breakup Model is based in part on a series of breakup tests – the most important being a 1992 hypervelocity impact test, the Satellite Orbital-debris Characterization Impact Test (SOCIT), on an un-flown U.S. Navy Transit satellite. This impact test served as the primary input to the Standard Breakup Model when the fully operational Iridium-33 satellite accidentally collided with a defunct Russian Kosmos satellite. When the resulting debris field was tracked by the SSN and compared with the results of the breakup model, a noticeable discrepancy appeared. While the Standard Breakup Model successfully predicted the population of the Kosmos debris field, it did not successfully predict the population of the Iridium-33 debris field. This discrepancy was attributed to the fact that the satellite used in the primary hypervelocity impact test, SOCIT, was an obsolete 1960’s satellite with outdated materials, components, and construction methods. Therefore, it was determined that the Standard Breakup Model should be updated to better predict the debris field of modern satellites and that a new ground based hypervelocity impact test needed to be conducted. This thesis addresses specific design considerations for a representative satellite currently being designed to improve the Standard Breakup Model and presents hypothetical hypervelocity impact test considerations that could be used to simulate an on-orbit fragmentation.
General Note: In the series University of Florida Digital Collections.
General Note: Includes vita.
Bibliography: Includes bibliographical references.
Source of Description: Description based on online resource; title from PDF title page.
Source of Description: This bibliographic record is available under the Creative Commons CC0 public domain dedication. The University of Florida Libraries, as creator of this bibliographic record, has waived all rights to it worldwide under copyright law, including all related and neighboring rights, to the extent allowed by law.
Statement of Responsibility: by Sheldon C Clark.
Thesis: Thesis (M.S.)--University of Florida, 2013.
Local: Adviser: Fitz-Coy, Norman G.

Record Information

Source Institution: UFRGP
Rights Management: Applicable rights reserved.
Classification: lcc - LD1780 2013
System ID: UFE0045626:00001


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1 DESIGN OF A REPRESENTATIVE LEO SATELLITE AND HYPERVELOCITY IMPACT TEST TO IMPROVE THE NASA STANDARD BREAKUP MODEL By SHELDON C OLBY CLARK A THESIS PRESENTED TO THE GRADUATE SCHOOL OF THE UNIVERSITY OF FLORIDA IN PARTIAL FULFILLMENT OF THE REQUIREMENTS FOR THE DEGREE OF MASTER OF SCIENCE UNIVERSITY OF FLORIDA 2013

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2 2013 Sheldon C. Clark

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3 To my family forever first and foremost for making me who I am today and supporting me unconditionally through all of my endeavors. To my friends, colleagues, and advisors who have accompanied me on my journey I am forever grateful.

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4 ACKNOWLEDGMENTS I would like to thank all those who have given me the opportunity to gro w to the person I am today. I thank my parents and family for their unconditional love and support. I am grateful to Dr. Norman Fitz Coy for the learning opportunities he has provided and the guidance he has shared with me. I also thank my longtime friends my colleagues in the Space Systems Group, and my peers at the University of Florida for their support and help through the years.

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5 TABLE OF CONTENTS page ACKNOWLEDGMENTS ................................ ................................ ................................ .. 4 LIST OF TABLES ................................ ................................ ................................ ............ 7 LIST OF FIGURES ................................ ................................ ................................ .......... 8 LIST OF ABBREVIATIONS ................................ ................................ ........................... 11 ABSTRA CT ................................ ................................ ................................ ................... 12 CHAPTER 1 INTRODUCTION ................................ ................................ ................................ .... 14 Motivation ................................ ................................ ................................ ............... 14 History and Backgr ound ................................ ................................ .......................... 14 Orbital Debris Risks ................................ ................................ .......................... 14 Previous On Orbit Events ................................ ................................ ................. 15 Solwind P78 1 ................................ ................................ ............................ 16 USA 193 ................................ ................................ ................................ .... 17 China FY 1C anti s at ellite t est ................................ ................................ ... 18 Iridium 33 and Cosmos 2251 c ollision ................................ ....................... 19 Additional Events ................................ ................................ .............................. 20 2 BREAKUP MODEL DEVELOPMENTS ................................ ................................ ... 22 History of Orbital Breakup Model Development ................................ ...................... 22 1970 Bess Experiments ................................ ................................ .................... 22 EVOLVE ................................ ................................ ................................ ........... 23 LEGEND ................................ ................................ ................................ ........... 24 Hypervelocity Impact Tests ................................ ................................ ..................... 25 Previous Hypervelocity Impact Tests ................................ ................................ ...... 25 SOCIT ................................ ................................ ................................ .............. 26 Low Velocity Impact Te sts ................................ ................................ ................ 27 Standard Breakup Model Improvements ................................ ................................ 28 Research Focus ................................ ................................ ................................ ...... 28 3 DEBRISAT: DESIGN ................................ ................................ .............................. 29 Design Motivation .. 2 9 System Level Design ................................ ................................ .............................. 31 LEO Satellite Study ................................ ................................ ................................ 34 DebriSat Subsystems ................................ ................................ ............................. 39 Electrical Power System ................................ ................................ ................... 40

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6 Telemetry, Tracking, and Command ................................ ................................ 48 Command and Data Handling ................................ ................................ .......... 54 Payload ................................ ................................ ................................ ............ 57 4 DEBRISAT PROPO SED HYPERVELOCITY TEST ................................ ............... 73 Test Objectives ................................ ................................ ................................ ....... 74 Test Range ................................ ................................ ................................ ............. 75 Projectile Selection ................................ ................................ ................................ 77 Test Des cription ................................ ................................ ................................ ...... 77 5 POST IMPACT CONSIDERATIONS ................................ ................................ ...... 83 Health and Safety Considerations ................................ ................................ .......... 83 Characteristic Analysis ................................ ................................ ............................ 83 6 CO NCLUSIONS AND FUTURE WORK ................................ ................................ 87 A PPENDIX: SPECTROLAB UTJ SOLAR CELLS ................................ ......................... 88 LIST OF REFERENCES ................................ ................................ ............................... 90 BIOGRAPHIC AL SKETCH ................................ ................................ ............................ 93

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7 LIST OF TABLES Table P age 3 1 Top level debrisat characteristics ................................ ................................ ....... 31 3 2 EPS mass ................................ ................................ ................................ .......... 48 3 3 List of TT&C components ................................ ................................ .................. 53 3 4 TT&C mass breakdown ................................ ................................ ..................... 54 3 5 Primary C&DH components ................................ ................................ ............... 57 3 6 C&DH component masses ................................ ................................ ................ 57 3 7 Optical imager mass characteristics ................................ ................................ .. 63 3 8 Spectrometer mass characteristics ................................ ................................ .... 70 3 9 Payload components ................................ ................................ ......................... 71 3 10 Total payload mass characteristics ................................ ................................ .... 72 5 1 Catch panel material comparison ................................ ................................ ...... 72

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8 LIST OF FIGURES Figure P age 1 1 The Solwind P 78 1 ................................ ................................ ........................... 17 3 1 External view of D ebr iS at ................................ ................................ ................. 30 3 2 DebriSat bay definitions ................................ ................................ .................... 31 3 3 C ompo nent layout ................................ ................................ ............................ 33 3 4 M ass distribution range of 467 satellites in UCS database ............................... 34 3 5 M ass distribution range of 50 selected satellites for UF study .......................... 35 3 6 ADCS sensor usage by mass ................................ ................................ ........... 36 3 7 ADCS actuator usage by mass ................................ ................................ ......... 37 3 8 P ropulsion system usage by mass ................................ ................................ ... 37 3 9 LEO communication bands by mass range ................................ ...................... 38 3 10 LEO batteries by mass range ................................ ................................ ........... 39 3 11 Spectrolab 28.3% UTJ CIC solar cells ................................ .............................. 41 3 12 D eployed solar panels ................................ ................................ ...................... 42 3 13 T ransparent view of li ion battery redesign ................................ ....................... 43 3 14 L i ion cell assembly ................................ ................................ .......................... 44 3 15 L i ion battery case dimensions ................................ ................................ ......... 44 3 16 P ower conditioning and distribution module model ................................ ........... 45 3 17 P ower conditioning and distribution module dimensions ................................ .. 46 3 18 B raided stainless steel shield and d sub connector ................................ .......... 47 3 19 Conceptual EPS component connections ................................ ........................ 47 3 20 Conceptual TT&C connection diagram ................................ ............................. 49 3 21 X band antenna design ................................ ................................ .................... 49 3 22 Helical S band antenna design ................................ ................................ ......... 50

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9 3 23 D igikey omni directional whip antenna ................................ ............................. 51 3 24 S band TT&C avionics model ................................ ................................ ........... 52 3 25 S band transceiver dimensions ................................ ................................ ........ 52 3 26 C ots signal manipulation components ................................ .............................. 53 3 27 DebriSat flight computer ................................ ................................ ................... 55 3 28 C&DH data recorder ................................ ................................ ......................... 56 3 29 Celestron nexstar 4SE maksutov cassegrain telescope ................................ ... 59 3 30 I nternal view of modified celestron telescope ................................ ................... 59 3 31 O ptical imager avionics box ................................ ................................ .............. 60 3 32 O ptical imager payload ................................ ................................ ..................... 61 3 33 E xploded view of optical imager ................................ ................................ ....... 62 3 34 N ear infrared spectrometer ................................ ................................ ............... 64 3 35 Internal view of the near infrared spectrometer ................................ ................ 65 3 36 D etailed dimensions of spectrometer assembly ................................ ............... 66 3 37 E xploded view of spectrometer optical bench ................................ .................. 67 3 38 E xploded view of 90 degree fold mirror and casing ................................ .......... 68 3 39 E xploded view of titanium lens mount ................................ .............................. 69 3 40 Spectrometer CCD housing ................................ ................................ .............. 69 3 41 P ayload support module ................................ ................................ ................... 72 4 1 P roposed hypervelocity impact target DebriSat ................................ ............. 74 4 2 T est range end view ................................ ................................ ...................... 76 4 3 T est range top view ................................ ................................ ....................... 76 4 4 T arget impact trajectory ................................ ................................ .................... 78 4 5 C atch panel assembly dimension ................................ ................................ ..... 81 5 1 VIS process flow ................................ ................................ ............................... 84

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10 5 2 Proposed VIS analysis station ................................ ................................ .......... 85 5 3 Hypothetical VIS software analysis using known objects ................................ 86

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11 LIST OF ABBREVIATIONS ASAT Anti Satellite Test BCR Battery Charge Regulator C&DH Command and Data Handling CIC Cell Interconnect Coverglass COTS Commercial Off the Shelf EPS Electrical Power System GEO Geosynchronous Orbit JSC Johnson Space Center LEGEND LEO to GEO Environment Debris model LEO Low Earth Orbit Li ion Lithium Ion Ni Cd Nickel Cadmium NORAD N orth American Aerospace Defense Command NRO National Reconnaissance Office ODPO Orbital Debris Program Office PARCS Perimeter Acquisition Radar Characterization System PPT Peak Power Tracker RF Radio Frequency SOCIT Satellite Orbital D ebris Characterization Impact Test SSN Space Surveillance Network TT&C Telemetry Tracking and Command UTJ Ultra Triple Junction VIS Visual Inspection System

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12 Abstract of Thesis Presented to the Graduate School of the University of Florida in Partial Fulfillment of the Requirements for the Degree of Master of Science DESIGN OF A REPRESENTA TIVE LEO SATELLITE AND HYPERVELOCITY IMPACT TEST TO IMPROVE THE NASA STANDARD BREAKUP MODEL By Sheldon C. Clark May 2013 Chair: Norman Fitz Coy Major : Aerospace Engineering The growth of the man made orbital debris population poses accelerating risk to the operational safety of both manned and unmanned space assets. Several government and private entities around the world identify, track, and monitor the debris population, though current tracking techn ology does not allow for an exhaustive mapping of the complete debris field. Currently the Space Surveillance Network (SSN), operated by the U.S. Strategic Command, is a recognized authority for tracking space objects and has the capability of identifying orbiting objects down to ap proximately 10cm in diameter. SSN tracks over 20,000 objects now in orbit. However, there is an estimated 500,000 objects 1cm to 10cm in diameter and over 100 million objects smaller then 1cm. These objects consist of intact oper ational and non operational spacecraft, upper stage rocket bodies, and natural materials such as meteoroids. The vast majority of concerning objects however are debris fragments resulting from on orbit breakups primarily explosions of upper stage rocket bodies and collisions with intact satellites To account for these fragmentation objects that are generated during an on orbit collision, orbital debris sci entists employ a leading statistical population model developed by NASA called the Standard Breakup Model. The Standard Breakup Model uses a combination of on orbit

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13 tracking data and terrestrial ground hypervelocity impact tests to predict the population a nd ejection velocity of collision debris fragments. The Standard Breakup Model is based in part on a series of breakup tests the most important being a 1992 hypervelocity impact test, the Satellite Orbital debris Characterization Impact Test (SOCIT), on an un flown U.S. Navy Transit satellite. This impact test served as the primary input to the Standard Breakup Model when the fully operational Iridium 33 satellite accidently collided with a defunct Russian Kosmos satellite. When the resulting debris fiel d was tracked by the SSN and compared with the results of the breakup model, a noticeable discrepancy appeared. While the Standard Breakup Model successfully predicted the population of the Kosmos debris field, it did not successfully predict the populatio n of the Iridium 33 debris field. This discrepancy was attributed to the fact that the satellite used in the primary hypervelocity impact test, SOCIT, was an obsolete methods. Therefore it was determined that the Standard Breakup Model should be updated to better predict the debris field of modern satellites and that a new ground based hypervelocity impact test needed to be conducted. This thesis addresses specific design consideration s for a representative satellite currently being designed to improve the Standard Breakup Model and presents hypothetical hypervelocity impact test considerations that could be used to simulate an on orbit fragmentation

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14 CHAPTER 1 INTRODUCTION Motivation The Orbital Debris Program Office (ODPO) at Johnson Space Center (JSC) tracks and monitors objects in orbit from low Earth orbit (LEO) to Geo synchronous Earth orbit (GEO). The office uses tracking measurements and statistical breakup models that are used in the current long term satellite population model LEO to GEO Environment Debris model ( LEGEND ) In addition to in situ tracking events the statistical breakup model is based upon terrestrial hypervelocity impact tests. Several impa ct tests have been used to improve the breakup model, though the SOCIT test series conducted in 1992 has been the primary impact test on which the breakup model is based [1] Following the breakup of the Iridium 33 and Cosmos 2 251 satellites in 2009, the ODPO has investigated the accuracy of its satellite breakup model a critical input to LEGEND, with regard to modern materials The O D P O, in cooperation with the University of Florida, The Aerospace Company and the Air Force Sp ace and Missile Command, seeks to improve the breakup model by conducting a new terrestrial hypervelocity impact test. History and Background Orbital Debris Risks Man made objects have been placed into Earth orbit since 1957 when the former Soviet Union s parked the space r ace with the launch of Sp utnik. Since the beginning of the s pace era in 1957, the number of artificial satellites launched into has steadily increased [2] Beginning with Sputnik, the North American Aerospace Defense Command (NORAD) began maintaining a database, the Space Objects Catalog, of all

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15 known man made ob jects launched into space. Upon the release of this database in the previous work detailed the collisional formation of asteroids in the asteroid belt, began apply his asteroid belt analysis to the objects tracked by NORAD database. Kessler and Burton Cour Palais concluded that a collisional instability, similar to the fo rmation of the asteroid belt, would occur in the near Earth orbit regimes over coming decades as the number of man made objects inserted into orbit increased [3] This theory, now popularly known as the Kessler syndrome, has si nce been continuously studied by the scientific community [4] The propagation of the Kessler syndrome presents the realization of exponentially increasing collision risks, both catastrophic and non catastrophic. As of 2011, th e NORAD Space Surveillance N etwork (SSN) tracks more than 22 ,000 objects with a diameter of 10 cm or greater. In addition, more than 500,000 objects with a diameter between 1 cm and 10cm as well as more than 100 million objects smaller than 1 cm are predic ted to exist in the near Earth orbital environment. While the number of objects is concentrated under the 1 cm diameter regime, the majority of the object mass is concentrated in approximately 2000 objects in the larger 10cm and up regime [5] The primary risk posed by orbital breakups stems from the explosion and fragmentation of spent upper stage rocket bodies. However, following the 2007 China anti satellite test and the 2009 accidental Iridium 33 and Cosmos 2251 collision, c ollision fragmentation has become a significant percentage of trackable orbital debris [4] [6] Previous On Orbit Events Historically, the primary contributor to the growth of orbital debris has been the explosion of spent upper stage rocket bodies. Recently however, collision fragmentation, both intentional and accidental, has become a significant contributor the

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16 debris object population with s everal recent on orbit collision events contributing to the problem [4] A non exhaustive selection of event s of particular relevance to this research work is detailed and given China, and former Soviet Union have researched numerous ant i satellite weaponry (ASAT) designed to eliminate targeted satellite c apabilities. Tests began with a host of ground and air based interceptors that explored the utilization of conventional, nuclear, and kinetic energy warheads. Though ASAT research has co by the U.S., U.S.S.R., and China only two successful on orbit intercept tests by the U.S. are known to have occurred the 1985 Solwind P78 1 ASAT test and the 2008 USA 193 takedown. Solwind P78 1 T he Solwind P78 1 shown in Figure 1 1 was a scient ific coronal research satellite constructed by Ball Aerospace as the main contractor and was used primarily for solar observation Launched in 1979, the 898 kg satellite suffered from Electrical Power Syst em battery power failures early into the mission that would often shut all but the most vital systems down due to under voltage warnings. The Solwind began inflating operation costs and termination of the program was considered prior to its selection as an ASAT target. The U.S. Air Force, concerned over recent successful Soviet Union testing, began research into a serious anti satellite weapon and developed an ASM 135 for anti satellite operations. On September 13 th 1985, a modified F 15 launched the ASM 1 35 and successfully destroyed the Solwind P78 1 at a closing velocity of 22,000 km/s.

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17 Figure 1 1 The Solwind P78 1 Observation of the resulting debris field began approximately twelve hours after the collision using the Perimeter Acquisition Radar Characterization System (PARCS) and SSN ; o ver time the collected data suggested approximately 300 objects larger than 10 cm were created by the ASAT test, which was significantly lower than t he 800 objects predicted by the NASA Standard Breakup Model [7] [8] Observation data from the Solwind ASAT was considered a significant source of breakup information when the NASA Standard Breakup Model was revised in the late [7] USA 193 The second known U.S. ASAT test o ccurred on February 22 nd 2008 when the satellite USA 193 was intercepted by a modified ship launched ballistic missile. The USA 193 was a classified National Reconnaissance Office (NRO) satellite with an unknown payload and mission. System failure occurred early into the mission which

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18 s atmosphe re. Due to the early failure and subsequent loss of spacecraft control, USA 193 contained a full fuel tank of approximately 450kg of the toxic and reactive fuel Hydrazine. This fuel tank was expected to survive reentry to the surface where, if landing in a populated area, could cause death and injury to people. Following this determination, the U.S. government publically announced the intent to conduct an ASAT takedown using a highly modified Standard Missile 3 anti ballistic ship launched missile to destro y the satellite and its fuel tank prior to reentry. To mitigate the consequences from the fragmentation of USA 193 the ASAT was conducted following orbit degradation from a periapsis of 349 km to a periapsis of 247 km just prior to reentry [7] The low altitude of USA 193 at the time of the takedown caused an unknown majority of the resulting fragmentation to immediately reenter the following the breakup even though the applied NASA Standard Breakup Model predicted approximately 1400 fragments larger than the threshold 10 cm in diameter [7] The classified nature of the ASAT as well as the extremely low orbit has made co llection for useful breakup data from the collision difficult. Observation data from the Haystack radar and SSN have been included in several attempts to characterize the breakup [7] China FY 1C Anti Satellite Test The th 2007 against a dysfunctional Chinese weather satellite the 960 kg Feng Yun 1C While much of the data regarding the technological specifics of the test has not been released by the Chinese government, some information regarding the ASAT test has been made

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19 available by release or intelligence gathering. A modified ballistic missile intercepted the Feng Yun 1C satellite at an altitude of 864 km with an approximate closing velocity of 9 km/ s [6] Within six months of the collision, operational instructions to multiple spacecraft, including the International Space Station, were given across the LEO regime to conduct avoidance maneuvers from the fragments of the ASAT test. The resulting fragmentation generated the largest single debris producing event in the history of space exploration [6] A debris fragmentation field of 20,000 to 40,000 objects greater than 1 cm in diameter was generated following the collision, of which over 3000 objects greater than 10 cm in diameter were tracked and cataloged by the SSN [9] This large number of tracked objects by SSN, in combination with Haystack radar cross section data and two line elements for each of the debris fragments, allowed for unprecedented study of orbital fragmentation effects. Johnson et. al. stated the Feng Yun 1C ASAT test generated more debris fragments then the typical orbital breakup event and can be considered the single worst fragmentation event in history [6] Further, the Feng Yun 1C fragment data showed the colli sion caused an immediate increase in the long term debris population model during the orbital lifetime of the fragments, with Iridium 33 and Cosmos 2251 Collision Prior to the Iridium 33 and C osmos 2251 collision, three accidental collisions between catalogued space objects were known to have occurred [4] : Cosmos 1934 and Cosmos 926 debris Cerise and Ariane rocket body fragment Thor rocket and Chinese explosion fragment The first accidental collision between two intact satellites occurred on February 10 th 2009 when the operational satellite Iridium 33 from the U.S. owned Iridium

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20 communication constellation collided with Comsos 2251, a non operational Russi an comm unication satellite approximately 770 km above Siberia. Iridium 33, with a mass of 556 kg and orbiting at an ascending node of 231 degrees, collided at a closing velocity of 10 km/s at a nearly perpendicular impact angle to the 900kg Cosmos 2251 [10] The collision of the two spacecraft created one of the largest debris creating events in LEO second only to the Feng Yun 1C ASAT test [9] Following the loss of signal from Iridium 33 the SSN and H aystack radars began tracking debris fragment clouds in the orbit planes of each satellite. Shortly thereafter the accidental collision was confirmed and observation of the fragments began. Approximately 1700 objects greater than 10 cm in diameter were cat alogued by the SSN and several thousand more identified over 1 cm in diameter with the orbital lifetime of the debris field expected to be greater than 40 years [9] The Iridium 33 Cosmos 2251 collision led to wide spread me dia coverage over the importance of orbital debris sciences. The event is seen to be a major highlight of the consequences of the Kessler syndrome in particular the self propagating nature of debris in the LEO environment [4] [9] In particular, the FY 1C ASAT test, as well as the collision of Iridium 33 and Cosmos 2251, created sharp increases in the number of man made objects in Earth orbit. Additional Events In addition to the major fragmentati on events discussed, there are numerous examples of additional fragmentation events ranging from upper stage rocket body explosions to collision with individual debris fragments An exhaustive description of fragmentation events is provided by the NASA ODPO [8] Examples of these, for further information to the reader, include:

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21 Briz M explosion [9] Cosmos 2241 breakup [8] Delta 18 0 explosion [7] Comsos 1934 [8] [11] NOAA 7 [8] [11] The implications of these collision events, along with analysis data taken fro m terrestrial hypervelocity impact tests, provides both justification and in situ observation data for the creation of breakup models used in long term object population estimation. These models are presented in Chapter 2.

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22 CHAPTER 2 BREAKUP MODEL DEVELOPMENTS History of Orbital Breakup Model Development arose early in the development of the space program over the characterization of natural meteoroids present in the space environment that could possibly threaten the operational safety of manned or unmanned vehicles. Following the first decade of space lau nch activity, concern over the increasing number of objects being inserted into orbit along with the implications of several explosion events began to concern scientists. As early as the ved to be greater than impact risks from na tural Earth orbiting debris [3] Several organizations including the ESA and NASA have developed modeling programs to determine the effects of on orbit explosions and breakups such as CHAIN E MASTER IDES and EVOLVE/LEGEND [12] [13] [14] [15] NASA has been developing breakup models for study of the orbital debris D ue to the objectives of this thesis research, only these NASA model developments are sum marized in detail 1970 Bess Experiments al conducted hypervelocity tests on simulated spacecraft walls to determine the breakup characteristics of the resulting fragmentation. The walls included simulated components such as power electronics boxes and solar cells and were impacted by an aluminum sphere f rom a .22 caliber light gas gun [ 16] The results of the impact were separated into three types of breakups low intensity, high intensity, and hypervelocity explosions. From these types of explosions it was

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23 discovered that fragmentation of objects in the high intensity energy regi me fit a power law function fit [16] This power law function generated a high number of small objects (<1 cm) which matched well with observations from the debris generated in the hypervelocity impact tests [16] The use of a power law to describe the generation of fragment debris was consequently verified by multiple sources and used in subsequent early developments of breakup models [17] It was also found that the distribution of fragments followed a highly irregular shape definition a difficulty studied in multiple revisions of breakup models [17] EVOLVE The EVOLVE model was a quasi empirical based debris environment evolution program developed during the early s, with the original version released in 1986, by NASA Johnson Space Center with the primary goal of predicting the long term population, special density, and evolution of man made objects in the LEO environment. The EVOLVE model has been subject to at least three major updates since the original release in 1986 [1] The model provides one dimension assessments of the population by altitude bins between 200 km and 2000 km [ 18] The program, which is intended to only estimate the long term (i.e., years to several decades) population of debris population, runs a computationally intensive study that bridges the gap between observable objects larger than 1 cm in diameter a nd unobservable objects smaller than 1 cm in diameter. The structure of the program primarily utilizes as input [18] [12] : Traffic models Breakup model Propagation model Decay model

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24 Spatial density model Historical, current, and projected near term object spatial densities are populated analytically using up to date traffic modeling which is inclusive of mass, size, and orbital characteristics. Random generation of collision fragment populations are populated into individual sub populations based on statistically determined diameter size bins using the breakup model, while the aggregated spatial density is propagated at each time step for future orbital characteristics with considerations for atmosph eric drag, solar radiation pressure, and other perturbations using the propagation and decay models. All of the input model considerations are run through a deterministic process and looped in a Monte Carlo simulation at a user specified time step to simul ate the random statistical generation of debris fragment clouds [5] [19] EVOLVE has been used extensively as a leading tool for the analysis of the LEO population e nvironment since its creation; t he program has been used extensively internationally and was a primary input to the creation of recognized space safety policies, such as the NASA defined 25 year orbital decay requirement [2] LEGEND The LEO to GEO Environme nt Debris (LEGEND) model was created as a successor environment modeling to three dimensions account for spatial density considerations up through Geosynchronous Earth Orbit (GEO) at 4 0,000 km and modernize the computer code from the previous EVOLVE model [19] The new dimensi ons considered include altitude, latitude, and longitude. The model closely follows the previous evolutionary model, EVOLVE, by using a combined empirical and deterministic approach to predict future object populations

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25 using traffic, breakup, and environment perturbation models to characterize object densities. The LEGEND program, written in FORTRAN, uses Monte Carlo app roaches to simulate collision events over specified time intervals and can be used to predict long term population models over time periods as long as centuries. LEGEND is currently the environment modeling program used by NASA to characterize the spatial density of objects in Earth orbit. Hyperveloc ity Impact Tests In addition to observable data obtained from on orbit collision events, terrestrial hypervel ocity impact tests are a critical input to the development of breakup models used in evolutionary mode ls. Capabilities for hypervelocity testing varies significantly based on flight characteristics and many facilities exist both domestically and internationally to recreate the conditions of hypersonic flow and flight. For hypervelocity characteristics in t he space flight regime, ballistic gas based projectile facilities are the most commonly used to recreate on orbit flight conditions. Several major facilities with capabilities to recreate hypervelocity flight conditions include [20] : AEDC Range G in Tullahoma, TN White Sands Test Facility Hypervelocity Test Laboratory in White Sands, NM Holloman High Speed Test Track at Holloman Air Force Base, NM Langley Research Center Temperature Tunnel in Hampton, VA Previous Hypervelocity Impact Tests NASA has conducted numerous hypervelocity impact tests over the past 40 years to determine the effects of extremely high velocity collisions on orbital debris propagation [1] These tests, ranging from hypervelocit y speeds in LEO environments to low speed impacts in GEO environments, have spanned from simple component mockups to fully intact, flight unit spacecraft. The NASA Standard Breakup Model in

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26 addition to the previously discussed on orbit collision incidents is based primarily off the results of the Satellite Orbital Debris Characterization Impact Test (SOCIT) series of hypervelocity impact tests conducted in 1992 at the AEDC Range G test facility. SOCIT The SOCIT tests were conducted primarily by the Depa rtment of Defense Orbital Debris Spacecraft Breakup Modeling program and involved NASA Johnson Space Center NASA Jet Propulsion Lab, and Wright Laboratory, as well as several private contract companies. The lack of extensive experimental testing and the n eed for high fidelity satellite fragmentation data in the development of breakup models drove the creation of the SOCIT tests. The SOCIT series conducted four hypervelocity impact tests with the fourth and final test being the most complex and critical of the series [21] The tests were conducted on a simulated spacecraft panel, rocket body mounting ring, Navy Transit OSCAR mockup, and a fully developed flight ready Navy Transit OSCAR satellite. The main target, the flight ready but un flown Navy Transit OSCAR, was the most complex and representative object to be used in terrestrial breakup model testing. The OSCAR satellite was a n octagonal core with a diameter of 46 cm, height of 30 cm, and mass of 35 kg [22] and used as the main SOCIT target for its availability. The satellite contained no multilayer insulation and little, if any, modern materials such as CIC solar cells and composite ma terials. The testing was conducted in 1992 at the AEDC Range G facility The fourth and final test of the series, targeting the Navy Transit satellite, was conducted at a n impact velocity of approximately 6 .1 km/s with a flight trajectory perpendicular to the spacecraft.

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27 The projectile, a 150g aluminum sphere with a diameter of 4.7 cm, released an impact energy to target mass ratio of 78 J/g above the catastrophic breakup threshold of 40 J/g [17] The resulting debris field gen erated approximately 4500 fragment s of which approximately 100 were large pieces collected on the floor of the Range G test chamber [23] The remaining objects were collected in a catch system of foam paneling. A majority of t he fragments were removed from the foam by high pressure washing to destroy the foam and remaining debris from the SOCIT test collected through a grated sieve for analysis by private contractors. Several foam panels were left intact and sent to JPL for rad ar analysis to determine velocity and flight trajectory data. The results of the debris characterization produced experimental data for characteristic fragment length, mass, cross sectional area, and area to mass ratio for various fragment size bins. These results were incorporated into the statistical population functions used in the NASA Standard Breakup Model to improve the accuracy of fragment sizes in the sub centimeter regime and provide comparison models to tracking data from observable on orbit frag ments larger than 1 cm in diameter. Low Velocity Impact Tests Following the update to the NASA Standard Breakup Model in 2001, a series of new low velocity impact tests were initiated to study both the breakup of modern materials and the fragmentation of c haracteristics of objects at low speeds. These series of tests, led by Kyushu Institute of Technology, were conducted on simulated velocity averaged approximately 1.7 k m/s and created catastrophic breakups above the 40 J/g threshold that completely fragmented the test targets [24]

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28 These low velocity impact tests were utilized in verifying the breakup of collisions at lower velocities, such as those occurring at or near GEO altitudes [25] As well, the tests indicated the need for addition hypervelocity impact testing to identify and account for modern materials skewing the population distribution of objects in the Standard breakup Model [24] [25] Standard Breakup Model Improvements The results of the terrestrial SOCIT test, along with observation data from multiple on orbit collision events, make up th e primary input to the NASA Standard Breakup Model inconsistencies observed in the Feng Yun 1 C ASAT and the Iridium 33 and Cosmos 2251 accidental collision as well as outcomes of recent made with low velocity impa ct testing, has led to the conclusion to conduct a new terrestrial hypervelocity impact test to account for changes in material and design of LEO satellites since the SOCIT series tests in 1992. These inconsistencies and conclusion have been confirmed by r ecent hypervelocity impact testing on small and simple structures that show a clear discrepancy in experimental results for composite and modern material with predictions by the current breakup model. Research Focus The pu rpose of the research in this work is to describe the design of a LEO satellite that is representative of modern materials and construction for use in a hypervelocity impact test. The results of this proposed test can be used for the improvement of the NASA Standard Breakup Mod el and, subsequently, the output of long term population models which rely on it.

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29 CHAPTER 3 DEBRISAT : DESIGN Design Motivation Recent collision events, in particular the 2009 Iridium/Cosmos collision described in Chapter 1, have encouraged the examinatio n of how the NASA Standard Breakup Model predicts the resulting fragmentation of on orbit collisions when modern materials and construction methods are considered. When the current version of the breakup model was applied to the resulting fragmentation fro m the Iridium/Cosmos collision, the prediction results did not match the in s itu fragment observation results from the more modern Iridium 33 satellite. Because of this and other fragmentation and testing events there currently exists a need to reexamine the breakup model to account for the breakup characteristics of modern materials being utilized on recent and future low Earth orbit satellites. The current edition of the Standard Breakup Model relies ext ensively on the 1992 SOCIT hyper velocity impact ground test of a fully functional Navy transit using hardware, material, and design techniques from the 1960's. These results incorporated into the breakup model are not representative of the materials and co nstruction techniques used in modern sat ellites. Therefore, a new hyper velocity ground test of a modern satellite is required to update the breakup model. The University of Florida is developing a 50 kg satellite that is representative of multiple modern L EO satellites. The satellite, referred to as DebriSat, will be composed of subsystems, components, and materials typical of modern LEO satellites. Since DebriSat is designed specifically for a hypervelocity ground impact test, its avionics (i.e., electrica l and software capabilities) will be non functional. However, every other aspect

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30 of the satellite will adhere to the typical standards of modern spacecraft design, fa brication, and assembly integration and testing (AI&T). The DebriSat design is intended to represent the overall system makeup and physical characteristics of unmanned LEO missions ranging from 1 5000 kg in mass. The completed system contains subsystems and components that are not typically found on a satellite of DebriSat's size but are found on satellites in other target mass ranges. Thus the design is not intended to be a functional representation of a 50 kg satellite but rather seeks to be as representative of component material and physical characteristics as possible across a broad range o f satellite platforms. This thesis specifically addresses the design considerations of the Electrical Power System (EPS), Telemetry Tracking and Command system (TT&C) the Command & Data Handling (C&DH) system, and the Payload system. Figure 3 1 shows an e xternal view of the DebriSat design model in a fully deployed configuration. Figure 3 1 External View of DebriSat

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31 System Level Design The DebriSat design is a non functioning satellite that emulates the materials and functionalities of modern LEO space missions. Basic system char acteristics are given in Table 3 1. Table 3 1 Top l evel debrisat c haracteristics Project Title: DebriSat Target Mass: 50 kg Physical Envelope (including deployed and protruding components): 76 cm (dia.) x 68 cm (ht.) Stabilization: 3 axis (non functional) Deployables: Yes (partially deployed solar panels) As shown in Figure 3 2 DebriSat is composed of a hexagonal prism structure containing six compartmentalized bays and a seventh cylindrical bay about the central axis. Two additional hexagonal panels are included and intended to serve as the nadir and zenith facing structural elements. Components are both panel and base mounted in addition to mounting on six internal honeycomb structural ribs. Figure 3 2 DebriSat bay definitions

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32 Each of the DebriSat side panels is constructed of aluminum honeycomb core sandwiched between two carbon composite face sheets. Individual components are mounted to these panels primarily through fastened inserts into the honeycomb core. In addition to the six side panel structures, six composite panel honeycomb core rib sections are included for structural support within the volume of DebriSat. Viewing from the zenith panel to the nadir hexagonal panel, the bay count begins with th e first Li ion battery box and proceeds clockwise through panel B and the propulsion module to bay 6 containing the power conditioning and distribution unit and spectrometer. Bay 7 contains the propulsion system and various other components mounted to the hexagonal panels and structural ribs. Mounting is primarily done on the six body side panels. A detailed overview of t his layout is given in Figure 3 3. In addition to the components mounted to the side paneling, several components are located and fa stened within the central volume of the structure. These components are considered included within bay 7. Mounting to the structure is done on the hexagonal base panels or the structural ribs to achieve either representing a component's most likely locatio n or for physical space considerations. Also included in bay 7 is the electrical distribution cabling and plumbing for the satellite's propulsion system. Cabling is run to each of the components and routed along the internal panel face sheet in harnessed t runk lines. These trunk lines are then routed around the major internal components and around the structural ribs to connect to the appropriate electrical connector box or flight computer avionics box. The routing of the propulsion tubing is completed in a similar manner to connect the thruster and solenoid components with the COPV pressure tank.

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33 Figure 3 3: Component Layout

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34 LEO Satellite Study A study of current and recent LEO unmanned missions was performed to determine typical components and characteristics of satellites that operate in the LEO environment This study, conducted at the University of Florida was used to drive the selection and design of components for DebriSat and is summarized here The base of the study is bui lt from a comprehensive database of currently active space based satellite missions maintained by the Union of Concerned Scientists [26] The comprehensive database, which includes all missions from LEO to Geosynchronous Earth Or bit (GEO), was parsed to identify only active unmanned LEO missions and reorganized by dry mass to determine an identifiable and quantifiable distribution of LEO missions. This distribution of 467 unmanned LEO satellite missions was then scaled to a select ion of 50 modern missions such that the ratio of number of satellites in any two mass ranges remains sufficiently constant. The distribution was then analyzed for physical characteristics including component selection, materials, and mission details. Figu re s 3 4 and 3 5 show the distribution of the satellites used in the study. Figure 3 4 Mass distribution r ange of 467 satellites in UCS database

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35 Figure 3 5 Mass Distribution Range of 50 s elected satellites for UF study The preliminary results of the s tudy concluded that there is correlation between the components that are implemented in a particular LEO mission and the dry mass of that mission. Therefore, the design of DebriSat was driven primarily by the components deemed either: standard among all ma ss ranges, significant to two or more mass ranges, a new design standard since 1992, trending towards increased use in the future Additional results from the study are outlined by Clark, Lane, and Strickland and the accompanying study results [27] The survey also was found to be deficient in representation of extremely detailed designs and of non U.S. or European backed missions (specifically Russia and China) due to the restrictions and limitations of informatio n available in the public domain. Therefore, it is important to the hypervelocity impact test and outcomes of breakup model improvements to note that DebriSat may be considered deficient in representing Russian and Chinese satellites. The 10 100 kg, 100 50 0 kg, 500 1000 kg, and 1000 2000 kg categories were the primary focus of this study because they held the majority of the LEO satellites studied.

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36 Analysis of the Attitude Determination and Control System (ADCS) components, as shown in Figure 3 6 revealed sun sensors and magnetometers were prevalent in the majority of all LEO satellites except the 2000 5000 kg range. While star trackers were not used on any satellites in the 10 100 kg category, they were heavily used in the thre e other mass ranges. Therefo re star trackers were considered representative components for LEO satellites A trend of increasing gyroscope use was also seen as the satellite mass increased. However, since this was a study created from public information, the terminology from one sate llite to the other might have been inconsistent. For example, the "gyroscope" and "inertial measurement unit (IMU)" terminology might have been used interchangeably. Regardless, due to the prevalence of both, an IMU will be included in the design. Figu re 3 6 ADCS Sensor U sage by Mass As seen in Figure 3 7 reaction wheels and magnetorquers were the actuators most used by larger satellites. Only smaller satellites used passive actuators, such as passive magnetic and reflection strips, therefore these wi ll not be included in the design.

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37 To represent actuators typically used in a LEO satellite, reaction wheels and magnetorquers were incorporated into the ADCS design, as well. Figure 3 7 ADCS Actuator Usage by Mass Figure 3 8 illustrates the results of an analysis of propulsion system. Propulsion systems are primarily used on larger satellites, particularly the 1000 2000 kg group. Note that for small satellites propulsion systems are relatively rare, including for DebriSat's mass category, 10 100 kg. Whi le uncommon on a satellite of DebriSat's size, a propulsion system is necessary to accurately represent typical LEO satellites and therefore was included in the design. Figure 3 8 Propulsion System Usage by Mass

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38 The study of representative telemetry, t racking, and command (TT&C) components was driven by the communication bands used on each of the selected rep resentative missions. Figure 3 9 shows the percent usage of the most commonly identified communication bands within each mass range. The results sh ow that S band frequencies are the most common between all mass ranges, while UHF and VHF bands are prevalent to smaller mass missions and X bands are used commonly throughout larger mass missions. Based on these results, a representative satellite would c onsider including TT&C components that operate in VHF, UHF, S and X band frequencies. Figure 3 9 LEO Communication Bands by Mass Range Design choices for the E lectrical Power System (EPS) were driven by the primary chemical composition of battery cells. A survey of common used batteries in LEO s atellites is shown in Figure 3 10 Li ion is most prevalent and has increased usage among smaller satellites, dominating the 1 10 kg range. Nickel hydro gen and nickel cadmium batteries have increasing usage in the larger mass ranges, with nickel hydrogen dominating usage in the 2000 5000 kg range. Li ion, while already in use on a number of LEO satellites, is also expected to become more common in future LEO satellites. Since Li ion is expected to be used more frequently in the future and because

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39 it is an increasingly common component in LEO satellites, Li ion was selected for use in the DebriSat design. Figure 3 10 LEO Batteries by Mass Range The payload survey classified each selected missions by their primary objectives: Earth observing, remote sensing, communication, and technology demonstration. The survey reveals that communication payloads and Earth observing are the most prevalent class of LEO payloads. Communication payloads were identified as common constellation payloads specifically that one common design is included over several dozen satellites, such as the Iridium and Orbcomm constellations. Earth observing missions typically use imagers that operate in either the near infrared or visible wavelength spectrum. In addition, the number of payload instruments per mission was typically greater than one. Thus, multiple payload instruments should be considered for a representative design The complexity and uniqueness of one mission payload compared to another also led to the conclusion that a truly representative payload contained in a single satellite is very difficult to identify. DebriSat Subsystems The DebriSat d e sign is broken into typical subsystems found on LEO satellite missions. While all of these standard subsystems are considered in the design of

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40 DebriSat, the scope of this thesis focuses specifically on the design of the Electrical Power Subsystem (EPS), Command and Data Handl ing (C&DH), Telemetry Tracking and Command (TT&C), and the Payload subsystems. In depth discussion of additional DebriSat subsystems, specifically propulsion, thermal, and structures, are discussed in the thesis work of Mark Werremeyer [28] Electrical Power System The EPS is a critical subsystem of all spacecraft and the challenges of the LEO environment drive the requirements for typical EPS design s. For DebriSat, the EPS design represents typical material and characteristics of power generation, power storage, and power distribution onboard unmanned LEO satellites Typically, all LEO unmanned spacecraft power generation designs are dominated by the inclusion of photovoltaic cells. These configurations include both deployable and body mounted cell designs. Traditionally for large satellite platforms, a deployable solar array mechanism is used to increase the power generating surface area of the satellite while smaller satellites utilize body mounted solar arrays to generate pow er While a majority of satellites utilize a deployable solar array configuration, DebriSat is limited to a partially deployed configuration with one solar panel deployed and another stowed due to the physical limits of the hypervelocity test facility. The partially deployed configuration will however, provide insight into how different deployed configurations might affect satellite breakups. Of particular conce rn to the research goals of the DebriSat test is the presence of modern cell Interconnect coverglass (CIC) solar cells in orbital breakups. The material utilized in CIC cells includes the completed photovoltaic cell, a cover glass material, a n d an interconnect that prevents the damage of cells from high energy electrons.

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41 Therefore, the power ge neration of DebriSat will be represented by the inclusion of CIC solar cells. The selected cells are 28.3% efficiency Ultra Triple Junction (UTJ) solar cells from Spectrolab, shown in Figure 3 11. Figure 3 11 Spectrolab 28.3% UTJ CIC Solar Cells These c onstituent materials used in the UTJ CIC cells are consistent with those found in modern CIC cells. The CIC's are constructed of layered photovoltaic cells and substrates composed primarily of gallium and germanium materials. Overall, the UTJ cell design f rom Spectrolab has logged over 2.6 million delivered units and thus is considered very representative of current LEO power generation designs. A data reference sheet for the UTJ cells is included in Appendix A. The CIC cells which are nonfunctional engin eering development units, are mounted to each of three composite solar panels, spaced evenly in the two planar directions. Each panel has 49 individual CIC cells mounted with a silicon elastomer t o the carbon fiber face sheet of each panel. This configurat ion provides for a n approximate total of 1300 cm 2 of theoretical power generating surface area per panel. Under a standard, beginning of life scenario in the LEO environment, t his surface area correlates to a theoretical power generation capability of appr oximately 50 W per pane l

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42 in direct sunlight if the cells and power subsystem were functional. Figure 3 12 presents the partially deployed solar panel configuration. Figure 3 12 Deployed Solar Panels To represent power storage capabilities, DebriSat cont ains three separate shielded battery racks to emulate the power storage capabilities found on LEO satellites. Historically, Nickel Cadmium (Ni Cd) batteries have been an industry standard and have a long flight heritage for energy storage in space. Over time this trend began to include Nickel Hydrogen based cells and most recently lithium ion (Li ion) based chemistry F ollowing the conclusion of the last comprehensive Standard Bre akup Model test in 1992, Li ion technology is quickly becoming a new standard for its beneficial charge c apacity, charge/discharge cycle performance and energy density. Thus to represent this modern trend in LEO satellites, DebriSat contains three Li ion based rechargeable battery cases in lieu of the more traditional Ni Cd design This

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43 approach is representative of both the battery material and the use of multiple cells for operational redundancy. Electrically, the batteries cases are non functional and consist of multiple prismatic cells grouped to create a string that represents the power storage capabilities typical of a 50 kg platform. A transparent view of the Li ion battery box without wiring is shown i n Figure 3 13 Figure 3 13 Transparent View of Li ion Battery Redesign Each individual prismatic cell case consists of wrapped layers of unpopulated aluminum, copper, and polyethylene foil all of which are typical mat e rials used as t he anode, cathode, and separator, respectfully, in functional battery designs. Each Li ion battery case consists of eight individual prismatic cells made of 3003 Aluminum and are electrically connected in a series and parallel combination. The cells containing the layers of copper, aluminum, and polyet hylene foils are inert and wrapped without the standard active chemical material s found in Li ion batteries. The use of inert materials allows for a reduction in health and safety hazards to personnel handling the resulting fragments following the breakup without significantly affecting the outcomes of the

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44 breakup Each of the eight aluminum cell cases are mounted to an aluminum mounting rack that is connected by a supporting aluminum truss, as shown in Figure 3 14. Figure 3 14 Li ion Cell Assembly Eac h battery is wired into a single shielded cable and that represents unregulated power run through a D sub connector to a battery charge regulator (BCR) in the power conditioning and distribution module is contained in a sh ielded 5 mm thick aluminum case fastened directly to the representative radiator panel of DebriSat. Inside, the completed battery cell rack is mounted to the bottom face of the battery avionics box. The battery box designs are shown in Figure 3 15. Figur e 3 15 Li ion Battery Case Dimensions

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45 The conditioning and distribution of power shown in detail in Figure 3 16 and Figure 3 17 is represented using a single integrated avionics module containing a set of emulated solid state power circuits that regulat es the incoming power generated by the solar panels to the battery racks and main power bus line and is representative of all the design's power circuitry, power conditioning, and power distribution electronics. Figure 3 16 Power Conditioning and Distribution Module Model The integrated design includes printed circuitry representative of a peak power tracker (PPT) that typically changes the operating point of the solar cell arrays to draw the correct power needed by the sys tem during operation. A set of battery charge regulators are also represented by commercial off the shelf ( COTS ) circuitry to include considerations of preventing overcharging and deep discharges of the battery cells to increase battery life and prevent E PS damage. The battery cells, solar cell arrays, BCR's, and PPT emulated circuitry are all connected to a set of emulated power regulation and distribution circuits contained in the shielded avionics box and routed out as regulated and conditioned power to the various onboard components. The power

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46 distribution module is the primary receiver and distributor of power in the DebriSat design. Each individual string of solar cell arrays is connected directly to the module as input where internal wiring routes p ower to the battery cells and main bus line. All power routing is accomplished using jacket shielded cable wiring and D sub connectors. Figure 3 17 Power Conditioning and Distribution Module Dimensions Wiring of power throughout the DebriSa t design is completed using shielded cabling and various sizes of D sub connect ors. Teflon coated space qualified copper wiring of varying gauge is soldered to the pins of the connector s and wrapped in br aided stainless steel jacket sleeves. A typical exp ected DebriSat cab le is shown below in Figure 3 18 T he exact length and mass of the cabling and harnesses is currently not accurately determinable. However, the mass of the combined wire, shield, and connectors is not insignificant. To estimate the additi onal mass to the overall design from electrical wiring, a combination of measurements and assumptions were made: The combined wire and shield contributes approximately 1.1 g/cm of cable The fully assembled D sub connector (pin sockets and casing) contribut es approximately 7.2 g per connection

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47 Cable lines existing in the design include both electrical power and data transfer Based on these assumptions, the mass of the cabling in DebriSat was determined using a mass tolerance of the overall EPS subsystem. The additional cable mass is bounded as +30% of the EPS subsystem mass. Figure 3 18 Braided Stainless Steel Shield and D Sub Connector Overall, the general conceptual wiring and connection of the EPS components are shown in Figure 3 19. Figure 3 19 Conc eptual EPS Component Connections The CIC solar cells are bonded to the external face sheet of the body panels and wired on the internal face sheet in seven series connected arrays per panel. Each array is wired into a single shielded panel cable line and c onnected to the power conditioning and distribution box. The battery cases are body panel mounted and distributed evenly

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48 120 degrees to balance distribution of mass. In addition, the battery cases are mounted in contact with the zenith facing hexagonal pan el for heat dissipation to the radiator. The power conditioning and distribution module is also body panel mounted and is located near the zenith hexagonal panel for heat dissipation. In summary, the EPS subsystem is a critical system of interest to the o verall research effort, particularly due to the growing inclusion of both the CIC solar cells and Li ion battery materials in modern designs. The mass breakdown of the EPS components is given in Table 3 2 Table 3 2 EPS m ass Component Mass (kg) Quantity Mass Subtotal (kg) Contingency (kg) CIC Solar Cell 0.0022 147 0.32 0.03 Li ion Battery Box 3.06 3 9.18 1.00 Power Distribution Module 1.81 1 1.81 0.15 Cabling (~ 0.30 ) 5.07 N/A 4.85 0.50 Telemetry, Tracking, and Command The DebriSat TT&C sub system consists of a variety of antenna designs and communication b ands as identified and selected according to the LEO satellite study. The most typical communication bands used in unmanned LEO applications for communications include UHF, VHF, S, and X band frequencies The use of these common bands were found to be prevalent across most of the mass ranges considered for driving the design selections for DebriSat and are therefore included as the communi cation bands used for the TT&C subsystem. Figure 3 20 shows a conceptual connection chart of the DebriSat TT&C components.

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49 Figure 3 20 Conceptual TT&C Connection Diagram Primary telemetry and data relay from data intensive component s and the payload instruments are represented using an emulated open feed horn X band antenna shown in Figure 3 21. The theoretical radio frequency (RF) beam of the X band antenna is high gain and narrow width to represent the capabilities of higher data delivery rates to target ground stations. Figure 3 21 X Band Ant enna Design The operational frequency of the antenna is intended to be in the 8.0 to 8.4 GHz range and thus the emulated antenna has a copper wire core of approximate ly 3.1 cm in

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50 length to match the theoretical wavelength expectations The transmission wire is connected by a standard RF coaxial adapter to the TT&C avionics box. The X band antenna is a transmit only communication node and does not represent the capability of simultaneous transmitting and receiving I n addition to the X ba nd feed horn antenna, an S band frequency conical antenna is also included to represent LEO missions that are inclusive of S band communication. The S band antenna design, as shown in Figure 3 22 represents the capability f or transmitting primary telemetr y and health monitoring data from various subsystems and for use a s a secondary payload communication an tenna. The S band antenna use s a copper helical shaped core housed in a converging coni cal beam cone that represents transmission at an operational freq uency between 2.5 and 3.5 GHz. Figure 3 22 Helical S Band Antenna Design The helical wire is mounted to the antenna housing and is connected to a standard male RF adapter at the base of the emulated design to route RF coaxial cable

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51 to the S ba nd tr ansceiver avionics module. The theoretical beam footprint is conical shaped and provides a wide beam diameter for moderate ly large gro und coverage capabilities. The S band antenna as designed for DebriSat is theoretically capable of both transmitting and r eceiving operations and can be used as a redundant A third communication band is included in the TT&C design as a VHF/UHF freque ncy communication system and is implemented using two omni directional antennas shown in Figure 3 23 The VHF/UHF antenna design represents both the primary communication designs of satellites in the lowest mass ranges and the redundancy provided to larger LEO satellites. Each of the three antenna designs are connected via RF coaxial cable to one of two shie lded telemetry avionics modules one for the S band and X band antennas and the other for the UHF/VHF antennas Each TT&C avionics module include s a set of communication circuitry boards and RF signal manipulators. Incoming command signal s are passed through a signal divider to the nonfunctional communication circuitry before being routed through a splitter to the intended transmitting /receiving antenna. Figure 3 23 Digikey Omni Directional Whip Antenna

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52 The TT&C avionics modules are single all inclusive shielded cased milled from 6000 series aluminum and contain both emulated RF circuitry and a COTS signal switch and divider. Each of the two avionics boxes are given a wall thickn ess of 3mm to represent sufficient shielding from typical LEO radiation. The avionics box design is shown in detail in Figure 3 24 and Figure 3 25 Figure 3 24 S band TT&C Avionics Model Figure 3 25 S Band Transceiver Dimensions

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53 In addition to communication circuitry and control boards, physical signal manipulators were included into the TT&C avionics design. These components include communi cation signals to a desired antenna. The COTS signal manipulators used in DebriSat are shown i n Figure 3 26 Figure 3 26 COTS Signal Manipulation Components For the selection of components to meet the desired design goals of the TT&C system, emp hasis for communication components is put on the geometry of the antennas and their corresponding RF elements. Table 3 3 details the primary components used in the TT&C subsystem. A finite list of antenna shapes and bands was determined and used to drive t he selection of emulation designs. Tabl e 3 3 List of TT&C c omponents Component Supplier Quantity S Band Helical Antenna Manufactured 1 X Band Conical Feed Horn Antenna Manufactured 1 VHF/UHF Omni Directional Antenna Digi Key 2 Shielded Aluminum Case Manufactured 2 MHX 910 Wired Communication Board MicroHard 4 Signal Divider/Combiner Mecca 2 Signal Switch DowKey 2 Male RF Connector Digi Key 8 emulated health monitoring, telemetry, commands, and data transfer in one of the four

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54 selected communication bands. The estimated TT&C subsystem mass is 2.77 kg Table 3 4 lists the estimated mass of the TT&C components. Table 3 4 TT&C mass b reakdown Component Mass (kg) Quantity Mass Subtotal (kg) Contingency (kg) X band Horn Antenna Manufactured 0.12 1 0.12 S band Helical Antenna Manufactured 0.15 1 0.15 UHF/VHF Omni directional Digi Key 0.15 2 0.30 Telemetry Avionics Module Manufactured 0.12 1 0.12 Command and Data Handling The control and processing of telemetry and command data is handled by an emulated integrated flight computer system and acco mpanying data recorder memory module. The flight computer is a fully integrated electronic component mounted to the satellite in a shielded aluminum avionics case. The computer is to provide onboard processing power and command protocol interfacing between components The flight computer design is based on an emulated shielded avionics case as shown in Figure 3 27 Shielding is accomplished by design of the avionics box wall thickness to 7mm of aluminum representative of the importance of shielding the critical C&DH electrical components from radiation. The flight computer avionics case represents the processing of command and telemetry data by the inclusion of standard COTS computer circuit components mounted within the case The use of COTS computer components represents the m aterials used in complex circuitry found in on board satellite flight computer s. In particular, a computer circuit component was selected as the representative component because of its variability in available sizes,

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55 extensive use of standard circui t r y com ponents such as t ransistors and micro controllers, its sufficient complexity of design, and its affordability. The emulated flight computer routes command and data handling to a majority of components included in the DebriSat design. The computer will co nnect the internal circuitry to multiple D sub type connectors (final D Sub number not shown in Figure ) and routed throughout the physical satellite l ayout. Figure 3 27 DebriSat flight computer Flight d ata recording is accomplished via an emulated flight recorder and memory unit s hown in Figure 3 28 T he data recorder like the flight computer, utilizes standard computer circuit component as its primary electrical component s The recorder is connected via D sub connectors directly to the flight computer. A shielded aluminum avionics case with 7 mm thickness is included to represent protec tion of the internal electronics from radiation.

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56 Figure 3 28 C&DH data recorder Both the flight computer and data recorder are considered thermally sensitive components both generating excessive heat and requiring controlled thermal conditions Therefore, the placement of the C&DH components is consistent with typical thermal control techniques Both units are mounted to the internal face sheet of a composite side panel containing thermal heat piping. This placement represents considerations to control and dissipate the heat generated by the internal electronic s. Thermal designs for DebriSat are discussed in further detail in research work by Mark Werremeyer. In summary, t he C&DH system design is predicated on the assumption that LEO spacecraft flight computer and data recording electronics can be successfully r epresented and emulated using stan dard computer electronics components Also of note the effects of advanced radiation hardening techniques, such as complex material liners and other radiation reducing methods are not accounted for in the design and is currently unknown if such considerations would significantly affect the material

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57 properties or size of the C&DH components following fragmentation from a hypervelocity impact. These considerations may be of intere st to future work of emulating C&DH electrical components. A list of the primary components used in the C&DH subsystem is shown in Table 3 5 and the mass breakdown of the C&DH components is presented in Table 3 6. Table 3 5 Primary C&DH c omponents Compone nt Supplier Quantity Flight Computer Avionics Box Manufactured 1 Data Recorder Avionics Box Manufactured 1 Computer Motherboard Various 3 D Sub Connector VARTA 12+ Table 3 6 C&DH component m asses Component Mass (kg) Quantity Mass Subtotal (kg) Contingency(kg) Flight Computer Avionics 1.0 1 1.0 0.1 Data Recorder Avionics 0.90 1 0.90 0.01 Total 1.90 0.11 Payload The DebriSat payload is a combination of three instruments that represent non communication payloads on LEO missions a visible spectrum optical imager and two identical near infrared spectrometers. The exclusion of communication payloads is motivated by the results of the LEO satellite study because w hile a representative satellite may include a communication payload, it w as determined that the components us ed in the TT&C and other subsystems were sufficiently representative of the materials found in communications payloads. However, Earth observing missions were not represented by any of the components included in other su bsystems. Therefore, two different optical instrument designs ar e included in the payload subsystem

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58 The primary payload for the DebriSat design is a n emulated visible wavelength optical imager. The design goal of the optical imager is to focus on the typical geometrical shape, physical size characteristic s and typical material of an imaging payload. The core of the design is a COTS Cassegrain telescope selecte d as an effective representation of larger space telescopes and is modified with additional components to represent a space application imager The COTS reflector directs incoming light rays from a back plane mirror to a CCD camera system added to backplan e of the telescope. The imaging internals are protected from stray light by a custom circular aluminum sunshade that extends out past the reflector telescope. An integrated avionics module, whi ch houses the CCD camera and all accompanying circuitry and cab ling, is fastened behind the backplane of the telescope. The assembl ed imager is then fastened to an aluminum cylindrical mounting case that provides increased stiffness to isolate the imaging components from random vibration In addition to the stiffening structure, a set of two dedicated reaction wheels are attached to the avionics module to include considerations for active momentum based e lements representative of precise jitter elimination To accomplish the design goals of the emulated optical imager a combination of manufactured and commercial components is selected to represent an imager that captures incident light in the visible spectrum. A commercial telescope model, the Celestron NexStar shown in Figure 3 29 is used as the core Cassegrain mode l for its material and size properties.

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59 Figure 3 29 Celestron NexStar 4SE Maksutov Cassegrain Telescope The telescope is stripped of its accompanying base mount as well as the externally attached optical lens. The resulting component is a 25 cm cylindr ical aluminum Cassegrain telescope containing a plano convex reflecting mirror and an angled reflecting focusing mirror. Figure 3 30 shows internal layout of the telescope with a modified support for the axial beam reflector. The modification is of typica l des igns found in space applications. Figure 3 30 Internal View of Modified Celestron Telescope A custom aluminum sunshade is included around the telescope to represent the imager. This manufactured sunshade is fastened to the telescope by a multi purpose adapter to a payload avionics box, shown in Figure 3 31, which mounts the telescope, sunshade, and attached payload avionics module.

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60 Figure 3 31 Optical Imager Avionics Box Attached in the payload avionics module is a focal lens to emulate the manipulation of light prior to collection by the CCD camera, which is represented with the internal circuitry of a COTS digital camera. The optical piece contains a combination of t hree plano lenses Several printed circuit cards are included to support the CCD camera and circuitry located just behind the focal lens. The entire electroni cs assembly is mounted inside an aluminum avionics box that is mounted to the multi purpose adapte r. The fully constructed optical imaging payload assembly is mounted coincident to the central body axis of the satellite by a custom circular adapter ring. The ring allows the assembly to protrude approximately 80 mm from the nadir hexagonal base panel.

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61 Figure 3 32 shows the completed representative imager design and an exploded view is shown in Figure 3 33 Table 3 7 outlines the main components of the optical imager. Not shown in the model is the included thermal multi layer insulation that will cover t he payload avionics box and the telescope to control and regulate the temperature of the assembly. Figure 3 32 Optical Imager Payload

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62 Figure 3 33 Exploded View of Optical Imager

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63 Table 3 7 O ptical imager mass characteristics Component Mass (kg) Qua ntity Mass Subtotal (kg) Payload Case 0.395 1 0.39 Cassegrain Telescope Assembly 1.32 1 1.32 Sunshade 0.31 1 0.31 Electronic Circuitry 0.1 N/A 0.1 Jitter Damper 0.15 2 0.30 Stiffening Case 1.67 1 1.67 TOTAL --4.09 In addition to the optical elements, the imager contains dedicated structural components to represent rigidity and motion control during image capture. A custom one quarter inch thick aluminum circular structural housing surrounds the optical imager design and is fastener mounted to the nadir hexagonal base plate of the satellite structure. The structure is intended to be representative of structural elements that reduce vibration both during launch and standard operation. In addition, a set of jitter react ion wheels are mounted to the payload avionics where each reaction wheel is a nonfunctional unit manufactured to include the exterior case, flywheel, and axis rod. While the overall optical payload design is non functional, the materials and assembly ori entation must be as representative as possible while eliminating constraining complexities in design. In particular, the most significant design assumption is the unimportance of the overall path of incident light rays over the identification and inclusion of optical and structural materials and thus the elimination of stringent tolerance and ray tracing considerations in the design. Also of note to the representatives of the payload system is t he use of cr yogenic cooling hardware is not included in the des ign. An emulated near infrared spectral imager shown in Figure 3 34, is included twice as a payload for the DebriSat design in addition to the optical imager The

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64 inclusion of the spectrometer represents missions that contain hardware and instrumentation used in Earth observation and remote sensing. The design attempts to capture the typical materials and thicknesses used for internal spectrometer components. Figure 3 34 Near Infrared Spectrometer Each of the two spectrometers is contained separately in a structural housing composed of an aluminum honeycomb core with two carbon fiber composite face sheet panels and is designed for increa sed stiffness consideration serving as the mounting location for each of the optical components. A baffle casing directs incoming light to a focusing optical lens before being routed by a 90 degree fold mirror and 45 degree bend mirror to a second focusing optical lens and CCD camera. The came ra system and accompanying dedicated image processing circuitry is located in a stiffened electronics

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65 housing that captures image information and runs data and commands to and from the spectrometer and other satellite subsystems through data cable assembli es A secondary set of static data recorder memory is panel mounted separate from the optical bench. An internal view of the assembly is show in Figure 3 35. Figure 3 35 Internal view of the near infrared spectrometer The component selection of each of the two identical spectrometers in the DebriSat design consists of completely emulated structural components and COTS optical elements. Each spectrometer is mounted to the nadir hexagonal base panel with the protruding baffle casing theoretically orientat ed towards the Earth.

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66 The baffle is made of a cylindrical Delrin tube and includes two planar focusing lenses mounted internally by an aluminum baffle casing. Internally, light ray diffraction is emulated by including a commercial 90 degree optical lens, a coated reflecting mirror, and a diffraction lens. An electronics module houses a CCD camera circuit card as well as standard computer memory that serve as emulated electronics components. As with the optical imager, the inclusion of particular component materials is of greater importance then the functional performance of the spectrometer and the design of the incident light ray paths. Emphasis is intended to be on the materials included and the general layout of internal components. For jitter control, a set of two reaction wheels are also included on the exterior of the aluminum housing Figure 3 36 presents detailed dimensions of the spectrometer external assembly. Figure 3 36 Detailed Dimensions of Spectrometer Assembly

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67 An emulated optical bench and mounting interface is included in each of the two spectrometers to provide consideration for the importance of vibrational stiffness in optical design s The bench is a three layer panel consisting of two carbon fiber composite face sheets made o f 1 mm M46 J carbon fiber and an internal aluminum 6063 honeycomb core a design similar to the structural panels of the DebriSat exterior The optical bench, shown in Figure 3 37, is mounted to the flat side of the open structural box of the spectrometer. All major optical elements are directly mounted to the optical bench using mounting inserts that are placed within the sandwiched panel. Figure 3 37 Exploded View of Spectrometer Optical Bench The incident light en ters the spectrometer through the circular baffle housing opening before coming into contact with a 3 mm plano convex lens made of N BK7 optical glass degree fold mirror that is emul ated using a standard COTS telescope prismatic fold lens. The folding mirror is contained by a custom manufactured aluminum casing to

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68 emulate structural and vibrational support for the mirror and is show as an exploded view in Figure 3 38. Figure 3 38 E xploded View of 90 Degree Fold Mirror and Casing Another primary emulated component in the spectrometer is the 3.8 mm uncoated convex lens mount that focuses the light ray just prior to entry into the CCD camera, shown in an exploded view in Figure 3 39 W hile titanium is no longer used in modern LEO structural elements, the material is still used as structural support in optical payload components. Thus, the primary structural element for the lens is a mounting bracket made of titanium to represent the hig h stiffn ess requirements demanded by standard LEO imaging payloads. For budgetary purposes, each spectrometer contains only one titanium mount. The 3.8 mm convex lens made of N BK7 optical glass rests in the mount and is held securely in place by an alumin um mounting face.

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69 Figure 3 39 Exploded View of Titanium Lens Mount The CCD camera is represented using the internal electronic components of a digital camera that are mounted into the holding braces of the aluminum electronics housing. Additional CO TS printed circuit boards are in included in the CCD housing to represent integrated data processing capabilities. The webbed housing in Figure 3 40 is mounted directly to the optical bench and has ports for D sub type connector cables (not shown) to be ro uted out of the spectrometer and to the payload support module. Figure 3 40 Spectrometer CCD Housing

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70 In addition to the CCD camera housing, three boards of standard computer RAM are included to emulate secondary internal payload data recording capabilit ies. Each board has a Wedge Lock attachment component and is mounted into an aluminum mounting frame on the spectrometer side panel to account for the increased stiffness requirements in the payload. Table 3 8 presents the mass breakdown of the main spect rometer payload components. Table 3 8 Spectrometer mass c haracteristics Component Mass (kg) Quantity Mass Subtotal (kg) Spectrometer Case 0.67 1 0.67 Spectrometer Baffle Assembly 0.13 1 0.13 CCD Housing 0.14 1 0.14 45 Degree Mirror Assembly 0.10 1 0.10 Titanium Mount Assembly 0.08 1 0.08 90 Degree Fold Mirror Assembly 0.53 1 0.53 Optical Bench Assembly 0.23 1 0.23 Memory Assembly 0.06 1 0.06 Jitter Damper 0.15 1 0.30 TOTAL --2.24 In addition to the dedicated internal circuitry and image processing components in each of the payload assemblies, an additional payload support module is included in the design of the DebriSat payload subsystem. The support module consist s of electronic c ircuitry shielde d in a dedicated, shielded avionics case that represents the capability for secondary on board post processing and data formatting before packaging and transmission to ground stations via the TT&C subsystem. The support module also routes C &DH data and electrical power to each of the three instruments and is mounted to the internal face sheets of the composite structure panel. Table 3 9 shows the primary components used in the DebriSat payloads.

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71 Table 3 9 Payload c omponents Component Supplier Quantity Cassegrain Telescope Celestron 1 Aluminum Sunshade Manufactured 1 Mounting Ring Manufactured 1 Avionics Adapter Ring Manufactured 1 Printed Circuitry Manufactured 3 CCD Camera Sparkfun 3 Optical Avionics Lens Mount Manufactured 1 Optical Lens Thor Labs 15 Jitter Flywheel Manufactured 6 Titanium Lens Mount Manufactured 2 CCD Housing Manufactured 2 Fold Mirror Casing Manufactured 2 Optical Stiffening Case Manufactured 1 Avionics Case Manufactured 1 Spectrometer Case Manufactured 2 Spectrometer Baffle Manufactured 2 Computer Memory Sparkfun 6 Payload Support Module Manufactured 1 Computer Motherboard ATI 10 D sub connectors Alvatek 15 The support module is designed to be a command and communication unit between all payload components. The box is mounted directly to an aluminum insert in the interior carbon fiber face sheet by twelve 5 mm mounting tabs and contains six 15 pin D sub type connectors that run to the payload components, power distribution module, and the flight computer. The module is a manufactured aluminum box with ten circuitry racks integrated into t he interior wall to hold ten COTS circuit board components. The support avionics design is to represent o n board computing, processing, command and control, and communication between the payload hardware and the other critical systems within DebriSat. The payload support module is shown below in Figure 3 41.

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72 Figure 3 41 Payload Support Module In summary the payload contains non functional emulated components that aim to be as representative as possible of material characteristics, shape, and size versus functional imaging capabilities. The design makes extensive use of a luminum components as well as titanium, optical glass, and printed electronics circuitry that Table 3 10 presents the mass of the DebriSat payload subsystem. Table 3 10 Total payload mass c haracteristics Component Mass (kg) Quantity Mass Subtotal (kg) Con tingency (kg) Optical Imager Assembly 4.09 1 4.1 0.40 Spectrometer Assembly 2.24 2 4.5 0.45 Payload Support Module 3.10 1 3.1 0.30 TOTAL --11.7 1.15

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73 CHAPTER 4 DEBRISAT PROPOSED HYPERVELOCITY TEST The previously described DebriSat satellite is currently being design to represent typical LEO satellite missions and will eventually be used to improve the NASA Standard Breakup Model A hypothetical hypervelocity impact test is now proposed to describe h ow an impact would occur to provide new fragmentation data for updating the empirical NASA Standard Breakup Model and, subsequently, improving the outcomes of long term orbital environment population models. The primary objective of this test would be to p rovide detailed information on the fragmentation of a modern spacecraft, giving primary consideration to advances in composite materials, CIC solar cells, and new construction methods. Need for the DebriSat test arose following the accidental collision of Iridium 33 and Cosmos 2251 in 2009 the first between two intact satellite bodies. Following application of the existing Standard Breakup Model results of the predicted fragment generation for the Cosmos 2251 fit well with the observed data from the SSN and Haystack monitoring telescopes. However, the predicted fragmentation of the Irdium 33 from the breakup model did not match with the observable data from the resulting trackable items. Th e principal shareholders of such a test may include: The satellite target construction team Hypervelocity facility operators Government stakeholders including NASA and the DoD Additional private entities identified as significant stakeholders

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74 Test Objectives The primary objective of the proposed hypervelocity impact test is to characterize the debris fragmentation field resulting from a catastrophic impact to a modern LEO satellite. Specifically, the test is to determine the resulting fragment population that is generated from the impact, the mass of each frag ment, and the cross sectional area of each fragment as averaged by its characteristic length. The secon dary objectives of the test include determination of the impact dynamics immediately following the impact collision and determination of secondary char acteristics that include the change in fragment velocity and their corresponding trajectories. The target for this proposed hypervelocity test will be the DebriSat satellite being designed for updating the NASA Standard Breakup Model shown in Figure 4 1 Figure 4 1 Proposed hypervelocity impact target DebriSat

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75 Test Range The proposed impact test must be conducted at a facility with sufficient capability to generate a hypervelocity impact. The facility equipment should be able to fire a projectile tha t generates an impact energy to target mass ratio of at least 40 J/g the threshold for determining if a collision is catastrophic or non catastrophic [1] In addition, the selected facility must have the capability to collect the resulting fragmentation objects in a manner that ensures their survivability and the measurement of flight velocities and trajectories. Several facilities exist in the U.S. with the capability to generate impact velocities and facilities to meet the requirements of the test. These facilities include [29] : White Sands Test Facility Hypervelocity Test Laboratory in White Sands, NM AEDC Range G in Tullahoma, TN Holloman High Speed Test Track at Holloman Air Force Base, NM Lan gley Research Center Temperature Tunnel in Hampton, VA A specific test facility is not explicitly selected for this proposed hypervelocity impact test Instead, the proposed test is based on typical characteristics of hypervelocity facilities and their capabilities An assumption is made that the impact will occur in a ballistic blast chamber immediately downrange of the hypervelocity gun. The range layout for this proposed test is shown in Figure 4 2 and Figure 4 3.

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76 Figure 4 2 Test range End view Figure 4 3 Test range Top view

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77 Projectile Selection The projectile to be used in t he test has been selected as a 160 g aluminum sphere with a diameter of appr oximately 4.9 cm. The projectile would be lau nched for a collision velocity target of 7.0 km/s that results in a n expected kinetic energy of 3.9 MJ. This corresponds to an expected energy to target mass of approximately 78 J/g much larger than the catastrophic energy threshold of 40 J/g. Test Description The proposed test will be cond ucted in two hypervelocity tests: a pre test impact test using a simulated composite spacecraft propulsion tank and the actual im pact with the completed target satellite. The impact pre test will occur on a simulated composite spacecraft propulsion tank in order to test that the selected facilities are correctly calibrated for capture of the actual impact data. In particular, the pre test will determine if the instrumentation selected with trigger times, sight line, and field of v iew prior to the actual tes t are well calibrated In addition to calibrating range instrumentation the pre test shot data can be used for important breakup characteristics of composite tanks as well as comparison to other hypervelocity impact tests on simulated spacecraft components conducted in the past The test on the target satellite will be the primary hypervelocity impact f or which fragmentation data would be collected and analyzed under the suggested test pa rameters given in this work to update the breakup model The DebriSat target satellite would be orientated in the range facility as shown in Figure 4 4, such that the projectile flight path will impact the center solar panel of the target at an angle of 90 degrees.

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78 Figure 4 4 Target impact trajectory The target will b e positioned directly in the flight path of the projectile such that the center of mass of the target is coincident with the flight path of the projectile. This will correspond to a direct impact to the middle of the target. Four stainless steel eyebolt sc rews will be placed 90 degrees along a 12 inch diameter circle along the zenith hexagonal panel of the target to connect the target with the range facility Four stainless steel braided cables will connect to the four eyebolts and secure the target to the ceiling of the facility such that the ce nter solar panel is parallel to the exit plane of the muzzle

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79 and perpendicular to the flight path of the projectile. The resulting orientation of the nadir hexagonal panel is parallel to the floor of the facility ran ge Fragment Catch System Successful capture of the resulting fragment objects following the collision is critical to the analysis and results of the impact test. A fragment catch system must be installed to the selected facility to allow the collection o f objects following the breakup of the target The catch system its orientation, and its design affect : Collection of fragment objects Determination of fragment velocities Determination of fragment impact angles and trajectories Thus, the proper design and placement of the catch system for the impact is critical to the outcomes of the proposed hypervelocity test. The catch system will provide as much of a bounded and enclosed physical envelope as possible, including considerations for both the projectile 's flight path and necessary instrumentation field of views Catch panels will be lined both parallel to the down range cross section and radially with respect to the down range centerline. Various materials were investigated for use as the catch panel cap ture medium and compared for both their material properties and test characteristics After compiling the material properties and characteristics into a design trade based on a five tier comparison system using poor, fair, good, excellent, and best, as shown in Table 5 1, each property of the foam material was characterized. T he choice of material was reduced to high density polyurethane f oam over commercial ceiling tiles and Styrofoam planks based on its high density and stiffness

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80 Table 5 1 Catch panel material c omparison Property Styrofoam HDPE Ceiling Tiles Breakoff Poor Best Good Size Availability Good Good Good Path Clarity Fair Good Fair Cost Best Fair Good Stiffness Poor Excellent Good Flammability Fair Good Good Toxicity Fair Good Good Density Poor Good Fair As shown in Figure 4 5 e ach individual catch panel assembly is composed of multiple sheets of foam measuring 48 inches length, 24 inches width, and 10.75 inches depth. An individual panel assembly is made of two pri mary materials polyurethane foam and shatter resistant Lexan. Both the Lexan and foam are cut to the proper dimensions and fastened together by four, 1 inch diameter, stainless steel bolts. Each individual panel is then mounted to the p roper position within the test range. The use of Lexan as backing for the panel provides structural rigidity for the panel as well as acting as prevention for debris fragments from passing completely through the foam panel. Because of its homogeneous high density, the HDPE foam allows the high velocity particles to enter the panel and leave a definitive angle and trajectory through the foam. Each foam panel can later be analyzed to determine the velocity, based on the depth into the foam, and the trajectory based on the angle and position in the foam. A total of 24 panel configurations will be used in the test, with 16 panels orientated radially with respect to the center axis of the test chamber and 8 panels orientated parallel to the exit plane of the gun Holes and cut aways will be added for considerations to line of sight of range instrumentation.

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81 Figure 4 5 Catch panel assembly dimension This proposed system is similar to the catch panel system used for the SOCIT tests described in Chapter 2 SOCIT utilized foam planks backed with plywood to collect the fragments produced by the hypervelocity test. These panels were analyzed in a similar method by JPL using laser radar to determine the ballistic trajectory and velocities of the fragments that were c ollected within the panels The proposed test for the DebriSat project at the AEDC facility however, currently utilizes ceiling tile panels. While this method collects debris fragments, based on the design study of panel materials, it may not be the best d esign choice for post test determination o f trajectories and velocities.

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82 Overall, the design of the DebriSat hypervelocity impact test depends on the finalization and detail of the test objectives and requirements specifically the study of ballistic frag ment velocities and trajectories. The inclusion of ballistic properties into the post test analysis of the results impacts the final design of the catch system and orientation of the target within the range facility. Advanced modern optical and radar instr umentation cannot independently and comprehensively determine ballistic properties. Use of material which leaves physically identifiable and measurable footprints must be considered for a comprehensive study of the fragment ballistic properties. This chapt er has discussed a proposed test for which post test analysis beyond the simple collection of fragments can be used to determine characteristics of the breakup.

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83 CHAPTER 5 POST IMPACT CONSIDERATION Following the delivery and actual impact test conducted on the DebriSat target, a nalysis of the fragm ents generated from the DebriSat hypervelocity impact will begin following the collection and release of the fragments to the University of Florida. Several considerations are now given to aspects of the project immediately following the test. Health and Safety Considerations Because of the nature of the impact test, extremely small particles, possibly on the order of nanometers, must be accounted for to prevent injury and illness to personnel tasked with analyzing the post impact debris fragments. Proper ventilation of the imp act area should occur at the test facility to eliminate airborne particles. Further, considerations to the cleanliness of the settled debris should be given to prevent the recontamination of small particles when being handled. Characteristic Analysis Th e primary objective of the DebriSat project is to identify the physical characteristics of individual hypervelocity collision fragments to include considerations for modern materials in the generation of statistical debris modeling. Specifically, the area to mass ratio of each fragment and the characterization of shape is of particular importance. In previous impact tests, characteristics of fragments were determined by utilizing small samples of fragments to find properties of common size bins. These size bins were typically determined using a characteristic length composed of x, y, and z dimensions, where the x dimension is the greatest point to point projected length in any plane of the fragment, the y dimension is the greatest point to point projected le ngth

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84 perpendicular to the x dimension plane, and the z dimension is the greatest point to point projected length in the plane perpendicular to both the x and y dimension planes. These lengths are then converted to a characteristic length before being again converted to an averaged cross sectional area. While this method has been successfully used in the past, it does not provide an actual cross sectional area for a given fragment and the cumbersome measuring process allows only a fraction of the generated d ebris to be measured. In order to efficiently catalogue the thousands of fragments generated during the impact test, a visual inspection system (VIS) is proposed that will reduce the time taken to analyze each fragment in future impact analysis With the proposed VIS, fragments are loaded on a rolling tra y underneath a mounted camera such that picture taking and measurement of two dimensions are performed in one step using custom software on a nearby computer. The rolling trays allow the n ext debris fragme nt to be loaded in a factory line fashion while the current debris fragment is analyzed under the camera. The diagram in Figure 5 1 illustrates the proposed steps of the process that the VIS will automate. Figure 5 1 VIS Process Flow The proposed VIS is intended to streamline the cataloguing of debris frag ments from a ground based hyper velocity impact test b y eliminating the need to analyze

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85 characteristic properties by hand with calip ers along three orthogonal axes. The proposed vision system could de monstrate an improvement in fragment analysis time, measurement accuracy, and repeatability. Furthermore, the VIS provides consistent measurements for the x and y dimensions, leaving only the thickness dimension to be measured via calipers. Debris fragmen ts are loaded onto a flat tray that is rolled under the camera and spring locked into place ensuring proper alignment each time. The camera focal plane is aligned with the tray surface so that the length and width dimensions are calculated via image proc essing. The proposed Vis setup is shown in Figure 5 2. Measurement of the thickness dimension is then quickly determined as the dimensions orthogonal to the length and width plane. In the traditional method, significant variance from the exact dimensions m ay be expected if all three dimensions are measured by calipers, which is notably problematic for highly irregular shaped fragments where the length, width, and thickness dimensions are not immediately obvious. Figure 5 2 Proposed VIS analysis station

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86 The proposed VIS also allows pertinent information, specifcially cross sectional area, to be determined quickl y by image processing. Figure 5 3 shows a possible method of measuring fragment cross sectional area by image analysis against a known reference a rea. Figure 5 3 Hypothetical VIS software analysis using known objects

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87 CHAPTER 6 CONCLUSIONS AND FUTURE WORK This work explains the development of a representative LEO satellite and a corresponding hypothetical terrestrial impact test for improvement of the NASA Standard Breakup Model The satellite design, specifically the EPS, C&DH, TT&C, and Payload subsystems are representative of material and design for typical modern day LEO satellite missions. A brief overview of the breakup model development and corresponding critical input events is given. Future work includes the development of a comprehensive hy pervelocity test to simulate the conditions of an actual on orbit catastrophic collision. In addition, additional work is required on considerations for analyzing debris fragments for physical characteristics using a computer vision based analysis approach

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88 APPENDIX SPECTROLAB UTJ SOLAR CELLS

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89

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90 LIST OF REFERENCES [1] Johnson, N. L.., Krisko, P. H., Liou, J. C., and Anz Meador, P. D., "NASA's New Breakup Model of EVOLVE 4.0," in Advanced Space Research 2001. [2] Reynolds, R., Eichler, P., Bade, A., Krisko P., and Johnson, N., "Sensitivity Analysis of the Orbital Debris Environment Using the EVOLVE 4.0 Model," 1999. [3] Kessler D. J., and Cour Palais, B. G., "Collision Frequency of Artificial Satellites: The Creation of a Debris Belt," Journal, vol. 83, no. A6, pp. 2637 2646, 1978. [4] Kessler, D. J., Nicholas, J. L., Liou J. C., and Matney, M., "The Kessler Syndrome: Implications to Future," in 33rd ANNUAL AAS GUIDANCE AND CONTROL CONFERENCE 2010. [5] Liou, J. C., "Orbital Debris Modeling and the Future Orbital Debris Environment," Boulder, 2012. [6] Nicholas, J. L., Stansberry, E., Liou, J. C., Horstman, M., Stokley C., and Whitlock, D., "The characteristics and consequences of the break up of the Fengyun 1 C spacecraft," Acta Astronautica, vol. 63, pp. 128 135, 2008. [7] Stansbery, G., Matney, M., Liou J.C., and Whitlock, D., "A Co mparison of Catastrophic On Orbit Collisions". [8] Orbital Debris Program Office, "History of On Orbit Satellite Fragmentations," NASA, 2008. [9] Liou J. C., and Anz Meador, P. D., "An Analysis of Recent Major Breakups in the Low Earth Orbit Region," in 61st International Astronautical Congress, 2010. [10] Wright, D., "Colliding Satellites: Consequences and Implications," Union of Concerned Scientists, 2009. [11] NASA, "Orbital Debris and Near Earth Environmental Management: A Chronology," 1993. [12] Reynolds R. C., and Eichler, P., "A Comparison of Debris Environment Projections Using the EVOLVE and CHAIN Models," Advanced Space Research, vol. 16, no. 11, pp. 127 135, 1995.

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91 [13] Jehn, W. R., "Comparison of Space Debris Models in the C entimetre Size Range," in Second European Conference on Space Debris 1997. [14] Klinkrad, H., Sdunnus H., and Bendisch, J., "Development Status of the ESA Space Debris Reference Model," Advanced Space Research, vol. 16, no. 11, pp. 93 102, 1995. [1 5] Rossi, A., Cordelli, A., Pardini, C., Anselmo L. and Farinella, P., "Modelling the Space Debris Evolution: Two New Computer Codes". [16] Bess, D. "Mass Distribution of Orbiting Man Made Space Debris," NASA, Washington, D.C., 1975. [17] Krisko, P. H., Horstman M., and Fudge, M. L., "SOCIT4 Collisional Breakup Test Data Analysis: With Shape and Materials Characterization," Advances in Space Research, vol. 41, pp. 1138 1146, 2008. [18] Yates K. W., and Jonas, F. M., "Assesment of the NASA EVOLVE Long Term Orbital Debris Evolution Model," Philips Laboratory, Kirtland Air Force Base, 1995. [19] Liou, J. C., Hall, D. T., Krisko P. H., and Opiela, J. N., "LEGEND A Three Dimensional LEO to GEO Debris Evolutionary Mod el," Advances in Space Research, vol. 34, pp. 981 986, 2003. [20] Zarchan, P. Advanced Hypersonic Test Facilities, AIAA, 2002. [21] Cunningham, T., "SOCIT Series Plan," General Research Corporation, 1991. [22] Liou J. C., "The Man Made Orbital Debris Problem and a New Satellite Impact Experiment to Characterize the Orbital Debris Properties". [23] Tsou, P., "Fragment Ballistics Determination," Jet Propulsion Lab, 1992. [24] Hanada, T., Liou, J. C., Nakajima T., and Stansbery, E., "Outcome of Recent Satellite Impact Experiments," Advances in Space Research, vol. 44, pp. 558 567, 2009. [25] Hanada T., and Liou, J. C., "Comparison of Fragments Created by Low and Hyper Velocity Impacts," Advances in Space Rese arch, vol. 41, pp. 1132 1137, 2008. [26] Union of Concerned Scientists, "Satellite Database," [Online]. Available:

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92 http://www.ucsusa.org/nuclear_weapons_and_global_security/space_weapons/tec hnical_issues/ucs satellite database.html. [27] Clark, S., La ne, K., Strickland, T., Fitz Coy N., and Liou, J. C., "Defining a Typical Low Earth Orbit Satellite Using Historical Mission Data to Aid Orbital Debris Mitigation," in AIAA Student R2 2012. [28] Werremeyer, M., "Design of Subsystems for a Representative Modern LEO Satellite," 2013. [29] Arnold Engineering Development Complex, "Impact and Lethality Testing," Arnold Air Forice Base, 2005. [30] Kessler, D. J., "Orbital Debris Environment for Spacecraft in Low E arth Orbit," Journal of Spacecraft and Rockets, vol. 28, no. 3, pp. 347 352, 1991.

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93 BIOGRAPHICAL SKETCH Sheldon Clark was born in 1988 in the central Florida region. A native of the Tampa area, he completed a Bachelor of Science degree in aerospace e ngineering from the University of Florida in 2011. He was immediately acc epted as a research a ssistant with the Space Systems Group under the direction of Dr. Norman Fitz Coy in pur suit of a Master of Science in aerospace e ngineering. Sheldon has been an active that have included project m anager of multiple engineering projects p resident of the Sm all Satellite Design Club, and graduate a dvisor. His prof essional experiences have included intern positions with Raytheon Missile Systems in Tucson, AZ and NASA Goddard Space Flight Center in Greenbelt, MD. His continuing research interests include spacecraft systems engineering and design as well as advancing concepts in small satellite mission utility.